As an important component of aircraft guidance system, non-metal radomes have multiple functions such as heat protection, wave transmission and load bearing. It is connected to the aircraft body through high-temperature resistant adhesive agent. Radome is frequently subjected to harsh mechanical and thermal loads during the period of service, and the joint structure is often the weakness of strength design. Therefore, accurate simulation for mechanical behavior of the joint structure and achieving accurate prediction of load-carrying capacity are crucial for the structural design and optimization of radome. Based on bilinear cohesive model, the failure behavior of radome adhesive interface is numerically described by ABAQUS. And then the damage evolution of adhesive layer is simulated, and the accurate prediction of bearing performance of radome joint structure under flight thermo-mechanical environment is achieved. The corresponding experimental study is carried out. The experimental results are in good agreement with the simulated ones, verifying the validity of numerical method. It provides an effective way to solve bearing problem of aircraft-level radome joint structure.

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.

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.

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.

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.

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.

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.

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