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.

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.

The optics window is usually faced with severe aerodynamic heating, and there is a prominent image saturation problem because of high temperature. Therefore, researchers often use film cooling to isolate the direct heat convection from the outflow. A tangential high-density cooling film generation structure is designed which is used for high pressure and heat flux. Through reasonable ground test, numerical simulation and engineering calculation, the main factors affecting the cooling efficiency are studied. The preliminary results show that tangential high density gas film can reliably reduce the surface heat at high pressure and high heat flux. The research effectively adapts to the increasingly harsh optical observation environment.

Currently, in fuze two levels of environmental force arming, high dynamic overload is often used as one insurance element. Generally, one arming method is used of the inertia mass block's movement trip under trajectory overload. Mass blocks are limited by mechanical movement and are difficult to integrate in fuze systems, and is poor in testability. A scheme based on MEMS sensor technology is proposed and designed to arm the fuze. The principle, composition, function of the system, software workflow and the test and verify situation are introduced. This method has a wide scope of use can improve the flexibility of fuze systems and is easy to system integration.

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.

Aerodynamic heating is a key issue in the research of hypersonic vehicles flying in near-space. It has an important influence on the aerodynamic, thermal environment and safety of the aircraft. Due to the limitation of experimental methods, wind tunnel experiments cannot simulate real flight conditions accurately. CFD is an important tool for studying aerothermodynamics problems. The format dissipativeness and grid are two important factors that affect aerothermodynamics simulation. The smaller the format dissipation, the better the CFD performance, but low dissipation will cause shock instability phenomena. A hybrid HLLCE format which has both the HLLE format stability and the low dissipativity of the HLLC format is constructed. This format exhibits the low dissipation properties of HLLC at a lower Mach number and can overcome shock instability phenomena at high speed. The thickness of the linear bottom layer of the boundary layer is used as the reference scale, and 1/10 of the thickness of the linear bottom layer calculated by the feature length is taken as the minimum grid scale for thermal environment calculation. The performance of low dissipation scheme is verified by hypersonic sphere example with the proposed grid scale

C/SiC composite material for hypersonic vehicle structure has a broad application in near-space area. The active / passive ablation performance of the C/SiC composite is studied numerically. An approach for active sublimation ablation performance up the ${2000}^{\circ}\mathrm{C}$ of $\mathrm{C}/\mathrm{{SiC}}$ composite materials is proposed and some wind tunnel experiments have been designed and completed. The results show that the ablation performance proposed has good accuracy compared with the wind tunnel results. The results can provide a reference for the structure and thermal protection design and safety assessment of the hypersonic vehicles based on C/SiC composite material.

SpaceX's SuperHeavy Starship interstellar transportation system project is developing rapidly and has now entered the orbital-level test flight stage, which may become another powerful launch vehicle to change the world's aerospace landscape. The different stages of the design iteration of the SuperHeavy Starship system are sorted out. The main technical changes and improvement motivations in each stage are summarized. The technical characteristics and the future application direction of SuperHeavy Starship are analyzed. It is hoped to provide enlightenment for the development of China's space launch vehicles.

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.