In order to achieve the measurement of unknown space optical environments for star sensors, the design of a wide and multi-spectral filter device integrated into the star sensor is proposed in this paper. The star point information and the spectral information of the space optical environment can be covered by the collected star images. On the basis of this process, experiments scheme of spectral response and radiation calibration and algorithm of brightness distribution inversion of the optical environment are conducted. The experimental results show that the inverted brightness of the optical environment is in good agreement with the measured values. The space optical environment for star sensor applications in real time can be estimated by applying this method conveniently based on a single star image. It can be served as a basis and reference for design optimization of subsequent star sensors and selection of observation windows.

An adaptive sliding mode control method is proposed for the control problem of a quadrotor UAV composed of six degrees of freedom under model uncertainties and unknown disturbances. This method can be used to realize position and attitude tracking control of quadrotor UAV. Firstly, based on the dynamics quadrotor system, the fully actuated and under-actuated subsystems are divided. The model uncertainties and unknown disturbances without upper and lower limit constraints are fully considered, and the lumped disturbance terms of each subsystem are extracted; Then, with the help of radial basis function neural network, the real-time approximation and estimation of the equivalent controller containing the extracted lumped disturbance terms are inplemented in progress. At the same time, adaptive control method is used to estimate the approximation error term. The sliding mode controller with the estimated values of the equivalent controller and approximation error terms and the corresponding adaptive update laws are designed. Based on Lyapunov theory, the reachability and convergence of the sliding mode surface where the state trajectories of each subsystem lies are analyzed and introduced. Finally, the comparative simulations are used to verify the effectiveness of the proposed method fairly well.

A differential games guidance law is proposed for the difficulty of intercept of hypersonic vehicle. Firstly, the model of air confrontation is established between friend and foe in three-dimensional space and the optimal performance indexes are proposed. The optimal strategies are resolved according to Riccati Equation updated in real-time that is based on state-dependent coefficient. In order to solve Riccati Equation, more accurate Schur method and approximate calculation that has but faster θ-D method are applied to obtaining two kinds of differential games guidance law. In the operational scenario of high speed serpentine maneuvering flight target, traditional proportional guidance and differential games guidance law are analyzed by using the simulation results. Compared with the performance, a conclusion is drawn that in face of high speed maneuvering target, the effect of differential games guidance law is superior to proportional guidance.

Aiming at unexpected the sudden change in inertia parameters of spacecraft on-orbit servicing, real-time identification methods of inertial parameters based on on-orbit attitude and control information are proposed, and recursive least squares (RLS) and extended Kalman filter (EKF) identification algorithms are designed. Adaptive forgetting factors are introduced into the RLS algorithm, which assign weights of prior data and current data during each iteration process to ensure the identification value tracking in time. In the EKF algorithm, the influence of parameter variation on the prior prediction covariance is defined, and it is substituted to update the prediction covariance matrix to solve the mutation of inertia parameters. Under consideration of the mutation in inertia parameters, the simulation results show that the identification accuracy of RLS and EKF can reach 1.5% and 1%, and the identification time is better than 30 s and 40 s respectively; In the scenario of time-varying inertia, both methods can achieve on-orbit identification of inertia parameters, and the identification accuracy fulfills the requirements of the attitude control system.

Regarding the controller design of the variant vehicle attitude control system, an integrated controller design method based on structure search and parameter optimization is proposed, which can realize the automation of the controller structure design work, searches for the most suitable controller structure and completes the parameter rectification in a shorter time. The simulation study on the attitude control of the pitch channel of a certain type of variant vehicle shows that the method can significantly improve the bandwidth of the system, and the number of iterations and the consumed time are both promoted by only about 38% than those of the traditional method.

In order to solve the problem that the filtering result of RIMU/GNSS integrated navigation system based on filtering method is unstable under the condition of large maneuvering of carrier, an estimation method based on factor graph is proposed. The state equation and measurement equation oriented to the redundant inertial navigation system are established, and the redundant inertial data are fused into the carrier coordinate system three axes. A data fusion method based on factor graph is established, and the fused inertial data and satellite information are abstracted as factor nodes, and the state information is abstracted as variable nodes. The cost function including inertia factor and GNSS factor is established and the state quantity is estimated in a nonlinear optimization way. The reslult of numerical simulation shows that performance of using the factor graph method can effectively reduce the position error of the carrier and the root mean square error is obviously degraded than that of the Kalman filter.

Aimed at the problem that traditional target tracking methods require prior position information of the target which are not suitable for non-cooperative target tracking, a target tracking method based on wide-angle camera image error for video satellite is proposed. A wide-angle camera imaging model is firstly established, and then the difference is calculated between the projected position of the target in the camera image plane and the expected position. The relative error, angular velocity and attitude error between the target and the small video satellite are calculated by image error, which are used as feedback terms to design a nonlinear controller. Finally, the stability of the system is demonstrated by using the Barbalat’s lemma. The simulation results show that the target driven to the desired position in the image plane can be effectively controlled in a wide field of view of a wide-angle camera and the effectiveness of this method is verified by Tiantuo-2 satellite on orbit data.

Regarding the issue of high angular velocity damping in satellites, a velocity damping strategy using attitude control thrusters is studied, which is token into account the design principles of fewer startup times and longer startup times for attitude control thrusters. A collaborative damping strategy is proposed, which combines the three-axis simultaneous damping method with the single axis sequential damping method. When the angular velocity of any two axes of the satellite exceeds the set angular velocity threshold, a Schmidt trigger is used for three-axis simultaneous damping to achieve effective damping of the satellite's large angular velocity. On the contrary, single axis damping is used in order of angular velocity to reduce the number of times that the thruster is turned on or off. The numerical simulation results show that the collaborative damping strategy can effectively achieve satellite high angular rate damping and effectively reduce the number and duration of thruster startup.

Regarding meeting the task requirement of rapid missile launch, a fast firing data solving method based on neural network is proposed. Firstly, the mapping relationship among terminal velocity, altitude, velocity inclination and firing data is established, and the LM optimization neural network model is derived. Then, based on Bayesian regularization theory, the BP neural network structure is designed, and the optimized network structure which can meet the accuracy requirements is obtained. Finally, Newton iteration method is used to generate the database as training set which is used to train the neural network, so that the network model with optimized parameters is obtained, and the simulation is implemented. The theoretical and simulation results show that the rapid calculation of the firing data before shooting can be achieved by applying this method.

The design scheme and in-orbit verification results of the high-precision star tracker is presented, which is used in the Shenzhou series of spacecraft in China. In the lens design, a front-placed diaphragm structure is adopted to meet the needs of miniaturization and long-term applications; In the overall structure design, a "frame combination wrap-around" integrated structure is adopted to focus on lightweight and stability; In the circuit design,radiation-resistant reinforced APS image sensors, processors and ASIC are applied to ensurance of high-sensitivity detection and high-reliable information processing throughout the service life; In the algorithm, clustering extraction, fast triangular recognition, curtain compensation and dynamic adjustment of exposure time, etc. are used to achieve key performance indicators such as precision, update rate, dynamic and capture; In addition, the thermal stability of the whole machine is improved to ensure that the performance of the star tracker can be maintained. The star tracker is successfully applied to a series of major space projects, including manned spaceflight, lunar exploration and Beidou-3 which have more than 200 units in orbit operation.

Aiming at using the attitude control scheme of taking rotator as ball-borne gondola, stepper motor is applied as the power source and a cascade controller based on active disturbance rejection control (ADRC) algorithm is designed, and the key parameters in the controller are optimized by using dung beetle optimization algorithm (DBO). The optimized controller parameters are then used to simulate the attitude control of the gondola, considering dropping ballast during the flight. The simulation results show that the proposed control algorithm can track static and dynamic targets with better control accuracy and anti-interference ability. Furthermore, the effectiveness of the DBO is also confirmed.

A pulse threshold relative orbit control method is proposed to address the issue of maintaining the formation configuration of inner formation satellite in case of collision due to the perturbation difference between internal and external satellite. A periodic-start pulse threshold controller is designed, and the applied pulse impulse is calculated based on the relative position and threshold of the inner and outer satellite centroids as well as the linearized relative orbital motion dynamics model. The Monte Carlo method is used to simulate the state estimation bias of the inner and outer satellite,which is compared with two continuous control methods under the same conditions. The results show that a good balance can be achieved by using this method regarding control accuracy, fuel consumption, and the influence of measurement error, which can effectively reduce the uncertainty caused by state deviation on fuel consumption and can be applied to the maintenance of formation configurations of inner formation satellite with low fuel consumption in orbit for a long time. The proposed method can also be applied to the maintenance of general formation satellite configurations and planetary companion flights.

A designed method is proposed in this paper for the parameter adjustment of three-loop autopilot with pseudo-angle. Ulteriorly, regarding the open-loop performances indexes of the control system, the stabilizing and the overload loops are successively disconnected at the rudder control, and the relationship is established between open-loop parameters and the three-loop autopilot parameters with pseudo-angle. Then, by taking the open-loop performance parameters as independent variable and the closed-loop performance parameter as optimization goal, particle swarm optimization algorithm structure and boundary conditions are designed, and the quickly parameters adjustment of three-loop autopilot with pseudo-angle is realized. The simulation results show that good open and closed loop performance of three-loop autopilot can be achieved through the designed method based on particle swarm optimization.

The requirements of the flexibility of aerospace control system are extensively analyzed in this paper，and the challenges are deeply discussed，which are concentrated on the system level and the device level. The design concept of the flexibility of the system level and the device level is proposed in innovation，and the key technical path of the flexibility of the system level and the device level is further elaborated，such as rapid assembly and interoperability management of system software and hardware，multi-level communication interconnection，flexible integrated conformal layout and analysis，and verification of the adaptability of flexible electronics to the aerospace environment，and the development trend of the flexibility of the aerospace control system is finally visioned in terms of integrating innovative technologies，strengthening cross domain cooperation，establishing standards and specifications，and expanding and deepening applications.

In this paper, a novel predefined-time line of sight (LOS) angle rate convergence guidance law is proposed for the diverse engagements in three-dimensional (3D) scenarios. By using this guidance law, the LOS angle rate can be converged to zero in an arbitrarily prespecified time, which is independent of the initial conditions and control parameters. Besides, there exists an exact upper bound of the convergence time. Firstly, a planar guidance law is designed to introduce guidance commands in two-dimensional (2D) scenarios by decoupling the 3D engagement model. Then, the guidance commands can be functional in 3D scenarios by the theoretical analysis. Finally, simulations are implemented to validate the guidance law and show the effectiveness of proposed method.

In order to fulfill the demands for rapid and precise attitude adjustment during the docking process of solar arrays, and to resolve the challenges associated with manual adjustment,this study proposes a leveling strategy, which can update the kinematic parameters of the AAM according to the ground leveling situation, uses the laser tracker to obtain the pose required for leveling, applies closed-loop joint position control based on kinematic inverse solution to realizing the leveling of the moving platform, avoids the false leg in the leveling process through implementation of the traditional outrigger lifting and leveling method, and realizes the automatic adjustment of the AAM from the state of the outrigger stowed away to the state of touching the ground as support and then to the state that the moving platform is automatically adjusted to the horizontal plane level. The feasibility of the leveling strategy is verified through simulation, which provides a solution to false leg-leveling in the engineering application of AAM during solar wing docking.

In response to the requirements of aerospace electromechanical servo systems for fast step response, no overshoot and high steady-state accuracy during tracking continuously changing signals, a composite control is involved on the basis of position-speed-current triple closed-loop feedback control with introduction of feedforward control. The tracking accuracy of continuously changing signals without reducing the dynamic performance and stability of the system is improved by using this control strategy. The design methods of closed-loop controller and feedforward controller are analyzed.Regarding the noise sensitivity of the feedforward function designed by invariance principle, a designed method based on error coefficient method is developed, and experimental verification is conducted. The experimental results show that the servo mechanism system can respond to step commands quickly without overshoot. The system has good tracking accuracy for continuously changing signals, and the system bandwidth has been expanded. The experimental results demonstrate the effectiveness and practical feasibility of the control method proposed in this paper.

In order to solve the low fault diagnosis accuracy of aircraft rudders under complex and dynamically changing operating conditions, an NDTL-CNN fault diagnosis method that combined convolutional neural network (CNN) with network-based deep transfer learning (NDTL) is proposed. Firstly, a simulation model of aircraft rudder fault is established to collect multi-dimensional sensor data under different operating conditions and health states; Then, a CNN is designed to adaptively extract deep features from the fixed operating conditions data, which can effectively capture the fault feature signals of the rudder; Finally, the pre-trained CNN under the fixed operating conditions is fine-tuned and transferred to variable operating conditions for fault diagnosis. The experimental results show that the proposed method improves the diagnostic accuracy of the CNN under variable operating conditions by 15% in a short time and the final diagnostic accuracy of the NDTL-CNN reaches 97.7%, which is capable of accurately recognizing the rudder's health state under complex and dynamically changing operating conditions.

Aiming at the collision rendezvous scene between the space vehicle and the potentially threatening object, a maneuvering and evading strategy of the space vehicle is proposed, which is based on genetic algorithm optimization and BP neural network prediction. By presenting the difference against the traditional maneuver strategy in the last phase of collision rendezvous, the evading strategy is selected after the warning and detection by the proposed strategy algorithm that expands the scope of time and space and takes the position and speed at both sides as input parameters, the time, direction and duration of maneuver as output parameters, miss distance and fuel consumption as performance indexes. The numerical simulation shows that the proposed strategy effectively reduces fuel consumption, increases miss distance and improving the possibility of successful evading, which is compared with the instantaneous maneuver strategy and provides a new idea of safety protection technology for space vehicle under security threat.

In order to avoid the huge workload caused by repeated iterations in the process of guidance and control system design, a three-dimensional integrated guidance and control design under impact angle and rudder surface saturation constraints is proposed, which is more applicable in practice. Based on the 6 DOF nonlinear model of missile and the 3 DOF missile-target relative motion model, an integrated guidance and control model with full-state coupling is established. Based on the dynamic surface control method, a four-stage sliding surface design is applied to establish a controller which satisfies the impact angle and rudder surface saturation constraints, and the closed-loop stability of the system is proven. The correctness, effectiveness and robustness of the proposed method are effectively verified by simulation results.

In order to achieve rapid and accurate recognition of combat intent in complex battlefield environments with fewer air combat data samples, a combat intent recognition model based on meta-metric learning framework is proposed.Regarding received small sample data during air combat intent recognition, a two-way gated recurrent unit network based on the air combat temporal data are developed to realize effective feature extraction, and the attention mechanism is introduced for promoting the network to fully extract the temporal core features of the air combat data when facing the small sample data are merely available for obtaining the inter-class differences which finally achieve a fairly higher recognition accuracy and recognition speed.Simulation results show that the model proposed in this paper has better accuracy and real-time performance for air combat target intent recognition, especially in the case of small sample data.

A method aiming at the formation control problem during the reentry of multi-unpowered vehicles is discussed in this paper. Based on second-order system consistency theory of virtual leader, a distributed formation keeping and control algorithm switching of adjustable assembly time is proposed by selecting optimal coordination variables and introducing a special transition link. The altitude of multi-vehicles are allowed to be changed in a certain range by using this proposed solution that considers the flight ability under the restraint of multi-vehicles formation mission and then realizes the accurate online control of the relative transverse and longitudinal range. Simulation results show that the proposed algorithm can achieve the desired control effect in scenarios of formation assembly and keeping, formation configuration switching and formation lateral maneuver. The algorithm takes into account the effects of the earth’s curvature and rotation on the reentry dynamics, and can adapt to random aerodynamic coefficients and atmosphere density pulling by good robustness.

According to calibration problem of ten redundant laser gyroscopes ‘three-autonomy’ strapdown inertial measurement units, A calibration model and calibration scheme are deduced, which includes five-axis gyros and five-axis accelerometers. The ten-axis redundant inertial measurement unit is divided into two groups known as straight assemble X\Y\Z and oblique assemble S\T. A system level calibration method is adopted for the straight assemble, by establishing a state equation about the calibration error model, taking navigation speed error as the measurement, and estimating the calibration error parameters through Kalman filtering. Using the calibration results of the straight assemble as reference, separate calibration models are established for the accelerometer, gyros scale coefficient and gyro bias. Finally, the calibration of ten redundant laser strapdown inertial units is verified by combining the 19-position system-level calibration method with the discrete calibration method. The calibration results are correct and the characteristics of the calibration errors are analyzed.

A design method for an active disturbance rejection controller based on complementary sliding mode is proposed to address the position and attitude control problem during the point cloud data collection process of quadrotor unmanned aerial vehicle （UAV） that carries a light detection and ranging （LiDAR） for power transmission lines. Regarding the problems of strong coupling，non-linearity，and external disturbance effects under the complex environment of quadrotor UAV，a finite time convergent extended state observer is adopted to estimate the state and lumped disturbances of the quadrotor UAV dynamic system，which introduces the observed lumped disturbances into the active disturbance rejection controller for feedforward compensation. At the same time，complementary sliding mode manifolds are established，and the exponential power function and the integral form of sign function are used to ensure the continuity of the active disturbance rejection controller. Finally，a rigorous proof of the asymptotic convergence of position and attitude tracking errors is provided，which is based on the Lyapunov analysis method，and the simulation results show that the proposed method has higher control precision and can effectively suppress different forms of external time-varying disturbances.

The transfer strategies for multi-objective missions around the Earth-Moon Lagrange points are studied and a two-impulse transfer strategy that allows spacecraft to move between near-rectilinear halo orbits and distant retrograde orbits is designed. Firstly，the spacecraft's orbital dynamics model is established in the synodic frame. Then，the overall design of the transfer strategy is implemented，the relevant optimization variables are analyzed，the objective function and constraints are determined，and the transfer strategy design problem is transfered into a trajectory optimization problem. Furthermore，the feasibility of genetic algorithms and particle swarm optimization algorithms for this specific problem is verified for solving the transfer trajectory. A numerical method is presented through this research by applying genetic and particle swarm optimization algorithms to transfer strategies design of halo orbits，which focuses on recent interest in near-rectilinear halo orbits and distant retrograde orbits. The proposed optimization method has ability of effectively resolving the orbital transfer design problem without prior information and can be applied to various transfer scenarios.

Aiming at the problem of multi-coverage of hot regions,a design scheme of common ground track constellation with regression orbit attribute is applied,and a design method based on genetic algorithm(GA)for optimization of the orbital parameters of a single satellite and analytic method for solving the orbital parameters of a constellation is proposed in this paper. Firstly, a regression orbit model for muti-coverage of hot region is established.Secondly,by taking the optimal coverage effect of a hot region as the performance index,genetic algorithm(GA)is used to optimize and solve the orbit parameters of a single satellite.Finally,the orbit parameters of the constellation are solved by analytical method. The results of theoretical analysis and simulation totally show that the design method can achieve periodic multi-coverage of hot regions and higher time resolution ratio, which present fairly good engineering application value.

Regarding the ascent trajectory of horizontal take-off and landing reusable launch vehicles，a simplex-pseudospectral loop optimization algorithm is proposed. By optimizing the control parameters of the horizontal run section，the initial value of the ascent section is adaptively generated，and taking a certain horizontal take-off and landing reusable launch vehicle as an example，the trajectory optimization design simulation of the launch vehicle from taking off to climbing to the handover point is completed by using the simplex-pseudospectral double-loop optimization algorithm，which verifies the feasibility and effectiveness of the proposed method. On this basis，the influences of engine thrust and sensitive parameters of the flight trajectory on air-breathing mode flight profile and remaining mass are studied. At the same time，a design method of air-breathing mode power compensation is proposed. By comparing with the conventional design method，fuel consumption is further reduced and the carrying capacity of the reusable launch vehicle is improved.

A study is conducted on the problem of achieving attitude safety control under low redundancy in micro-nano satellite. A resource efficient attitude safety control scheme is proposed, which aims to cut off all potential fault sources as much as possible when diagnostic information is not in complete. Only magnetometer and magnetic torquer are used to reconstruct the satellite attitude control system. The proposed scheme can realize angular velocity damping under any initial attitude condition of the satellite, and can control the satellite maneuvering to the sun oriented attitude, which ensures the efficiency of the solar array charging and thus extends the satellite’s lifespan under fault conditions. Mathematical simulation shows that even when the main control components such as gyroscopes, star sensors and momentum wheels are not available, the proposed scheme can still achieve stable single axis safe attitude control towards the sun, which achieves attitude determination precision by better than 2 degrees and the sun orientation precision by better than 5 degrees that greatly improves the satellite’s safety operation ability in orbit.

According to inertial uncertainties and multiple types of actuator faults of spacecraft formation attitude cooperative control, the generalized disturbances and actuator faults of the formation system are estimated by using the approximation property of the broad learning system, in the meantime, the hysteresis quantizer is also used to quantize the control torque signal, thereby reducing the communication speed requirements and jitter phenomenon. On this basis, a composite structure fault-tolerant controller based on model predictive control and fast non-singular integral terminal sliding mode is proposed. The stability of the closed-loop attitude system is analyzed by involving algebraic graph theory and Lyapunov theory. Finally, simulation results show the superiority of the proposed control method compared with existing methods.

Aiming at the problem of multi-aircraft game confrontation under the condition of network communication delay, a cooperative differential guidance law is proposed. Firstly, based on the differential countermeasure theory on the basis of the traditional norm-type performance index, the weight coefficient is designed to increase the weight coefficient and a collaborative differential guidance method is derived under “bang-bang” structure. Secondly, by considering the time delay in the network communication among interceptors, the target acceleration value obtained is an attainable domain instead of an accurate value, and the target acceleration range can be reduced by introducing the observed target acceleration direction information, and the error between the target acceleration and the target actual acceleration can be reduced, thus, the guidance precision of the interceptor is improved. Finally, numerical simulation is implemented to verify the effectiveness of this algorithm, and the simulation results show that this guidance law does not need to assume the maneuvering law of the target in advance and can effectively intercept the maneuvering target under the information time delay.

A solid strap-on launch vehicle joint swing control scheme is proposed, which is dedicated to handle large range thrust variation and the non-synchronous thrust among solid boosters. Firstly, the combustion chamber pressure is used to estimate the real-time thrust of each solid booster, and the controller parameters are accordingly modified on-line due to the thrust variation. Secondly, by adopting the feed-forward control to eliminate the loss of thrust synchronism. Finally, the real time swing subsidence angle compensation for solid booster is adopted to solve the effect of pivot point excursion. The simulation results show that the proposed adaptive control scheme is able to efficiently solve control matters caused by the thrust uncertainties and the non-synchronization thrust among solid boosters, which takes advantage of increasing the attitude control precision with large scale and improving the launch vehicle’s adaptive capability in the real flying environment.

In order to solve the path planning of unmanned aerial vehicles (UAVs) in three-dimensional space under threats, an improved snake optimization (ISO) algorithm is proposed for the challenging optimization of UAVs path planning in three-dimensional complex environments. Firstly, the total cost function including range, threat, altitude and smoothing costs is established, and the path planning of UAVs is transformed into an optimization issue that takes into account the safe operation requirements of UAV. In the framework of snake optimization (SO), Tent chaotic map is used to initialize the population, and the temperature threshold is dynamically adjusted to improve the local optimization ability of the algorithm, and Lévy flight strategy is introduced in the exploitation phase to improve the global optimization ability of the algorithm. The simulation results show that the ISO algorithm is superior to the SO algorithm by solving the path planning of UAVs.

A guidance law with variable control weight is proposed，which is based on the optimal control method for intercept different types of maneuvering targets by using air-defense missile. A state equation of relative motion between missile and target in the longitudinal plane is set up，and an indicator function based on variable control weight and line of sight angular velocity constraint is designed，furthermore，the optimal guidance law related to weight coefficients is derived by using the minimum numerical principle. Simulation experiments are implemented for different types of maneuvering targets，and the results indicate that the performance of the optimal guidance approach proposed in this paper is superior to traditional proportional guidance method. The line-of-sight angular velocity convergence is fairly quick，which can form preferable reverse orbit situation. Furthermore，regarding targets under different maneuvering modes，the weight of the optimal guidance law can be adjusted to match the optimal control parameters of each maneuvering target.

In order to solve the problem of positioning and reliable recovery of some important structures of aircraft after flight test, a highly reliable positioning system based on Beidou and GPS dual-mode module is designed. The system uses the BDM 910 module integrated with GPS/RNSS (Radio navigation satellite system) module and RDSS (Radio determination satellite system) module as the core positioning device is used in the system, and the BDS (Beidou navigation satellite system) frequency point and GPS frequency point navigation messages are received through dual links and four antennas redundancy. While effective satellite navigation signals are captured, the system own position information is resoved in real time. The location information is sent to the search device in real time in the form of Beidou short message, and the landing point coordinates of the aircraft are finally determined. The problem of positioning and recovery of relevant important structures is solved in the flight test in a complex environment. The positioning system has the ability to analyze the positioning data and the communication information of the positioning device in real time, the data analysis rate is less than 5 s, the positioning accuracy is less than 10 m, the communication success rate is more than 95%, the system is continuously running and lasts more than 4 hours, and the positioning is highly reliable.

The algorithm of the integrated celestial navigation combined accelerometers with X-ray pulsar sensor is researched. Firstly，the inertial navigation algorithm is provided for spacecraft based on the heliocentric inertial reference frame，and the accelerometer drift is estimated by the phase difference of the inertial navigation and X-ray pulsar sensor in pulsar direction based on PI filter. Then，the algorithm of the integrated celestial navigation combined X-ray pulsar sensor with accelerometers based on the time of arrival （TOA） of the pulsar pulse is proved. Lastly，simulation results show that the algorithm of the integrated celestial navigation is effective.

A dual-arm space robot finite time compliant capture strategy for non-cooperative targets is proposed. Firstly, the end-effector dynamic model of generalized coordinates presented by end-effector pose degree of freedom based on the free-floating space manipulator is established. On the basis of this, a global fast non-singular terminal sliding mode controller has been designed to achieve the finite-time trajectory tracking control of the end-effector. Furthermore, a three-fingered end-gripper mechanism has been introduced and a sub-system dynamic model consists of the end-effector and the target has been established. Based on this model, a composite compliant grasping control strategy has been designed, which is based on the compliant control theory to resolve the compliant grasping control of the end-effector towards the target. The simulation results show that the compliant capture of a tumbling target by a dual-arm space manipulator can be achieved by the proposed control strategy.

Based on the upper wind benchmark of the sounding balloon，the upper winds and the precision of corresponding maximum aerodynamic load from the wind profile radar and numerical weather prediction model forecasts from 1st day to 4th day are compared and analyzed. The results show that: the precision of the upper wind from low to high is presented by the wind profile radar and the forecast of the 4th day to the forecast of the 1st day. The precision of the wind profile radar at the altitude of 7.6 km and above is obviously low，and absolute difference is more than 5 m/s； The precision of maximum aerodynamic load from low to high is wind profile radar and the forecast of the 4th day to the forecast of the 1st day. The average absolute differences of the maximum aerodynamic load from wind profile radar and the the forecast of the 1st day to the forecast of the 4th day is respectively showed by 326.72 Pa∙rad，126.53 Pa∙rad，162.26 Pa∙rad，183.15 Pa∙rad and 212.59 Pa∙rad，and the correlation coefficient values are separately recorded by 0.76，0.98，0.96，0.95 and 0.92. Therefore，the precision of the maximum aerodynamic load from the wind profile radar is low，which cannot be used for the safety guarantee of rocket flight and needs to be further improved.

Aiming at the orbit determination of non-cooperative continuous low-thrust maneuvering spacecraft，a rapid pre-identification method of thrust acceleration based on single-arc orbit determination is proposed. Based on the relationship between satellite orbit parameters and acceleration，the single-arc orbit determination results of two radar observations with a certain time interval are used to inversely solve the tangential acceleration of the spacecraft under continuous tangential thrust，and fairly smaller solution error is remaining under the condition of sparse data in a short time by using this method Which is applied respectively to the orbital climb of Starlink，OneWeb and the Qianfan constellation. The results show that when the observation interval is greater than 11 h and 15 h separately，the proposed method can realize the rapid pre-identification of the tangential thrust acceleration of Starlink and OneWeb satellites，and the solving error is less than 2%. The calculation results can be used as initial values for precise orbit determination of continuous low-thrust maneuvering spacecraft.

Regarding the planning of on-orbit refueling tasks for spacecraft in geosynchronous orbit (GEO), an on-orbit refueling strategy based on hierarchical refueling is proposed, which can effectively improve the refueling efficiency compared with the traditional “one-to-many” refueling method. In the proposed hierarchical on-orbit refueling strategy, a sub-service spacecraft (SSc) is defined to jointly accomplish the on-orbit refueling task with the primary service spacecraft (PSSc). On the basis of this, the spacecraft orbital transfer model based on the multi-turn Lambert orbital transfer, the relationship between orbital maneuvering speed increment and transfer time is obtained, and the fuel cost and transfer time of spacecraft orbital transfer are taken as the objective function under consideration of the constraints such as spacecraft load and mission time, and the spacecraft refueling sequence and orbital transfer time are optimally solved by using an improved genetic algorithm. Finally, the effectiveness of the theory is verified through numerical simulation.

A hybrid control method combining feed-forward compensation with closed-loop feedback based on the target relative motion trajectory is proposed in this paper, which is aiming at significantly improving the satellite dynamic pointing tracking control precision. Firstly, the dynamic error coefficient method is adopted to analyze the attitude tracking errors, and power transfer function is established to analyze attitude angular velocity errors caused by sensor noise. Next, a feed-forward compensation control strategy calculating the desired line-of-sight angular velocity and angular acceleration by estimation of the target relative motion is proposed, which resolves the challenge of space-borne detection equipment that can only measure target pointing deviations and doesn’t directly perform hybrid control. Finally, the system stability and attitude tracking performance are validated through mathematical simulations, and ground-based air-bearing experiments are implemented to test the controller's tracking performance. This attitude controller reaches high-precision during dynamic tracking by different input commands without altering the closed-loop feedback bandwidth.