In order to solve the problem of vulnerability of the missile-loaded motor shaft and bearing in the high overload environment of 23 000g guided artillery projectile, the anti-high overload brushless DC motor using a combination of disc spring and special steel ball for buffering and vibration damping is designed. Firstly, the acceleration load curve of the missile-borne motor is obtained by calculation. The bearing and shaft of the brushless DC motor are analyzed to be the failure prone parts. Secondly, according to the energy absorption characteristics of dish spring and steel ball, a composite damping structure of brushless DC motor combining dish spring and special steel ball is designed in the case of limited structural space. And based on this structure, a high speed and fast response anti-high overload motor is designed. Finally, the simulation and experimental analysis of the designed motor show that the design of the composite buffer structure can effectively reduce the stress and deformation of the motor shafting parts under the high overload environment. The structure of the anti-high overload motor has no obvious damage after the Marshall's drop hammer test, and the working performance meets the requirements of precision guidance control.

As a critical component of the space transportation system, solid launch vehicle has advantages such as full-vehicle storage, versatility for sea-based and land-based launch, rapid response, and low requirements for launch support. The development process and overall technical scheme of Smart Dragon-3 solid rocket are introduced. In addition, the practical experience in the development of solid launch vehicles is drawn as a brief summery. Furthermore, the series development of solid launch vehicles is claimed, based on the analysis of the future positioning of solid rocket development. The carrying capacity of medium-sized solid launch vehicle needs to be improved, through research innovation, scale development and optimization of comprehensive cost. In this case, the demands of various launch orbits can be satisfied by the style of offhand sea launch.

In response to the problems of large deformation, multi-contacts, and time-varying folding stiffness during the deployment of flexible folding spacecraft, research has been conducted on the dynamic modeling methods of the deployment process of large-flexibility folding structures. Based on the folding method of such large-flexibility structures, a typical folding model is simplified and extracted, and a three-dimensional finite-segment discrete model suitable for the deployment of large-flexibility folding structures is established using the finite segment method. Utilizing vector mechanics, a mechanical model of the connecting forces between adjacent units is derived from the Newton-Euler equations, and the stiffness matrix of the connection forces is provided. Nonlinear equivalent contact spring damping and a continuous friction coefficient Coulomb friction force model are introduced to simulate the contact and friction states between units, ensuring the stability and efficiency of the solution. A time-varying nonlinear elastic connection model at the crease is proposed, along with a parameter identification method for bending stiffness. Using ADAMS software, dynamic simulations of deployment are conducted for typical folding models under four working conditions, and corresponding deployment experiments are designed. The comparison between the two results shows that the change in tensile force during the deployment process of typical folding models matches well, validating the feasibility and accuracy of dynamic modeling for such structures, and laying the foundation for subsequent simulations of the overall structure's unfolding.

As the key interface between launch vehicle and ground support equipment, the cryogenic connector is used for filling and venting of cryogenic propellant, and it falls off before or after launch. Affected by the high temperature and high humidity environment of Hainan launch site, the low temperature surface of the rocket-ground interface is more prone to frost and ice after long-time filling, which hinders the action of the separation and increases the resistance of the connector to fall off. Quality problems affecting the launch process occurred in historical missions. An improved scheme for the rocket-ground interface is proposed, which can reduce or avoid icing at the unlocking part and reduce unlocking resistance through thermal design optimization.

Selective laser melting (SLM) technology is a key technology for parts manufacturing in the aerospace field. With integrating the overall structural optimization design and SLM technology, the overall lightweight manufacturing of complex components can be achieved. In view of the high-performance manufacturing requirements of parts in the aerospace field, the progress in the manufacturing of complex integral components and lightweight structures in the aerospace field by SLM technology in recent years is introduced, and the future development direction is pointed out.

This study investigates the combustion performance of gas-oxygen and gas-methane pintle injectors, which discusses the influence of pintle injector structure design parameters and thrust chamber combustion design parameters on the combustion performance. Theoretical analysis is carried out by numerical simulation and other methods. The simulation results show that the oxidizer central pintle injector has a higher combustion performance. A larger angle of the pintle head and a smaller distance of the intermediate sleeve have a higher combustion performance. A smaller pintle head radius can significantly play a role in cooling the pintle head, but there is a certain loss of combustion performance. When the speed ratio is about 1.1, the combustion performance is higher, and the temperature of the pintle head is lower. When the mixing ratio of oxygen and methane is around 3, the combustion performance is higher.

A 5B70 aluminum alloy sheet with ${1.5}\mathrm{\;{mm}}$ thickness is welded by tungsten inert gas (TIG) welding with 5B71 filler wire. The micro-structure evolution and refinement characteristics of the welded joints are investigated. The research results indicated that weld joints with excellent formation can be obtained by adopting reasonable welding parameters. The internal structure of the weld seam shows that the weld zone is mainly composed of equiaxed crystals forming a cast structure, the size of the grain structure is uneven, there are widely larger grains with a diameter of ${40}\sim {50\mu}\mathrm{m}$, as well as ultrafine grain areas with a diameter of about ${20\mu }\mathrm{m}$ which distributes in fine strips and small blocks. The area of the ultrafine grain region is obviously smaller than other areas. It mainly distributes in the direction parallel to the fusion line, with a few distributed on the weld surface. Sc and Zr were used as modificator to refine the grains. The ${\mathrm{{Al}}}_{3}\left({\mathrm{{Sc}},\mathrm{{Zr}}}\right)$ second phase particles are precipitated in the welding pool during solidification. The function of such particle is to form heterogeneous nucleation particles, reduce nucleation power and increase the number of crystal nuclei. The area where Sc elements are enriched had a higher degree of grain refinement. The “undercooling” zone is formed along the front of the solid-liquid boundary, which promotes the formation of equiaxed grains. It results in a higher degree of grain refinement, which size is only about one half of other zones.

When the design input is limited during the initial stage of scheme argumentation, in order to solve the problem of how to conduct control capability analysis quickly and effectively, the adaptability of traditional control capability analysis method is analyzed. A set of controllability analytical method for reentrant launch vehicles is researched by dynamic modeling and original method improving. And under the premise of limited control capability, the constraint conditions for allowable flight conditions is limited preliminary in order to quickly locate and feedback the closure of the overall unit control capability.

Aiming at the characteristics of multiple hypersonic vehicles cooperative combat, a time cooperative reentry guidance scheme based on deep deterministic policy gradient and linear quadratic regulator (DDPG-LQR) is proposed. Firstly, the sequential convex programming method is used to generate the time cooperative reentry trajectory satisfying multiple constraints and its corresponding steady-state control quantity. The Radau pseudospectral method is used to discretize the motion equations to improve the discretization accuracy of trajectory optimization. Secondly, the linear quadratic regulator (LQR) is used to track the time cooperative reentry trajectory. In order to improve the cooperative guidance accuracy and guidance effect, the deep deterministic policy gradient (DDPG) is used to optimize the weight matrix coefficients of the LQR online. In the DDPG algorithm, the optimization performance of the algorithm is improved by introducing an appropriate reward function. The simulation results show that the cooperative guidance scheme proposed has better cooperative guidance effect and guidance accuracy than the traditional LQR controller in the case of initial state error and uncertainty.

The engine nozzle is prone to work with a state of over-expansion in a high-pressure underwater environment, which will influence the jet flow field and thrust. Experiments of over-expansion supersonic underwater gas jets at different co-flow velocities are conducted using a circulating water tunnel, various forms of wake cavity and their corresponding thrust time-frequency characteristics are analyzed. It is shown that at low velocities, the over-expanded gas jet is difficult to form a cavity attached to the wake or a pulsating foam-type wake cavity. Under such flow conditions, the vehicle experiences pronounced thrust oscillations with a rich frequency spectrum (36 to ${743}\mathrm{\;{Hz}}$). As the Froude number at the co-flow velocity increases to ${Fr}= {8.57}$, the jet eventually forms an intact tail cavity, at which point the amplitude of the thrust oscillations decreases significantly and the mean thrust value increases significantly. The internal gas reflux phenomenon in the jet-induced tail cavity is also directly observed, and the observed reflux velocity at ${Fr}= {8.57}$ is about ${1.13}\mathrm{\;m}/\mathrm{s}$.