A concept design method for the return of the first sub-stage of a launch vehicle is proposed. It provides a solution for the overall design of a recoverable Launch vehicle. Firstly, a motion model for return of first stage is built. In order to calculate the thrust and its adjustment range, two parameters that named thrust ratio and thrust adjustment factor are defined in the model. Secondly, relation between recovery thrust and left propellant is discussed. A constraint of left propellant is introduced. Finally, a return scheme of a first stage is planned through a simulation example. The method proposed answers such questions as how much propellant should be reserved to realize the return of the first stage, how much thrust should be used, what is the adjustment range of thrust during the return process, the landing point of the stage, the max velocity during the return, and height and velocity before the stage enters the landing phase. The example shows that this method can be used to plan a complete return scheme for the first stage of a launch vehicle.

The low orbit internet constellation represented by "Starlink" is a hot topic in the current development of the aerospace field. Multi-satellite stack technology can significantly improve the utilization rate of fairing space, thereby accelerating the construction speed of low orbit constellations, and is one of the important directions of future multi-satellite launching technology. This article analyzes the requirements and significance of multi-satellite stack technology, summarizes the existing connection forms of multi-satellite stack, and then points out the suitable connection forms for large-scale constellation construction through comparative analysis. The mature connection schemes of multi-satellite stack are interpreted, and the composition principle and technical characteristics of this technology are obtained. Finally, the difficulties and challenges in the stucture design and connection technology of multi-satellite stack are analyzed and prospects are proposed, providing reference for future multi-satellite stack structure design.

This article focuses on the dynamic model of air rudder rotation around the rudder shaft, considering a dry friction model with rotational clearance and Stribeck effect, and establishing the forced vibration equation of air rudder rotation with clearance and dry friction. The average method is applied to analyze the amplitude frequency characteristics of the rotational direction. The influence of different friction torque coefficients and external excitation frequencies on the amplitude frequency response of axial vibration is studied. The results show that when the external excitation frequency changes, the amplitude frequency curve of the vibration around the axis exhibits hysteresis nonlinear characteristics of different softness and hardness. Multiple solution frequency bands appear in the frequency domain, and the frequency domain multiple solution frequency bands increase with the increase of the friction torque coefficient.

Aiming at solving the SBIRS three satellites detection trajectory estimation problem, a data fusion trajectory estimation algorithm based on the GEO satellite and the HEO satellite detection is proposed. According to the SBIRS constellation composition and detection mechanism, the STK is used to analyze the SBIRS coverage capability to a certain area, calculation shows that over three satellites can fully cover it in about 43% of the simulation time. Establishing the three satellites detection data fusion estimation algorithm model to estimate the missile target motion state in real time, the current statistical model is adopted to describe the missile motion state, the centralized structure is adopted to achieve detection data fusion, in addition, the unscented Kalman filter is used as trajectory estimation filter. Simulation results show that, compared with the binary detection trajectory estimation error, the three satellites detection trajectory estimation error is significantly reduced.

In order to enhance the delivery and survival capability of unmanned aerial vehicle group in specific areas in long-range mission scenarios, and effectively complete various missions, a missile borne unmanned aerial vehicle group delivery scheme is proposed. The unmanned aerial vehicle group is delivered to the mission area through missile carriers, utilizing the rapid reentry advantage of the missile to improve the delivery and survival capability of the missile borne unmanned aerial vehicle group. Taking the end interception system "Dense Array" as a scenario, the survival capability of the unmanned aerial vehicle group is simulated and analyzed under two schemes of missile based delivery and parachute based delivery. The effectiveness of the proposed scheme is verified, meeting the requirement of the unmanned aerial vehicle group entering the mission area with a high survival probability in actual mission scenarios.

Interstage thermal separation is one of the core technologies in the flight process of solid rocket. The traditional "three in one" test usually includes three main subsystems: upper stage motor, servomechanism and separation system. For practical engineering development, it is costly, expensive and long cycle. It can only be used for system level performance verification test, and is not suitable for exploratory developmental test. Based on the new ground test method of cold air simulated rocket interstage separation test, combined with theoretical model, numerical simulation and ground test verification, the internal flow field in the initial pressure holding stage of interstage separation is simulated, the law of gas pressure change in interstage section is revealed, and the test mechanism of cold air simulated rocket interstage separation is expounded; Through numerical simulation analysis, the effects of different initial pressure of pressure accumulator, different volume of pressure accumulator, different pipe area and different volume of interstage section on the change of gas pressure are studied, which provides a solution for the design of interstage separation test of cold air simulation.

In order to study the mixing and explosion characteristics of liquid oxygen and kerosene, a partial confined space test system is established to simulate the application scenarios of the launch site and the tests are carried out. It is demonstrated that when liquid oxygen and kerosene are discharged and mixed at the same time, the concentration of kerosene and oxygen increases at first and then decreases. When kerosene temperature is ${65}^{\circ}\mathrm{C}$ and ignition excitation is ${5.9}\mathrm{\;J}$, even if liquid oxygen leaks, the ignition and explosion condition are still not reached at the measuring points arranged above the liquid level. In the test when the kerosene temperature is raised to ${80}^{\circ}\mathrm{C}$, explosion occurs when liquid oxygen is released only for about 20 seconds. At ignition time, the concentration of kerosene, oxygen, and nitrogen are ${1.31}\%,{40.05}\%$ and ${58.64}\%$ respectively, and the ratio of kerosene concentration to oxygen concentration is 0.033 .

For solid-engine aircraft, closed-loop guidance methods with energy matching or angle constraints are generally used in the powered flight segment, which have high control accuracy. However, its accuracy is greatly affected by engine performance deviations. Therefore, a zero-range Orientation closed-loop guidance method based on neural network is proposed to reduce the impact of engine performance deviation on the guidance accuracy. Firstly, the motion model of the powered flight segment of the aircraft is established, and the closed-loop guidance of the zero-range Orientation is analyzed and deduced. Secondly, a multi-input neural network algorithm is designed, the input and output parameters are determined, the residual energy, the velocityto be increased and the angle of the zero-range Orientation are trained. Then, the training results of the above neural network with the zero-range Orientation closed-loop guidance are combined. This method enables feedback of different zero-range Orientation angles under different engine deviations. Finally, different deviation states for simulation verification are chosen. The simulation results show that this method can effectively reduce the influence of engine deviation on the guidance accuracy, and has strong anti-bias ability and high guidance accuracy.

To study the safety of a tower slewing platform structure, a three dimensional numerical model is established with actual sizes and operating conditions. The structural force, pressure margin of hydraulic system with different wind loads, and the influence of lift platforms’ amount are researched which are concerned closely in use. In working condition, the maximum stress of structure is ${223.98}\mathrm{{MPa}}$, and the structure safety factor is 1.58 . The pressure margin of hydraulic system is calculated with the worst wind direction and the maximum windward area. When the platform is opened with permitted wind speed, the hydraulic system pressure can not be lower than ${5.40}\mathrm{{MPa}}$, and the pressure margin is 2.96. The structure safety factor is 1.35 when the number of lift platforms is 9. According to the results of simulating calculation, the security of the slewing platform can be confirmed. The simulating model can provide a reference for the slewing platform’s application.

Thermal protection is a regional protection measure covering the outer surface of the launch pad to prevent the bearing structure of the launch pad from being eroded by the rocket launch gas flow. It is a key link to ensure that the launch pad can reliably carry the rocket safely. In recent years, as the launch frequency of the conventional liquid carrier rocket represented by the Long March 2C rocket has increased year by year, the recovery time of the launch pad after the launch of the rocket has also been greatly reduced. Therefore, a rapid recovery scheme for thermal protection of the launch pad is proposed, which combines the ablation type heat protection technology with the absorption type heat protection technology, and the modular thermal protection components are designed to realize rapid replacement through simulation, test and use. The results show that the repair time of the new thermal protection is shortened from 3 days to 1 day, and the protection effect is good, which can meet the needs of high-density rocket launch.