In order to meet the design requirements of small size,light weight and low cost of single soldier rocket,the model of pitch and roll dual-channel individual rocket controlled by two steering engines was established,and roll control by differential,pitch control by deflection in the same direction. The inversion method was used to design the pitch channel controller,and compared with PID control. To compare the proposed method with three-channel control,the yaw control was introduced at the same time. The results show that the decoupling performance of the controller is better than PID control,and the controller has good tracking effect. In the case of aerodynamic parameter perturbation,it can maintain certain robustness and anti-interference ability. The inversion controller can meet the control requirements of dual channel signle solidier rocket. Compared with the three-channel control,the proposed two-way actuator control model has more flexible internal layout,and it is more suitable for small-caliber and low-cost requirements of individual rocket.

Aiming at the poor impact point accuracy and low strike accuracy of traditional mortar shell,the trajectory correction algorithm of mortar shell was designed. The trajectory was corrected by perturbation impact point deviation prediction method,adaptive proportional guidance method and adaptive proportional differential guidance method. The six degree of freedom trajectory model and control model were established,and the perturbation impact point deviation prediction method and proportional guidance method were proposed. In view of the fact that the constant proportional coefficient of proportional guidance law does not conform to the actual trajectory change,the adaptive proportional guidance law was designed in the longitudinal plane,and the adaptive proportional differential guidance law was designed in the transverse plane. Monte-Carlo simulation was used to verify the correction ability of guidance law in longitudinal plane,transverse plane and compound guidance. The simulation results show that the adaptive proportional navigation law was the best in the longitudinal plane,and the adaptive proportional differential navigation law was the best in the transverse plane. The composite guidance effect of three guidance methods was simulated and analyzed. The simulation results show that the composite guidance and correction ability of perturbed impact-point deviation prediction guidance method in the longitudinal plane of ascending arc,adaptive proportional guidance law in the longitudinal plane of descending arc and adaptive proportional differential guidance law in the transverse plane,were effectively improved,and the circular probable error of impact points of mortar was reduced from 126.317 m to 0.965 5 m. Under the conditions of large shooting angle and small range,the circular probable error is 1.864 3 m.

Due to the high revolution of the spin-stabilized projectile with spoiler,there is a deviation between the motion response direction and the direction of the control force,which brings great difficulties to the design of guidance method. Aiming at this problem,a guidance law based on velocity tracking and parameter optimization was proposed in order to develop a guidance method for spin-stabilized projectiles with spoiler. According to the six degree of freedom trajectory model of the spin-stabilized projectile with spoiler,the ballistic characteristics and hit accuracy of the projectile under the guidance law were simulated and analyzed. The results show that the designed guidance law can guide the projectile to hit the target,meet the control accuracy,effectively reduce the impact point dispersion under the interference condition and improve the firing density. The change of the azimuth angle of the spoiler is smooth in the control process,which is easy to be realized.

In order to maintain a better aerodynamic shape of the guided projectile in different flight states,improve the flight efficiency and increase the range,a morphing guided projectile was designed. The aerodynamic parameters calculation,aerodynamic characteristics analysis,ballistic simulation were carried out. Firstly,the aerodynamic layout of the morphing guided projectile was determined according to the design parameters,and the specific shape parameters of the wing,canard and tail fin of projectile were determined through iterative optimization,the control modes and ballistic characteristics of the morphing guided projectile were described. Then the aerodynamic data of the morphing guided projectile was calculated by using engineering algorithms,the relationship between lift coefficient,drag coefficient,static stability and the change of the projectile wing shape was analyzed. Finally,a morphing scheme was developed for the designed morphing guided projectile based on the results of aerodynamic characteristics analysis. Hp-adaptive pseudo-spectrum method was used to optimize the ballistics of the morphing guided projectile and the fixed-profile guided projectile with the goal of maximum range. The results show that the designed morphing guided projectile meets the design parameters and has good aerodynamic characteristics and maneuverability,and the ballistic calculation results show that the range can be increased by 10% to 22.5% by introducing the morphing flight technology.

Aiming at the problem of impact angle constraint and the control input saturation caused by the physical constraint of the actuator,an integrated guidance and control design method with impact angle constraint and input saturation was proposed. An integrated three-channel independent design model of the guidance and control system with impact angle constraint was established,and an extended state observer was used to estimate and compensate the model error and the coupling relationship between channels. On this basis,the terminal sliding mode control and backstepping control were used to design the integrated guidance and control system,and the Nussbaum function and an auxiliary system were introduced to deal with the input saturation problem. The six-degree-of-freedom simulation of the missile was carried out to verify the effectiveness of the proposed integrated guidance and control algorithm. Compared with the existing integrated guidance and control algorithms,the guidance accuracy and angle constraint accuracy are higher,and the guidance and control effect is better.

Aiming at the terminal guidance with impact angle constraint against maneuvering targets,a guidance law with impact angle constraint was proposed. Based on the relative motion model of the missile and target,the impact angle constraint was transformed into the terminal line of sight(LOS)angle constraint. A new nonsingular fast terminal sliding-mode-surface was designed. Combined with the improved super-twisting algorithm,an adaptive nonsingular fast terminal second-order sliding-mode guidance law was proposed. The guidance law can make the LOS angle and LOS angular rate converge in finite-time,and a parameter adaptive-law was designed to effectively compensate the unknown disturbance. The simulation results indicate that by the proposed guidance law,the target can be accurately intercepted with expected impact angle. Compared with the existing guidance laws,the proposed guidance law has faster convergence speed,higher guidance precision and less energy consumption.

In the field of air defense,the speed of the intercepted target is high,and the control time we have is short. The laser terminal guidance loop was built as well as the adjoint system of the loop based on Simulink model,and the variation law of miss distance with terminal guidance time which was caused by measurement error,angular velocity noise and angle constraint of guidance law under the conditions of the approach speed of projectile and target continuously increasing and the laser seeker at different response speeds. The results of simulation show that it can effectively reduce the effect of the increase of approach speed of projectile and target to increase the response speed of the laser seeker. If the Mach number of the projectile approaching the target reaches 4 or more,the seeker bandwidth must be more than 80Hz so that the guidance accuracy will not decrease further with the increase of the approach speed.

In the process of guided bomb hardware-in-loop simulation test,the difference among pitch angle simulation results of the same bomb is larger,and the difference between pitch angle telemetering results and simulation results is also larger by one batch guided bomb,but the convergence of pitch angles is consistent,and the tested miss quantity and the simulated miss quantity are also consistent. Aiming at this problem,the error propagation theory of guided bomb hardware-in-loop simulation system was established based on the covariance analysis describing function technique(CADFT). The simulated attitude angle errors were calculated as well as the errors of attitude angle velocity. The influence caused by simulated attitude angle errors and attitude-velocity errors on simulation results was analyzed. The result shows that the aerodynamic force error and aerodynamic torque error are the main reasons causing the difference among pitch angle simulation results. The dynamic verification method by attitude angle isn’t suit for guided bomb simulation model,but the static verification method by miss quantity is suitable.

To solve the problem of precise guidance of guided missile(GM)under multi-constraint,the optimal sliding mode guidance law with relative distance weight function was applied. The weight function was introduced to the optimal guidance law. The guidance precision can be improved by changing the motion trajectory and motion time of GM. The single weight function is difficult to satisfy the guidance accuracy,the constraint of line-of-sight angle of seeker and overload constraint. The piecewise weighting method of the function was applied to solve the guidance problem caused by weighted optimal guidance. Combined with the sliding mode variable-structure control theory,the segmented weighted optimal sliding-mode guidance law was designed to enhance the anti-interference ability of guidance system. The simulation result shows that the terminal guidance law can solve the problem of accurate guidance under multi-constraint conditions such as overload,seeker line-of-sight angle and falling angle,and it has certain robustness.

Aiming at the problem that multi-hypersonic missiles can realize saturation attack to the target with a specified impact angle,a terminal guidance law with constraints of impact angle and impact time was proposed,and the feasible initial position domain of terminal guidance of missiles was studied. A relative motion model of missiles in the three-dimensional space was established. The vertical and lateral guidance command was used to realize the impact angle control and the impact time control respectively,and a three-dimensional cooperative terminal guidance law was formed. Adopting the law,the constraints of final velocity and available overload of missiles were considered,and the initial position of terminal guidance of missiles was taken as a design variable,and the difference between the actual impact time and the ideal impact time was taken as the performance index function,and an optimization model was established. The model was solved to obtain one-dimensional and two-dimensional feasible initial position domains of terminal guidance of missiles. Simulation results show that the three-dimensional cooperative guidance law can make multiple-missiles hit the target at the designated impact time with the required impact angle,and the position domain solving-method can be used to obtain the feasible initial position domains for each missile under given constraints.

In order to meet the requirements of impact angle constraint,a guidance law considering the second-order dynamics of missile autopilot was designed. A guidance system model with impact angle constraint considering the second-order autopilot dynamics was established. The super-twisting algorithm was used to improve the extended state observer. A super-twisting extended state observer was proposed to estimate the target acceleration. A nonsingular fast terminal-sliding-surface with impact angle constraint was selected. Combined with dynamic surface control,a new guidance law was proposed. The guidance law can achieve global finite-time convergence of the system states and compensate the autopilot dynamics. The comparison of simulation results shows that the proposed observer has high estimation accuracy. The proposed guidance law can achieve finite-time convergence of line-of-sight angular rate and impact angle,and it has better guidance performance.

The irrelevant parameters of proportional navigation(PN)for spinning missiles may result in a divergent coning motion due to the cross-coupling between the pitch and yaw induced by spinning,which can cause large miss distance. Aiming at homing spinning missile guided by one pair of control fins,the guidance equations,control equations and dynamics equations were derived,which included cross-coupling elements. The convergence condition of angular rate of line of sight(LOS)was analyzed,as well as the optimal navigation ratio and the effect on miss distance. The phase-delay of the seeker leads to a larger navigation ratio for angular rate of LOS convergence,and the optimal navigation ratio increases at the same time. The missiles fly in form of coning motion with the divergent angular rate of LOS,and a large miss distance generates. Numerical simulation of a case shows that the miss distance increases exponentially with the increase of the spinning rate. The upper limit of spinning rate and navigation ratio should be designed in justified range.

To solve the problem of impact time and impact angle control under varying missile velocity,a novel impact time and impact angle control guidance law was proposed for the impact time and impact angle control under varying missile velocity by using shaping theory and nonsingular terminal sliding-mode theory. It was proved that the guidance law can satisfy the Lyapunov stability criterion. Based on the relative motion model of missile-target,the guidance problem under varying missile velocity was transformed into the guidance problem under constant missile velocity. Through the shaping theory,the line-of-sight(LOS)angle polynomial was constructed,and the coefficients were calculated by numerical method,and the ideal line-of-sight angle satisfying the impact time and impact angle constraints was obtained. Based on the nonsingular terminal sliding-mode theory,the lateral acceleration of the missile was designed so that the actual LOS angle of the missile can change according to the ideal LOS angle to satisfy impact time and impact angle control. The numerical simulation results under different conditions validate the effectiveness of the proposed guidance law.

Aiming at the characteristics of high speed and high maneuvering of near space hypersonic-target,the midcourse guidance off-line trajectory optimization based on grid division was combined with neural network,and an on-line trajectory generation method for midcourse guidance was proposed to improve the midcourse-guidance performance of interceptor missile. Taking the grid points obtained by meshing the intercept area as constrain,the trajectory data base covering intercept area was constructed by means of off-line trajectory optimization. Based on the property of BP neural network approximating any nonlinear functions,the approximation of trajectory data base was achieved,and the BP neural network model was constructed. The on-line generation of midcourse-guidance trajectory was realized. Simulation results show that in the process of midcourse guidance,the proposed generation method of midcourse-guidance trajectory has good performance with and without adjusting terminal constrain.

In order to study the flow field and aerodynamic characteristics of the canard ramjet extended-range guided projectile,according to the aerodynamic characteristics of the projectile under the ramjet state and passive flight state,the 3D flow field simulation and numerical calculation analysis were carried out by using the sliding meshing technique and Realizable k-ε model respectively. The flow field and aerodynamic characteristics of projectile at different Mach numbers were studied. The results show that under supersonic conditions,both the drag coefficient and the lift coefficient decrease with the increase of Mach number at the same angle of attack. Under the same working condition,compared with the canard layout guided projectile with the same shape parameters but without ramjet structure(reference projectile),the drag coefficient is about 50.5% higher and the lift coefficient is about 35.7% smaller in the ramjet state,42.9% higher and 11.9% smaller in the passive flight state. The forward thrust form of the central cone assembly in the passive flight state is beneficial to reduce the drag. The research results provide a theoretical basis and reference for the aerdynamic shape design and performance analysis of canard ramjet extended-range guided projectile.

To meet the requirement of adaptability of the ascent guidance of combined-cycle vehicle under multi-constraints,a closed-loop guidance method based on receding horizon pseudo-spectral optimization was studied. The guidance was carried out under the framework of the hierarchy-structure receding horizon optimization,which applied segmented pseudo-spectral method to generate the online real-time command,and the command can be updated iteratively with the state feedback. The mesh update strategy was introduced to improve the efficiency of the algorithm,and the guidance cycle was adjusted adaptively based on real-time state deviation. Simulation results show that the proposed guidance method is feasible,and it can effectively suppress the influence of model parameter uncertainty and external interference. The guidance accuracy meets the mission requirements. The calculation time is in the order of hundred milliseconds. Thus,the method is of certain value for engineering application.

Compared with single aircraft detection system,aircraft collaborative detection has the advantages of higher accuracy and enhanced ability of anti-jamming. The traditional guidance control design process is usually decoupled into the target position tracking and guidance algorithm. The guidance algorithm,which requires real-time estimated target position as input parameters,can also affect the establishment of state and measurement equations and observation conditions of target tracking guidance algorithm. The decoupling design may cause information loss and reduce the estimation and guidance accuracy. To solve this problem,an integrated design method of collaborative detection and guidance based on unscented transformation(UT)was proposed in this paper. On the one hand,the collaborative detection target positioning algorithm was introduced into the guidance law,and the mean and variance of the fused target position estimation were used as the input of the guidance algorithm. At the same time,the guidance process was regarded as a nonlinear propagation process of a random vector,and the mean and variance of flight control angle were obtained by UT. Compared with the traditional velocity direction control filtering algorithm,it was verified that this method can not only improve the detection accuracy,but also meet the requirements of lower overload,and better control the velocity direction and attitude of the aircraft.

In order to improve the ground-seeking accuracy of guided projectile(GP),a new ground-seeking method based on pressure sensor was proposed by using pressure sensor to measure the high-spin projectile body. The ground-seeking principle of GP was analyzed theoretically. The test scheme and its circuit were designed to improve measurement precision. The accurate measurement of ground-seeking and rotational rate can be achieved through flight-test verification. The experimental results show that the proposed method can be used for GP ground-seeking and rotational rate measurement,and it has high sensitivity and high precision,and the effect of external factors is small. The circuit has some advantages such as simple structure,small volume and good performance. The method can be widely used for attitude control and guidance system of rotating projectile.

Aiming at the limit of three-point guidance applied in 6-DOF trajectory simulation,a sliding mode control(SMC)guidance law was proposed based on three-point intercept geometry. By using direction cosine matrix(DCM),the three dimensional intercept geometry of missile-target was modelled in the ground coordinate frame. Taking target maneuver as the model disturbance,the state-space model of three-point intercept geometry was given. An appropriate sliding surface was chosen,and its parameters were determined by pole configuration method,and the guidance law was derived based on SMC theory. Based on Lyapunov stability theory,the stability of the proposed guidance law was proved. The numerical simulation results show that,the origin point together with the position of missile and target are always on the same line in the ground coordinate frame,which satisfies the three-point intercept geometry. The guidance command is relatively smooth during the whole flying process,and it meets the actual requirements.

To ensure the firing accuracy and warhead density of guided multi-warhead,a numerical navigation and relative guidance method was proposed. Based on the zero apparent acceleration theory,the numerical navigation method was adopted in unpowered phase,which effectively eliminated the influence of drift error of accelerometer on the navigation accuracy and reduced the drop-point deviation of warhead. The relative guidance method was also studied based on perturbation guidance. The attitude control equation and guidance cutoff equation were designed. The state parameters of the separation moment of the main warhead and the subordinate warhead were used as the standard shutdown amount,and the shutdown amount of the subordinate warhead was calculated in real time and compared with the standard shutdown amount,and the shutdown was implemented when the relative distance condition was met. Simulation shows that,through the numerical navigation and relative guidance,the guided multi-warhead can effectively achieve the formation flight and hit the target with a high precision. The subordinate warhead shielding the main warhead has a good attack effect.

In order to improve the effectiveness of cooperative air-combat,the method of selecting cooperative guidance platform was given,and the guidance handover scheme was designed. The maximum loss of the enemy and the minimum loss of our side was taken as the objective function,and the air combat model of cooperative guidance was proposed. The stepwise-estimation immune algorithm was proposed to improve the solving efficiency. The crossover and mutation based on marginal product model were introduced into the immune algorithm to prevent the disrupt of the superior model,and the fitness sharing was incorporated in the new algorithm to enhance the diversity of population. The stepwise-estimation immune algorithm was used to solve the proposed air-combat model. Simulation results show that the combat effectiveness obviously improves by cooperative guidance,and the stepwise-estimation immune algorithm is superior to others.

In order to improve the precision of tracking low altitude target,the line-of-sight(LOS)angle is needed to be restricted near the Brewster angle to reduce the influence of multipath interference on the detection accuracy of the radar seeker. Based on the principle of global sliding mode control,a global sliding-mode guidance law was designed. This guidance law can make the LOS angle restricted near the Brewster angle in the process of tracking and intercepting the low altitude target. The guidance law can eliminate the stage of approaching the sliding surface,which makes the system equipped with robustness in the whole process. Aiming at the possible failure of the missile autopilot executing guidance command,a virtual command was designed to compensate the failure and disturbance. Combining with the designed virtual command,a composite global sliding mode guidance law was designed. The simulation results show that the composite guidance law can greately improve the anti-interference capability and maintain the stability of the system.

In order to achieve impact time control with field-of-view constraint,the problem of impact time control guidance with field-of-view constraint was studied based on relative motion model of missile-target. The model was established according to the criterion of the impact time error and leading angle of missle converging to zero. Based on the proportional navigation,a nonsingular impact-time-control guidance law with field-of-view constraint was designed by using the form of additional bias control. The guidance law has certain optimality. The impact time error converges to zero before the leading angle converges to zero,and the guidance commande finally converges to zero. The stability of impact-time-control guidance law was proved by using Lyapunov stability theory. Simulation results show that the guidance law can effectively achieve the impact time control under the constraint of field-of-view.

A bias proportional guidance law with constrained impact angle was proposed,which was used to attack moving target. According to the relative motion of missile and target,the moving target was considered as a stationary virtual target. The actual guidance instruction can be transformed into virtual guidance instruction,namely the design problem of the relative flight path angle rate. Based on the relative relation,the instantaneous variation and small angle hypothesis was used to deduce the bias instruction item. The relative flight-trajectory-angle rate is proportional to the LOS angle and contains an additional bias item,and the constraint control for the actual trajectory angle can be indirectly realized by controlling relative flight-trajectory-angle. Due to the time-to-go required for the guidance scheme,a time-to-go estimation method for moving targets was derived considering proportional and bias terms. The simulation results show that the moving target can be omnidirectionally striked by the guidance law,and the guidance law has good performance.

To intercept hypersonic aircraft effectively,a 3-dimensional differential game-based guidance law using predictive control was proposed. The terminal interception for high-speed maneuvering target was formulated as a zero-sum differential game with line-of-sight angle rate and fuel consumption as performance index. Under the framework of nonlinear model predictive-control,the decision-constrained finite-horizon differential game was solved in real time by gradient descent algorithm,and the optimal action was obtained. The guidance law enables accurately intercepting target in compliance with highly dimensional nonlinear dynamic systems,maneuverability constraints and fuel consumption limits. With fuel consumption similar to that of proportional guidance law,the proposed guidance law has the advantages of straighter trajectory and higher guidance accuracy.

In order to further improve the maneuverability of rockets with small length-diameter ratio while ensuring the static stability and balancing the contradiction between the static stability and the maneuverability,the anti-stable wings were installed on the basis of the original projectile model,and the numerical simulation of the two models was carried out,and the influence of anti-stable wings on the flow field and aerodynamic characteristics of whole projectile was analyzed. The results show that the normal force coefficient of the whole projectile is increased by 2%-3%,and the static stability is decreased,and the pressure center is close to the centroid. The normal overload is increased due to the normal force of whole projectile. After installing the anti-stable wings,the normal force of the warhead is increased,and the upper tail rudder is affected seriously by the down wash flow within a certain angle of attack,and the normal force efficiency of the upper tail rudder is reduced. The tail lift is reduced,and the static stability is reduced. The maneuverability of the whole projectile is increased significantly when the two parts work at the same time.

Smoke, as the most common product of combat in modern warfare, reduces the visibility of combat scenarios inevitably, which in turn affects the performance of downstream military intelligence systems. Therefore, it is very important to restore the smoke-containing images. Existing algorithms usually ignore both the high-level semantic information in the image，and the degraded image itself can provide valuable supervision information for improving the smoke removal ability of network. Accordingly, a semantic-guidance and contrastive learning-based generative adversarial network (SCLGAN) is proposed to remove smoke from battlefield images. Specifically, semantic information is regarded as guidance to help the network better recover the structural and color information of images incorporating the high-level semantic features in low-level visual tasks. The contrastive learning paradigm is used to adopt clear image and smoke-containing image as positive and negative samples, and the contrastive regularization ensures that the restored image is pulled in closer to the clear image and pushed far away from the smoke-containing image. In addition, a smoke-containing battlefield dataset is first constructed to simulate the real smoke-containing battlefield scene, which promotes the development of related research. Experiments demonstrate that, compared with the existing smoke removal algorithms, the proposed algorithm can surpass the previous state-of-the-art methods in both quantitative and qualitative assessment.

Aiming at the problem that clouds have a great attenuation effect on laser signal and affect the performance of terminal guidance in cloudy weather, a trajectory programming method of gliding extended range laser-guided projectile considering the attenuation effect of clouds on laser is proposed. A cloud approximation model is established based on the reasonable assumption, and an optimal compound evasion objective function (CEOF) of cloud evasion-control energy is designed. A trajectory programming model is established according to the working characteristics of laser seeker and the flight characteristics of guided projectile,and the trajectory programming problem is transformed into a nonlinear programming problem (NLP) by using Radaupseudospectral method. Finally, the NLP solver SNOPT is used to solve the NLP. Several groups of cloud examples are selected for simulation analysis, and the simulated results are used to verify the effectiveness of the proposed trajectory programming method. Compared with the traditional trajectory programming methods, the simulated results show that the trajectory programming method proposed in this paper reduces the attenuation of clouds on laser greatly at the cost of increasing less kinetic energy loss, and controlling energy consumption and attack time, which verifies its superiority.

Weapon-target assignment (WTA) and relay guidance of air-to-ground missiles are difficult problems to be solved urgently in long-range air-to-ground multi-target attack, and are characterized by complex parameters, multiple constraints, and strong nonlinearity. A multi-target and multi-constraint optimization model of WTA and guidance sequence is established, in which the objective includes the minimum comprehensive survival probability of targets and the minimum number of used missiles, and the constraints involve missile configuration of the attack aircraft, damage capability of the missiles, damage requirements of the targets, performance of the guidance stations, etc. A double sequence coding multi-population non-dominated sorting genetic algorithm Ⅱ (DSMPNSGA-Ⅱ) is proposed by improving NSGA-Ⅱ, which optimizes the scheme of WTA and every missile’s guidance sequence through optimizing the WTA sequence and guidance stations sequence. In DSMPNSGA-Ⅱ, the depth-first search Dijkstra (DFS-DJ) algorithm is used to search for missiles’ guidance sequences and improve crossover and mutation operations so as to reduce the production of infeasible solutions, and the multi-population strategy is introduced to improve the performance of DSMPNSGA-Ⅱ. The simulation results show that the DSMPNSGA-Ⅱ can obtain effective schemes of WTA and air-to-ground missile relay guidance, and that its solution quality is better than that of the single-population NSGA-Ⅱ and the multi-objective discrete particle swarm optimization (MODPSO) algorithm.

To solve the problem of missile impact time control, a two-stage impact time control guidance law is designed based on proportional navigation guidance and the data-driven method, which is independent of time-to-go information. The first stage is the impact time control stage. The relationship data set between the flight path angle and the flight state of the missile is constructed by simulation under the framework of proportional navigation guidance method, and then the corresponding mapping network model is trained offline by using the neural network method. According to the mapping network, the ideal flight path angle corresponding to the impact time can be calculated in real time during missile flight, and the guidance command will control the actual flight path angle converging to the ideal one. In the second stage, the proportional navigation guidance law is directly applied, and the missile impact time control is finally realized. The simulation results under different conditions verify the feasibility and effectiveness of the proposed guidance law. Compared with the existing similar guidance laws, the proposed one can control the impact time with higher precision and less control energy. In addition, the theoretical analysis shows that the guidance law can be extended to impact angle control by changing the mapping network.

A 3-dimensional (3D) differential game guidance law based on adaptive dynamic programming (ADP) is proposed to solve the attack and defense problem in hypersonic interceptors attacking hypersonic maneuvering targets. Firstly, the 3D guidance problem in hypersonic attack and defense game is decomposed into the longitudinal plane and lateral plane. Then, the problem model of a hypersonic two-person zero-sum differential game is established in the longitudinal plane. A non-quadratic cost function is designed to satisfy the control constraints. The optimal differential game strategy of the hypersonic interceptor and the target is obtained by the Hamilton-Jacobi-Isaacs (HJI) equation. Since the HJI equation can not be solved analytically, the ADP method is used and a critic neural network is introduced to approximate the optimal cost function to obtain the bank angle amplitudes of both sides. Finally, the lateral guidance is realized by establishing the bank angle reversal logic in the lateral plane by the parallel approach. The simulation results show that the guidance law can achieve precision strike by a hypersonic interceptor under the condition of a maneuvering hypersonic target avoiding interception within the control constraints of both sides of the attack and defense game.

Cooperative strike is an important mode to improve the overall combat effectiveness of aircraft clusters. To address the problem of mutual collision or interference in the process of cooperative attack on targets by dense aircraft clusters, a typical mission scenario model of multi-target cooperative attack is established, and the maneuver law in the process of cooperative movement of aircraft cluster is revealed. Then, based on the optimal theory of utility function, the coordinated optimal maneuver strategy of aircraft cluster to avoid collision is studied, ensuring collision avoidance between aircrafts. Considering the time/space synchronous arrival constraint in aircraft cluster cooperative guidance, a multi-constraint cooperative guidance strategy based on time control is adopted, which meets the requirements of arrival time and miss distance at the same time. On this basis, a cooperative collision avoidance mechanism based on conflict triggering is introduced, and a cooperative adjustment strategy that considers collision avoidance and cooperative attack is proposed. And a cooperative guidance method satisfying collision avoidance and spatio-temporal synchronization is obtained. The simulation results show that the proposed aircraft cluster cooperative guidance method considering collision avoidance constraints ensures safe distance constraints between aircrafts and has strong aircraft cluster cooperative time control ability and high guidance accuracy. It requires less iterative calculation, reduces the requirements for airborne processors, and provides a solution to the collision avoidance problem in aircraft cluster cooperative attack. This method has strong engineering application value.

To address the problem of coning motion instability of spinning missiles induced by the introduction of the strapdown seeker and attitude pursuit guidance law, the dynamic model of the strapdown seeker in the non-spinning missile coordinate system is derived, and the mathematical model of the attitude pursuit guidance and control system in complex form is established. The response delay of the seeker and the gyro scale-factor error are considered under different spinning rates and damping loop gains. The stability of the spinning missile’s control and guidance system is analyzed, and the range of the corresponding characteristic parameters are solved by numerical methods. The results show that the larger the delay angle of the seeker is, the smaller the upper limit of the guidance loop gain that can stabilize the system is. When the gyro scale-factor error coefficient is greater than 1, the upper limit of the guidance loop gain become larger; when the scale-factor error coefficient is less than 1, the upper limit becomes smaller.

To improve the engineering applicability of aircraft when pursuing a target in the presence of dynamic obstacles, the relative motion relationship between the aircraft and obstacles, as well as between the aircraft and the target are analyzed. Leveraging the control barrier function (CBF), the velocity obstacle approach is used to constrain the velocity of the aircraft, and a quadratic programming (QP) problem is constructed with the existing proportional guidance law. Then, a closed-form obstacle avoidance guidance law is proposed by solving the QP problem. Numerical simulation and analysis of the trajectory properties and overload show that the proposed obstacle avoidance and guidance law can avoid a moving obstacle in real time and pursue the target with favorable global properties. Compared to existing methods, the overload curve is smoother, and the saturation overload ratio is smaller.

To deal with the large truncation errors and defects of being unable to give quick and accurate solutions in calculating the firing data of laser terminal guidance projectiles by the firing table, a method based on multi-model network is proposed. Firstly, a six-degree-of-freedom trajectory model of the guided projectile is established, and the influence of the rear sight on the range in the initial conditions is analyzed. Secondly, considering the influence of air temperature, air pressure, wind interference and other factors on the solution, the least square method is used to fit the air temperature and pressure, and the layer weight method is adopted to obtain the accurate ballistic wind as part of the input of the network model. Thirdly, samples are generated in batches through simulations, which are applied to train each model in the deep neural network. Finally, the proposed method is verified by simulations and flight tests. The simulation results show that the proposed method has a higher precision and an error rate of less than 0.2% compared with the traditional firing table.

Aiming at the problem of three-dimensional multiple missiles intercepting a maneuvering target simultaneously, a three-dimensional distributed cooperative guidance law considering time and space constraints. The proposed guidance law is based on fixed-time stable theory and adaptive sliding mode theory. Along the light-of-sight (LOS) direction, an improved robust adaptive cooperative guidance law is presented based on fixed-time consensus, which can drive a group of missiles to attack maneuvering target at a desired impact time without inherent chattering of sliding mode control. Along the direction vertical to the line-of-sight, a robust adaptive guidance law is proposed based on nonsingular fixed-time terminal sliding mode technique, which can drive missiles attack target with desired impact angle and avoid collision between missiles. Through the Lyapunov stability theory, the fixed-time stability of closed-loop system is demonstrated. The simulation results show that the proposed approach can be implemented to intercept maneuvering targets with desired impact time and impact angle.

In the line structured light three-dimensional measurement system, the high-precision laser stripe center-line extraction is the key to improve the measurement accuracy. Aiming at the existing laser center-line problems such as low extraction accuracy and poor retention of details, a laser center-line extraction algorithm based on normal guidance was proposed. The specific implementation steps of the algorithm were as follows: Firstly, the image was pre-processed, and the laser line was preliminarily extracted by combining edge detection and geometric center method. Then, the principal component analysis (PCA) was used to obtain its normal line, divided the angle eight neighborhood at the laser center point, and searched for the effective point set through the normal angle guidance. Finally, the gray centroid method was used to extract the sub-pixels from the point set. The experimental results show that the root mean square error of the algorithm is improved by 0.233 9 pixel compared with that of the gray centroid method, which can retain the light strip details better than the Steger algorithm and the directional template method, and can extract the light strip center more accurately and achieve the sub-pixel level accuracy.

Eddy current asymmetry and transonic aerodynamic transition usually occure at a large angle-of-attack during the cross-domain flight process of guided missile, which may cause control instability and further increase miss distance. To solve the problems above, a missile roll channel dyanamics model is established, in which the strong nonlinearity of aerodynamic model, strong parameter uncertainty and strong external disturbance, and then an “observer+controller” composite control framework is proposed. This roll stabilization control framework contains a novel sliding mode disturbance observer and nonsingular terminal sliding mode. Based on this framework, a backstepping controller is proposed to compensate actuator dynamics. The stability and finite-time convergence properties of closed-loop system are verified through Lyapunov theory. Finally，the numerically simulated results demonstrate the superiority and universality of the proposed control method.

To realize the precise tracking and safe navigation control of the unmanned surface vehicle (USV) in the complex ocean environment, firstly, the line-of-sight (LOS) guidance strategy used in USV trajectory tracking control is studied in detail, and the algorithms are divided into two categories: LOS guidance strategy based on look-ahead distance and on encircling circle. Then, the mathematical model of the USV is built, and the USV trajectory tracking control algorithm is designed based on the LOS guidance strategy using the look-ahead distance and the PID control idea. Finally, the simulation results show that the proposed algorithm has good trajectory tracking control effects, which can provide insights into research on the autonomous navigation control technology and the construction of the USV equipment in China.

A distributed cooperative guidance law is proposed to solve the problem of multiple aircrafts cooperatively attacking a stationary target with field-of-view constraints. Firstly, the nonlinear motion model is transformed into a second-order agent model through the motion equations for engagement in the velocity coordinate system, and the field-of-view constraints are transformed into agent velocity constraints. Combined with the velocity constraint consistency protocol, an arrival time cooperative guidance law is designed to realize the cooperative attack of multiple aircrafts with field-of-view constraints. The proposed guidance law overcomes the problem of numerical singularities in the traditional cooperative guidance law, and there is no need to switch the guidance law in the cooperative process. Finally, through numerical simulation and comparative study, the effectiveness of the proposed guidance law is verified, which still has stable guidance performance in the presence of measurement noise and communication delay.