55 results about "Aircraft dynamics" patented technology

The invention discloses a low-altitude traffic flow collaborative awareness and conflict prediction method and system based on a multi-modal large model. The method comprises the following steps: collecting multi-source heterogeneous data of an aircraft, fusing and mapping the multi-source heterogeneous data to a space-time grid with dynamically adjustable granularity based on a real-time traffic situation, and realizing adaptive characterization; the method comprises the following steps of: constructing a robust special prediction model by introducing aircraft dynamics constraint and airspace rule knowledge as priori and performing fine adjustment by adopting a loss function containing an adversarial regularization item based on a pre-trained large model of the body-equipped agent; inputting spatial-temporal characteristics into the model, outputting behavior intentions and probability trajectories in an end-to-end manner, constructing an incomplete information game deduction model based on predicted trajectories, and calculating a comprehensive collision risk index fusing trajectory uncertainty, a motion state and a game deduction result; and generating a hierarchical early warning and collaborative avoidance strategy, and performing network distribution. According to the method, the accuracy, the robustness and the decision intelligence of conflict prediction under the low-altitude dense traffic flow are remarkably improved.

The invention provides a civil aircraft safety perception method and system based on virtual flight, and relates to the technical field of aircraft safety perception. A dynamic and hierarchical three-dimensional wind shear field model is constructed through space-time alignment and fusion of multi-source meteorological data; deep coupling of wind disturbance and aircraft dynamic response is realized based on a small disturbance theory and a flight dynamics model; continuous and dynamic quantification of risks is realized by combining and analyzing flight parameters and a safety threshold value dynamically adjusted based on wind disturbance intensity and utilizing multi-parameter risk level weighted integration; according to the method, the risk coefficient is coupled with the three-dimensional terrain and the virtual flight trajectory, hierarchical color mapping is adopted to carry out visual space visualization, and a complete closed loop from environment modeling to risk early warning is formed, so that the identification accuracy and timeliness of the flight risk in the wind shear environment are greatly improved, and the scientificity and practicability of security situation awareness are enhanced.

The invention provides an airport flying area crossing traffic command system based on an A-SMGCS system, and belongs to the technical field of intelligent traffic, and the system comprises a sensor layer which is used for collecting aircraft dynamic data, to-be-passed authorized object data, environment meteorological data, and dynamic obstacle data; the data processing and decision-making layer is used for carrying out standardization processing and space-time mapping processing on the collected data, judging existing potential conflicts in combination with a dynamic threshold adjustment mechanism, and generating an issuing instruction based on a multi-target optimization decision; and the execution layer is used for controlling crossing traffic management and control equipment according to the issued instruction and feeding back an execution result to the data processing and decision-making layer in real time to form closed-loop control. The traffic conflict risk of the airport flight area crossing is effectively reduced, the crossing passing efficiency is improved, the vehicle waiting time and the aircraft taxiing delay are reduced, the fine scheduling requirement for scene movement is met, and reliable support is provided for safe and efficient traffic operation of the airport flight area.

This invention discloses an airborne display and control method for enhancing situational awareness and flight operations, comprising: providing onboard operational information and basic traffic, weather, and terrain displays based on flight phases and crew operational requirements, configuring display content and layout, and establishing basic situational awareness; providing a list of traffic aircraft dynamic information, setting aircraft of interest, and acquiring operational information of aircraft of interest to establish enhanced traffic situational awareness; receiving conflict detection, conflict prediction, and conflict resolution information at short, medium, and long ranges, and outputting alarm information in visual, auditory, or tactile forms according to alarm priority to establish enhanced conflict alarm situational awareness; applying for and receiving control clearance for enhanced flight operations through CPDLC, calling the flight operation setup program, and inputting the information required for enhanced flight operations; and generating target flight paths and target flight states in real time based on enhanced flight operation requirements, and implementing enhanced flight operations by the crew or flight management system.

The application provides a kind of aircraft servo elastic dynamics model correction method considering operating surface gap, first aircraft structure is divided into substructure, and the aircraft dynamics model containing aileron gap nonlinearity is established according to substructure branch modal;And additional inertia force caused by aircraft operating aileron deflection is introduced in dynamics model, and the aircraft open-loop servo elastic motion equation containing aileron gap nonlinearity under independent generalized coordinate is obtained;According to sensor output equation, rudder transfer function and flight control system equation, the aircraft nonlinear servo elastic closed-loop system model containing aileron gap nonlinearity is established, and the model is corrected based on ground vibration test, rudder frequency response test and gap measurement test result, and the stability of the model is judged based on simulation analysis result and open-loop frequency response test result and is further corrected, which improves the accuracy of gap nonlinearity aircraft aerodynamic servo elastic characteristic evaluation, and is helpful to guide aircraft servo elastic ground test engineering practice.

The invention provides an aircraft dynamic derivative test device in a strong interference environment, which relates to the technical field of aircraft aerodynamic characteristic tests and comprises an interference flow field establishing device, a dynamic derivative measuring device and a synchronous control module. The disturbing flow field establishing device is composed of a shock wave generator, a supporting connecting piece and a wind tunnel attack angle mechanism, the shock wave generator is connected with the wind tunnel attack angle mechanism through the supporting connecting piece, + / -15-degree pitching and + / -10-degree yawing motion is achieved based on the wind tunnel attack angle mechanism, and multiple types of strong disturbing flow fields can be established by replacing the shock wave generator; the dynamic derivative measuring device is installed in a plug-in mechanism of a wind tunnel, based on a forced vibration test method, a model is driven by a high-precision servo motor to vibrate slightly, precise acquisition is realized by matching with a high-frequency dynamic balance and a data filtering unit, and the displacement of the model relative to a shock wave generator can be adjusted through the plug-in mechanism. According to the scheme, a real strong interference environment can be accurately reproduced, and the dynamic derivative measurement accuracy is improved.

The invention relates to the technical field of high-speed aircraft health management, and particularly discloses an interpretable aircraft state sensing and predicting system and method. The method comprises the following steps: constructing and training a lightweight anomaly diagnosis network with an expansion perception layer; aircraft dynamics, kinematics and fault mechanism formulas are used as soft constraints to be embedded into the future state estimation network; an SHAP interpretation framework is used for sorting the importance of the input features, and a network structure is optimized accordingly; and finally, accurate diagnosis of abnormal types and positions and rapid prediction of aircraft states at a plurality of moments in the future are realized. According to the method, by introducing three innovation points of an expansion perception layer, physical information embedding and feature importance sorting, the diagnosis accuracy and the model interpretability are remarkably improved, and the high-performance calculation requirement in a scene with limited calculation power is met.

The application provides a kind of aircraft dynamics modeling method containing multi-degree-of-freedom actuator, the modeling method returns general mechanics and advanced theory and thought of mechanics basic disciplines, aircraft is regarded as multi-rigid body system, based on virtual power principle and multi-body system topological configuration recursive modeling method, the deformation multi-body system dynamics model suitable for aircraft containing multi-actuator is established, and the motion characteristics of multi-degree-of-freedom, strong coupling between the motion characteristics of aircraft external body attitude motion and internal component relative motion are accurately described.

The invention relates to the technical field of aircraft power control, and discloses an artificial intelligence optimized vertical fixed wing power control method and system, and the method comprises the steps: obtaining the original airspeed data of an airspeed sensor, the ground speed data of a global positioning system and the acceleration data of an inertial measurement unit, and generating the ground speed of an aircraft based on the data fusion. Under the low-wind-speed condition, the ground speed of the aircraft is used as the reference airspeed, the difference between the original airspeed data of the airspeed sensor and the reference airspeed is continuously compared, and the sliding average value of the difference is calculated. When the sliding average value continuously exceeds the preset tolerance range, the method can identify that the airspeed sensor has continuous calibration drift. And then, according to the identified continuous calibration drift, performing real-time compensation on original airspeed data of an airspeed sensor to obtain calibrated airspeed data, and inputting the calibrated airspeed data into flight mode switching judgment logic to trigger an action instruction of a tilting mechanism.

The invention belongs to the field of design of systems for preventing out-of-control accidents of aircrafts, and particularly relates to an integrated platform and method for preventing out-of-control accidents of aircrafts. The comprehensive platform comprises an airborne modeling and simulation unit used for describing changes of aircraft dynamics and control characteristics caused by aircraft damage and external threats; the aircraft health management unit is used for monitoring and predicting the state of the aircraft and continuously evaluating and managing the health condition and the remaining life of the aircraft; the elastic guidance and control unit is used for aircraft out-of-control prevention and recovery and external threat suppression; and the man-machine interface unit is used for realizing man-machine information exchange and coordination under the condition that the airplane is out of control. The aircraft dynamics and control characteristics can be effectively represented under the condition that the aircraft is abnormal, aircraft health and flight safety are evaluated in real time, and effective context awareness and decision support are provided for flight crew, so that the aircraft out-of-control accident risk is reduced, and related accidents are prevented.

The embodiment of the invention provides an attitude control method and system of an aircraft, the aircraft, a medium and a program product, and relates to the technical field of flight control. The method comprises the following steps: acquiring an actual state observed quantity of an aircraft; estimating a disturbance estimation value of the aircraft dynamics model; wherein the estimation error converges within a preset observation time; generating an expected angular velocity instruction by using a preset time attitude controller; wherein the attitude tracking error converges within a preset first time; generating an expected torque instruction by using a preset time angular velocity controller; wherein the angular velocity tracking error converges within a preset second time; and solving the distribution problem of the expected torque instruction to obtain control instructions of a plurality of execution mechanisms so as to perform attitude control on the aircraft. According to the embodiment of the invention, by setting the related convergence time, the control model can realize convergence of the related state error within the preset time, so that the accuracy and the stability of aircraft attitude control are effectively improved.

The invention provides an active control aeroelastic modeling method and device for a flap of a half-unfolded tilt-rotor aircraft, and the method comprises the steps: integrating a flap model into a cantilever wing model through a trailing edge flap interface and flap structure parameters reserved at a rotor structure assembly through the dynamic comprehensive analysis software of the rotor aircraft, establishing a cantilever wing model with flap active control; establishing a tilt rotor model; the cantilever wing model with the flap active control function and the tilt rotor model are combined together, and a flap active control model of the half-unfolded tilt rotor aircraft is obtained; carrying out aeroelastic stability initialization calculation on the flap active control model of the half-spread tilt-rotor aircraft, and extracting geometric shape data of a flap active control rotor blade of the half-spread tilt-rotor aircraft, so that the flap active control model can be integrated on a wing of a tilt rotor; the semi-expanded tilt rotor active control scaling model is created, and theoretical support is provided for follow-up tilt rotor flap active control aeroelastic design.

The invention discloses a dynamic model of a fixed-wing aircraft, and relates to the technical field of flight training. Comprising an aircraft mass center kinetic equation of the fixed wing aircraft, a mass center rotation kinetic equation, an aircraft aerodynamic characteristic calculation model, an aircraft kinetic equation resolving model, an abnormal flight state trajectory and attitude inverse solution calculation model and a ski-jump take-off and carrier landing model. Wherein the aircraft aerodynamic characteristic calculation model is used for calculating aerodynamic force and aerodynamic moment of the fixed-wing aircraft; the aircraft kinetic equation resolving model is used for resolving the aircraft mass center kinetic equation and the mass center rotation kinetic equation; the abnormal flight state trajectory and attitude inverse solution calculation model is used for inverse solution calculation of stress and torque of the fixed wing aircraft in an abnormal flight state; and the ski-jump takeoff and carrier landing model is used for solving the carrier landing arresting longitudinal load problem of the fixed-wing aircraft. The technical problems that in the prior art, the flight training process is abstract in theory and large in difficulty are solved.

This invention discloses a method, apparatus, device, and readable storage medium for determining controller parameters, applied in the field of flight controller technology. The method includes: acquiring an ideal model based on flight quality design, a nonlinear aircraft control model, and an aircraft dynamics model; determining the initial controller parameters of the nonlinear aircraft control model; and using a global optimization algorithm to determine the optimal values ​​of the controller parameters for the nonlinear aircraft control model. Furthermore, it enables real-time flight quality evaluation after optimization, ensuring that the optimal controller parameter values ​​meet Level 1 flight quality requirements. Compared to current manual adjustment of controller parameters, this invention utilizes a frequency-domain-dependent ideal model based on flight quality, and optimization objectives based on the time and complex frequency domains, allowing simultaneous optimization of controller parameters in both the time and frequency domains. Moreover, it intelligently determines the optimal controller parameter values ​​based on a global optimization algorithm, improving the comprehensiveness and intelligence of the optimization process.

The invention belongs to a flight control system test technology, and provides a semi-physical simulation and test method and system for a redundancy flight control system, and the method comprises the steps: constructing a semi-physical simulation test environment, constructing a redundancy sensor simulation model integrated with a high-fidelity error model, an execution mechanism model and an aircraft dynamics model in a simulation computer; the simulation computer generates an instruction and drives the model to operate, performs instruction-response-feedback closed-loop data interaction with flight control hardware, and sequentially executes model correction, integration test, scene test, robustness test and stepped fault injection test; quantitative indicators are calculated based on the test data to assess control performance and redundancy management capability of the flight control system. According to the method, the problems of unreal modeling, incomplete test coverage and weak fault verification capability of a traditional simulation method for a redundancy system are solved, and efficient and high-fidelity ground verification of the redundancy flight control system from functions to reliability is realized.

The invention discloses an aircraft trajectory optimization method based on a fixed thrust engine, and belongs to the field of guidance control, and the method comprises the steps: S1, building an aircraft dynamics model, and determining the working parameters of the engine; s2, determining a drop point range and drop point parameters; s3, determining a power-on point area and a power-off point area which can meet the drop point range and the drop point parameters through Monte Carlo simulation, and obtaining a flight control domain; and S4, making a guidance strategy of an engine working section based on the flight control domain. For the engine which can only be ignited for one time, the trajectory is only adjusted in the working period of the engine, the situation that attitude control consumption is too much due to frequent attitude adjustment is avoided, the attitude control total impulse and the attitude control service life requirement are optimized, and the drop point precision is guaranteed.

This invention discloses a ground-based test platform and method for integrated research on aircraft dynamics and control, relating to fields such as mechanics, aircraft dynamics and control technology, information technology, and testing technology. The ground-based test platform for integrated research on aircraft dynamics and control includes a support base, moving components, sensor components, and a control component. The sensor components are primarily force-measuring components. The support base is used to mount the force-measuring components, and the force-measuring components are used to mount the moving components. The moving components are used to mount the element under test (DUT). The control component receives signals transmitted from the sensor components and generates control commands to control the DUT. Based on this test platform, the required moving components can be assembled, thereby enabling testing of different DUTs.

This invention belongs to the field of reusable aircraft control technology, and relates to an intelligent fault diagnosis and control method for reusable aircraft. This method establishes a control-oriented aircraft dynamics model, integrates a neural network adaptive fault observer and a particle swarm optimization algorithm, to achieve high-precision tracking control of the aircraft's attitude and trajectory. When faced with faults such as actuator performance loss and unmeasurable deviations, it can quickly and accurately estimate fault information and integrate the reconstructed fault information into the control law design, ensuring stable flight even under severe fault conditions. Simultaneously, this method can effectively cope with complex disturbances such as aerodynamic uncertainties, elastic modal coupling, and external environmental interference. Through high-precision joint online estimation, it improves the real-time performance and accuracy of fault diagnosis. Furthermore, the particle swarm optimization algorithm further enhances the overall performance and adaptive capability of the control system, making the aircraft operate more efficiently and stably during flight.

A computer-implemented method for predicting behavior of aircraft is provided. The method comprises inputting a current state of a number of aircraft into a number of hidden layers of a neural network, wherein the neural network is fully connected. An action applied to the aircraft is input into the hidden layers concurrently with the current state. The hidden layers, according to the current state and current action, determine a residual output that comprises an incremental difference in the state of the aircraft resulting from the current action. A skip connection feeds forward the current state of the aircraft, and the residual output is added to the current state to determine a next state of the aircraft.

The application belongs to the technical field of aircraft planning, and provides an aircraft online trajectory re-planning method based on meta-learning hot start sequence convex optimization, which comprises the following steps: step one: establishing an aircraft dynamics model and an offline trajectory library; a three-degree-of-freedom high-speed aircraft dynamics model under a spherical coordinate system is established, and an offline trajectory library is constructed; step two: establishing a model-independent meta-learning training framework fused with nearest neighbor residuals; data features in the offline trajectory library are learned by fusing the nearest neighbor residuals, and optimal common initial parameters are trained and obtained, so that MAML can quickly adapt to a new environment; step three: establishing an online trajectory re-planning framework based on MAML hot start sequence convex optimization; the hot start initial value of the sequence convex optimization is used to guide the sequence convex optimization iteration to reach convergence, and an obstacle avoidance trajectory is generated under the premise of meeting the aircraft dynamics constraints. The application realizes the dual improvement of computational efficiency and control smoothness while ensuring the avoidance of threat areas.

The embodiment of the disclosure provides an imaging quality constrained flight control system, in which: a quality evaluation module is used for generating an imaging quality vector containing signal-to-noise ratio, phase noise, strip overlap degree and point cloud sparsity based on millimeter wave echo and point cloud data; a constraint mapping module is used for generating cost weight, soft constraint penalty coefficient, hard constraint threshold and safety boundary tightening coefficient through a mapping function based on the imaging quality vector; a rolling optimization module is used for constructing a quality-aware cost function based on an aircraft dynamics model and the output of the constraint mapping module, and solving an optimal control sequence in a limited time domain combined with geometric consistency and coverage constraint; and an execution interface module is used for converting the first control quantity in the optimal control sequence into an executable driving command and issuing the executable driving command to a UAV propulsion system and a gimbal. The system can achieve the effects of imaging stability improvement, coverage integrity improvement, robustness improvement and efficiency improvement.

This invention discloses a nonlinear control law evolution method based on a large language model and control effect evaluation index, including initializing the control law; converting the weights of the control law operators into sampling probabilities; sampling each control law operator according to the probability distribution of the control law operators in the large language model to generate a control law formula; inputting the set of control law formulas into an aircraft dynamics simulation model to obtain a tracking error sequence and determining the evaluation index matrix of the tracking error sequence; analyzing the relationship between the control law operators and the evaluation index and outputting the update gradient of each control law operator; updating the weights of each control law operator and generating multiple new control law operators, and initializing the weights of the new control law operators; ending the iteration when the number of iterations reaches the maximum limit and outputting the control structure and control parameters of the control law, otherwise resampling and iterating; this invention solves the problem that large language models cannot be embedded in the dynamics simulation environment and evaluate the effect of control laws.

The invention discloses a semi-physical simulation method and system based on an FPGA neural network control method, and the method comprises the steps: operating an aircraft dynamics model at a simulation end, and transmitting state variables, such as an attitude angle, a speed and an angular rate, to an FPGA end in real time through Socket communication; a fixed-point neural network controller is deployed at the FPGA end, a control instruction is reasoned in a parallel pipeline mode and is transmitted back to the simulation end through a TCP / IP protocol, the system adopts a PS-PL cooperative bus to realize zero-copy data transmission, hardware-level reasoning acceleration is realized in combination with a Viis HLS and PYNQ framework, and the communication delay is less than 400 s. Through closed-loop verification, the error between the FPGA reasoning result and the floating point simulation is smaller than 4.5 * 10 <-5 >, and microsecond-level real-time response and high consistency of the flight control algorithm are achieved.

The application provides an airport flight area crossing traffic command system based on an A-SMGCS system, and belongs to the technical field of intelligent traffic, and comprises: a sensor layer, which is used for collecting aircraft dynamic data, to-be-passed authorization object data, environmental meteorological data and dynamic obstacle data; a data processing and decision-making layer, which is used for standardizing processing and space-time mapping processing of the collected data, judging existing potential conflicts in combination with a dynamic threshold adjustment mechanism, and generating and issuing instructions based on multi-objective optimization decision-making; and an execution layer, which is used for controlling crossing traffic control equipment according to the issued instructions, and feeding back execution results to the data processing and decision-making layer in real time to form a closed-loop control. The application effectively reduces the traffic conflict risk of the crossing of the airport flight area, improves the crossing efficiency, reduces the vehicle waiting time and the aircraft taxi delay, adapts to the fine scheduling demand of the airport scene movement, and provides reliable support for the safe and efficient operation of the airport flight area traffic.

The application provides a kind of low airspace collaborative management method and system based on digital twinning, comprising: based on the real-time state data of aircraft obtained, analyzing aircraft maneuverability, flight envelope characteristic parameters, characteristic parameters include maximum speed, maximum acceleration, minimum turning radius, utilize kinematics and dynamics model, depict different aircraft movement law;Combined with real-time flight data of aircraft, through digital twinning technology, real-time synchronization aircraft motion trajectory and attitude change in virtual environment, accurately reproduce aircraft dynamic behavior, provide high-fidelity digital model for flight conflict detection, flight path prediction;For aircraft motion characteristic difference, construct multi-level, multi-granularity digital twinning model, digital twinning model includes airspace structure model at macro level, aircraft group behavior model at mesoscopic level and single aircraft dynamic model at micro level.

The invention relates to a carrier-based aircraft dynamic boundary coupling simulation method and system based on a neural network model, and belongs to the technical field of computational fluid mechanics simulation. The method comprises the steps of training an engine neural network agent model based on test or simulation data; initializing a CFD model and obtaining initial flow information; cFD transient calculation is carried out, and air inlet channel outlet flow field parameters are extracted; inputting air inlet outlet parameters and time into the agent model, and predicting an engine boundary condition update value; the predicted value is given to a CFD corresponding boundary, and the geometric profile of the spray pipe is updated according to a preset rule or parameter back calculation; and carrying out loop iteration until the end. According to the method, real-time dynamic closed-loop updating of the boundary condition and the geometric profile of the spray pipe is achieved, the problems that in traditional simulation, the boundary is rigid, dynamic feedback is ignored, and efficiency and precision are contradictory are effectively solved, the authenticity and efficiency of simulation of the air inlet duct-engine coupling system are remarkably improved, and the method is suitable for temperature distortion prediction and safety evaluation.

The invention relates to the technical field of air traffic management, in particular to an airport control conflict resolution auxiliary method and system, and the method comprises the steps: obtaining a multi-source heterogeneous air traffic control signal, carrying out the preprocessing and weighted fusion of the multi-source heterogeneous air traffic control signal, and generating aircraft dynamic data under a unified coordinate system; acquiring historical conflict resolution success trajectory data, and generating an airport scene exclusive AEM sub-model; based on the aircraft dynamic data, real-time conflict analysis and pre-judgment are carried out according to the airport scene exclusive AEM sub-model, and risk early warning information is generated; and generating a control conflict resolution scheme according to the risk early warning information, and outputting an expected effect of the control conflict resolution scheme. According to the method and the device, the problem of deviation of core data such as the position and the speed of the aircraft due to data isomerism in the prior art can be solved, and the problem of poor adaptability to complex conflict scenes in the re-flight stage and the take-off stage due to manual operation can be solved.

The application relates to a virtual-real combined multi-machine cooperative semi-physical simulation system and method; the system at least comprises a multi-model parallel simulation machine, an interface device, a real physical guidance prototype and a real-time network; the multi-model parallel simulation machine is used for real-time solution of mathematical software of aircraft dynamics, kinematics equations and control systems; data interaction between the real physical guidance prototype and the multi-model parallel simulation machine is realized through the real-time network and the interface device. The simulation method provided by the application has strong expansibility and high universality; by using the simulation method and system provided by the application, only adaptive development of model quantity in the multi-model parallel simulation machine and different model requirements is needed, so that different model verification requirements can be matched.

The embodiment of the invention provides an imaging quality constrained model prediction flight control system, and the system comprises a quality evaluation module which is used for generating an imaging quality vector comprising a signal-to-noise ratio, phase noise, a stripe overlapping degree and point cloud sparseness based on millimeter wave echoes and point cloud data; the constraint mapping module is used for generating a cost weight, a soft constraint penalty coefficient, a hard constraint threshold value and a security boundary tightening coefficient through a mapping function based on the imaging quality vector; the rolling optimization module is used for constructing a quality perception cost function based on the aircraft dynamics model and the output of the constraint mapping module, and solving in a limited time domain by combining geometric consistency and coverage constraint to obtain an optimal control sequence; and the execution interface module is used for converting the first control quantity in the optimal control sequence into an executable driving command and issuing the executable driving command to an unmanned aerial vehicle propulsion system and a holder. The system can achieve the effects of improving the imaging stability, improving the coverage integrity, improving the robustness and improving the efficiency.

This invention discloses a method for estimating the remaining average velocity and controlling the flight time of a variable-speed missile, comprising the following steps: constructing a missile motion model and a differential geometry guidance model; predicting the remaining flight path of the missile under proportional guidance, and constructing a flight path error dynamic equation in the form of biased proportional guidance; designing a flight path control guidance law based on the optimal error dynamics method; predicting the missile velocity profile and estimating the remaining average velocity; converting the desired remaining flight time into the desired remaining flight path, and achieving flight time control through the flight path control guidance law. This invention relates to the field of aircraft dynamics, guidance, and control, and can effectively improve the accuracy of existing missile remaining average velocity estimation algorithms and reduce flight time control errors. It can provide an effective solution to the problem of time-varying multi-missile cooperative guidance.