Auxiliary decision-making system and method for landing of manned aircraft

Abstract

The invention provides a manned aircraft landing aid decision-making system and method, and relates to the technical field of aviation. The method specifically comprises the following steps: acquiring current state data of an aircraft in real time, planning a landing trajectory of the aircraft by adopting an index leveling trajectory method, and determining an optimal glide path of the aircraft; based on the current state data of the aircraft and the optimal glide path of the aircraft, calculating an expected attitude angle and an expected thrust variation of the aircraft by adopting an instruction decision algorithm, performing attitude control on the aircraft according to the expected attitude angle of the aircraft, and generating a flight decision suggestion; wherein the attitude control is divided into two parts: one part is used for calculating an expected three-dimensional angle and a three-axis attitude angular velocity by utilizing angle control; according to the other part, the three-axis attitude angular acceleration is calculated through angular velocity control; and the pilot controls the flight path of the aircraft according to the generated flight decision suggestion to ensure that the aircraft lands along the optimal glide path of the aircraft. The invention aims to solve the problem that the landing trajectory is inaccurate when the aircraft lands.

Description

Technical Field

[0001] The present invention relates to the field of aviation technology, and in particular to a manned aircraft landing auxiliary decision system and method. Background Art

[0002] With the rapid development of the aviation transportation industry, aircraft flight safety has become increasingly critical; ensuring the stability and safety of every flight has become a top priority. During the landing phase of the aircraft, accurate control of the trajectory is crucial. In the traditional aircraft landing process, the landing of the aircraft often relies solely on the pilot's experience and operation, which has certain uncertainties and is easily affected by the external environment and the pilot's own state. When landing, the pilot needs to concentrate highly and operate the control stick frequently, which not only increases the pilot's workload, but also may cause untimely or inaccurate operations under complex weather conditions. This method of relying on manual operation is not only inefficient, but may also cause unstable flight attitude due to human factors, increasing the risk of landing. At the same time, in some advanced flight control systems, although there is a certain degree of automated assistance for landing, it is still not perfect and accurate enough, and cannot fully adapt to various complex and changeable landing scenarios. Summary of the invention

[0003] In view of the above-mentioned deficiencies in the prior art, the present invention proposes a manned aircraft landing auxiliary decision system and method, which aims to solve the problem of inaccurate landing trajectory during aircraft landing. It can accurately plan the flight landing trajectory during the aircraft landing process and make aircraft command decisions on the related landing trajectory, and can achieve precise landing of the flight landing through manual intervention of the pilot to ensure a smooth and safe landing.

[0004] The first aspect of the present invention provides a manned aircraft landing auxiliary decision system, the system comprising: an aircraft state and environmental information perception module, a landing trajectory planning and command decision module and a data interaction module;

[0005] The aircraft status and environmental information perception module is used to monitor the aircraft's attitude and environmental data in real time and perform fusion collection to obtain aircraft status information including aircraft altitude information, aircraft three-dimensional position information, meteorological information and aircraft posture information, and send it to the landing trajectory planning and command decision module and the data interaction module;

[0006] The landing trajectory planning and command decision module is used to plan the landing trajectory of the aircraft according to the received aircraft status and environmental information by using the exponential flattening trajectory method to determine the optimal glide path of the aircraft; use the command decision algorithm to calculate the expected attitude angle of the aircraft and perform attitude control on the aircraft to determine the expected attitude angular velocity and expected angular acceleration; generate flight decision suggestions, and transmit the optimal glide path and flight decision suggestions of the aircraft to the data interaction module;

[0007] The data interaction module is used to display the aircraft status information, the optimal glide path and the flight decision suggestion to the pilot in real time through the data visualization interface, so that the pilot can control the aircraft to land along the optimal glide path according to the flight decision suggestion;

[0008] The aircraft status and environmental information perception module includes: a location and weather module, a machine vision module and a data acquisition module;

[0009] The position and weather module is used to obtain the aircraft's altitude information, the aircraft's three-dimensional position information and weather information in real time, and send them to the data acquisition module;

[0010] The location and weather module includes: an air pressure sensor module, a laser sensor module, an ultrasonic sensor module, a global positioning system module, an atmospheric data system module and a microcontroller;

[0011] The air pressure sensor module is used to calculate the altitude of the aircraft by measuring the atmospheric pressure at the altitude of the aircraft and send the information to the microcontroller;

[0012] The laser sensor module is used to emit laser pulses to the ground and receive reflected lasers from the ground, and record the time from the emission of the laser pulse to the reception of the laser pulse, calculate the altitude of the aircraft based on the recorded time and send it to the microcontroller;

[0013] The ultrasonic sensor module is used to transmit ultrasonic waves to the ground and receive returned ultrasonic waves, and record the round-trip time difference of the ultrasonic waves, calculate the height of the aircraft according to the recorded round-trip time difference and send it to the microcontroller;

[0014] The global positioning system module is used to obtain the three-dimensional position information of the aircraft and send it to the microcontroller; the three-dimensional position information of the aircraft includes: longitude and latitude;

[0015] The atmospheric data system module is used to obtain meteorological information and send it to the microcontroller; wherein the meteorological information includes: runway wind direction, runway wind speed, atmospheric temperature and atmospheric humidity;

[0016] The microcontroller is used to use a weighted average algorithm to average the heights of the aircraft calculated by the air pressure sensor module, the laser sensor module and the ultrasonic sensor module to obtain the height information of the aircraft; and send the height information of the aircraft, the three-dimensional position information of the aircraft and the meteorological information to the data acquisition module;

[0017] The machine vision module is used to obtain the position and posture information of the aircraft in real time; wherein the position and posture information of the aircraft includes: the pitch angle of the aircraft, the roll angle of the aircraft, the heading angle of the aircraft and the runway position;

[0018] The machine vision module comprises: a CCD camera, an image acquisition card and a visual processing module; wherein the CCD camera is connected to an external signal interface of the image acquisition card;

[0019] The CCD camera is used to capture the ground sign image of the aircraft landing point in real time;

[0020] The image acquisition card is used to transmit the ground sign image captured by the CCD camera to the visual processing module;

[0021] The visual processing module is used to perform image preprocessing on the received ground mark image, perform feature extraction and feature matching on the processed ground mark image, and estimate the position and posture of the aircraft according to the matching result to obtain the position and posture information of the aircraft;

[0022] The data acquisition module is used to integrate the received aircraft altitude information, aircraft three-dimensional position information, weather information and aircraft posture information, and send the integrated data information to the landing trajectory planning and command decision module;

[0023] The data interaction module is implemented through a data visualization interface; the data visualization interface is divided into three display areas, namely: an aircraft status information area, an environmental information area, and a decision suggestion area;

[0024] The aircraft status information area is used to display the current altitude, current speed, current acceleration, current pitch angle and current roll angle of the aircraft in real time;

[0025] The environmental information area is used to display the runway information and weather information during the aircraft landing process in real time;

[0026] The decision suggestion area is used to display the aircraft's glide path and decision suggestions in real time, and to issue a warning when the aircraft's trajectory deviates. The pilot controls the aircraft to correct the flight trajectory to the optimal glide path according to the flight decision suggestions.

[0027] A second aspect of the present invention provides a manned aircraft landing auxiliary decision method, which is implemented by a manned aircraft landing auxiliary decision system. The method comprises the following steps:

[0028] Acquire the current status data of the aircraft in real time, including: the current altitude of the aircraft, the current three-dimensional position of the aircraft, weather information and the current posture of the aircraft; wherein the weather information includes: runway wind direction, runway wind speed, atmospheric temperature and atmospheric humidity; the current posture of the aircraft includes: current pitch angle, current roll angle, current heading angle and runway position;

[0029] According to the current state data of the aircraft, the exponential flattening trajectory method is used to plan the landing trajectory of the aircraft and determine the optimal glide path of the aircraft;

[0030] Based on the current state data of the aircraft and the optimal glide path of the aircraft, a command decision algorithm is used to calculate the expected attitude angle and the expected thrust change of the aircraft, and the aircraft is attitude controlled according to the expected attitude angle of the aircraft, and a flight decision suggestion is generated; wherein the attitude control is divided into two parts: one part is to calculate the expected three-dimensional angle and three-axis attitude angular velocity by angle control; the other part is to calculate the three-axis attitude angular acceleration by angular velocity control; the flight decision suggestion includes: the expected thrust change, the expected three-dimensional angle, the three-axis attitude angular velocity and the three-axis attitude angular acceleration;

[0031] The desired attitude angle includes: a desired roll angle and a desired pitch angle; the command decision algorithm includes: a lateral command and a longitudinal command;

[0032] The pilot controls the flight trajectory of the aircraft based on the generated flight decision suggestions to ensure that the aircraft lands along the aircraft's optimal glide path;

[0033] The method for obtaining the current altitude of the aircraft is: using an air pressure sensor to collect the atmospheric pressure of the aircraft at the current altitude in real time, and calculating the altitude of the aircraft h according to the atmospheric pressure of the aircraft at the current altitude 1 ;

[0034]

[0035] Where T b Indicates height H b The atmospheric temperature at that time; β is the vertical temperature change rate; Indicates the height is h 1 Atmospheric pressure at that time; P b Indicates height H b Atmospheric pressure at 1000 s; R is the air specific gas number; g is the gravitational acceleration;

[0036] The laser sensor is used to emit laser pulses towards the ground. After being reflected by the ground, part of the laser pulses is received by the laser sensor. The time from the emission to the reception of the laser pulses is recorded, and the height of the aircraft h is calculated based on the recorded time. 2 ;

[0037] The ultrasonic sensor transmits ultrasonic waves to the ground and receives the returned ultrasonic waves, records the time difference of the ultrasonic waves and calculates the height h of the aircraft. 3 ;

[0038]

[0039] Where c represents the speed of sound waves in air; t 3 Indicates the time difference of the ultrasonic wave's round trip;

[0040] The weighted average algorithm is used to calculate the aircraft's altitude h 1 、h 2 and h 3 The average value is taken as the current height h of the aircraft;

[0041]

[0042] where k 1 is the height h 1 The weight coefficient of k 2 is the height h 2 The weight coefficient of k 3 is the height h 3 The weight coefficient of

[0043] The method of planning the landing trajectory of the aircraft using the exponential flattening trajectory method is as follows: the landing trajectory of the aircraft is divided into a flattening section and a glide section, and the flattening angle θ of the aircraft when landing is obtained. 1 , glide angle θ 2 , flatten the trajectory asymptote distance h jd , the aircraft’s approach speed V and the aircraft’s approach time t 0 ;

[0044] First, plan the landing trajectory of the aircraft in the leveling section, specifically: use the leveling angle θ when the aircraft lands 1 Calculate the horizontal distance of the aircraft in the leveling section. The horizontal distance of the aircraft in the leveling section includes two parts: the horizontal distance L in the first half of the leveling section 2 The horizontal distance L from the second half of the pull 1 ;

[0045] The horizontal distance flown by the aircraft in the leveling section is expressed as:

[0046]

[0047] Where L f It indicates the horizontal distance that the aircraft flies in the leveling phase, and L f =L 1 +L 2 ; V 0 Indicates the average speed of the aircraft during the leveling phase; t 1 is the leveling time of the aircraft in the leveling section; τ is the time constant of the exponential leveling curve; H 2 To level the decision-making height; Indicates the vertical speed at which the aircraft can land safely;

[0048] The horizontal distance L that the aircraft flies in the leveling section f With the leveling decision height H 2 The relationship between them is:

[0049]

[0050] Then calculate the horizontal distance L of the first half of the leveling 2 The horizontal distance L from the second half of the pull 1 , expressed as:

[0051]

[0052] Then, the landing trajectory of the aircraft in the descent phase is planned. Specifically, since the end of the descent phase is connected to the beginning of the leveling phase, the speed of the aircraft at the end of the descent phase is determined according to the landing trajectory of the aircraft in the leveling phase, and the landing speed V and the descent angle θ are calculated according to the landing speed V and the descent angle θ. 2 The aircraft's descent time is calculated by the aircraft's speed at the end of the descent phase, and then the aircraft's descent decision altitude H is calculated. 1 , using the aircraft's descent to determine the altitude H 1 Calculate the horizontal distance L of the aircraft during the glide path 3 , expressed as:

[0053]

[0054] Finally, the optimal glide path of the aircraft is determined according to the landing trajectory of the aircraft in the leveling phase and the landing trajectory of the aircraft in the glide phase, which is expressed as:

[0055]

[0056] Among them, H D It represents the altitude of the aircraft during landing; L represents the horizontal distance of the aircraft during landing; t represents time; e is a natural constant;

[0057] The lateral instruction is: using the L1 guidance law, calculating the desired roll angle of the aircraft according to the current position of the aircraft and the optimal glide path of the aircraft;

[0058] The method for calculating the desired roll angle is as follows: select a reference point P on the aircraft's optimal glide path according to the aircraft's current position and flight speed. R , connect the reference point and the current position of the aircraft, and record the length of the connection as L 1 , the angle between the current speed direction of the aircraft and the connecting line is recorded as η, according to the current position of the aircraft and the reference point P R The centripetal acceleration of the aircraft is calculated based on the geometric relationship between them, which can be divided into the following two cases:

[0059] The first case: When the reference point is located in the arc section of the optimal glide path, take 2η as the center angle, and use the current position of the aircraft and the reference point P R Construct an arc and then calculate the centripetal acceleration of the aircraft based on the geometric relationship. The expression is:

[0060]

[0061] where a Scmd is the centripetal acceleration of the aircraft; v is the horizontal speed of the aircraft; R 1 is the arc radius;

[0062] The second case: When the reference point is located on the straight line segment of the optimal glide path, the aircraft's glide path is l R , the current speed direction of the aircraft and the glide path l R The angle between 1 , the glide path l R The angle between the reference point and the current position of the aircraft is denoted as η 2 , and then calculate the centripetal acceleration of the aircraft based on the geometric relationship, the expression is:

[0063]

[0064] Where d represents the vertical distance between the aircraft's current position and the glide path; is the rate of change of the vertical distance between the aircraft's current position and the glide path;

[0065] The expected roll angle of the aircraft is calculated based on the centripetal acceleration of the aircraft. The expression is:

[0066]

[0067] Where m is the weight of the aircraft; g is the acceleration due to gravity; φ sp is the desired roll angle;

[0068] The longitudinal instruction is: based on the aircraft's track angle, the aircraft's current actual airspeed, the aircraft's current acceleration and the optimal glide slope in the current flight state, the thrust change control law and the elevator control law are respectively constructed by the TECS control method, and the thrust change control law and the elevator control law are used to solve the aircraft's expected thrust change and the aircraft's expected pitch angle;

[0069] The method for calculating the desired thrust change and the desired pitch angle is as follows: obtaining the flight path angle γ of the aircraft in the current flight state, the current actual airspeed V of the aircraft n and the aircraft's current acceleration And determine the expected flight path angle γ of the aircraft based on the aircraft's optimal glide path c and the desired airspeed, and use the desired flight angle of the aircraft to calculate the flight angle feedback error γ e , using the aircraft's airspeed and desired airspeed to calculate the acceleration control amount And according to the current acceleration of the aircraft Calculate the acceleration feedback error

[0070] The track angle feedback error γ e and acceleration feedback error It is expressed as:

[0071]

[0072] Based on the track angle feedback error γ e and acceleration feedback error The thrust variation control law and elevator control law of the TECS controller are constructed and expressed as:

[0073]

[0074] Where T c is the expected thrust change; K tp is the proportional coefficient of thrust control; K is the total energy change rate feedback of the aircraft; ti The integral coefficient of thrust control; s is the integral variable; Indicates the total energy change rate deviation of the aircraft; δ ec is the desired elevator change; K ep is the proportional coefficient of elevator control; K is the energy distribution rate feedback of the aircraft; ei is the integral coefficient of elevator control; It is the energy distribution rate deviation of the aircraft;

[0075] The desired thrust change T is calculated using the thrust change control law and elevator control law of the TECS controller. c and the desired elevator change δ ec , assuming that the change in the aircraft angle of attack is negligible, the expected elevator change is directly used as the expected pitch angle;

[0076] The method of calculating the desired three-dimensional angle and three-axis attitude angular velocity by angle control comprises the following steps:

[0077] Get the angular velocity and acceleration of the aircraft in the current flight state, and calculate the actual roll angle and actual pitch angle of the aircraft in the current flight state;

[0078] According to the actual roll angle, the Kalman filter method is used to calculate the roll angle estimate According to the actual pitch angle, the Kalman filter method is used to calculate the estimated pitch angle

[0079] Calculate the expected roll angle φ sp and the estimated roll angle The difference between them is taken as the rolling angle error φerr , and then the roll angle error φ err Perform proportional control to obtain the desired rolling angular velocity

[0080] Calculate the desired pitch angle θ sp and the pitch angle estimate The difference between them is taken as the pitch angle error θ err , and then the pitch angle error θ err Perform proportional control to obtain the desired pitch angular velocity

[0081] According to the desired roll angle φ sp and the desired pitch angle θ sp , using the coordinated turning control law to calculate the desired yaw rate And according to the expected yaw rate Calculate the desired yaw angle ψ sp ;

[0082] The coordinated turning control law is expressed as:

[0083]

[0084] Where V T Indicates true airspeed;

[0085] The desired roll angle φ sp , desired pitch angle θ sp and the desired yaw angle ψ sp as the desired three-dimensional angle;

[0086] Using the expected roll velocity Expected roll angular velocity and the desired yaw rate Construct the three-axis attitude angle change rate, expressed as:

[0087]

[0088] in Indicates the rate of change of the three-axis attitude angle;

[0089] Then the three-axis attitude angle change rate is converted into a coordinate system to obtain the three-axis attitude angular velocity in the body coordinate system

[0090]

[0091] in Indicates the three-axis attitude angular velocity; p sp represents the expected rolling angular velocity; q sp represents the expected pitch angular velocity; γ sprepresents the desired yaw rate; φ represents the roll angle;

[0092] The method of calculating the three-axis attitude angular acceleration by using angular velocity control comprises the following steps:

[0093] The three-axis attitude angular velocity As the feedforward input of angular velocity control, and through feedforward control calculation feedforward control three-axis attitude angular acceleration

[0094] Get the actual value of the three-axis attitude angular velocity It is expressed as:

[0095]

[0096] in Indicates the actual value of the rolling angular velocity; Indicates the actual value of the pitch angular velocity; Indicates the actual value of the track angular velocity;

[0097] Using three-axis attitude angular velocity And the actual value of the three-axis attitude angular velocity Calculate the angular velocity error of the three-axis attitude It is expressed as:

[0098]

[0099] where per r Indicates the rolling angular velocity error; qer r Indicates the pitch angular velocity error; γer r represents the track angular velocity error;

[0100] The three-axis attitude angular velocity error Input PI controller to get PI controlled angular acceleration

[0101] Using feedforward to control the three-axis attitude angular acceleration and PI control angular acceleration Calculate the three-axis attitude angular acceleration It is expressed as:

[0102]

[0103] Among them FF Scaler is the feedforward control airspeed scaling factor; PI Scaler is the PI control airspeed scaling factor; and:

[0104]

[0105] Where V Tis the true airspeed; V I is the indicated airspeed; and They are the set true airspeed and indicated airspeed respectively.

[0106] The beneficial effects of adopting the above technical solution are:

[0107] The system of the present invention can use high-precision sensors and advanced data processing technology to quickly and accurately obtain the aircraft's attitude information through the coordinated work of the aircraft status and environmental information perception module and the landing trajectory planning and command decision algorithm module, and accurately plan the landing trajectory based on the real-time flight data, and then perform intelligent angle correction on the aircraft according to the optimal glide path, and display the correction results to the pilot in real time, so that the pilot can adjust the flight attitude in a timely and accurate manner during the landing process. This intelligent angle correction function can adapt to various complex and changeable landing scenarios, not only improving the accuracy of landing, but also reducing the risks caused by trajectory deviation, thereby improving the stability and safety of the landing process.

[0108] The system of the present invention, through the data interaction module, can enable the pilot to perform precise operations according to the obtained flight decision suggestions, reduce the risks caused by human judgment errors, and can reduce the burden on the pilot and improve the comfort during the flight.

[0109] In summary, the system of the present invention has the ability to respond to abnormal situations in a timely manner under the coordination of various modules, ensuring accurate control of the aircraft and achieving accurate landing of the aircraft. When encountering sudden weather changes or abnormal aircraft status, the system can respond quickly, ensure flight safety through intelligent algorithms and real-time data processing, and has good stability, reducing the probability of accidents and improving flight reliability.

[0110] The method of the present invention adopts the L1 guidance law to control the lateral flight of the aircraft and calculates the desired roll angle. By selecting a reference point to calculate the centripetal acceleration, the desired roll angle can be quickly generated and a smooth transition of the aircraft in the subsequent attitude control process can be ensured, thereby improving the stability and safety of the flight, and having good anti-interference performance, and being able to effectively cope with the influence of external environmental factors such as wind disturbance.

[0111] The method of the present invention adopts a TECS controller to control the flight speed and altitude of the aircraft and calculate the expected pitch angle and the expected thrust change. By reasonably allocating the control amount of thrust and pitch angle, the change of total energy is mainly adjusted by the throttle, while the conversion between kinetic energy and potential energy is controlled by the elevator, so that the control of the altitude and speed of the aircraft is no longer coupled with each other and can be adjusted independently.

[0112] In summary, the method of the present invention can obtain relevant real-time data more accurately through the L1 and TECS guidance algorithms, and provide corresponding flight decision suggestions for pilots, thereby optimizing the fuel consumption and maintenance costs of the aircraft and improving the utilization efficiency of aviation resources. BRIEF DESCRIPTION OF THE DRAWINGS

[0113] Figure 1 This is a structural diagram of a manned aircraft landing auxiliary decision system in this embodiment;

[0114] Figure 2 This is a structural diagram of an aircraft state and environment information perception module in a fixed-wing manned aircraft in this embodiment;

[0115] Figure 3 This is a functional schematic diagram of a data visualization interface corresponding to the data interaction module in this embodiment;

[0116] Figure 4 This is a layout diagram of the data visualization interface in this implementation mode;

[0117] Figure 5 This is a flow chart of a manned aircraft landing auxiliary decision method in this embodiment;

[0118] Figure 6 is a schematic diagram of a manned aircraft landing auxiliary decision method in this embodiment;

[0119] Figure 7 Schematic diagram of the landing process of the fixed-wing manned aircraft in this embodiment;

[0120] Figure 8 This is a schematic diagram of an aircraft path analysis in which the target route is an arc in this implementation manner;

[0121] Fig. 9 This is a schematic diagram of an aircraft path analysis in which the target route is a straight line in this embodiment;

[0122] Fig.10 This is a schematic diagram of the lateral force analysis of the aircraft in this embodiment;

[0123] Fig.11 Schematic diagram of the core algorithm of the TECS controller in this embodiment;

[0124] Fig.12 This is a control principle diagram of the posture control in this embodiment;

[0125] Fig.13 This is the outer ring angle control principle diagram in this implementation mode;

[0126] Fig.14 This is a schematic diagram of the inner ring angular velocity control principle in this embodiment. DETAILED DESCRIPTION

[0127] In order to facilitate the understanding of the present application, the specific embodiments of the present invention are further described in detail below in conjunction with the accompanying drawings and embodiments. The following embodiments are used to illustrate the present invention, but are not intended to limit the scope of the present invention. On the contrary, the purpose of providing these embodiments is to make the disclosure of the present application more thoroughly understood.

[0128] A manned aircraft landing auxiliary decision system of this embodiment, such as Figure 1 As shown, the system includes: an aircraft status and environmental information perception module, a landing trajectory planning and command decision module, and a data interaction module.

[0129] The aircraft status and environmental information perception module is used to monitor the aircraft's attitude and environmental data in real time and perform fusion collection to obtain aircraft status information including aircraft altitude information, aircraft three-dimensional position information, meteorological information and aircraft posture information, and send it to the landing trajectory planning and command decision module and the data interaction module.

[0130] In this embodiment, fusion acquisition refers to the process of integrating and processing data from different sensors to provide a more accurate and comprehensive data view to facilitate subsequent analysis and decision-making. The landing trajectory planning and command decision module performs trajectory planning and correction based on the acquired aircraft attitude information and environmental information, and in terms of the flight trajectory correction function, an intelligent attitude control algorithm is used to ensure that the aircraft trajectory can be adjusted in a timely and accurate manner. In the data interaction module, visualization technology is used to provide pilots with an intuitive and clear operating interface.

[0131] The aircraft status and environmental information perception module includes: a position and weather module, a machine vision module and a data acquisition module.

[0132] For fixed-wing manned aircraft, such as Figure 2 As shown in the figure, the state and environmental information perception is mainly composed of two parts: the position and meteorological module and the machine vision module. The position and meteorological module is used to measure the aircraft's altitude, specific location and environmental meteorological information. The aircraft's altitude is mainly measured by the air pressure sensor module, ultrasonic sensor module and laser sensor module, the aircraft's specific location is mainly measured by the global positioning system module, and the environmental meteorological information is mainly provided by the atmospheric data system module.

[0133] The position and weather module is used to obtain the aircraft's altitude information, the aircraft's three-dimensional position information and weather information in real time, and send them to the data acquisition module.

[0134] The location and weather module includes: an air pressure sensor module, a laser sensor module, an ultrasonic sensor module, a global positioning system module, an atmospheric data system module and a microcontroller.

[0135] The air pressure sensor module is used to calculate the altitude of the aircraft by measuring the atmospheric pressure at the altitude of the aircraft and send the result to the microcontroller.

[0136] The laser sensor module is used to emit laser pulses to the ground and receive reflected lasers from the ground, while recording the time from the emission of the laser pulse to its reception, and calculating the altitude of the aircraft based on the recorded time and sending it to the microcontroller.

[0137] The ultrasonic sensor module is used to transmit ultrasonic waves to the ground and receive returned ultrasonic waves, and record the round-trip time difference of the ultrasonic waves. The height of the aircraft is calculated according to the recorded round-trip time difference and sent to the microcontroller.

[0138] The global positioning system module is used to obtain the three-dimensional position information of the aircraft and send it to the microcontroller; the three-dimensional position information of the aircraft includes: longitude and latitude.

[0139] In this embodiment, the global positioning system GPS receives signals transmitted by multiple satellites and determines the geographical location of the aircraft, including longitude and latitude, based on the received signals using the triangulation positioning principle.

[0140] The atmospheric data system module is used to obtain meteorological information and send it to the microcontroller.

[0141] The meteorological information includes: runway wind direction, runway wind speed, atmospheric temperature and atmospheric humidity.

[0142] In this embodiment, the atmospheric data system module includes data acquisition equipment such as anemometers, wind vanes, temperature sensors, and humidity sensors. The anemometers and wind vanes are used to measure wind direction and wind speed near the runway, and the temperature sensors and humidity sensors are used to measure the temperature and humidity of the atmosphere.

[0143] The microcontroller is used to use a weighted average algorithm to average the heights of the aircraft calculated by the air pressure sensor module, the laser sensor module and the ultrasonic sensor module to obtain the height information of the aircraft; and send the height information of the aircraft, the three-dimensional position information of the aircraft and the meteorological information to the data acquisition module.

[0144] In this embodiment, three altitude measurement sensor modules, namely, the air pressure sensor module, the laser sensor module and the ultrasonic sensor module are all connected to the microcontroller. The microprocessor uses a weighted average algorithm to calculate the altitude information of the aircraft based on the altitude data measured by the three altitude sensors, thereby obtaining high-precision data suitable for autonomous landing.

[0145] The machine vision module is used to obtain the position and posture information of the aircraft in real time.

[0146] The aircraft's position information includes: the aircraft's pitch angle, the aircraft's roll angle, the aircraft's heading angle, and the runway position.

[0147] The machine vision module comprises: a CCD camera, an image acquisition card and a vision processing module; wherein the CCD camera is connected to an external signal interface of the image acquisition card.

[0148] The CCD camera is used to capture the ground sign image of the aircraft landing point in real time.

[0149] The image acquisition card is used to transmit the ground sign image captured by the CCD camera to the visual processing module.

[0150] The visual processing module is used to perform image preprocessing on the received ground mark image, and perform feature extraction and feature matching on the processed ground mark image, and estimate the aircraft's posture according to the matching result to obtain the aircraft's posture information.

[0151] In this embodiment, the CCD camera is mainly made of a CCD charge coupled device, which can capture the ground mark image of the landing point in real time. The image acquisition card has an external signal interface connected to the CCD camera. The image acquisition card can collect the ground image captured by the CCD camera and send it to the visual processing module. The visual processing module is based on the STM32F427CPU, has rich hardware resources and expansion interfaces, performs image preprocessing such as image filtering on the collected ground recognition mark image, and then performs feature extraction, feature matching and posture estimation on the ground recognition mark image. The machine vision module finally solves the pitch angle, roll angle and heading angle and other attitude information and position information of the aircraft.

[0152] The data acquisition module is used to integrate the received aircraft altitude information, aircraft three-dimensional position information, weather information and aircraft posture information, and send the integrated data information to the landing trajectory planning and command decision module.

[0153] In this embodiment, the position and weather module calculates the altitude, three-dimensional position and weather information, and the machine vision module calculates the pitch angle, roll angle, heading angle and other attitude information of the fixed-wing manned aircraft as well as the runway information, and transmits the above information to the data acquisition module for data integration, and sends the integration result to the landing trajectory planning and command decision module.

[0154] The landing trajectory planning and command decision module is used to plan the landing trajectory of the aircraft according to the received aircraft status and environmental information by using the exponential flattening trajectory method to determine the optimal glide path of the aircraft; to calculate the expected three-dimensional angle of the aircraft by using the command decision algorithm and perform attitude control on the aircraft to determine the expected attitude angular velocity and expected angular acceleration; to generate flight decision suggestions, and transmit the optimal glide path and flight decision suggestions of the aircraft to the data interaction module.

[0155] In this embodiment, the landing trajectory planning part in the landing trajectory planning and command decision module analyzes and calculates the collected data, comprehensively considers factors such as aircraft performance and runway conditions, determines the optimal glide path, and intelligently corrects and adjusts the angle during the aircraft landing process. The command decision algorithm in the landing trajectory planning and command decision module refers to obtaining the desired roll angle and pitch angle through lateral and longitudinal commands, and then inputting them into the attitude control unit to obtain the final three-dimensional velocity, angular velocity, and angular acceleration.

[0156] In the flight control of fixed-wing aircraft, when the aircraft deviates from the desired track, if the aircraft is to return to the track, that is, to reach the next desired point, it is necessary to decouple the lateral and longitudinal directions of the fixed-wing aircraft to make corrections separately, in order to more effectively control the lateral movement of the aircraft, that is, the roll angle, and the longitudinal movement, that is, the pitch angle and speed, to achieve more accurate path tracking and stable flight. Specifically: after obtaining the current state data of the aircraft, the desired roll angle and pitch angle are first calculated through the L1 guidance rate and the total energy control system (TECS) controller, and then substituted into the attitude control to obtain the specific three-axis angular velocity and angular acceleration, and then the data is input into the pilot operation interface so that the pilot can intuitively see the relevant parameters that need to be adjusted before performing specific operations.

[0157] The data interaction module is used to display the aircraft status information, the optimal glide path and the flight decision suggestions to the pilot in real time through the data visualization interface, so that the pilot can control the aircraft to land along the optimal glide path according to the flight decision suggestions.

[0158] In this embodiment, the data visualization interface can display the aircraft status information, the aircraft's optimal glide path, and flight decision suggestions to the pilot in real time. When the aircraft's optimal glide path is determined, the aircraft's landing glide path and the real-time attitude data required for landing according to the path will be displayed, and the pilot will operate the aircraft's control console according to the information and flight decision suggestions displayed in real time on the data visualization interface.

[0159] In this embodiment, the data interaction module is implemented through a data visualization interface. The human-machine interaction interface is designed using Beckhoff's TwinCAT HMI software, and the communication with the Beckhoff controller is based on the ADS communication protocol. TwinCAT HMI is Beckhoff's new generation HMI framework. TwinCAT HMI integrates the human-machine interface directly into the familiar Visual Studio environment. The interface style uses existing HTML5 templates to enable developers to achieve more complex but ergonomic interfaces with less work. The interface logic can be implemented on the client side through JavaScript or as an extension of the server, and can be implemented through C++ or .net. The ADS communication protocol is a special communication protocol defined by Beckhoff, which implements the protocol for communication between the TwinCAT controller and the upper user interface by calling a dynamic link library file.

[0160] In this embodiment, if Figure 3 As shown, the data visualization interface is divided into three display areas: aircraft status information area, environmental information area and decision suggestion area.

[0161] The aircraft status information area is used to display the current altitude, current speed, current acceleration, current pitch angle and current roll angle of the aircraft in real time, including: altitude trend indicator, speed trend indicator, accelerometer, attitude indicator, fuel quantity and remaining time; the altitude trend indicator is located on the right side of the data visualization interface, and is displayed in the form of an altitude trend indicator diagram. The current altitude of the aircraft is intuitively displayed through the scale and numbers on the altimeter. The scale color is white. If the current altitude exceeds the safety range, the digital display will turn red and flash. Figure 4 As shown, the altitude trend indicator includes: an altimeter, a numerical display of the target altitude, and the current altitude. The speed trend indicator is located on the left side of the data visualization interface, and intuitively displays the current speed of the aircraft in the form of numbers and scales on the airspeed indicator. The scale color is white. If the current speed exceeds the safety range, the digital display changes to red and flashes. The speed trend indicator includes: an airspeed indicator, a current speed, and a speed alarm. The accelerometer is located in the upper left portion of the data visualization interface, and intuitively displays the current acceleration of the aircraft through the scales and numbers on the accelerometer. The attitude indicator is located in the middle of the data visualization interface, and displays the pitch angle and roll angle of the aircraft at this time in the form of a scale from the first-person perspective of the aircraft. When the attitude changes, the scale pointer deflects accordingly. As shown Figure 4 As shown, the attitude indicator includes: a sky-ground artificial line, a pitch attitude scale, a sideslip indicator, a roll pointer and a roll scale. The fuel quantity and remaining time are located at the lower left of the data visualization interface, and the fuel quantity is displayed in the form of a number and a progress bar, and the remaining flight time is estimated based on the current speed and fuel consumption of the aircraft.

[0162] The environmental information area is used to display the runway information and weather information in real time during the aircraft landing process; the runway information displays the length and width of the runway and the relative position of the aircraft and the runway in a combination of text and graphics in the data visualization interface. The weather information is displayed in the lower right corner of the data visualization interface, namely the temperature and air pressure conditions, and the wind direction, wind speed and other weather information are displayed in the form of images and numbers.

[0163] The decision suggestion area is used to display the aircraft's glide path and decision suggestions in real time, and to issue a warning when the aircraft's trajectory deviates. The pilot controls the aircraft to correct the flight trajectory to the optimal glide path based on the flight decision suggestions. In the middle of the data visualization interface, the optimal glide path is displayed as a virtual line, and the icon of the aircraft symbol moves along the optimal glide path. When the aircraft deviates from the optimal glide path, the line turns red and flashes. The decision suggestions respectively represent the target speed, target altitude, target acceleration, target roll and target pitch in the form of scales on the airspeed indicator, altimeter, accelerometer, roll scale and pitch attitude scale.

[0164] A manned aircraft landing auxiliary decision method of this embodiment is implemented by using a manned aircraft landing auxiliary decision system, such as Figure 5 and Figure 6 As shown, the method comprises the following steps:

[0165] Obtain the current status data of the aircraft in real time, including: the current altitude of the aircraft, the current three-dimensional position of the aircraft, weather information and the current posture of the aircraft.

[0166] The method for obtaining the current altitude of the aircraft is: using an air pressure sensor to collect the atmospheric pressure of the aircraft at the current altitude in real time, and calculating the altitude of the aircraft h according to the atmospheric pressure of the aircraft at the current altitude 1 .

[0167]

[0168] Where T b Indicates height H b The atmospheric temperature at that time; β is the vertical temperature change rate; Indicates the height is h 1 Atmospheric pressure at that time; P b Indicates height H b is the atmospheric pressure at that time; R is the gas number for air; g is the acceleration due to gravity.

[0169] The laser sensor is used to emit laser pulses towards the ground. After being reflected by the ground, part of the laser pulses is received by the laser sensor. The time from the emission to the reception of the laser pulses is recorded, and the height of the aircraft h is calculated based on the recorded time. 2 .

[0170] The ultrasonic sensor transmits ultrasonic waves to the ground and receives the returned ultrasonic waves, records the time difference of the ultrasonic waves and calculates the height h of the aircraft. 3 .

[0171]

[0172] Where c represents the propagation speed of sound waves in the air, which is generally 340m / s; t 3 Indicates the time difference of the ultrasonic wave's round trip.

[0173] The weighted average algorithm is used to calculate the aircraft's altitude h 1 、h 2 and h 3 and takes the average value as the current height h of the aircraft.

[0174]

[0175] where k 1 is the height h 1 The weight coefficient of k 2 is the height h 2 The weight coefficient of k 3 is the height h 3 The weight coefficient is used to indicate the proportion of the data behind the weight coefficient in the entire statistical data, and also to indicate the degree of influence of this data on the statistical results.

[0176] In this embodiment, the air pressure sensor can indirectly obtain the altitude of the aircraft by measuring the atmospheric pressure at the altitude of the aircraft. The laser sensor is composed of a laser transmitter, a laser receiver and an avalanche photodiode. The height measurement principle of the laser sensor is as follows: the laser transmitter emits a laser pulse toward the ground, and the laser is partially received by the laser receiver after being reflected by the ground, imaged on the avalanche photodiode and amplified, so that the weak laser signal can be detected, and the time from the emission of the laser pulse to the return reception is recorded and processed, so that the flight altitude of the aircraft can be determined. The ultrasonic sensor is mainly composed of a transmitter part, a receiver part and a control part. The height measurement principle of the ultrasonic sensor is to emit ultrasonic waves to the ground through an ultrasonic transmitter, and calculate the height of the aircraft from the ground based on the time difference of the ultrasonic wave.

[0177] The current three-dimensional position of the aircraft includes: longitude, latitude and current altitude of the aircraft.

[0178] In this embodiment, the current three-dimensional position of the aircraft includes directly acquired longitude and latitude, and also includes the calculated current altitude of the aircraft.

[0179] The meteorological information includes: runway wind direction, runway wind speed, atmospheric temperature and atmospheric humidity.

[0180] The current posture of the aircraft includes: current pitch angle, current roll angle, current heading angle and runway position.

[0181] According to the current state data of the aircraft, the exponential flattening trajectory method is used to plan the landing trajectory of the aircraft and determine the optimal glide path of the aircraft.

[0182] In this embodiment, if Figure 7 As shown in the figure, the entire landing process of a fixed-wing manned aircraft includes four stages: approach flight, straight descent, leveling and falling, and ground taxiing. The control purpose of the entire process is to create a suitable aircraft attitude and sinking rate conditions for the moment when the aircraft's main wheels touch the ground. The design of the landing trajectory is a reverse design process, which divides the landing trajectory into two parts: the descent section and the leveling section. The aircraft's descent decision height H 1 and leveling decision height H 2 It is usually predetermined based on factors such as aircraft performance, runway conditions, and weather conditions. First, the trajectory of the leveling section is designed, mainly based on the landing weight of the aircraft to determine the speed, touchdown attitude angle, and touchdown vertical speed at touchdown, and then calculate the trajectory inclination of the fixed-wing manned aircraft during landing, and determine the leveling height based on the flight distance of the terminal leveling section. Secondly, based on the leveling section design, the glide section trajectory is designed, mainly based on the requirements of the glide trajectory angle, glide speed, and lifting speed, to determine the glide trajectory angle and glide speed, and then determine the glide decision height based on the glide flight distance, and then determine the distance between the glide decision point and the landing point.

[0183] In this embodiment, the trajectory design of the terminal leveling section uses the method of exponential leveling trajectory. The design idea of ​​the leveling trajectory is to reduce the vertical speed of the aircraft accordingly as the altitude decreases, so that the aircraft's descent speed at each moment is proportional to its current altitude, and mainly adopts the vertical speed control. Ideally, when the speed is 0 during descent, the altitude is also exactly 0, that is, the following homogeneous differential equation is satisfied:

[0184]

[0185] in represents the vertical speed of the aircraft at time t; t represents time; a is a constant; H(t) represents the altitude of the aircraft at time t; H 2 is the flattening decision height; τ is the time constant of the exponential flattening curve.

[0186] When time t reaches the moment when the aircraft trajectory just flattens, by solving the above differential equation, the height of the aircraft at the flattening moment is expressed as:

[0187]

[0188] Where e is a natural constant.

[0189] From expression (4), we can see that when the runway plane is used as the asymptotic plane for leveling, the aircraft can only touch the ground when time t tends to infinity, and the descent speed is 0, which is unrealistic. Considering that the aircraft can land within the allowable safe touchdown speed, expression (4) can be rewritten as expression (6), and the equation for the change of aircraft height with time when the aircraft is in the leveling stage is established:

[0190]

[0191] in Indicates the vertical speed at which an aircraft can land safely.

[0192] The solution of equation (6) is:

[0193]

[0194] When the aircraft's altitude drops to 0, the aircraft's leveling time t 1 It is expressed as:

[0195]

[0196] Assuming that the flight speed of the aircraft in the leveling section is constant, the horizontal distance flown by the aircraft in the leveling section is:

[0197]

[0198] Where L f V is the horizontal distance the aircraft flies in the leveling phase; 0 Represents the average speed of the aircraft during the leveling phase; θ 1 Indicates the flattening angle.

[0199] Formula (9) shows that when L f →∞, the leveling ends, which is obviously not allowed. When the aircraft lands, due to the ground resistance brought by the ground runway, the aircraft continues to sail uphill without the action of other forces after landing, so that the horizontal speed of the aircraft is 0. Considering the ground resistance of the aircraft when landing, the runway plane is set to be higher than the asymptote of the leveling trajectory during the landing process to simulate the ground resistance of the aircraft when landing. At this time, the equation of the change of aircraft height over time is rewritten as:

[0200]

[0201] Where h(t) represents the landing altitude of the aircraft at time t; is the specified landing speed of the aircraft, and h jd is the distance to the asymptote of the flattened trajectory; And the time constant τ, we can determine the asymptotic distance h of the flattened trajectory jd .

[0202] Therefore, when designing the leveling trajectory, first determine the landing speed based on the landing weight of the drone. Then determine h jd , by flattening the distance L 1 +L 2 and the flare trajectory angle, the flare decision altitude can be obtained.

[0203] When designing the glide path, the key is to select the glide path angle, which will affect whether the glide speed can be maintained through attitude control and altitude control. The glide speed is based on the speed change during the terminal leveling process plus the required landing speed of the aircraft.

[0204] According to the glide angle, leveling angle, glide height and leveling decision height, combined with Figure 7 The landing image shown, the glide distance and flare distance can be calculated by the following formula:

[0205]

[0206] Where L 3 Indicates the horizontal distance of the downslope; L 2 Indicates the horizontal distance of the first half of the leveling; L 1 Indicates the horizontal distance of the second half of the leveling; θ 2 is the glide angle.

[0207] According to the trajectory design of the leveling section and the glide section, the landing trajectory of the aircraft is planned, which is expressed as:

[0208]

[0209] Among them, H D Indicates the altitude of the aircraft during landing; L indicates the horizontal distance flown by the aircraft during landing; V indicates the approach speed of the aircraft; t 0 Indicates the entry time.

[0210] The method of planning the landing trajectory of the aircraft using the exponential flattening trajectory method is as follows: the landing trajectory of the aircraft is divided into a flattening section and a glide section, and the flattening angle θ of the aircraft when landing is obtained. 1 , glide angle θ 2 , flatten the trajectory asymptote distance h jd , the aircraft’s approach speed V and the aircraft’s approach time t 0 ;

[0211] First, plan the landing trajectory of the aircraft in the leveling phase. Specifically, according to formula (9), use the leveling angle θ when the aircraft lands1 Calculate the horizontal distance L that the aircraft flies during the leveling phase f The horizontal distance of the aircraft in the leveling section includes two parts: the horizontal distance L in the first half of the leveling section 2 The horizontal distance L from the second half of the pull 1 ; According to formula (11), solve the leveling decision height H 2 , and then calculate the horizontal distance L of the first half of the leveling 2 The horizontal distance L from the second half of the pull 1 , as shown in formula (12) and formula (13).

[0212] Then, the landing trajectory of the aircraft in the descent phase is planned. Specifically, since the end of the descent phase is connected to the beginning of the leveling phase, the speed of the aircraft at the end of the descent phase is determined according to the landing trajectory of the aircraft in the leveling phase, and the landing speed V and the descent angle θ are calculated according to the landing speed V and the descent angle θ. 2 The aircraft's descent time is calculated by the aircraft's speed at the end of the descent phase, and then the aircraft's descent decision altitude H is calculated. 1 , using the aircraft's descent to determine the altitude H 1 Calculate the horizontal distance L of the aircraft during the glide path 1 , that is, formula (14).

[0213] Finally, the optimal glide path of the aircraft is determined according to the landing trajectory of the aircraft in the leveling phase and the landing trajectory of the aircraft in the glide phase, as shown in formula (15).

[0214] Based on the current state data of the aircraft and the optimal glide path of the aircraft, a command decision algorithm is used to calculate the expected attitude angle and the expected thrust change of the aircraft, and the aircraft is attitude controlled according to the expected attitude angle of the aircraft, and a flight decision suggestion is generated; wherein the attitude control is divided into two parts: one part is to calculate the expected three-dimensional angle and three-axis attitude angular velocity by angle control; the other part is to calculate the three-axis attitude angular acceleration by angular velocity control; the flight decision suggestion includes: the expected thrust change, the expected three-dimensional angle, the three-axis attitude angular velocity and the three-axis attitude angular acceleration.

[0215] The desired attitude angle includes: a desired roll angle and a desired pitch angle.

[0216] The instruction decision algorithm includes: lateral instructions and longitudinal instructions.

[0217] The lateral instruction is: using the L1 guidance law, calculating the desired roll angle of the aircraft according to the current position of the aircraft and the optimal glide path of the aircraft.

[0218] In this embodiment, the lateral command usually uses the L1 guidance law to control the lateral flight of the aircraft, that is, to control the roll angle of the aircraft. The basic principle of the L1 guidance law is to select a reference point on the aircraft's glide path, and call the distance between the reference point and the current horizontal position of the aircraft L, and calculate the horizontal expected acceleration of the aircraft based on L and the current horizontal speed of the aircraft.

[0219] The method for calculating the desired roll angle is as follows: select a reference point P on the aircraft's optimal glide path according to the aircraft's current position and flight speed. R , connect the reference point and the current position of the aircraft, and record the length of the connection as L 1 , the angle between the current speed direction of the aircraft and the connecting line is recorded as η, according to the current position of the aircraft and the reference point P R The centripetal acceleration of the aircraft is calculated based on the geometric relationship between them, which can be divided into the following two cases:

[0220] The first case: When the reference point is located in the arc section of the optimal glide path, take 2η as the center angle, and use the current position of the aircraft and the reference point P R Construct an arc and then calculate the centripetal acceleration of the aircraft based on the geometric relationship. The expression is:

[0221]

[0222] where a Scmd is the centripetal acceleration of the aircraft; v is the horizontal speed of the aircraft; R 1 is the arc radius.

[0223] In this embodiment, if Figure 8 As shown, the aircraft's gliding arc trajectory is l R , the length of the line between the reference point and the current position of the aircraft is L, the angle between the current speed of the aircraft and the line between the reference point and the current position of the aircraft is η, and from the geometric relationship, it can be seen that the angle corresponding to the arc is 2η.

[0224] The second case: When the reference point is located on the straight line segment of the optimal glide path, the aircraft's glide path is l R , the current speed direction of the aircraft and the glide path l R The angle between 1 , the glide path l R The angle between the reference point and the current position of the aircraft is denoted as η 2 , and then calculate the centripetal acceleration of the aircraft based on the geometric relationship, the expression is:

[0225]

[0226] Where d represents the vertical distance between the aircraft's current position and the glide path; It is the rate of change of the vertical distance between the aircraft's current position and the glide path.

[0227] In this embodiment, if Fig. 9 As shown, the straight line trajectory of the aircraft is l R , the length of the line between the reference point and the current position of the aircraft is L 1 At this time, the angle between the current speed of the aircraft and the glide path is η 1 The angle between the glide path and the line connecting the reference point and the current aircraft position is η 2 According to the geometric relationship, assuming that η 1 and η 2 The angles are very small, then sin(η 1 +η 2 )≈η 1 +η 2 .

[0228] Considering η 2 In rare cases, there are:

[0229]

[0230] in It is the rate of change of the aircraft's vertical speed from its current position to the glide path.

[0231] In this embodiment, the vertical distance d between the current position of the aircraft and the glide path is differentiated to obtain the first-order distance change rate, that is, Then we get the second-order velocity change rate, that is

[0232] The expected roll angle of the aircraft is calculated based on the centripetal acceleration of the aircraft. The expression is:

[0233]

[0234] Where m is the weight of the aircraft; g is the acceleration due to gravity; φ sp is the desired roll angle.

[0235] In this embodiment, if Fig.10 As shown, the lateral force analysis of the aircraft is performed, and the expected roll angle of the aircraft is calculated based on the force analysis results, as shown in expression (19).

[0236] The longitudinal instruction is: based on the aircraft's track angle, the aircraft's current actual airspeed, the aircraft's current acceleration and the optimal glide slope in the current flight state, the thrust change control law and the elevator control law are respectively constructed using the TECS control method, and the thrust change control law and the elevator control law are used to solve the aircraft's expected thrust change and the aircraft's expected pitch angle.

[0237] In this embodiment, the longitudinal command uses a TECS controller, i.e., a total energy controller, to control the flight speed and altitude. The elevator and throttle lever of a fixed-wing aircraft have a coupling characteristic in their response to the aircraft, i.e., the elevator swing and throttle change will affect the pitch angle and airspeed of the aircraft, especially in the low dynamic pressure flight state of the aircraft, the coupling of flight speed and track is more intense. The TECS controller controls from the perspective of energy, dividing the total energy of the aircraft into kinetic energy and potential energy, wherein the kinetic energy depends on the flight speed and the potential energy depends on the flight altitude. The change in the total energy of the aircraft is mainly controlled by the change in the thrust of the aircraft engine. The desired throttle amount is calculated by the total energy demand, and the kinetic energy and potential energy are balanced by adjusting the pitch angle, thereby achieving the decoupling of the throttle and the elevator.

[0238] In this embodiment, in the process of controlling the flight speed and altitude using the TECS controller, first, the motion equation of the aircraft longitudinal plane is constructed:

[0239]

[0240] in is the forward position derivative; V n is the current actual airspeed of the aircraft; γ is the track angle; V w is the wind speed; is the height derivative, i.e. the rate of change of height; is the current acceleration of the aircraft; T is the engine thrust; D is the air resistance;

[0241] Since the general track angle is relatively small, there is:

[0242] sinγ=γ (21)

[0243] Then the thrust during the flight is calculated as:

[0244]

[0245] In this embodiment, it is considered that the engine pulling force is completely equal to the thrust during the flight process.

[0246] The total energy of the aircraft is calculated based on its current airspeed and altitude:

[0247]

[0248] Where E represents the total energy of the aircraft; E k Indicates the kinetic energy of the aircraft; E p represents the gravitational potential energy of the aircraft; H represents the current altitude of the aircraft.

[0249] Define the total energy per unit mass E 1 for:

[0250]

[0251] The total rate of change of energy per unit mass is for:

[0252]

[0253] According to the equation of motion of the aircraft in the longitudinal plane, substituting the height differential:

[0254]

[0255] It can be seen that the total energy change of the aircraft is mainly controlled by the thrust change of the engine, and the thrust change has the following relationship:

[0256]

[0257] Where ΔT represents the change in thrust; Represents the acceleration feedback error; γ e Represents the track angle feedback error;

[0258] Then we have:

[0259]

[0260] Where ∝ indicates a proportional relationship; Represents the total energy rate of change deviation of the aircraft.

[0261] Therefore, when the aircraft thrust changes, the total energy change rate of the aircraft also changes at a certain ratio, so that the sum of the aircraft's acceleration and track angle changes. Therefore, the aircraft thrust can be used as the main control variable for the aircraft's total energy change. Based on this, the control law for controlling the thrust change by controlling the aircraft's total energy change can be obtained:

[0262]

[0263] Where T c is the expected thrust change; s is the integral variable; K tp is the proportional coefficient of thrust control; K ti is the integral coefficient of thrust control; integral control is used to eliminate the steady-state error of tracking total energy changes.

[0264] The deflection of the elevator can mainly cause the change of the pitch moment of the aircraft, change the flight attitude of the aircraft, and have little effect on the thrust and drag. When the thrust remains unchanged, the kinetic energy and potential energy of the aircraft are converted by manipulating the elevator. Therefore, the elevator can be used as a distribution controller for the kinetic energy and potential energy of the aircraft. The energy distribution rate is used as the control quantity in the elevator control channel, and is also defined as the difference between the rate of change of potential energy and kinetic energy:

[0265]

[0266] in It is the difference between the rate of change of potential energy and kinetic energy.

[0267] Therefore, for the control of the elevator, the elevator control law is constructed and expressed as:

[0268]

[0269] where δ ec is the desired elevator change; K is the energy distribution rate deviation of the aircraft; ep K is the proportional coefficient of elevator control; ei is the integral coefficient of elevator control.

[0270] To prevent the control response overshoot caused by the unexpected zero point in the control process, the proportional part is directly replaced by the rate of change, and the thrust change control law and elevator control law of the TECS controller are constructed, which can be expressed as:

[0271]

[0272] in is the total energy change rate feedback of the aircraft; It is the energy distribution rate feedback of the aircraft.

[0273] like Fig.11 As shown, the control response of the TECS controller tends to be smooth. In the core algorithm of total energy control, the control quantity is formed, that is, and The input of track angle control, i.e. the desired track angle γ, is introduced into the core algorithm of the TECS controller. c , the output of track angle control, that is, track angle γ, the input of acceleration control, that is, acceleration control amount The output of acceleration control, i.e. acceleration Based on the above, the target instructions for throttle and pitch angle are finally obtained, that is, the expected throttle and expected pitch angle of the aircraft. The expected throttle of the aircraft is represented by the expected thrust change.

[0274] The method for calculating the desired pitch angle is as follows: obtaining the flight path angle γ of the aircraft in the current flight state, the current actual airspeed V of the aircraft n and the aircraft's current acceleration And determine the expected flight path angle γ of the aircraft based on the aircraft's optimal glide path c and the desired airspeed, and use the desired flight angle of the aircraft to calculate the flight angle feedback error γ e , using the aircraft's airspeed and desired airspeed to calculate the acceleration control amount And according to the current acceleration of the aircraft Calculate the acceleration feedback error The track angle feedback error γ e and acceleration feedback error As shown in formula (43).

[0275] Based on the track angle feedback error γ e and acceleration feedback error The thrust variation control law and elevator control law of the TECS controller are constructed as shown in formula (32).

[0276] The desired thrust change T is calculated using the thrust change control law and elevator control law of the TECS controller. c and the desired elevator change δ ec , assuming that the change in the aircraft's angle of attack is negligible, the expected elevator change is directly used as the expected pitch angle.

[0277] like Fig.11 As shown, the total energy change rate deviation of the aircraft By feedbacking the track angle error γ e and acceleration feedback error The total energy change rate feedback of the aircraft is calculated by summing By calculating the aircraft's track angle γ and the aircraft's current acceleration The sum is used to solve; the aircraft's energy distribution rate feedback By calculating the aircraft's track angle γ and the aircraft's current acceleration The energy distribution rate deviation of the aircraft is calculated by difference. By feedbacking the track angle error γ e and acceleration feedback error Solve by taking the difference.

[0278] In this embodiment, the control process of the TECS controller is divided into two parts: the inner loop design of the TECS controller and the outer loop design of the TECS controller. Specifically, in the core algorithm of the TECS controller, the elevator control is the main control amount of the energy distribution channel. Through simulation and experiments, it is found that the process of using the TECS controller alone will produce an undesirable dynamic response of the short-period pitch mode. Therefore, it is necessary to introduce a pitch angle control inner loop in the TECS controller to increase damping to counteract the pitch angle oscillation generated by the TECS system.

[0279] The TECS controller must provide a pitch angle control variable θ c , that is, the desired pitch angle. Considering that the TECS controller changes the aircraft's flight path with the elevator while the thrust remains unchanged, the elevator control law is established as:

[0280]

[0281] Since the thrust remains unchanged at this time, we have:

[0282]

[0283] in Indicates the rate of change of the aircraft's speed; Indicates the desired rate of change of aircraft speed.

[0284] Therefore, when the thrust remains unchanged, the energy distribution rate feedback of the aircraft and the energy distribution rate deviation of the aircraft are expressed as:

[0285]

[0286] Then, formula (33) is rewritten as:

[0287]

[0288] In the case of steady-state control, due to the effect of integration, γ e will tend to zero, at which point the flight path is directly related to the elevator deflection angle, that is, the expected elevator change δ ec Proportional, that is:

[0289] δ ec =K ep γ (38)

[0290] Assuming that the aircraft angle of attack does not change much, the aircraft's elevator control amount is directly used as the control input of the pitch angle, that is, the desired pitch angle, so as to obtain effective control of the track angle. In order to complete the conversion from elevator to pitch angle control input, the proportional coefficient K of the elevator control is taken ep is 1, we get:

[0291] θ c =γ c (39)

[0292] After converting the TECS control input from the elevator to the pitch angle control input, in order to avoid the adverse dynamic response of the short-period pitch mode, it is necessary to design the pitch angle control inner loop. Simply introducing the pitch angle signal into the inner loop is not enough to reduce the pitch oscillation, so it is necessary to introduce the pitch rate signal to increase the pitch damping.

[0293] The aircraft pitch angle inner loop, i.e., attitude loop control law in the TECS controller inner loop design is expressed as:

[0294]

[0295] where δe represents the attitude loop control law; K θ is the feedback gain of the pitch angle feedback channel; k q is the feedback gain of the pitch angle rate feedback channel; θ is the actual pitch angle; is the pitch angle change rate; θ e is the difference between the expected pitch angle and the actual pitch angle.

[0296] Considering the TECS core control law, the pitch angle control input in equation (40) is expressed as:

[0297]

[0298] Converting thrust to the position of the throttle lever or the percentage of throttle lever rotation requires a certain conversion relationship, that is, a proportional relationship, which is also a problem that needs to be considered in the design of the inner loop of the TECS controller. Regarding the design of the inner loop of the TECS control system, each aircraft is different. The aircraft inner loop design is independent of the overall concept of total energy control, because total energy control mainly studies the flight performance of the aircraft in the long period, and the aircraft inner loop control is to improve the aircraft's short-term flight performance. Generally, aircraft need to control their inner loop. There are many methods of inner loop control, and the use of these methods is independent of the design of the TECS outer loop control parameters.

[0299] In this embodiment, based on the structure, core algorithm and inner loop design of the TECS controller, the outer loop design of the TECS controller should satisfy the requirements for the track angle feedback γ and the acceleration feedback The control input and feedback input of the outer loop are respectively the track angle control quantity γ c , acceleration control amount Track angle feedback γ and acceleration feedback The output of the outer loop is the thrust required for aircraft control and the pitch angle control quantity of the pitch angle control loop.

[0300] For the flight control system of an aircraft, the track angle feedback γ cannot generally be obtained directly. It can be calculated through the feedback of the aircraft's pitch angle θ and angle of attack α. Without considering sideslip, the solution formula is as follows:

[0301] γ=θ-α(42)

[0302] Similarly, the acceleration feedback It cannot be obtained directly, but can be obtained by taking the differential or difference of the flight speed feedback

[0303] In order to form the core algorithm control law of the TECS controller, it is necessary to introduce the error signal between the control quantity and the feedback quantity for integral control, where the acceleration control quantity error The error of the track angle feedback γ e They are:

[0304]

[0305] The method of calculating the desired three-dimensional angle and three-axis attitude angular velocity by angle control comprises the following steps:

[0306] Get the angular velocity and acceleration of the aircraft in the current flight state, and calculate the actual roll angle and actual pitch angle of the aircraft in the current flight state.

[0307] According to the actual roll angle, the Kalman filter method is used to calculate the roll angle estimate According to the actual pitch angle, the Kalman filter method is used to calculate the estimated pitch angle

[0308] Calculate the expected roll angle φ sp and the estimated roll angle The difference between them is taken as the rolling angle error φ err , and then the roll angle error φ err Perform proportional control to obtain the desired rolling angular velocity

[0309] Calculate the desired pitch angle θ sp and the pitch angle estimate The difference between them is taken as the pitch angle error θ err , and then the pitch angle error θ err Perform proportional control to obtain the desired pitch angular velocity

[0310] According to the desired roll angle φ sp and the desired pitch angle θ sp , using the coordinated turning control law to calculate the desired yaw rate And according to the expected yaw rate Calculate the desired yaw angle ψ sp .

[0311] The coordinated turning control law is expressed as:

[0312]

[0313] Where V T Indicates true airspeed.

[0314] The desired roll angle φ sp , desired pitch angle θ sp and the desired yaw angle ψ sp as the desired 3D angle.

[0315] Using the expected roll velocity Expected roll angular velocity and the desired yaw rate Construct the three-axis attitude angle change rate, expressed as:

[0316]

[0317] in Indicates the rate of change of the three-axis attitude angle.

[0318] Then the three-axis attitude angle change rate is converted into a coordinate system to obtain the three-axis attitude angular velocity in the body coordinate system

[0319]

[0320] in Indicates the three-axis attitude angular velocity; p sp represents the expected rolling angular velocity; q sp represents the expected pitch angular velocity; γ sp represents the desired yaw rate; φ represents the roll angle.

[0321] In this embodiment, if Fig.12 As shown in the figure, the overall attitude controller design adopts a dual-loop cascade PID, with the outer loop control being angle control and the inner loop control being angular velocity control. The closed-loop control quantity needs to obtain the state quantity for feedback control, and the state quantity acquisition requires the use of a filtering estimation algorithm.

[0322] (1) Outer ring angle control

[0323] like Fig.13 As shown, the desired roll angle φ sp and the estimated roll angle The error φ is obtained by subtraction err , error φ err After the proportional control link, the desired rolling attitude angular velocity is obtained Similarly, the desired pitch angle θ sp and the pitch angle estimate The error θ is obtained by subtraction err , error θ err After the proportional control link, the desired pitch attitude angular velocity is obtained

[0324] Expected yaw rate The acquisition of adopts the coordinated turning control law, and its calculation is shown in formula (44). In summary, the desired attitude angular velocity three-axis vector, that is, the three-axis attitude angle change rate, can be obtained.

[0325] because It is the rate of change of the attitude angle, not the three-axis roll angular velocity in the body coordinate system. Considering that the object that directly generates torque from rudder deflection is the aircraft body, the reasonable intermediate control quantity here should be the three-axis roll angular velocity in the aircraft body coordinate system The origin of the body coordinate system is usually located at the center of mass of the aircraft, with the X-axis pointing to the nose of the aircraft and coinciding with the longitudinal axis of the fuselage; the Y-axis points to the right side of the aircraft; and the Z-axis is perpendicular to the plane of the fuselage and points upward, forming a right-handed coordinate system.

[0326] The method of calculating the three-axis attitude angular acceleration by using angular velocity control comprises the following steps:

[0327] The three-axis attitude angular velocity As the feedforward input of angular velocity control, and through feedforward control calculation feedforward control three-axis attitude angular acceleration

[0328] Get the actual value of the three-axis attitude angular velocity It is expressed as:

[0329]

[0330] in Indicates the actual value of the rolling angular velocity; Indicates the actual value of the pitch angular velocity; Indicates the actual value of the track angular velocity.

[0331] Using three-axis attitude angular velocity And the actual value of the three-axis attitude angular velocity Calculate the angular velocity error of the three-axis attitude It is expressed as:

[0332]

[0333] where p err represents the rolling angular velocity error; q err represents the pitch angular velocity error; γ err Represents the track angular velocity error.

[0334] The three-axis attitude angular velocity error Input PI controller to get PI controlled angular acceleration

[0335] Using feedforward to control the three-axis attitude angular acceleration and PI control angular acceleration Calculate the three-axis attitude angular acceleration It is expressed as:

[0336]

[0337] Among them FF Scaleris the feedforward control airspeed scaling factor; PI Scaler is the PI control airspeed scaling factor; and:

[0338]

[0339] Where V T is the true airspeed; V I is the indicated airspeed; and They are the set true airspeed and indicated airspeed respectively.

[0340] (2) Inner ring angular velocity control

[0341] like Fig.14 As shown, in order to obtain the three-axis attitude angular acceleration Here, the expected value feedforward FF and error value PI control are used, specifically:

[0342] Will As the feedforward input, the feedforward control three-axis attitude angular acceleration is obtained

[0343] Will and Subtract and get the three-axis attitude angular velocity error The three-axis attitude angular velocity error As the PI controller input, the PI controlled angular acceleration is obtained

[0344] Using feedforward to control the three-axis attitude angular acceleration and PI control angular acceleration Calculate the three-axis attitude angular acceleration:

[0345]

[0346] However, the same rudder deflection will produce different torques in high-speed and low-speed flight. Therefore, based on this link, airspeed scaling is added, and the feedforward control output and PI control output are multiplied by a scaler coefficient, as shown in formulas (50) and (51), where V T is the true airspeed; V I is the indicated airspeed; here V T and V I These are the values ​​obtained after sensor data processing. and The true airspeed and indicated airspeed are set for trim conditions, which refer to a set of parameters in a stable flight state that ensures the aircraft maintains the current flight state without additional control input. and These are all set parameter values, generally set to the cruising speed or the median value between the minimum flight speed and the maximum flight speed.

[0347] The final three-axis attitude angular acceleration is shown in formula (49).

[0348] In summary, after a series of calculations by the command decision algorithm, the expected thrust change, expected three-dimensional angle, three-axis attitude angular velocity and three-axis attitude angular acceleration in the next stage of the fixed-wing aircraft can be obtained and displayed to the pilot as a flight decision recommendation.

[0349] The pilot controls the aircraft's flight trajectory based on the generated flight decision suggestions to ensure that the aircraft lands along the aircraft's optimal glide path.

[0350] In this embodiment, the pilot adjusts the current state data of the aircraft through manual control operations according to the generated flight decision suggestions, thereby adjusting the flight trajectory of the aircraft to ensure that the aircraft can land according to the optimal glide path.

[0351] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some or all of the technical features therein. However, these modifications or replacements do not cause the essence of the corresponding technical solutions to deviate from the scope defined by the claims of the present invention.

Claims

1. A manned aircraft landing auxiliary decision system, characterized in that: The system includes: aircraft status and environmental information perception module, landing trajectory planning and command decision module and data interaction module; The aircraft status and environmental information perception module is used to monitor the aircraft's attitude and environmental data in real time and perform fusion collection to obtain aircraft status information including aircraft altitude information, aircraft three-dimensional position information, meteorological information and aircraft posture information, and send it to the landing trajectory planning and command decision module and the data interaction module; The landing trajectory planning and command decision module is used to plan the landing trajectory of the aircraft according to the received aircraft status and environmental information by using the exponential flattening trajectory method to determine the optimal glide path of the aircraft; use the command decision algorithm to calculate the expected attitude angle of the aircraft and perform attitude control on the aircraft to determine the expected attitude angular velocity and expected angular acceleration; generate flight decision suggestions, and transmit the optimal glide path and flight decision suggestions of the aircraft to the data interaction module; The data interaction module is used to display the aircraft status information, the optimal glide path and the flight decision suggestions to the pilot in real time through the data visualization interface, so that the pilot can control the aircraft to land along the optimal glide path according to the flight decision suggestions.

2. The manned aircraft landing decision support system according to claim 1, characterized in that: The aircraft status and environmental information perception module includes: a location and weather module, a machine vision module and a data acquisition module; The position and weather module is used to obtain the aircraft's altitude information, the aircraft's three-dimensional position information and weather information in real time, and send them to the data acquisition module; The location and weather module includes: an air pressure sensor module, a laser sensor module, an ultrasonic sensor module, a global positioning system module, an atmospheric data system module and a microcontroller; The air pressure sensor module is used to calculate the altitude of the aircraft by measuring the atmospheric pressure at the altitude of the aircraft and send the information to the microcontroller; The laser sensor module is used to emit laser pulses to the ground and receive reflected lasers from the ground, and record the time from the emission of the laser pulse to the reception of the laser pulse, calculate the altitude of the aircraft based on the recorded time and send it to the microcontroller; The ultrasonic sensor module is used to transmit ultrasonic waves to the ground and receive returned ultrasonic waves, and record the round-trip time difference of the ultrasonic waves, calculate the height of the aircraft according to the recorded round-trip time difference and send it to the microcontroller; The global positioning system module is used to obtain the three-dimensional position information of the aircraft and send it to the microcontroller; the three-dimensional position information of the aircraft includes: longitude and latitude; The atmospheric data system module is used to obtain meteorological information and send it to the microcontroller; wherein the meteorological information includes: runway wind direction, runway wind speed, atmospheric temperature and atmospheric humidity; The microcontroller is used to use a weighted average algorithm to average the heights of the aircraft calculated by the air pressure sensor module, the laser sensor module and the ultrasonic sensor module to obtain the height information of the aircraft; and send the height information of the aircraft, the three-dimensional position information of the aircraft and the meteorological information to the data acquisition module; The machine vision module is used to obtain the position and posture information of the aircraft in real time; wherein the position and posture information of the aircraft includes: the pitch angle of the aircraft, the roll angle of the aircraft, the heading angle of the aircraft and the runway position; The machine vision module comprises: a CCD camera, an image acquisition card and a visual processing module; wherein the CCD camera is connected to an external signal interface of the image acquisition card; The CCD camera is used to capture the ground sign image of the aircraft landing point in real time; The image acquisition card is used to transmit the ground sign image captured by the CCD camera to the visual processing module; The visual processing module is used to perform image preprocessing on the received ground mark image, perform feature extraction and feature matching on the processed ground mark image, and estimate the position and posture of the aircraft according to the matching result to obtain the position and posture information of the aircraft; The data acquisition module is used to integrate the received aircraft altitude information, aircraft three-dimensional position information, weather information and aircraft posture information, and send the integrated data information to the landing trajectory planning and command decision module.

3. The manned aircraft landing decision support system according to claim 2, characterized in that: The data interaction module is implemented through a data visualization interface; the data visualization interface is divided into three display areas, namely: an aircraft status information area, an environmental information area, and a decision suggestion area; The aircraft status information area is used to display the current altitude, current speed, current acceleration, current pitch angle and current roll angle of the aircraft in real time; The environmental information area is used to display the runway information and weather information during the aircraft landing process in real time; The decision suggestion area is used to display the aircraft's glide path and decision suggestions in real time, and to issue a warning when the aircraft's trajectory deviates. The pilot controls the aircraft to correct the flight trajectory to the optimal glide path according to the flight decision suggestions.

4. A manned aircraft landing auxiliary decision method, implemented by using a manned aircraft landing auxiliary decision system as claimed in claim 1, characterized in that: The method comprises the following steps: Acquire the current status data of the aircraft in real time, including: the current altitude of the aircraft, the current three-dimensional position of the aircraft, weather information and the current posture of the aircraft; wherein the weather information includes: runway wind direction, runway wind speed, atmospheric temperature and atmospheric humidity; the current posture of the aircraft includes: current pitch angle, current roll angle, current heading angle and runway position; According to the current state data of the aircraft, the exponential flattening trajectory method is used to plan the landing trajectory of the aircraft and determine the optimal glide path of the aircraft; Based on the current state data of the aircraft and the optimal glide path of the aircraft, a command decision algorithm is used to calculate the expected attitude angle and the expected thrust change of the aircraft, and the aircraft is attitude controlled according to the expected attitude angle of the aircraft, and a flight decision suggestion is generated; wherein the attitude control is divided into two parts: one part is to calculate the expected three-dimensional angle and three-axis attitude angular velocity by angle control; the other part is to calculate the three-axis attitude angular acceleration by angular velocity control; the flight decision suggestion includes: the expected thrust change, the expected three-dimensional angle, the three-axis attitude angular velocity and the three-axis attitude angular acceleration; The desired attitude angle includes: a desired roll angle and a desired pitch angle; the command decision algorithm includes: a lateral command and a longitudinal command; The pilot controls the aircraft's flight trajectory based on the generated flight decision suggestions to ensure that the aircraft lands along the aircraft's optimal glide path.

5. The method for assisting decision-making in landing of a manned aircraft according to claim 4, characterized in that: The method for obtaining the current altitude of the aircraft is: using an air pressure sensor to collect the atmospheric pressure of the aircraft at the current altitude in real time, and calculating the altitude h1 of the aircraft according to the atmospheric pressure of the aircraft at the current altitude; Where T b Indicates height H b The atmospheric temperature at that time; β is the vertical temperature change rate; Indicates the atmospheric pressure at an altitude of h1; P b Indicates height H b Atmospheric pressure at 1000 s; R is the air specific gas number; g is the gravitational acceleration; A laser sensor is used to emit a laser pulse toward the ground. After being reflected by the ground, a portion of the laser pulse is received by the laser sensor. The time from the emission to the reception of the laser pulse is recorded, and the height h2 of the aircraft is calculated based on the recorded time. Use the ultrasonic sensor to transmit ultrasonic waves to the ground and receive the returned ultrasonic waves, record the time difference between the ultrasonic waves and the return trip, and calculate the height h3 of the aircraft; Where c represents the propagation speed of sound waves in the air; t3 represents the time difference of the ultrasonic wave's round trip; The weighted average algorithm is used to calculate the average of the aircraft's heights h1, h2 and h3, and the average is used as the aircraft's current height h; Among them, k1 is the weight coefficient of height h1; k2 is the weight coefficient of height h2; k3 is the weight coefficient of height h3.

6. The manned aircraft landing auxiliary decision method according to claim 5, characterized in that: The method for planning the landing trajectory of an aircraft using the exponential flattening trajectory method is as follows: the landing trajectory of the aircraft is divided into a flattening section and a gliding section, and the flattening angle θ1, the gliding angle θ2, and the flattening trajectory asymptote distance h are obtained when the aircraft lands. jd , the aircraft's approach speed V and the aircraft's approach time t0; First, the landing trajectory of the aircraft in the leveling section is planned, specifically: the horizontal distance of the aircraft in the leveling section is calculated using the leveling angle θ1 of the aircraft during landing. The horizontal distance of the aircraft in the leveling section includes two parts: the horizontal distance L2 of the first half of the leveling and the horizontal distance L1 of the second half of the leveling; The horizontal distance flown by the aircraft in the leveling section is expressed as: Where L f It indicates the horizontal distance that the aircraft flies in the leveling phase, and L f =L1+L2; V0 represents the average speed of the aircraft during the leveling phase; t1 represents the leveling time of the aircraft during the leveling phase; τ represents the time constant of the exponential leveling curve; H2 represents the leveling decision altitude; Indicates the vertical speed at which the aircraft can land safely; The horizontal distance L that the aircraft flies in the leveling section f The relationship between the leveling decision height H2 is: Then, the horizontal distance L2 of the first half of the leveling and the horizontal distance L1 of the second half of the leveling are calculated respectively, which can be expressed as: Then, the landing trajectory of the aircraft in the glide phase is planned. Specifically, since the end of the glide phase is connected to the beginning of the leveling phase, the speed of the aircraft at the end of the glide phase is determined according to the landing trajectory of the aircraft in the leveling phase, and the aircraft's glide time is calculated according to the aircraft's approach speed V, the glide angle θ2 and the aircraft's speed at the end of the glide phase, and then the aircraft's glide decision height H1 is calculated. The aircraft's horizontal distance L3 in the glide phase is calculated using the aircraft's glide decision height H1, which is expressed as: Finally, the optimal glide path of the aircraft is determined according to the landing trajectory of the aircraft in the leveling phase and the landing trajectory of the aircraft in the glide phase, which is expressed as: Among them, H D It represents the altitude of the aircraft during landing; L represents the horizontal distance of the aircraft during landing; t represents time; e is a natural constant.

7. The method for assisting decision-making in landing of a manned aircraft according to claim 6, characterized in that: The lateral instruction is: using the L1 guidance law, calculating the desired roll angle of the aircraft according to the current position of the aircraft and the optimal glide path of the aircraft; The method for calculating the desired roll angle is as follows: select a reference point P on the aircraft's optimal glide path according to the aircraft's current position and flight speed. R , connect the reference point and the current position of the aircraft, and record the length of the connection as L1, record the angle between the current speed direction of the aircraft and the connection as η, according to the current position of the aircraft and the reference point P R The centripetal acceleration of the aircraft is calculated based on the geometric relationship between them, which can be divided into the following two cases: The first case: When the reference point is located in the arc section of the optimal glide path, take 2η as the center angle, and use the current position of the aircraft and the reference point P R Construct an arc and then calculate the centripetal acceleration of the aircraft based on the geometric relationship. The expression is: where a Scmd is the centripetal acceleration of the aircraft; v is the horizontal speed of the aircraft; R1 is the radius of the arc; The second case: When the reference point is located on the straight line segment of the optimal glide path, the aircraft's glide path is l R , the current speed direction of the aircraft and the glide path l R The included angle is η1, and the glide path l R The angle between the reference point and the current position of the aircraft is recorded as η2, and then the centripetal acceleration of the aircraft is calculated based on the geometric relationship. The expression is: Where d represents the vertical distance between the aircraft's current position and the glide path; is the rate of change of the vertical distance between the aircraft's current position and the glide path; The expected roll angle of the aircraft is calculated based on the centripetal acceleration of the aircraft. The expression is: Where m is the weight of the aircraft; g is the acceleration due to gravity; φ sp is the desired roll angle.

8. The manned aircraft landing auxiliary decision method according to claim 7, characterized in that: The longitudinal instruction is: based on the aircraft's track angle, the aircraft's current actual airspeed, the aircraft's current acceleration and the optimal glide slope in the current flight state, the thrust change control law and the elevator control law are respectively constructed by the TECS control method, and the thrust change control law and the elevator control law are used to solve the aircraft's expected thrust change and the aircraft's expected pitch angle; The method for calculating the desired thrust change and the desired pitch angle is as follows: obtaining the flight path angle γ of the aircraft in the current flight state, the current actual airspeed V of the aircraft n and the aircraft's current acceleration And determine the expected flight path angle γ of the aircraft based on the aircraft's optimal glide path c and the desired airspeed, and use the desired flight angle of the aircraft to calculate the flight angle feedback error γ e , using the aircraft's airspeed and desired airspeed to calculate the acceleration control amount And according to the current acceleration of the aircraft Calculate the acceleration feedback error The track angle feedback error γ e and acceleration feedback error It is expressed as: Based on the track angle feedback error γ e and acceleration feedback error The thrust variation control law and elevator control law of the TECS controller are constructed and expressed as: Where T c is the expected thrust change; K tp is the proportional coefficient of thrust control; K is the total energy change rate feedback of the aircraft; ti The integral coefficient of thrust control; s is the integral variable; Indicates the total energy change rate deviation of the aircraft; δ ec is the desired elevator change; K ep is the proportional coefficient of elevator control; K is the energy distribution rate feedback of the aircraft; ei is the integral coefficient of elevator control; It is the energy distribution rate deviation of the aircraft; The desired thrust change T is calculated using the thrust change control law and elevator control law of the TECS controller. c and the desired elevator change δ ec , assuming that the change in the aircraft's angle of attack is negligible, the expected elevator change is directly used as the expected pitch angle.

9. The method for assisting decision-making in landing of a manned aircraft according to claim 8, characterized in that: The method of calculating the desired three-dimensional angle and three-axis attitude angular velocity by angle control comprises the following steps: Get the angular velocity and acceleration of the aircraft in the current flight state, and calculate the actual roll angle and actual pitch angle of the aircraft in the current flight state; According to the actual roll angle, the Kalman filter method is used to calculate the roll angle estimate According to the actual pitch angle, the Kalman filter method is used to calculate the estimated pitch angle Calculate the expected roll angle φ sp and the estimated roll angle The difference between them is taken as the rolling angle error φ err , and then the roll angle error φ err Perform proportional control to obtain the desired rolling angular velocity Calculate the desired pitch angle θ sp and the pitch angle estimate The difference between them is taken as the pitch angle error θ err , and then the pitch angle error θ err Perform proportional control to obtain the desired pitch angular velocity According to the desired roll angle φ sp and the desired pitch angle θ sp , using the coordinated turning control law to calculate the desired yaw rate And according to the expected yaw rate Calculate the desired yaw angle ψ sp ; The coordinated turning control law is expressed as: Where V T Indicates true airspeed; The desired roll angle φ sp , desired pitch angle θ sp and the desired yaw angle ψ sp as the desired three-dimensional angle; Using the expected roll velocity Expected roll angular velocity and the desired yaw rate Construct the three-axis attitude angle change rate, expressed as: in Indicates the rate of change of the three-axis attitude angle; Then the three-axis attitude angle change rate is converted into a coordinate system to obtain the three-axis attitude angular velocity in the body coordinate system in Indicates the three-axis attitude angular velocity; p sp represents the expected rolling angular velocity; q sp represents the expected pitch angular velocity; γ sp represents the desired yaw rate; φ represents the roll angle.

10. The manned aircraft landing auxiliary decision method according to claim 9, characterized in that: The method of calculating the three-axis attitude angular acceleration by using angular velocity control comprises the following steps: The three-axis attitude angular velocity As the feedforward input of angular velocity control, and through feedforward control calculation feedforward control three-axis attitude angular acceleration Get the actual value of the three-axis attitude angular velocity It is expressed as: in Indicates the actual value of the rolling angular velocity; Indicates the actual value of the pitch angular velocity; Indicates the actual value of the track angular velocity; Using three-axis attitude angular velocity And the actual value of the three-axis attitude angular velocity Calculate the angular velocity error of the three-axis attitude It is expressed as: where p err represents the rolling angular velocity error; q err represents the pitch angular velocity error; γ err represents the track angular velocity error; The three-axis attitude angular velocity error Input PI controller to get PI controlled angular acceleration Using feedforward to control the three-axis attitude angular acceleration and PI control angular acceleration Calculate the three-axis attitude angular acceleration It is expressed as: Among them FF Scaler is the feedforward control airspeed scaling factor; PI Scaler is the PI control airspeed scaling factor; and: Where V T is the true airspeed; V I is the indicated airspeed; and They are the set true airspeed and indicated airspeed respectively.

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