In the industrial field, tailored blank forming with aluminum alloy (Al-alloy) has developed fast to meet the demands for large size integrated components with curved surfaces of high precision and with uniform mechanical properties. Traditional forming methods for tailored blank components faced challenges with uneven deformation behaviors and coexistence of rupture and wrinkling defects occuring during the forming process. In this paper, a new manufacturing procedure is proposed with advanced welding and forming technologies for forming integrated shell components. Friction stir welding with post-weld heat treatment was employed to prepare the tailor welded blank and improve its formability prior to forming. A double-sided pressure sheet hydroforming process was introduced to fabricate the Al-alloy tailored blank into a curved surface shell. Finite element modeling was established to analyze the effect of the weld line position and loading paths of stress distributions during the double-sided sheet hydroforming (DSHF) process. A large double-action CNC sheet hydroforming press with tonnage of 150 MN and high pressure liquid volume of 5 m3  was developed in China. As an application case of the proposed process and equipment, a full-scale tank dome with a diameter of 3 m was successfully hydroformed with a large size Al-alloy tailored blank. It was shown that the DSHF process has the advantages in controlling rupture and wrinkling defects with an Al-alloy tailored blank, and the novel manufacturing procedure enables the production of integrated thin-walled component more competitively than traditional methods.

The slug rivet is widely used in wing assembly due to its longer fatigue life and better sealing performance compared with other connection technologies. As a countersink with dual-angle is widely adopted for this type of connection, the countersink diameter and depth are key factors that affect assembly quality. Therefore, it is of great importance to efficiently inspect the countersink quality to ensure high accuracy. However, contact measurements are susceptible to the loss of accuracy due to cutting debris and lube build-up, while the hole-scanning method using laser profilometry is time consuming and complex. In this paper, a non-contact method for countersink diameter and depth measurement based on a machine vision system is proposed. The countersink diameter can be directly measured by the machine vision system, while the countersink depth is determined through the countersink diameter indirectly. First, by means of image processing technology together with an improved edge detection algorithm, the countersink diameter can be obtained. Then, a 3D microscope is employed to measure the countersink depth, which helps to model the countersink. As a result, once the countersink diameter is measured, so is the depth. The experimentation demonstrated that this method has strong easibility and enables time saving, which is conducive to improve the riveting 
efficiency.

A virtual thermal test system was built through the co-simulation using Simulink and Comsol to realize the complete virtualization of the thermal test. Using the co-simulation technology, comprehensive simulation analysis of the control system, electric field and thermal field was realized. The data state of each observation point could be directly observed at one time, including the output state information of the power amplifier, the output state information of the heater, and the thermal state information of the test unit. The virtual thermal test system has a predictive and guiding role for engineering thermal tests, and can realize thermal environment simulation beyond the existing thermal environment ground simulation capabilities, providing a basis for the development of future models.

In this paper, the properties of different low-density heat-resistant coating prescriptions were compared, and a heat-resistant coating with a density of 0.65 g/cm3, a tensile strength of 2.04 MPa, and a thermal conductivity of 0.126 W/(m·K) was obtained. The thermal performance of the coating was characterized by an oxygen-acetylene ablation test, a hot radiation test using a quartz lamp and in an arc-heated wind tunnel test. The results indicated that the low-density heat-resistant coating prescription has advantages of high temperature resistance, erosion resistance, ablative resistance and excellent heat resistance, which can satisfy the heat resistance requirements of new-generation launch vehicle fairings.
The successful application of this heat-resistant coating technology resolved the problems of low efficiency and poor adhesive performance of the traditional hand-pasted cork process, and realized the rapid spray application of the thermal protection layer, effectively improved the manufacturing efficiency and production quality of the heat-resistant layer of new-generation launch vehicle fairings.

In order to meet the urgent need for diversified and multi-functional deployable antennas in many major national aerospace projects, such as interstellar exploration, the fourth phase of lunar exploration project, and the industrial application of BeiDou, a deployable antenna structure composed of hexagonal prism and pentagonal prism modules is proposed. Firstly, the arrangement and combination rules of pentagonal prism and hexagonal prism modules on the plane were analyzed. Secondly, the spatial geometric model of the deployable antenna composed of pentagonal prism and hexagonal prism modules was established. The influence of module size on the antenna shape was then analyzed, and the kinematic model of the deployable antenna established by coordinate transformation. Finally, the above model was verified using MATLAB software. The simulation results showed that the proposed modular deployable antenna structure can realize accurate connection between modules, complete the expected deployment and folding functional requirements. It is hoped that this research can provide reference for the basic research and engineering application of deployable antennas in China.

Cure and decomposition reaction kinetics of typical organic materials in aerospace applications are introduced. From the data of dynamic differential scanning calorimetry (DSC) experiments, and based on changes of the peak temperatures (Tp) with different heating rates (β), a linear equation, Tp=T1+∆Tlnβ, has been obtained more reasonably. The above equation can be used to explain some laws of higher or lower of apparent activation energies (Ea), by which the apparent activation energy (Ea) is nearly equal to RT12/∆T. A number of kinetic investigations of typical thermosetting resins and energetic materials in aerospace applications were chosen to validate the above equations.

Noise characteristic is one of the important factors to be considered during the design of a launch vehicle system. In this paper, the acceptance conditions for the external noise environment of the instrument cabin are given based on multi-source data, including the measured data of the launch vehicle lift-off noise, the mechanical environment of the launch vehicle equipment, and the external noise environment of the instrument cabin deduced from empirical formula. Then an acoustic and vibration transfer model is established based on the response data of the instrument equipment used to conduct measurements in the noise test reverberation room. By using an external noise environment and a transfer model, the response of the instrument and equipment for the acceptance condition can be obtained. The acoustic and vibration prediction technology introduced in this paper can provide technical support in the environmental prediction analysis of heavy launch vehicles for the future.

For the final stage of a launch vehicle which cannot start multiple times, the main constraint of a highly ec-centric orbit launch mission is the argument of perigee, which is conditioned by the range. This paper studies the pay-load capacity at different launch sites for a highly eccentric orbit (e.g. GTO) and gives a method to improve the payload capacity with an argument of perigee constraint by increasing the sliding time before the final ignition and simultaneously adjusting the launch azimuth and yaw at the final stage. An example of launching to GTO orbit on a rocket from Hainan is given, which proves that the method has strong engineering value.

For the final stage of a launch vehicle which cannot start multiple times, the main constraint of a highly ec-centric orbit launch mission is the argument of perigee, which is conditioned by the range. This paper studies the pay-load capacity at different launch sites for a highly eccentric orbit (e.g. GTO) and gives a method to improve the payload capacity with an argument of perigee constraint by increasing the sliding time before the final ignition and simultaneously adjusting the launch azimuth and yaw at the final stage. An example of launching to GTO orbit on a rocket from Hainan is given, which proves that the method has strong engineering value.

For the final stage of a launch vehicle which cannot start multiple times, the main constraint of a highly ec-centric orbit launch mission is the argument of perigee, which is conditioned by the range. This paper studies the pay-load capacity at different launch sites for a highly eccentric orbit (e.g. GTO) and gives a method to improve the payload capacity with an argument of perigee constraint by increasing the sliding time before the final ignition and simultaneously adjusting the launch azimuth and yaw at the final stage. An example of launching to GTO orbit on a rocket from Hainan is given, which proves that the method has strong engineering value.

For the final stage of a launch vehicle which cannot start multiple times, the main constraint of a highly ec-centric orbit launch mission is the argument of perigee, which is conditioned by the range. This paper studies the pay-load capacity at different launch sites for a highly eccentric orbit (e.g. GTO) and gives a method to improve the payload capacity with an argument of perigee constraint by increasing the sliding time before the final ignition and simultaneously adjusting the launch azimuth and yaw at the final stage. An example of launching to GTO orbit on a rocket from Hainan is given, which proves that the method has strong engineering value.

The divergence angle of laser beam used in space laser communication is usually no more than 100 μrad. Using laser beam with small divergence angle to achieve acquisition and tracking for space laser link has al-ways been a difficult problem. In addition, the random nature of the atmosphere will affect the satellite-ground laser link, which increases the difficulty of the acquisition and stable tracking for the laser link. Thus, taking into account the above challenges for satellite-ground laser communication, an acquisition and tracking scheme of using both beacon beam and signal beam was designed for the Laser Communication Terminal (LCT) of Shijian 20 satellite. In-orbit test results indicated that under the condition of moderate atmospheric turbulence (atmospheric coherence length r0≈3cm), the process of acquisition and tracking for the satellite-ground laser link can be completed within 1s after the initial pointing between the LCT and Optical Ground Station (OGS) is performed, and the tracking error was less than 1 μrad (3σ). In addition, the laser link can be re-established quickly once being interrupted by unsteady atmospheric turbulence, and can be maintained for a long time under moderate twurbulence conditions, which lays a foundation for future application of satellite-ground laser communication.

Large lift-to-drag ratio, high maneuverability, and good controllability are the basic performance for flight vehicles. Studying the rolling stability problems of a high lift-to-drag ratio aircraft is of great significance to the safety and control in maneuvering flight. The research on the static stability in the rolling direction of a HTV-2 like shape under a typical Mach number and attack angle was carried out. Similarly, by using Euler, laminar, turbulence and transition models, investigations of the dynamic stability in a single degree of freedom rolling motion of the same shape structure were executed. The numerical results show that the dynamic derivative of Euler is the largest, and the dynamic derivatives resulting from laminar, turbulence, and transition models are not much different.

This paper first introduces the technical requirements for autonomous flight, with a brief review of the International Academy of Astronautics (IAA) study group, “autonomous dynamic trajectory optimization control of launch vehicle”. Two research scenarios, ascent rescue and powered descent, are compared from the viewpoint of optimal control. On this basis, the technologies on the autonomous trajectory planning and control under the thrust-drop failures in the ascending phase, and the autonomous guidance method during the powered landing for the recovery of the rockets are discussed respectively. For the ascending problem, the characteristics of different solutions, including the iterative guidance method (IGM)-based residual carrying capacity evaluation, the state-triggered indices (STI), the joint planning with the payload’s performance, and the multiple graded ptimization (MGO), are analyzed for comparison. For the landing problem, the challenges such as the feasible region reduction caused by high thrust weight ratio (HTWR) and the disturbance adaptability brought by the limited feasible region, are studied in detail, as well as the onboard planning demonstration flight in China are introduced. Finally, the foundations supporting the above methods are summarized, which play an important role in promoting the flight autonomy.

Large constellations have developed rapidly in recent years because of their unique advantages, but they will inevitably have a major negative impact on the space debris environment, leading to its deterioration. The key to mitigate the impact is the success rate and duration of the post-mission disposal (PMD) process. Aiming at solving this problem, this paper further studies the impact of large constellations on other space assets under different PMD strategies through simulation, and proposes corresponding strategies and suggestions for mitigation.
According to OneWeb’s large constellation launch plan, the dangerous intersection of the large constellation with existing space assets at different stages of the constellations life cycle is calculated by simulation. Based on this, the influence of the large constellation operation on existing space assets at different times and strategies of PMD is analyzed. The conclusion shows that in the PMD stage, large constellations have the greatest impact on existing space assets; the PMD duration and number of satellites performing PMD at the same time are key factors to the degree of negative impact. The faster the PMD is, the less threat it poses to other spacecraft. More results and conclusions are still being analyzed.

satellite of the China national geophysical field observation series of satellite missions, the China Seismo-Electromagnetic Satellite (CSES) was designed upon an optimized CAST2000 platform for a sun synchronous orbit. Onboard CSES, there are total eight types of scientific payloads including the Search-coil Magnetometer, Electric Field Detector, High Precision Magnetometer, GNSS Occupation Receiver, Plasma Analyzer, Langmuir Probe, Energetic Particle Detector Package, and a Three-band Transmitter to individually acquire the global electromagnetic field, electromagnetic waves, ionospheric plasma parameters as well as energetic particles. Up to now, CSES has been operating normally in orbit for 2 years. By using the various sensor data acquired by CSES, we have achieved scientific research in the areas of the global geomagnetic field modeling, space weather, earthquake event analysis, the Lithosphere-Atmosphere-Ionosphere coupling mechanism and so on.

As the first satellite of the China national geophysical field observation series of satellite missions, the China Seismo-Electromagnetic Satellite (CSES) was designed upon an optimized CAST2000 platform for a sun synchronous orbit. Onboard CSES, there are total eight types of scientific payloads including the Search-coil agnetometer, Electric Field Detector, High Precision Magnetometer, GNSS Occupation Receiver, Plasma Analyzer, Langmuir Probe, Energetic Particle Detector Package, and a Three-band Transmitter to individually acquire the global electromagnetic field, electromagnetic waves, ionospheric plasma parameters as well as energetic particles. Up to now, CSES has been operating normally in orbit for 2 years. By using the various sensor data acquired by CSES, we have achieved scientific research in the areas of the lobal geomagnetic field modeling, space weather, earthquake event analysis, the Lithosphere-Atmosphere-Ionosphere coupling mechanism and so on.

Along with the increase of the number of failed satellites, plus space debris, year by year, it will take considerable manpower and resources if we rely just on ground surveillance and early warning. An alternative effective way would be to use autonomous long-range non-cooperative target relative navigation to solve this problem. For longrange non-cooperative targets, the stereo cameras or lidars that are commonly used would not be applicable. This paper studies a relative navigation method for long-range relative motion estimation of non-cooperative targets using only a monocular camera. Firstly, the paper provides the nonlinear relative orbit dynamics equations and then derives the discrete recursive form of the dynamics equations. An EKF filter is then designed to implement the relative navigation estimation. After that, the relative “locally weakly observability” theory for nonlinear systems is used to analyze the observability of monocular sequence images. The analysis results show that by relying only on monocular sequence images it has the possibility of deducing the relative navigation for long-range non-cooperative targets. Finally, numerical simulations show that the method given in this paper can achieve a complete estimation of the relative motion of longrange non-cooperative targets without conducting orbital maneuvers.

Rapid development of Chinese commercial launch vehicles brings new challenges under the traditional systems engineering (TSE) working method. A new model-based systems engineering (MBSE) working method was proposed for Smart Dragon 1 (SD-1), which is a low-cost commercial launch vehicle developed by the China Academy of Launch Vehicle Technology (CALT). Based on the characteristics of a commercial launch vehicle, the system model based on information cards was established. Through a problem-oriented working method, risk identification and management, the process of Card-MBSE was utilized and verified by the success of the maiden flight of SD-1. This paper introduces a new method and reference for the development of low-cost and high-reliability launch vehicles.