A method for constructing a flight simulation environment for a robot pilot and a simulation platform

Abstract

The present invention relates to the technical field of semi-physical simulation of aircraft, and in particular to a method for constructing a flight simulation environment for a robot pilot and a simulation platform. The construction method includes configuring a retired aircraft and original aircraft support equipment; removing the seats in the cockpit of the retired aircraft, and installing the robot aircraft on the original seat back guide rail position; using the original aircraft support equipment to lift the retired aircraft so that the tires are off the ground by a preset distance, and connecting the hydraulic pump and power supply; carrying a simulation system on the retired aircraft, and establishing a communication connection between the simulation system and the robot pilot; the simulation system includes a simulation computer, a data acquisition component, and a data cable and a power cable for connecting them. Using retired aircraft to configure the original aircraft support equipment and carry the simulation system can provide a realistic simulation environment for the robot pilot and shorten the process of the robot pilot forming an independent driving ability. It makes full use of existing equipment and adopts non-invasive assembly, with strong versatility and low cost.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft hardware-in-the-loop simulation, and in particular to a method for constructing a flight simulation environment for a robot pilot and a simulation platform. Background Art

[0002] Achieving unmanned flight of retired aircraft without extensive modifications has important civil and military strategic significance. With the rapid development of robotics, using robot pilots with aircraft-flying capabilities to fly retired aircraft is a conceptually advanced and technically feasible option.

[0003] The robot pilot needs to have dimensions roughly equivalent to those of a seated adult, allowing it to fit inside a cockpit. Furthermore, the robot pilot must possess visual and auditory acquisition, recognition, analysis, and judgment capabilities, as well as the ability to sense three-dimensional position, attitude, velocity, angular velocity, and overload for decision-making, judgment, and control. Furthermore, the robot pilot must possess the ability to manipulate the aircraft's throttle, control control surfaces, and operate special mission buttons, replacing humans.

[0004] Like human pilots, robot pilots must undergo training and modification in a realistic simulation environment of their aircraft before actually flying it, equipping them with the necessary skills to fly it. Therefore, the construction of a simulation environment for robot pilots is essential for their development and can significantly improve the efficiency of robot pilot development. However, existing technologies do not disclose methods for constructing flight simulation environments for robot pilots. Summary of the Invention

[0005] The purpose of the present invention is to provide a method for constructing a flight simulation environment for a robot pilot and a simulation platform for making full use of retired aircraft to realize physical simulation and thus improve the efficiency of robot pilot development.

[0006] To achieve the above object, the present invention provides the following technical solutions:

[0007] In a first aspect, the present invention provides a method for constructing a flight simulation environment for a robot pilot, comprising the following steps:

[0008] S10. Configure retired aircraft and original aircraft support equipment;

[0009] S11. Remove the seat in the decommissioned aircraft cockpit and install the robot pilot in the original seat back rail position;

[0010] S12. Use the original aircraft support equipment to lift the decommissioned aircraft so that the tires are off the ground a preset distance, and connect the hydraulic pump and power supply;

[0011] S13. Install a simulation system on the retired aircraft and establish a communication connection between the simulation system and the robot pilot; the simulation system includes a simulation computer, a data acquisition component, and a data cable and a power cable for connecting the simulation system and the robot pilot; wherein the data acquisition component includes an elevator angle collector, a rudder angle collector, an aileron angle collector, a discrete state video collector, and a throttle handle scale video collector installed at a preset position of the retired aircraft; the simulation computer is configured with an aerodynamic force and torque solution model, a dynamic solution model, and a kinematic solution model; the simulation computer receives the elevator angle, rudder angle, aileron angle, throttle handle position, and discrete state information, and sequentially uses the aerodynamic force and torque solution model, the dynamic solution model, and the kinematic solution model to solve them, and obtains the three-dimensional position, three-dimensional speed, three-dimensional angle and angular velocity, weight, and fuel quantity information of the retired aircraft through variable conversion, and outputs the above information to the robot pilot through the data interface.

[0012] As a possible implementation method, before executing the simulation system, the simulation system is initialized, specifically including: collecting the zero position of the elevator, rudder and aileron control surfaces; calibrating the throttle handle position; setting the initial three-dimensional position, three-dimensional speed, three-dimensional angle, angular rate, weight, fuel volume, external plug-in model and quantity of the retired aircraft.

[0013] As a possible implementation method, the aerodynamic force and torque solution model includes a lift solution operator model, a drag solution operator model, a side force solution operator model, a pitch moment solution operator model, a rolling moment solution operator model, a yaw moment sub-model, an engine thrust sub-model, a fuel consumption quantum model, a center of gravity solution operator model, a moment of inertia solution operator model and an atmospheric data solution operator model.

[0014] As a possible implementation method, the lift solver model is:

[0015] Y=qs×C y

[0016]

[0017] Where Y is lift; qs is dynamic pressure, which represents the pressure generated when the air flows, and the unit is Pa, qs = ρ × v 2 / 2, ρ is the air density at the flight altitude, υ is the flight speed; C y is the lift coefficient; C y0 、 They are zero lift coefficient, angle of attack lift coefficient, and horizontal tail lift coefficient, which are obtained based on the difference of longitudinal aerodynamic parameters; C yjy is the flap lift coefficient. If there are flaps, different values ​​are taken according to the flap angle. The value is calculated based on the wind data or aerodynamic theory. If there are no flaps, C yjy is 0; Cydx is the influence of ground effect, and the ground effect is considered when the aircraft height from the ground is less than the height threshold; α is the aircraft angle of attack; δ z is the horizontal tail deflection angle;

[0018] The resistance solution operator model is:

[0019] X=qs×C x

[0020]

[0021] Where X is the resistance; qs is the dynamic pressure; C x is the drag coefficient; C x0 、 is the zero-lift drag coefficient and the induced drag coefficient, which are obtained based on the difference in longitudinal aerodynamic parameters; α is the aircraft angle of attack; C is the altitude correction for zero lift resistance; xqlj is the landing gear drag coefficient; C xjy is the flap drag coefficient, which takes different values ​​according to the flap angle. The data value is determined based on the wind data or aerodynamic theory. If there is no flap, this value is 0; C xwg is the external drag coefficient, calculated based on aerodynamic theory. Additionally, when the engine is stopped, the above drag coefficients increase by 0.0074.

[0022] The lateral force solution operator model is:

[0023] Z=qs×C z

[0024]

[0025] Where Z is the lateral force; qs is the dynamic pressure; C z is the lateral force coefficient; is the side force coefficient of the sideslip angle, is the rudder side force coefficient, According to the original aerodynamic data table obtained from the wind test, the difference is obtained by querying the flight altitude and flight speed as input; C zf is the engine side force coefficient, β is the side slip angle of the aircraft; δ y For the aircraft rudder.

[0026] As a possible implementation method, the pitching moment solver model is:

[0027] M z =qs×b A ×m z

[0028]

[0029] Among them, M z is the pitching moment; qs is the dynamic pressure; b A is the mean aerodynamic chord length; m z is the pitch moment coefficient; α is the aircraft angle of attack, δ z is the horizontal tail deviation angle of the aircraft, m z0 is the pitching moment coefficient generated by the aircraft shape when the angle of attack is 0°, is the pitching moment coefficient generated by the aircraft's angle of attack; It is the ground effect coefficient caused by the aircraft's angle of attack. According to engineering test data, it is included when the aircraft is less than 2.2 meters above the ground. α is the aircraft angle of attack; is the pitching moment coefficient generated by the rate of change of the aircraft's angle of attack; is the rate of change of the aircraft's angle of attack; is the pitch moment coefficient generated by the aircraft elevator; δ z is the horizontal tail deflection angle; is the pitch moment coefficient generated by the pitch angular rate of the aircraft; ω z is the pitch angular rate; m zf is the effect of the aircraft on the pitch moment coefficient, m zs is the moment of lift on the center of gravity; m zwg The data is calculated based on the shape of the plug-in through aerodynamics theory; m zjy The flap angle takes different values. The data value is determined based on wind data or aerodynamic theory. If there is no flap, this value is 0. is the rate of change of the aircraft's angle of attack; ω z is the pitch angle rate; b A is the mean aerodynamic chord length;

[0030] The rolling moment solver model is:

[0031] M x =qs×l×m x

[0032]

[0033] Among them, M x is the rolling moment; qs is the dynamic pressure; l is the wingspan; m x is the transverse torque coefficient; is the rolling moment coefficient generated by the roll angle; β is the side slip angle of the aircraft; is the rolling moment coefficient generated by the rudder; δ y For the aircraft rudder; is the rolling moment coefficient generated by the aileron rudder; δ x It is the aileron rudder of the aircraft; is the rolling moment coefficient generated by the rolling angular rate; ω x is the roll angular rate; is the rolling moment coefficient generated by the yaw angular rate; ω y is the yaw rate; is the rolling moment coefficient generated by the yaw angle acceleration; α is the aircraft's angle of attack; v is the true airspeed; According to the original aerodynamic data table obtained from the wind test, the difference is obtained by querying the flight altitude and flight speed as input; ω y is the yaw rate; l is the wingspan;

[0034] The yaw moment sub-model is:

[0035] M y =qs×l×m y

[0036]

[0037] Among them, M y is the yaw moment; qs is the dynamic pressure; l is the wingspan; m y is the yaw moment coefficient; is the yaw moment coefficient generated by the roll angle; β is the side slip angle of the aircraft; is the yawing moment coefficient generated by the rudder; δ y For the aircraft rudder; is the yaw moment coefficient generated by the yaw angular rate; ω y is the yaw rate; is the yaw moment coefficient generated by the roll angular rate; ω x is the roll angular rate; v is the true airspeed; m yf is the effect of the engine on the yaw moment coefficient, According to the original aerodynamic data table obtained from the wind test, the difference is obtained by querying the flight altitude and flight speed as input.

[0038] As a possible implementation, the engine thrust sub-model is:

[0039] P=k p ×T h

[0040] Where P is the effective thrust of the engine; k p is the inlet loss correction coefficient. According to engineering experience data, when the flight Mach number Ma>0.3, k p=0.905, when Ma≤0.3, k p =0.88+0.0833×Ma; the engine thrust is T h , unit is N;

[0041] The quantum model of fuel consumption is:

[0042] S cf =k p ×scf

[0043] Among them, S cf is the fuel consumption; k p is the intake loss correction factor; the fuel consumption rate is scf;

[0044] The centroid solution operator model is:

[0045] Oil weight (L) 0 620 1360 1600 Center of gravity (%) 43.18 38.45 40.26 38.71 Oil weight (L) 2378 2770 2847 Center of gravity (%) 41.18 40.39 40.17

[0046] The moment of inertia solver model is:

[0047] When two 760-liter auxiliary fuel tanks are hung and the fuel capacity is greater than 2378 liters, the three-axis rotational inertia is:

[0048]

[0049] When two 760-liter auxiliary fuel tanks are hung and the fuel capacity is less than 2378 liters and greater than 1600 liters, the three-axis rotational inertia is:

[0050]

[0051] When there is no external plug-in, the three-axis moment of inertia is:

[0052]

[0053] Among them, I x is the x-axis moment of inertia of the aircraft, I y is the moment of inertia of the aircraft along the y-axis, I z is the moment of inertia of the aircraft along the z axis, I x , I y , I z The unit is N·m·s 2 ; G y is the fuel quantity; after the landing gear is lowered, I z Increase by 18.5 N·m·s 2 ;

[0054] The atmospheric data solution operator model is:

[0055] When the flight altitude is less than 11,000 meters,

[0056] When the altitude is greater than 11,000 meters and less than 20,000 meters,

[0057] Where T is the atmospheric temperature in °C; ρ is the atmospheric density in kg / m 3 ; p is atmospheric pressure, unit is Pa; v s is the speed of sound, in m / s; h is the flight altitude.

[0058] As a possible implementation method, the dynamics solution model includes a dynamics solution model for the aerial flight segment, a dynamics solution model for the takeoff and run segment, and a dynamics solution model for the landing and run segment;

[0059] Among them, the dynamics solution model of the aerial flight segment is:

[0060]

[0061] The dynamic solution model of the takeoff roll phase is:

[0062]

[0063] Among them, ∑F x is the total resistance; ∑F y is the total lift; ∑F z is the total lateral force; ∑M x is the total rolling moment; ∑M y is the total yaw moment; ∑M z is the total pitching moment; P is the effective thrust of the engine; X is the drag; Y is the lift; Z is the side force; φ e is the angle between the thrust and the aircraft axis; α is the aircraft angle of attack; β is the aircraft sideslip angle; θ is the aircraft pitch angle; γ is the aircraft roll angle; mg is the aircraft gravity; f is the aircraft rolling resistance; N is the rolling support force;

[0064] The dynamic solution model of the landing roll phase is:

[0065]

[0066] Among them, ∑F x is the total resistance; ∑F y is the total lift; ∑F z is the total lateral force; ∑M x is the total rolling moment; ∑M y is the total yaw moment; ∑M z is the total pitching moment; v x is the horizontal velocity of the aircraft; v y is the longitudinal velocity of the aircraft; v z is the lateral velocity of the aircraft; ω x is the roll angular rate; ωy is the yaw angular rate; ω z is the pitch angle rate; I x I is the moment of inertia of the aircraft along the x-axis; y I is the moment of inertia of the aircraft along the y-axis; z I is the moment of inertia of the aircraft along the z axis; xy is the vehicle's moment of inertia.

[0067] As a possible implementation method, the kinematic solution model is:

[0068]

[0069] Among them, γ is the roll angle of the aircraft; φ is the true heading; θ is the pitch angle of the aircraft; ω x is the roll angular rate; ω y is the yaw angular rate; ω z is the pitch angular rate; is the x-axis component of the aircraft velocity in the earth axis system; is the y-axis component of the aircraft velocity in the earth axis system; is the z-axis component of the aircraft velocity in the earth axis system; v x is the horizontal velocity of the aircraft; v y is the longitudinal velocity of the aircraft; v z is the lateral velocity of the aircraft.

[0070] As a possible implementation method, the three-axis velocity of the earth axis system and the earth velocity are converted into variables in the following way:

[0071]

[0072] Among them, v xg is the x-axis component of the aircraft velocity in the earth axis system; v yg is the y-axis component of the aircraft velocity in the earth axis system; v zg is the z-axis component of the aircraft velocity in the earth axis system; v d is the aircraft ground speed;

[0073] The airspeed is converted into a variable as follows:

[0074]

[0075] Among them, u x is the x-axis component of wind speed in the body axis system; u y is the y-axis component of wind speed in the body axis system; u z is the z-axis component of the wind speed in the body axis system; u xg is the x-axis component of wind speed in the earth's axis system; u yg is the y-axis component of wind speed in the earth's axis system; u zg is the z-axis component of the wind speed in the earth's axis system; vx is the horizontal velocity of the aircraft; v y is the longitudinal velocity of the aircraft; v z is the lateral velocity of the aircraft; φ is the true heading; is the pitch angle of the aircraft; γ is the roll angle of the aircraft; v is the true airspeed of the aircraft.

[0076] In a second aspect, a robotic pilot flight simulation platform includes a retired aircraft, original aircraft support equipment, and a simulation system. The simulation system is mounted on the retired aircraft and is in communication with the robotic pilot. The original aircraft support equipment is used to raise the retired aircraft to a certain distance above the ground and provide power to the retired aircraft, the robotic pilot, and the simulation system.

[0077] The simulation system includes a simulation computer, a data acquisition component, and a data cable and a power cable for connecting them; wherein the data acquisition component includes an elevator angle collector, a rudder angle collector, an aileron angle collector, a discrete state video collector, and a throttle handle scale video collector installed at a preset position of the retired aircraft; the simulation computer is configured with an aerodynamic force and torque solution model, a dynamic solution model, and a kinematic solution model; the simulation computer receives the elevator angle, rudder angle, aileron angle, throttle handle position, and discrete state information, and uses the aerodynamic force and torque solution model, the dynamic solution model, and the kinematic solution model to solve them in turn, and obtains the three-dimensional position, three-dimensional speed, three-dimensional angle and angular velocity, weight, and fuel quantity of the retired aircraft through variable conversion, and outputs the above information to the robot pilot through the data interface.

[0078] Compared with the prior art, the present invention has the following beneficial effects:

[0079] 1. The present invention utilizes retired aircraft, configures original aircraft support equipment, and carries a simulation system, which can provide a realistic simulation environment for robot pilots, greatly shortening the process of robot pilots developing independent driving capabilities.

[0080] 2. The present invention makes full use of existing equipment and adopts a non-invasive mounting method, which has the advantages of strong versatility, low cost and easy implementation. BRIEF DESCRIPTION OF THE DRAWINGS

[0081] The accompanying drawings described herein are used to provide a further understanding of the present invention and constitute a part of the present invention. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention. In the accompanying drawings:

[0082] Figure 1 A flowchart of a construction method provided in an embodiment of the present invention;

[0083] Figure 2The working principle of the simulation environment provided by the embodiment of the present invention;

[0084] Figure 3 A flowchart of a simulation computer execution provided by an embodiment of the present invention;

[0085] Figure 4 A schematic diagram of the structure of a simulation platform provided by an embodiment of the present invention;

[0086] Figure 5 This is a flow chart for identifying the throttle handle scale position provided by an embodiment of the present invention.

[0087] Reference numerals:

[0088] 10- jack, 11- hydraulic pump;

[0089] 20-discrete state video collector, 21-throttle handle scale video collector, 22-elevator angle collector, 23-rudder angle collector, 24-aileron rudder angle collector. DETAILED DESCRIPTION

[0090] To facilitate a clear description of the technical solutions of the embodiments of the present invention, the words "first" and "second" are used in the embodiments of the present invention to distinguish between identical or similar items with substantially the same functions and effects. For example, the first threshold and the second threshold are merely used to distinguish between different thresholds and do not limit their order. Those skilled in the art will understand that the words "first" and "second" do not limit the quantity or execution order, and the words "first" and "second" do not necessarily mean different.

[0091] It should be noted that, in the present invention, words such as "exemplary" or "for example" are used to indicate examples, illustrations, or descriptions. Any embodiment or design described as "exemplary" or "for example" in the present invention should not be construed as being preferred or advantageous over other embodiments or designs. Rather, the use of words such as "exemplary" or "for example" is intended to present the relevant concepts in a concrete manner.

[0092] In the present invention, "at least one" refers to one or more, and "more" refers to two or more. "And / or" describes the association relationship of associated objects, indicating that three relationships may exist. For example, A and / or B can represent: the existence of A alone, the existence of A and B at the same time, and the existence of B alone, where A and B can be singular or plural. The character " / " generally indicates that the previous and next associated objects are in an "or" relationship. The following at least one item (item) or similar expressions thereof refer to any combination of these items, including any combination of single items (items) or plural items (items). For example, at least one item (item) of a, b or c can represent: a, b, c, the combination of a and b, the combination of a and c, the combination of b and c, or the combination of a, b and c, where a, b, c can be single or multiple.

[0093] First, see Figures 1 to 4 The present invention provides a method for constructing a robot pilot flight simulation environment, comprising the following steps:

[0094] S10. Configure the decommissioned aircraft and original aircraft support equipment. A near-realistic simulation environment is established based on the decommissioned aircraft for robot pilot simulation training. This provides the robot pilot with a realistic simulation environment, such as a realistic cockpit, joystick, throttle, and switches, as well as highly realistic flight dynamics based on the simulation system. The original aircraft support equipment must at least include a jack 10, a hydraulic pump 11, or a power supply to ensure the normal operation of the decommissioned aircraft's joystick, pedals, throttle, and switches during the simulation.

[0095] S11. Remove the seat from the decommissioned aircraft cockpit and install the robot aircraft in the original seat back rail position. Alternatively, install the robot pilot in the cockpit by installing a conversion bracket.

[0096] S12. Lift the decommissioned aircraft by jacking it up to a certain distance from the ground. Simultaneously, connect the power supply, hydraulic pump 11, and simulation system.

[0097] S13. Install a simulation system on the decommissioned aircraft and establish a communication connection between the simulation system and the robot pilot. The simulation system includes a simulation computer, a data acquisition component, and data and power cables for connecting them. The data acquisition component includes an elevator angle collector 22, a rudder angle collector 23, an aileron angle collector 24, a discrete state video collector 20, and a throttle handle scale video collector 21, installed at a preset location on the decommissioned aircraft. The simulation computer is equipped with an aerodynamic force and torque solver model, a dynamic equation solver model, and a kinematic equation solver model. The simulation computer receives information on the elevator angle, rudder angle, aileron angle, throttle handle position, and discrete state, and sequentially performs calculations using the aerodynamic force and torque solver model, the dynamic equation solver model, and the kinematic equation solver model. Through variable conversion, the simulation computer obtains the decommissioned aircraft's three-dimensional position, three-dimensional velocity, three-dimensional angle and angular rate, weight, and fuel level information. This information is then output to the robot pilot via a data interface.

[0098] The simulation computer consists of hardware and software. The hardware includes a CPU board, video signal processing board, discrete quantity board, power supply board, and chassis. The aerodynamic force and torque calculations, dynamics calculations, and kinematics calculations are implemented using software algorithms.

[0099] Scales, pointers, and cameras are installed at the fixed ends of the elevator, rudder, and aileron. The cameras capture images of the pointer changes on the scales. Image processing methods are then used to calculate the corresponding angles of the elevator, rudder, and aileron, which serve as input for the aerodynamic force and torque calculation model. The simulation of the three control surface angles requires high real-time performance, with a data update rate of milliseconds.

[0100] The discrete state video collector 20 is used to capture the status of landing gear retraction and extension, flaps retraction and extension, and external equipment. Specifically, a ground-based camera can be installed at the rear of the decommissioned aircraft to capture real-time video of the retraction and extension of the landing gear, flaps retraction and extension, and external equipment. This video can then be processed using video processing methods. The discrete state update rate requirement is not high; a data update rate of seconds is sufficient.

[0101] As a possible implementation method, the video processing method for the landing gear retracted or extended, flaps retracted or extended, and external load status uses the YOLOv5 algorithm, which specifically includes the following steps:

[0102] S130. Data Collection and Preprocessing. Use a binocular camera to collect images of the aircraft under various angles and lighting conditions, with the landing gear retracted or extended, flaps retracted or extended, and external attachments present. These images are annotated to construct a training dataset and a test dataset.

[0103] S131. Model training and optimization. Use the YOLOv5 framework for model training. Optimize the model's target detection performance by adjusting hyperparameters such as network structure and anchor point size.

[0104] The above technical solution achieves multi-scale detection enhancement. This addresses the issue of changes in the size of detected objects due to changes in camera placement by leveraging YOLOv5's multi-scale detection capabilities to enhance detection capabilities. It also optimizes real-time performance. While maintaining detection accuracy, it reduces model size through techniques like model pruning and quantization, improving inference speed and ensuring real-time performance.

[0105] Throttle handle scale video collector 21: A camera is installed above the throttle handle in the cockpit. The throttle handle scale position is calculated through image processing, and converted into engine speed data through conversion, and converted into engine thrust and fuel consumption rate data. The calculation needs to be combined with the altitude data of the retired aircraft.

[0106] As a possible implementation method, the throttle handle scale position recognition collects the current position image of the throttle handle through the camera, and then Figure 5 The steps are processed:

[0107] S140. Pointer Extraction: Throttle Grip Position Pointer Extraction uses image processing methods to separate the pointer image from the complete instrument image. Using the subtraction method to reduce background interference from the dial background can improve pointer reading accuracy. The subtraction method works by subtracting a template image, showing only the background and without the pointer, from the image under test to achieve background reduction. The template image can be obtained by performing an AND operation on two dial images with the pointer in different positions.

[0108] S141. Preprocessing the pointer image. After the pointer is extracted, it needs to be preprocessed. The goal is to extract the pointer edge from the pointer image to prepare for the subsequent Hough transform to extract straight lines. Methods used in this stage include grayscale conversion, Gaussian smoothing, and binary edge detection.

[0109] Among them, the image captured by the camera is an RGB true color format image. The color information is not very meaningful for pointer recognition. On the contrary, the true color image will increase the image processing time. Therefore, the image is first grayscaled and used as the basis for the next step of recognition.

[0110] Since digital images generate noise during acquisition and transmission, which affects the accuracy of pointer recognition, the image needs to be smoothed before further processing. The 5-neighborhood Gaussian filtering algorithm is used to smooth the image.

[0111] After the pointer image is grayed and Gaussian transformed, the brightness of the pointer part is relatively high. At this time, a binarization operation can be performed to remove other useless information, and the binarization threshold is set to 100.

[0112] Edge detection can reduce the number of straight lines obtained during Hough transform, making the reading calculation results more accurate.

[0113] S142. Hough transform is performed on the pointer image after edge detection to find the straight line, and the straight line equation of the pointer edge can be obtained.

[0114] S143. Pointer reading: Identify the dial scale using the scale method, read the instrument based on the relative position of the pointer and scale, and ultimately obtain the throttle handle scale data.

[0115] The installation of each component unit of the simulation system is temporary when in use, without changing the aircraft status, realizing a non-intrusive design.

[0116] The solution cycle of the aerodynamic and torque solution model, dynamic solution model and kinematic solution model configured by the simulation computer must be guaranteed to reach the millisecond level.

[0117] See also Figure 2 and Figure 3 The data acquisition component sends the elevator angle, rudder angle, aileron angle, throttle handle position, and discrete state information to the aerodynamic force and torque calculation model configured in the simulation computer. Combined with the original aerodynamic data of the retired aircraft, it calculates the retired aircraft's aerodynamic forces and torques, specifically lift, drag, side force, pitching moment, engine thrust, fuel consumption rate, center of gravity, moment of inertia, and atmospheric data. Based on this, dynamic and kinematic calculations are then performed, and through transformation, information such as the retired aircraft's longitude, latitude, altitude, northeast celestial velocity, pitch angle, bank angle, true heading, pitch rate, bank rate, and yaw rate is obtained. This information is then sent to the robot pilot, who then operates the retired aircraft. The data acquisition component then collects relevant data in real time before entering the next calculation cycle and repeating the above process.

[0118] For ease of understanding, the solution method is introduced in detail below:

[0119] The lift solution operator model is:

[0120] Y=qs×C y

[0121]

[0122] Among them, Y is lift; qs is dynamic pressure, which represents the pressure generated when the air flows, and the unit is Pa, qs = ρ × υ2 / 2, ρ is the air density at the flight altitude, υ is the flight speed; C y is the lift coefficient; C y0 、 They are zero lift coefficient, angle of attack lift coefficient, and horizontal tail lift coefficient, which are obtained based on the difference of longitudinal aerodynamic parameters; C yjy is the flap lift coefficient. If there are flaps, different values ​​are taken according to the flap angle. The value is calculated based on the wind data or aerodynamic theory. If there are no flaps, C yjy is 0; C ydx The ground effect is considered when the aircraft is less than the height threshold. For example, the ground effect is taken into account when the aircraft is less than 2.2 meters above the ground. The calculation formula is: α is the aircraft angle of attack; δ z is the horizontal tail deflection angle;

[0123] The resistance solution operator model is:

[0124] X=qs×C x

[0125]

[0126] Where X is the resistance; qs is the dynamic pressure; C x is the drag coefficient; C x0 、 is the zero-lift drag coefficient and the induced drag coefficient, which are obtained based on the difference in longitudinal aerodynamic parameters; α is the aircraft angle of attack; C is the altitude correction for zero lift resistance; xqlj is the landing gear drag coefficient; C xjy is the flap drag coefficient, which takes different values ​​according to the flap angle. The data value is determined based on the wind data or aerodynamic theory. If there is no flap, this value is 0; C xwg is the external drag coefficient, calculated based on aerodynamic theory. Additionally, when the engine is stopped, the above drag coefficients increase by 0.0074.

[0127] The lateral force solution operator model is:

[0128] Z=qs×C z

[0129]

[0130] Where Z is the lateral force; qs is the dynamic pressure; C z is the lateral force coefficient; is the side force coefficient of the sideslip angle, is the rudder side force coefficient, According to the original aerodynamic data table obtained from the wind test, the difference is obtained by querying the flight altitude and flight speed as input; C zf is the engine side force coefficient, β is the side slip angle of the aircraft; δ y For the aircraft rudder.

[0131] The pitching moment solver model is:

[0132] M z =qs×b A ×m z

[0133]

[0134] Among them, M z is the pitching moment; qs is the dynamic pressure; b A is the mean aerodynamic chord length; m z is the pitch moment coefficient; α is the aircraft angle of attack, δ z is the horizontal tail deviation angle of the aircraft, m z0 is the pitching moment coefficient generated by the aircraft shape when the angle of attack is 0°, is the pitching moment coefficient generated by the aircraft's angle of attack; It is the ground effect coefficient caused by the aircraft's angle of attack. According to engineering test data, it is included when the aircraft is less than 2.2 meters above the ground. α is the aircraft angle of attack; is the pitching moment coefficient generated by the rate of change of the aircraft's angle of attack; is the rate of change of the aircraft's angle of attack; is the pitch moment coefficient generated by the aircraft elevator; δ z is the horizontal tail deflection angle; is the pitch moment coefficient generated by the pitch angular rate of the aircraft; ω z is the pitch angular rate; m zf is the effect of the aircraft on the pitch moment coefficient, m zs is the moment of lift on the center of gravity; m zwg The data is calculated based on the shape of the plug-in through aerodynamics theory; m zjy The flap angle takes different values. The data value is determined based on wind data or aerodynamic theory. If there is no flap, this value is 0. is the rate of change of the aircraft's angle of attack; ω z is the pitch angular rate;

[0135] The rolling moment solution operator model is:

[0136] M x =qs×l×mx

[0137]

[0138] Among them, M x is the rolling moment; qs is the dynamic pressure; l is the wingspan; m x is the transverse torque coefficient; is the rolling moment coefficient generated by the roll angle; β is the side slip angle of the aircraft; is the rolling moment coefficient generated by the rudder; δ y For the aircraft rudder; is the rolling moment coefficient generated by the aileron rudder; δ x It is the aileron rudder of the aircraft; is the rolling moment coefficient generated by the rolling angular rate; ω x is the roll angular rate; is the rolling moment coefficient generated by the yaw angular rate; ω y is the yaw rate; is the rolling moment coefficient generated by the yaw angle acceleration; α is the aircraft's angle of attack; v is the true airspeed; According to the original aerodynamic data table obtained from the wind test, the difference is obtained by querying the flight altitude and flight speed as input;

[0139] The yaw moment sub-model is:

[0140] M y =qs×l×m y

[0141]

[0142] Among them, M y is the yaw moment; qs is the dynamic pressure; l is the wingspan; m y is the yaw moment coefficient; is the yaw moment coefficient generated by the roll angle; β is the side slip angle of the aircraft; is the yawing moment coefficient generated by the rudder; δ y For the aircraft rudder; is the yaw moment coefficient generated by the yaw angular rate; ω y is the yaw rate; is the yaw moment coefficient generated by the roll angular rate; ω x is the roll angular rate; v is the true airspeed; m yf is the effect of the engine on the yaw moment coefficient, According to the original aerodynamic data table obtained from the wind test, the difference is obtained by querying the flight altitude and flight speed as input.

[0143] The engine thrust sub-model is:

[0144] P=k p ×T h

[0145] Where P is the effective thrust of the engine; k p is the inlet loss correction coefficient. According to engineering experience data, when the flight Mach number Ma>0.3, k p =0.905, when Ma≤0.3, k p =0.88+0.0833×Ma; the engine thrust is T h , unit is N;

[0146] It is necessary to further explain that the engine thrust and fuel consumption rate are divided into two states: maximum and rated. The thrust and fuel consumption rate of other throttle states have a certain proportional relationship with the rated state. According to the flight altitude and speed, the difference between the engine thrust and fuel consumption rate at maximum and rated states can be obtained, which is recorded as T d 、T e 、scf d 、scf e When the altitude or speed exceeds the range, the difference is calculated based on the three most recent data sets. Assume that the thrust ratios of cruise 1, cruise 2, cruise 3, cruise 4 and other states to the rated state are k t1 、k t2 、k t3 、k t4 The ratio of fuel consumption rate to rated fuel consumption rate is k s1 、k s2 、k s3 、k s4 , the thrust and fuel consumption rate of each state are shown in the following table:

[0147] Thrust and fuel consumption rate table at different throttle states

[0148] Throttle status Thrust (kg) Fuel consumption rate (kg / h) maximum <![CDATA[T d ]]> <![CDATA[scf d ]]> Rated <![CDATA[T e ]]> <![CDATA[scf e ]]> Cruise 1 <![CDATA[k t1 T e ]]> <![CDATA[k s1 scf e ]]> Cruise 2 <![CDATA[k t2 T e ]]> <![CDATA[k s2 scf e ]]> Cruise 3 <![CDATA[k t3 T e ]]> <![CDATA[k s3 scf e ]]> Cruise 4 <![CDATA[k t4 T e ]]> <![CDATA[k s4 scf e ]]> local 70-h / 300 0.167×3600 parking 0 0

[0149] The fuel consumption quantum model is:

[0150] S cf =k p ×scf

[0151] Among them, S cf is the fuel consumption; k p is the intake loss correction factor; the fuel consumption rate is scf;

[0152] The center of gravity of a retired aircraft changes continuously during flight as fuel is consumed. The center of gravity of specific nodes during this change is shown in the table below. The change between two adjacent points is linear and can be calculated by difference. The center of gravity solution submodel is:

[0153] Table of changes in center of gravity with fuel consumption

[0154] Oil weight (L) 0 620 1360 1600 Center of gravity (%) 43.18 38.45 40.26 38.71 Oil weight (L) 2378 2770 2847 Center of gravity (%) 41.18 40.39 40.17

[0155] It should be further explained that when the landing gear is extended, the center of gravity moves back 0.9. The calculation of the center of gravity also needs to take into account the weight and position of the external loads.

[0156] The moment of inertia solver model is:

[0157] When two 760-liter auxiliary fuel tanks are hung and the fuel capacity is greater than 2378 liters, the three-axis rotational inertia is:

[0158] When two 760-liter auxiliary fuel tanks are hung and the fuel capacity is less than 2378 liters and greater than 1600 liters, the three-axis rotational inertia is:

[0159]

[0160] When there is no external plug-in, the three-axis moment of inertia is:

[0161]

[0162] Among them, I x is the x-axis moment of inertia of the aircraft, I y is the moment of inertia of the aircraft along the y-axis, I z is the moment of inertia of the aircraft along the z axis, I x , I y , I z The unit is N·m·s 2 ; G y is the fuel quantity; after the landing gear is lowered, I z Increase by 18.5 N·m·s 2 ;

[0163] The atmospheric data solution operator model is:

[0164] When the flight altitude is less than 11,000 meters,

[0165] When the altitude is greater than 11,000 meters and less than 20,000 meters,

[0166] Where T is the atmospheric temperature in °C; ρ is the atmospheric density in kg / m 3 ; p is atmospheric pressure, unit is Pa; vs is the speed of sound, in m / s; h is the flight altitude.

[0167] The dynamics solution model includes a dynamics solution model for the aerial flight segment, a dynamics solution model for the takeoff and run segment, and a dynamics solution model for the landing and run segment;

[0168] Among them, the dynamics solution model of the aerial flight segment is:

[0169]

[0170] The dynamic solution model of the takeoff roll phase is:

[0171]

[0172] Among them, ∑F x is the total resistance; ∑F y is the total lift; ∑F z is the total lateral force; ∑M x is the total rolling moment; ∑M y is the total yaw moment; ∑M z is the total pitching moment; P is the effective thrust of the engine; X is the drag; Y is the lift; Z is the side force; φ e is the angle between the thrust and the aircraft axis; α is the aircraft angle of attack; β is the aircraft sideslip angle; θ is the aircraft pitch angle; γ is the aircraft roll angle; mg is the aircraft gravity; f is the aircraft rolling resistance; N is the rolling support force;

[0173] The dynamics solution model of the landing roll phase is:

[0174]

[0175] Among them, ∑F x is the total resistance; ∑F y is the total lift; ∑F z is the total lateral force; ∑M x is the total rolling moment; ∑M y is the total yaw moment; ∑M z is the total pitching moment; v x is the horizontal velocity of the aircraft; v y is the longitudinal velocity of the aircraft; v z is the lateral velocity of the aircraft; ω x is the roll angular rate; ω y is the yaw angular rate; ω z is the pitch angle rate; I x I is the moment of inertia of the aircraft along the x-axis; y I is the moment of inertia of the aircraft along the y-axis; z I is the moment of inertia of the aircraft along the z axis; xyis the vehicle's moment of inertia.

[0176] The kinematic solution model is:

[0177]

[0178] Among them, γ is the roll angle of the aircraft; φ is the true heading; θ is the pitch angle of the aircraft; ω x is the roll angular rate; ω y is the yaw angular rate; ω z is the pitch angular rate; is the x-axis component of the aircraft velocity in the earth axis system; is the y-axis component of the aircraft velocity in the earth axis system; is the z-axis component of the aircraft velocity in the earth axis system; v x is the horizontal velocity of the aircraft; v y is the longitudinal velocity of the aircraft; v z is the lateral velocity of the aircraft.

[0179] The three-axis speed of the earth axis system and the earth speed are converted into variables in the following way:

[0180]

[0181] Among them, v xg is the x-axis component of the aircraft velocity in the earth axis system; v yg is the y-axis component of the aircraft velocity in the earth axis system; v zg is the z-axis component of the aircraft velocity in the earth axis system; v d is the aircraft ground speed;

[0182] The airspeed is converted into a variable as follows:

[0183]

[0184] Among them, u x is the x-axis component of wind speed in the body axis system; u y is the y-axis component of wind speed in the body axis system; u z is the z-axis component of the wind speed in the body axis system; u xg is the x-axis component of wind speed in the earth's axis system; u yg is the y-axis component of wind speed in the earth's axis system; u zg is the z-axis component of the wind speed in the earth's axis system; v x is the horizontal velocity of the aircraft; v y is the longitudinal velocity of the aircraft; v z is the lateral velocity of the aircraft; φ is the true heading; θ is the pitch angle of the aircraft; γ is the roll angle of the aircraft; and v is the true airspeed of the aircraft.

[0185] See also Figure 3Before the first solution cycle, the simulation system needs to collect the zero position of the three control surfaces to determine the initial attitude and position of the retired aircraft. The throttle handle position is calibrated, and the initial values ​​of the retired aircraft's initial 3D position, 3D velocity and angular rate, weight, fuel volume, external plug-in model and quantity are set.

[0186] The simulation computer executes the following process in a 1ms calculation cycle: first, it collects the positions of the three control surfaces and throttle handle, as well as the status of the landing gear, flaps, and external attachments. It then calculates the aircraft's aerodynamic forces and torques, solving the dynamic equations for both ground and air segments. Then, it solves the kinematic equations. Finally, through variable conversion, it calculates the aircraft's 3D position, 3D velocity, 3D angle and angular rate, weight, and fuel level, and outputs these information to the robot pilot via a data interface. A new calculation cycle begins after the set timer expires. A shorter calculation cycle increases simulation fidelity. This creates a closed-loop simulation environment, providing the necessary support for training the robot pilot's driving skills.

[0187] The present invention utilizes retired aircraft, configures original aircraft support equipment and carries a simulation system, and can provide a realistic simulation environment for robot pilots, greatly shortening the process of robot pilots developing independent driving capabilities.

[0188] The present invention makes full use of existing equipment and adopts a non-invasive mounting method, and has the advantages of high versatility, low cost and easy implementation.

[0189] See also Figure 4 An embodiment of the present invention further provides a robot pilot flight simulation platform, comprising a retired aircraft, original aircraft support equipment, and a simulation system; wherein the simulation system is mounted on the retired aircraft and is communicatively connected to the robot pilot; the original aircraft support equipment is used to raise the retired aircraft to a certain distance from the ground and provide power for the retired aircraft, the robot pilot, and the simulation system. The simulation system includes a simulation computer, a data acquisition component, and data and power cables for connecting them. The data acquisition component includes an elevator angle collector 22, a rudder angle collector 23, an aileron angle collector 24, a discrete state video collector 20, and a throttle handle scale video collector 21, installed at a preset position on the decommissioned aircraft. The simulation computer is equipped with an aerodynamic force and torque solution model, a dynamic solution model, and a kinematic solution model. The simulation computer receives the elevator angle, rudder angle, aileron angle, throttle handle position, and discrete state information, and sequentially calculates it using the aerodynamic force and torque solution model, the dynamic solution model, and the kinematic solution model. Through variable conversion, the simulation computer obtains the decommissioned aircraft's three-dimensional position, three-dimensional velocity, three-dimensional angle and angular rate, weight, and fuel level information, and outputs this information to the robot pilot via a data interface. The original aircraft support equipment includes at least a jack 10, a hydraulic pump 11, and a power supply.

[0190] Although the present invention is described herein in conjunction with various embodiments, in the process of implementing the claimed invention, those skilled in the art can understand and implement other variations of the disclosed embodiments by viewing the drawings, the disclosure, and the drawings, etc. In the specification, the word "comprising" does not exclude other components or steps, and "one" or "an" does not exclude multiple situations. A single processor or other unit can implement several functions listed in the specification. Certain measures are recorded in different embodiments, but this does not mean that these measures cannot be combined to produce good results.

[0191] Although the present invention has been described with reference to specific features and embodiments thereof, it will be apparent that various modifications and combinations thereof may be made without departing from the spirit and scope of the invention. Accordingly, this specification and drawings are merely illustrative of the present invention and are deemed to cover any and all modifications, variations, combinations or equivalents within the scope of the invention. It will be apparent that various modifications and variations of the present invention may be made by those skilled in the art without departing from the spirit and scope of the invention. Thus, the present invention is intended to include such modifications and variations as fall within the scope of the invention and its equivalents.

Claims

1. A method for constructing a flight simulation environment for a robot pilot, characterized in that: The steps include: S10. Configure retired aircraft and original aircraft support equipment; S11. Remove the seats from the decommissioned aircraft cockpit and install the robot pilot in the original seat back rail position. S12. Use the original aircraft support equipment to lift the decommissioned aircraft so that the tires are off the ground a preset distance, and connect the hydraulic pump and power supply; S13. A simulation system is mounted on the retired aircraft, and the simulation system establishes a communication connection with the robot pilot; the simulation system includes a simulation computer, a data acquisition component, and a data cable and power cable for connecting them; The data acquisition component includes an elevator angle collector, a rudder angle collector, an aileron angle collector, a discrete state video collector, and a throttle handle scale video collector installed at a preset position on the retired aircraft; The simulation computer is configured with an aerodynamic force and torque solution model, a dynamic solution model, and a kinematic solution model; the aerodynamic force and torque solution model includes a lift solution operator model, a drag solution operator model, a side force solution operator model, a pitch moment solution operator model, a rolling moment solution operator model, a yaw moment sub-model, an engine thrust sub-model, a fuel consumption quantum model, a center of gravity solution operator model, a moment of inertia solution operator model, and an atmospheric data solution operator model; The lift solution operator model is: in, for lift; is the dynamic pressure, which represents the pressure generated when the air flows, and its unit is Pa. , is the air density at the flight altitude, is the flight speed; is the lift coefficient; 、 、 They are zero lift coefficient, angle of attack lift coefficient, and horizontal tail lift coefficient, which are obtained based on the difference of longitudinal aerodynamic parameters; is the flap lift coefficient. If there are flaps, the value is different according to the flap angle. The value is calculated based on the wind data or aerodynamic theory. If there are no flaps, is 0; For the influence of ground effect, the ground effect is considered when the aircraft altitude is less than the altitude threshold; is the aircraft angle of attack; is the horizontal tail deflection angle; The resistance solution operator model is: in, For resistance; is the dynamic pressure, is the drag coefficient; 、 is the zero-lift drag coefficient and the induced drag coefficient, which are obtained based on the difference in longitudinal aerodynamic parameters; is the aircraft angle of attack; Altitude correction for zero lift drag; is the landing gear drag coefficient; The flap drag coefficient takes different values ​​according to the flap angle. The data value is determined according to the wind data or aerodynamic theory. If there is no flap, this value is 0. is the external drag coefficient, calculated based on aerodynamic theory. Additionally, when the engine is stopped, the above drag coefficients increase by 0.0074. The lateral force solution operator model is: in, is the lateral force; is dynamic pressure; is the lateral force coefficient; is the side force coefficient of the sideslip angle, is the rudder side force coefficient, 、 According to the original aerodynamic data table obtained from the wind test, the difference is obtained by querying the flight altitude and flight speed as input; is the engine side force coefficient, ; is the aircraft sideslip angle; For the aircraft rudder; The simulation computer receives elevator angle, rudder angle, aileron angle, throttle handle position and discrete state information, and sequentially performs calculations using an aerodynamic force and torque solution model, a dynamic solution model and a kinematic solution model, and obtains the three-dimensional position, three-dimensional speed, three-dimensional angle and angular rate, weight and fuel quantity information of the retired aircraft through variable conversion, and outputs the above information to the robot pilot through a data interface.

2. The method for constructing a robot pilot flight simulation environment according to claim 1, characterized in that: Before executing the simulation system, the simulation system is initialized, specifically including: Collect the zero position of the elevator, rudder and aileron surfaces; Calibrate the throttle handle position; set the retired aircraft's initial 3D position, 3D speed, 3D angle, angular rate, weight, fuel volume, external attachment model and quantity.

3. The method for constructing a robot pilot flight simulation environment according to claim 1, characterized in that: The pitching moment solver model is: in, is the pitching moment; is dynamic pressure; is the mean aerodynamic chord length; is the pitching moment coefficient; is the aircraft angle of attack, is the tail deflection angle of the aircraft, is the pitching moment coefficient generated by the aircraft shape when the angle of attack is 0°, is the pitching moment coefficient generated by the aircraft's angle of attack; It is the ground effect coefficient caused by the aircraft's angle of attack. According to engineering test data, it is included when the aircraft is less than 2.2 meters above the ground. ; is the aircraft angle of attack; is the pitching moment coefficient generated by the rate of change of the aircraft's angle of attack; is the rate of change of the aircraft's angle of attack; is the pitching moment coefficient produced by the aircraft elevator; is the horizontal tail deflection angle; is the pitch moment coefficient generated by the pitch angular rate of the aircraft; is the pitch angular rate; is the effect of the aircraft on the pitch moment coefficient, ; is the moment of lift about the center of gravity; The data is calculated based on the shape of the external plug-in through aerodynamics theory; The flap angle takes different values. The data value is determined based on wind data or aerodynamic theory. If there is no flap, this value is 0. The rolling moment solution operator model is: in, is the rolling moment; is dynamic pressure; for wingspan; is the transverse torque coefficient; is the rolling moment coefficient generated by the roll angle; is the aircraft sideslip angle; is the rolling moment coefficient produced by the rudder; For the aircraft rudder; is the rolling moment coefficient generated by the aileron rudder; It is the aileron rudder of the aircraft; is the rolling moment coefficient generated by the rolling angular rate; is the roll angular rate; is the rolling moment coefficient generated by the yaw rate; is the yaw rate; is the rolling moment coefficient generated by the yaw angle acceleration rate; is the aircraft angle of attack; is the true airspeed of flight; 、 、 、 、 、 According to the original aerodynamic data table obtained from the wind test, the difference is obtained by querying the flight altitude and flight speed as input; The yaw moment sub-model is: in, is the yaw moment; is dynamic pressure; for wingspan; is the yaw moment coefficient; is the yaw moment coefficient generated by the roll angle; is the aircraft sideslip angle; is the yawing moment coefficient generated by the rudder; For the aircraft rudder; is the yaw moment coefficient generated by the yaw angular rate; is the yaw rate; is the yaw moment coefficient generated by the roll angular rate; is the roll angular rate; is the true airspeed of flight; is the effect of the engine on the yaw moment coefficient, ; 、 、 、 According to the original aerodynamic data table obtained from the wind test, the difference is obtained by querying the flight altitude and flight speed as input.

4. The method for constructing a robot pilot flight simulation environment according to claim 3, characterized in that: The engine thrust sub-model is: in, is the effective thrust of the engine; is the inlet loss correction coefficient. According to engineering experience data, when the flight Mach number hour, , hour, The thrust of the engine is , unit is N; The fuel consumption quantum model is: in, is the fuel consumption; is the intake duct loss correction factor; the fuel consumption rate is ; The center of gravity solution operator model is: The moment of inertia solver model is: When two 760-liter auxiliary fuel tanks are hung and the fuel capacity is greater than 2378 liters, the three-axis rotational inertia is: When two 760-liter auxiliary fuel tanks are hung and the fuel capacity is less than 2378 liters and greater than 1600 liters, the three-axis rotational inertia is: When there is no external plug-in, the three-axis moment of inertia is: in, For aircraft Shaft moment of inertia, For aircraft Shaft moment of inertia, For aircraft Shaft moment of inertia, 、 The units are ; is the fuel quantity; after the landing gear is lowered, Increased by 18.5 ; The atmospheric data solution operator model is: When the flight altitude is less than 11,000 meters, , When the altitude is greater than 11,000 meters and less than 20,000 meters, , in, is the atmospheric temperature, in °C; is the atmospheric density in kg / m 3 ; is atmospheric pressure, the unit is Pa; is the speed of sound, in m / s; For the flight altitude.

5. The method for constructing a robot pilot flight simulation environment according to claim 4, characterized in that: The dynamics solution model includes a dynamics solution model for the aerial flight segment, a dynamics solution model for the takeoff and run segment, and a dynamics solution model for the landing and run segment; Among them, the dynamics solution model of the aerial flight segment is: The dynamic solution model of the takeoff roll phase is: in, is the total resistance; is the total lift force; is the sum of the lateral forces; is the effective thrust of the engine; For resistance; for lift; is the lateral force; is the angle between the thrust and the aircraft axis; is the aircraft angle of attack; is the aircraft sideslip angle; is the pitch angle of the aircraft; is the aircraft roll angle; is the aircraft gravity; is the aircraft's rolling resistance; Provide support for sliding; The dynamics solution model of the landing roll phase is: in, is the total resistance; is the total lift force; is the sum of the lateral forces; is the total rolling moment; is the total yaw moment; is the total pitching moment; is the horizontal speed of the aircraft; is the longitudinal velocity of the aircraft; is the lateral velocity of the aircraft; is the roll angular rate; is the yaw rate; is the pitch angular rate; For aircraft Shaft moment of inertia; For aircraft Shaft moment of inertia; For aircraft Shaft moment of inertia; is the vehicle's moment of inertia.

6. The method for constructing a robot pilot flight simulation environment according to claim 5, characterized in that: The kinematic solution model is: in, is the aircraft roll angle; is the true heading; is the pitch angle of the aircraft; is the roll angular rate; is the yaw rate; is the pitch angular rate; is the aircraft speed in the Earth axis system Axis component; is the aircraft speed in the Earth axis system Axis component; is the aircraft speed in the Earth axis system Axis component; is the horizontal speed of the aircraft; is the longitudinal velocity of the aircraft; is the lateral velocity of the aircraft.

7. The method for constructing a robot pilot flight simulation environment according to claim 6, characterized in that: The three-axis speed of the earth axis system and the earth speed are converted into variables in the following way: in, is the aircraft speed in the Earth axis system Axis component; is the aircraft speed in the Earth axis system Axis component; is the aircraft speed in the Earth axis system Axis component; is the aircraft ground speed; The airspeed is converted into a variable as follows: in, is the wind speed in the body axis Axis component; is the y-axis component of wind speed in the body axis system; is the wind speed in the body axis Axis component; is the wind speed in the earth axis Axis component; is the wind speed in the earth axis Axis component; is the wind speed in the earth axis Axis component; is the horizontal speed of the aircraft; is the longitudinal velocity of the aircraft; is the lateral velocity of the aircraft; is the true heading; is the pitch angle of the aircraft; is the aircraft roll angle; is the true airspeed of the aircraft.

8. A robot pilot flight simulation platform capable of implementing the method for constructing a robot pilot flight simulation environment according to any one of claims 1 to 7, characterized in that: The system comprises a retired aircraft, original aircraft support equipment, and a simulation system; wherein the simulation system is mounted on the retired aircraft and is in communication with the robot pilot; the original aircraft support equipment is used to raise the retired aircraft to a certain distance from the ground and provide power to the retired aircraft, the robot pilot, and the simulation system; The simulation system includes a simulation computer, a data acquisition component, and a data cable and a power cable for connecting them; wherein, the data acquisition component includes an elevator angle collector, a rudder angle collector, an aileron rudder angle collector, a discrete state video collector, and a throttle handle scale video collector installed at a preset position of the retired aircraft; the simulation computer is configured with an aerodynamic force and torque solution model, a dynamic solution model, and a kinematic solution model; the simulation computer receives the elevator angle, rudder angle, aileron rudder angle, throttle handle position and discrete state information, and uses the aerodynamic force and torque solution model, the dynamic solution model, and the kinematic solution model to solve them in turn, and obtains the three-dimensional position, three-dimensional speed, three-dimensional angle and angular velocity, weight and fuel quantity of the retired aircraft through variable conversion, and outputs the above information to the robot pilot through the data interface.

Images