A method of aircraft rapid formation flight control based on direct force

Abstract

The present invention discloses a method for controlling the rapid formation flight of aircraft based on direct force, which relates to the field of aviation flight control technology, including: designing a torque control channel based on dynamic inversion, a direct force control channel based on dynamic inversion, and a lead aircraft track tracking guidance law; designing an aircraft formation constraint; based on the aircraft formation constraint, designing a wingman consistency formation coordination control law, and completing the rapid formation flight control of aircraft based on direct force. The present invention solves the problem that conventional control methods have large phase lag and equivalent delay, resulting in low control accuracy and efficiency.

Description

Technical Field

[0001] The invention relates to the technical field of aviation flight control, and in particular to a method for controlling rapid formation flight of aircraft based on direct force. Background Art

[0002] Formation flying refers to an organizational model in which multiple aircraft are arranged in formation and assigned tasks according to mission requirements. In order to meet mission requirements, multiple aircraft need to be arranged in a certain way and keep the formation unchanged during the mission execution. Aircraft formation flying has its unique advantages in dealing with diverse tasks and complex environments. It can improve the overall efficiency of mission completion through collaboration between multiple aircraft, division of tasks, resource sharing, etc. Therefore, aircraft formation flight control technology has become a research hotspot in the field of flight control. At the same time, formation flying needs to maintain formation, which brings about the problem that the aircraft needs to accurately and quickly achieve attitude and track control.

[0003] Conventional torque control can achieve precise control of attitude and track angle, but track control based on torque control usually requires first changing the pitch moment and roll moment to adjust the attitude, thereby indirectly changing the track of the aircraft. Therefore, the track control system based on torque control will have a large phase lag and equivalent delay, which is slightly insufficient in terms of the rapidity of track control. Therefore, on the premise of achieving precise track control, how to improve the rapid response of the aircraft track is of great research significance for achieving rapid consistency of formation aircraft. Summary of the invention

[0004] In view of the above-mentioned deficiencies in the prior art, the present invention provides a direct force-based aircraft rapid formation flight control method which solves the problem of low control accuracy and efficiency caused by large phase lag and equivalent delay in conventional control methods.

[0005] In order to achieve the above-mentioned invention object, the technical solution adopted by the present invention is: a method for controlling rapid formation flight of aircraft based on direct force, comprising the following steps:

[0006] S1: Design the aircraft's torque control channel based on dynamic inversion, direct force control channel based on dynamic inversion and long aircraft track tracking guidance law;

[0007] S2: Design aircraft formation constraints;

[0008] S3: Based on the aircraft formation constraints, design the wingman consistency formation collaborative control law to complete the aircraft rapid formation flight control based on direct force.

[0009] Furthermore, the design of the torque control channel based on dynamic inversion in S1 includes the following sub-steps:

[0010] A1: The aircraft state variables are divided into angular rate loop and angle loop according to the singular perturbation theory;

[0011] A2: Take the angular rate loop as the inner loop and the angle loop as the outer loop, design dynamic inverse control laws for the angular rate loop and the angle loop respectively, and complete the design of the torque control channel based on dynamic inversion.

[0012] Furthermore, in A1, the aircraft state variables are divided into an angular rate loop and an angle loop according to the singular perturbation theory, and the formula is:

[0013]

[0014] in, is the aircraft state variable, and are the angular rate state variable and the angle state variable, respectively. , and are the roll angular rate, pitch angular rate and yaw angular rate, respectively. , and are pitch angle, sideslip angle and roll angle respectively, with superscript is the transpose of the matrix;

[0015] The dynamic inverse control law for the diagonal rate loop design in A2 is specifically:

[0016] The differential equation for angular rate is expressed as:

[0017]

[0018] in, is the angular rate state function, is the system angular rate state output function, is the nonlinear coupling torque, is the control torque generated by the rudder surface, is the input vector function, For time, superscript represents the first-order derivative;

[0019]

[0020] in, is the nonlinear coupled rolling moment, is the nonlinear coupled pitching moment, is the nonlinear coupled yaw moment, is the forward moment of inertia of the aircraft, is the lateral moment of inertia of the aircraft, is the vertical moment of inertia of the aircraft, is the product of inertia, , and They are respectively the rolling moment, pitching moment and yaw moment in addition to the moment caused by the deflection of the rudder surface;

[0021]

[0022] in, is the rolling control moment generated by the aileron rudder, is the rolling control torque generated by the rudder, is the pitch control moment produced by the elevator, is the yaw control torque generated by the aileron rudder, is the yaw control torque generated by the rudder;

[0023] Then the dynamic inverse control law of the angular rate loop is:

[0024]

[0025]

[0026] in, is the input vector, is the expected dynamic characteristic of the roll angular rate, is the expected dynamic characteristics of the pitch rate, is the expected dynamic characteristics of the yaw rate, Roll rate bandwidth, is the pitch angle rate bandwidth, is the yaw rate bandwidth, The roll rate control command generated by the dynamic reverse angle loop control law, The pitch rate control command generated by the dynamic reverse angle loop control law, The yaw rate control command generated by the dynamic reverse angle loop control law;

[0027] The dynamic inverse control law for the diagonal loop design in A2 is specifically:

[0028] The angular differential equation is expressed as:

[0029]

[0030] in, is the angular state function, is the system angular state output function, is the nonlinear coupling force on the sideslip angle The impact of is the angular rate state variable and angle state variables The kinematic relationship between is the angular rate state variable changes;

[0031]

[0032]

[0033] in, is the effect of nonlinear coupling force on pitch, is the effect of nonlinear coupling force on sideslip, is the effect of nonlinear coupling force on roll, is the aircraft mass, is the aircraft speed, The dynamic pressure of the aircraft, is the wing area, is the side force coefficient affected by the sideslip angle, is the acceleration due to gravity, is the track inclination angle, is the track roll angle, is the angle of attack, For thrust;

[0034] Then the dynamic inverse control law of the corner loop is:

[0035]

[0036]

[0037] in, is the expected dynamic characteristic of the pitch angle, is the expected dynamic characteristic of the sideslip angle, is the expected dynamic characteristic of the roll angle, Elevation bandwidth, Sideslip angle bandwidth, is the roll angle bandwidth, is the pitch angle control command of the dynamic reverse angle loop control law, The sideslip angle control command is based on the dynamic reverse angle loop control law. is the roll angle control command of the dynamic reverse angle loop control law.

[0038] Furthermore, the design of the direct force control channel based on dynamic inversion in S1 includes the following sub-steps:

[0039] B1: Select the aircraft state track inclination angle and state track azimuth as the control state variables, and the control surfaces are the trailing edge flaps and vertical canards. The formula is:

[0040]

[0041]

[0042] in, To control the state variables, is the state track inclination angle, is the state track azimuth, To control the rudder, For the trailing edge flap, It is a vertical canard;

[0043] Control state variables The differential equation is expressed as:

[0044]

[0045] in, is the control state function, is the system track angle output function, is the nonlinear coupling force affecting the track, To directly control the control force generated by the rudder surface, To control the rudder function;

[0046]

[0047]

[0048] in, is the nonlinear coupling force affecting the vertical track, is the nonlinear coupling force affecting the lateral track, is the lateral force, For lift, To remove the residual lift of the aircraft after the flaps generate lift, In order to remove the residual side force of the aircraft caused by the side force generated by the canard, is the control force generated by the flaps, is the control force generated by the vertical canard, is the flap lift coefficient, is the canard side force coefficient;

[0049] B2: Use nonlinear dynamic inversion theory to calculate the control command signal of the direct force control channel and complete the design of the direct force control channel based on dynamic inversion. The formula is:

[0050]

[0051]

[0052] in, is the expected track tilt angle response, is the expected track azimuth response, is the track tilt angle response bandwidth, is the track azimuth response bandwidth, is the track tilt angle command, is the track azimuth instruction.

[0053] Furthermore, the design of the S1 mid- and long-range aircraft trajectory tracking guidance law includes the following sub-steps:

[0054] C1: Select a track point through the L1 guidance algorithm, calculate the L1 distance between the current position of the aircraft and the track point, and obtain the lateral acceleration instruction based on the speed and position of the aircraft. , the formula is:

[0055]

[0056] in, is the aircraft speed and The angle between the distances;

[0057] C2: According to the lateral acceleration command , calculate the roll angle command required for the aircraft to track the desired trajectory , and the roll angle command Convert to track azimuth command , the formula is:

[0058]

[0059]

[0060] in, and is the proportionality coefficient, is the roll angle command signal of the previous beat, is the Laplace variable;

[0061] C3: Use proportional conversion to obtain altitude instructions , complete the design of the lead aircraft's track tracking guidance law, the formula is:

[0062]

[0063] in, is the proportionality coefficient, is the height instruction, It is the aircraft altitude signal.

[0064] Furthermore, the aircraft formation constraints designed in S2 are specifically:

[0065] S21: Taking the lead aircraft as the reference point, convert the position information of all aircraft in the ground coordinate system into the position information in the lead aircraft body coordinate system. The formula is:

[0066]

[0067] in, For the The forward position of the aircraft in the ground coordinate system, For the The lateral position of the aircraft in the ground coordinate system, No. The forward position of the aircraft in the lead aircraft system, For the The lateral position of the aircraft in the lead aircraft system, is the yaw angle;

[0068] S22: According to the expected formation, position constraint information relative to the lead aircraft is given to obtain Constraint matrix for a formation of aircraft , the formula is:

[0069]

[0070] in, For the The position constraint matrix of the aircraft, , For the The first aircraft The forward position constraint of the aircraft, For the The first aircraft The lateral position constraint of the aircraft, For the The first aircraft The vertical position constraints of the aircraft, ;

[0071] S23: Based on Constraint matrix for a formation of aircraft , design the cooperative control law to realize the control of aircraft formation, the formula is:

[0072]

[0073] in, For the The first aircraft The forward relative position distance of the aircraft, For the The first aircraft The lateral relative position distance of the two aircraft, For the The first aircraft The vertical relative distance between the two aircraft.

[0074] Furthermore, the wingman consistency formation collaborative control law in S3 includes a height consistency formation control law, a lateral consistency formation control law and a forward consistency formation control law.

[0075] Furthermore, the highly consistent formation control law is:

[0076]

[0077] in, For the The track angle command of the aircraft, and is the proportionality coefficient, For the The first aircraft The communication relationship between aircraft, For the The aircraft's flight path inclination angle, For the The flight path inclination angle of the aircraft.

[0078] Furthermore, the lateral consistency formation control law is:

[0079]

[0080] in, For the The track azimuth command of the aircraft, and is the proportionality coefficient, For the The aircraft's track bearing, For the The flight path angle of the aircraft.

[0081] Furthermore, the forward consistency formation control law is:

[0082]

[0083] in, For the The speed command of the aircraft, and is the proportionality coefficient, For the The speed of the aircraft, For the The speed of the aircraft.

[0084] The beneficial effects of the present invention are as follows: the present invention proposes a fast consistency formation flight control method based on direct force. According to the generation and control characteristics of direct force, a nonlinear dynamic inverse control method is used to construct a direct lift control law and a direct side force control law, and the design of the torque control loop and the design of the direct force control loop are realized, thereby realizing the decoupling control between attitude and track, so as to improve the rapidity and accuracy of track response; the "leader-winger" formation control structure is used for formation flight control, and the leader track tracking guidance law is designed. Secondly, according to the formation flight formation constraints, a corresponding wingman consistency formation flight control protocol is designed by a consistency protocol to realize the formation and maintenance of the formation formation. Finally, according to the designed direct force formation flight control system structure, a fast consistency formation flight control based on direct force is realized, the rapidity of the formation formation and the accurate maintenance of the formation are improved, and it is ensured that the aircraft performs formation flight along the corresponding route in a specified formation. BRIEF DESCRIPTION OF THE DRAWINGS

[0085] Figure 1 The present invention is a flow chart of the aircraft rapid formation flight control method based on direct force.

[0086] Figure 2 This is a structural diagram of the aircraft rapid formation flight control system based on direct force of the present invention.

[0087] Figure 3 This is a three-dimensional view of the simulation results of formation flight of six aircraft in a triangle formation.

[0088] Figure 4 This is a two-dimensional view of the simulation results of formation flight of 6 aircraft in a triangle formation.

[0089] Figure 5 This is a diagram of the lateral distances of the five wingmen relative to the lead aircraft.

[0090] Figure 6 This is the altitude and distance diagram of the five wingmen relative to the leader aircraft.

[0091] Figure 7 This is a diagram of the forward distance of the five wingmen relative to the lead aircraft.

[0092] Figure 8 This is the track azimuth response diagram of 6 aircraft during formation flying.

[0093] Fig. 9 This is the track tilt angle response diagram of 6 aircraft during formation flying.

[0094] Fig.10 This is the speed response diagram of 6 aircraft during formation flying. DETAILED DESCRIPTION

[0095] The present invention will be further described below in conjunction with the accompanying drawings and specific embodiments.

[0096] like Figure 1 As shown, a method for controlling rapid formation flight of aircraft based on direct force comprises the following steps:

[0097] S1: Design the aircraft's torque control channel based on dynamic inversion, direct force control channel based on dynamic inversion and long aircraft track tracking guidance law;

[0098] S2: Design aircraft formation constraints;

[0099] S3: Based on the aircraft formation constraints, design the wingman consistency formation collaborative control law to complete the aircraft rapid formation flight control based on direct force.

[0100] The design of the torque control channel based on dynamic inversion in S1 includes the following steps:

[0101] A1: The aircraft state variables are divided into angular rate loop and angle loop according to the singular perturbation theory;

[0102] A2: Take the angular rate loop as the inner loop and the angle loop as the outer loop, design dynamic inverse control laws for the angular rate loop and the angle loop respectively, and complete the design of the torque control channel based on dynamic inversion.

[0103] As the demand for aircraft performance continues to increase, the nonlinear dynamic characteristics of aircraft are becoming more and more complex. Continuing to use traditional linear control methods to design control systems will lose the nonlinear characteristics of the aircraft and cannot meet the needs of complex control tasks. The present invention introduces a nonlinear dynamic inverse control method to solve the problem of nonlinear system control law design. This method is an effective method for solving nonlinear multivariable decoupling control problems, which can meet the high-performance control requirements of aircraft and realize inter-channel decoupling control.

[0104] For aircraft flight control systems, the condition that the input vector and the state vector have the same dimension cannot be met. Therefore, it is necessary to combine the nonlinear dynamic inversion method with the singular perturbation theory and use the partial inverse approximation to solve the problem. The input vector is selected as , determined by the elevator deflection angle , Aileron deflection angle 、Rudder deflection angle When using the dynamic inversion method for design, it is necessary to first separate the state variables so that they meet the dynamic inversion requirements. According to the different time characteristics of the aircraft state variables and the requirements of the direct force control torque channel, the state variables According to the singular perturbation theory, the design of the angular loop and the angular rate loop is divided into two parts.

[0105] In A1, the aircraft state variables are divided into an angular rate loop and an angle loop according to the singular perturbation theory, and the formula is:

[0106]

[0107] in, is the aircraft state variable, and are the angular rate state variable and the angle state variable, respectively. , and are the roll angular rate, pitch angular rate and yaw angular rate, respectively. , and are pitch angle, sideslip angle and roll angle respectively, with superscript is the transpose of the matrix;

[0108] The dynamic inverse control law for the diagonal rate loop design in A2 is specifically:

[0109] The differential equation for angular rate is expressed as:

[0110]

[0111] in, is the angular rate state function, is the system angular rate state output function, is the nonlinear coupling torque, is the control torque generated by the rudder surface, is the input vector function, For time, superscript represents the first-order derivative;

[0112]

[0113] in, is the nonlinear coupled rolling moment, is the nonlinear coupled pitching moment, is the nonlinear coupled yaw moment, is the forward moment of inertia of the aircraft, is the lateral moment of inertia of the aircraft, is the vertical moment of inertia of the aircraft, is the product of inertia, , and They are respectively the rolling moment, pitching moment and yaw moment in addition to the moment caused by the deflection of the rudder surface;

[0114]

[0115] in, is the rolling control moment generated by the aileron rudder, is the rolling control torque generated by the rudder, is the pitch control moment produced by the elevator, is the yaw control torque generated by the aileron rudder, is the yaw control torque generated by the rudder;

[0116] Then the dynamic inverse control law of the angular rate loop is:

[0117]

[0118]

[0119] in, is the input vector, is the expected dynamic characteristic of the roll angular rate, is the expected dynamic characteristics of the pitch rate, is the expected dynamic characteristics of the yaw rate, Roll rate bandwidth, is the pitch angle rate bandwidth, is the yaw rate bandwidth, take , The roll rate control command generated by the dynamic reverse angle loop control law, The pitch rate control command generated by the dynamic reverse angle loop control law, The yaw rate control command generated by the dynamic reverse angle loop control law;

[0120] The dynamic inverse control law for the diagonal loop design in A2 is specifically:

[0121] The angular differential equation is expressed as:

[0122]

[0123] in, is the angular state function, is the system angular state output function, is the nonlinear coupling force on the sideslip angle The impact of is the angular rate state variable and angle state variables The kinematic relationship between is the angular rate state variable changes;

[0124]

[0125]

[0126] in, is the effect of nonlinear coupling force on pitch, is the effect of nonlinear coupling force on sideslip, is the effect of nonlinear coupling force on roll, is the aircraft mass, is the aircraft speed, The dynamic pressure of the aircraft, is the wing area, is the side force coefficient affected by the sideslip angle, is the acceleration due to gravity, is the track inclination angle, is the track roll angle, is the angle of attack, For thrust;

[0127] The control command signal of the angular rate loop can be obtained by using nonlinear dynamic inverse theory;

[0128] Then the dynamic inverse control law of the corner loop is:

[0129]

[0130]

[0131] in, is the expected dynamic characteristic of the pitch angle, is the expected dynamic characteristic of the sideslip angle, is the expected dynamic characteristic of the roll angle, Elevation bandwidth, Sideslip angle bandwidth, is the roll angle bandwidth, take , is the pitch angle control command of the dynamic reverse angle loop control law, The sideslip angle control command is based on the dynamic reverse angle loop control law. is the roll angle control command of the dynamic reverse angle loop control law.

[0132] The design of the direct force control channel based on dynamic inversion in S1 includes the following steps:

[0133] B1: Select the aircraft state track inclination angle and state track azimuth as the control state variables, and the control surfaces are the trailing edge flaps and vertical canards. The formula is:

[0134]

[0135]

[0136] in, To control the state variables, is the state track inclination angle, is the state track azimuth, To control the rudder, For the trailing edge flap, It is a vertical canard;

[0137] Control state variables The differential equation is expressed as:

[0138]

[0139] in, is the control state function, is the system track angle output function, is the nonlinear coupling force affecting the track, To directly control the control force generated by the rudder surface, To control the rudder function;

[0140]

[0141]

[0142] in, is the nonlinear coupling force affecting the vertical track, is the nonlinear coupling force affecting the lateral track, is the lateral force, For lift, To remove the residual lift of the aircraft after the flaps generate lift, In order to remove the residual side force of the aircraft caused by the side force generated by the canard, is the control force generated by the flaps, is the control force generated by the vertical canard, is the flap lift coefficient, is the canard side force coefficient;

[0143] B2: Use nonlinear dynamic inversion theory to calculate the control command signal of the direct force control channel and complete the design of the direct force control channel based on dynamic inversion. The formula is:

[0144]

[0145] Assume that the closed-loop dynamic system expected by the angular state subsystem is a first-order link, that is:

[0146]

[0147] in, is the expected track tilt angle response, is the expected track azimuth response, is the track tilt angle response bandwidth, is the track azimuth response bandwidth, is the track tilt angle command, is the track azimuth instruction.

[0148] The L1 guidance algorithm can be used for both straight and curved track tracking, and has a good tracking effect. The guidance control instructions it obtains are related to speed, and different guidance control instructions can be obtained according to the speed changes of the aircraft in actual flight. Therefore, the track tracking guidance law of the long aircraft is designed using the L1 algorithm.

[0149] The design of the S1 mid- and long-range aircraft trajectory tracking guidance law includes the following steps:

[0150] C1: Select a track point through the L1 guidance algorithm, calculate the L1 distance between the current position of the aircraft and the track point, and obtain the lateral acceleration instruction based on the speed and position of the aircraft. , the formula is:

[0151]

[0152] in, is the aircraft speed and The angle between the distances;

[0153] The relationship between the drone's own gravity and centripetal force can be obtained through coordinated turning:

[0154]

[0155] C2: According to the lateral acceleration command , calculate the roll angle command required for the aircraft to track the desired trajectory , and the roll angle command Convert to track azimuth command , the formula is:

[0156]

[0157]

[0158] in, and is the proportionality coefficient, is the roll angle command signal of the previous beat, is the Laplace variable;

[0159] Since direct force control is directly on the track To control, the above roll angle command needs to be Convert to track azimuth command . Track azimuth command It consists of two parts: the roll angle command and the integral of the difference between the roll angle command and the previous beat. The integral signal of the command change is added to avoid the problem that when there is only proportional control, after the aircraft tracks the track, the roll angle command is calculated to be 0, making the track angle command also 0, causing the aircraft to deviate from the route repeatedly;

[0160] Direct force control for altitude control is also a direct control of the track. For control, only simple proportional conversion is used here;

[0161] C3: Use proportional conversion to obtain altitude instructions , complete the design of the lead aircraft's track tracking guidance law, the formula is:

[0162]

[0163] in, is the proportionality coefficient, is the height instruction, It is the aircraft altitude signal.

[0164] When multiple aircraft are performing flight missions, reasonable formation design is crucial to the successful execution of the mission. The formation not only affects the relative positions and communication links between aircraft, but is also directly related to the overall mission execution efficiency and environmental adaptability. During formation flight, it is necessary to maintain the stability of the formation structure and be able to coordinate the positions of each aircraft according to mission requirements. Therefore, formation control is the core of formation control, and how to accurately describe the position relationship between each aircraft in the formation is the primary issue.

[0165] The relative positions between aircraft are described using Method, through the aircraft and the formation reference point axis, axis, At the same time, in order to avoid the problem that the formation constraints cause the actual formation to change when the formation coordinated flight turns, the lead aircraft is used as a reference point to convert the position information of all aircraft in the ground coordinate system into the position information in the lead aircraft body coordinate system, so as to achieve the matching of the actual formation and the formation constraints.

[0166] The aircraft formation constraints designed in S2 are specifically:

[0167] S21: Taking the lead aircraft as the reference point, convert the position information of all aircraft in the ground coordinate system into the position information in the lead aircraft body coordinate system. The formula is:

[0168]

[0169] in, For the The forward position of the aircraft in the ground coordinate system, For the The lateral position of the aircraft in the ground coordinate system, No. The forward position of the aircraft in the lead aircraft system, For the The lateral position of the aircraft in the lead aircraft system, is the yaw angle;

[0170] S22: According to the expected formation, position constraint information relative to the lead aircraft is given to obtain Constraint matrix for a formation of aircraft , the formula is:

[0171]

[0172] in, For the The position constraint matrix of the aircraft, , For the The first aircraft The forward position constraint of the aircraft, For the The first aircraft The lateral position constraints of the aircraft, For the The first aircraft The vertical position constraints of the aircraft, ;

[0173] S23: Based on Constraint matrix for a formation of aircraft , design the cooperative control law to realize the control of aircraft formation, the formula is:

[0174]

[0175] in, For the The first aircraft The forward relative position distance of the aircraft, For the The first aircraft The lateral relative position distance of the two aircraft, For the The first aircraft The vertical relative position distance of the two aircraft.

[0176] For the aircraft formation system, the consistency control protocol represents the information transmission rules between aircraft. Under the effect of the consistency control protocol, the status and output of all aircraft in the formation will eventually converge to the same. Under the "leader-winger" control structure, the leader is responsible for tracking the route and information transmission. The wingman forms the formation and realizes the consistency of the track and attitude based on the information transmitted by the leader and the information exchanged between the wingmen. Figure 2 shown.

[0177] The wingman consistency formation cooperative control law in S3 includes a height consistency formation control law, a lateral consistency formation control law and a forward consistency formation control law.

[0178] The altitude consistency control law consists of two parts: 1. Consistency under altitude formation constraints; 2. Consistency constraints on the track inclination angles between formation aircraft.

[0179] The highly consistent formation control law is:

[0180]

[0181] in, For the The track angle command of the aircraft, and is the proportionality coefficient, For the The first aircraft The communication relationship between aircraft, For the The aircraft's flight path inclination angle, For the The flight path inclination angle of the aircraft.

[0182] The lateral consistency control law consists of two parts: 1. Consistency under lateral formation constraints; 2. Consistency constraints on the track azimuths between formation aircraft.

[0183] The lateral consistency formation control law is:

[0184]

[0185] in, For the The track azimuth command of the aircraft, and is the proportionality coefficient, For the The aircraft's track bearing, For the The flight path angle of the aircraft.

[0186] The forward consistency control law consists of two parts: 1. Consistency under the forward formation constraint; 2. Consistency constraint on the speed between formation aircraft.

[0187] The forward consistency formation control law is:

[0188]

[0189] in, For the The speed command of the aircraft, and is the proportionality coefficient, For the The speed of the aircraft, For the The speed of the aircraft.

[0190] In one embodiment of the present invention, a simulation is performed to verify the rapid coherent formation flight based on direct force:

[0191] Set the formation, give the corresponding formation constraints, and the communication adjacency matrix between the formations, run the program to perform a fast consistency formation flight simulation verification based on direct force, obtain the position and state information of each aircraft in the formation flight, and verify the effect and performance of the formation consistency flight under direct force control. Figure 3 and Figure 4 It can be seen that under the direct force formation control system, the formation aircraft can quickly gather to form a triangular formation at the beginning, and maintain this formation for four-sided flight, and can still maintain the corresponding formation at the turn. Figure 5 It can be seen that under direct side force control, the wingman can quickly reach the lateral formation constraint, and although the wingman has a certain deviation from the expected formation at the turn, it is quickly corrected; at the same time, Figure 6 It can be seen that under direct lift control, the six aircraft can still accurately maintain their altitude even when performing turning maneuvers, and achieve the formation and maintenance of the longitudinal formation. Figure 8 and Fig. 9 It can be seen that under direct force control, the six aircraft can achieve rapid consistency of track inclination angle and track azimuth angle. It can be seen that the aircraft rapid formation flight control method based on direct force can achieve rapid formation of formation, precise maintenance of formation and rapid consistency of track angle. Figure 7 and Fig.10 It can be seen that the formation aircraft can form and maintain a forward formation and maintain speed consistency, but the effect is slightly worse when the formation turns. This is because the control of the forward formation is achieved by controlling the speed through the engine. However, after completing the turning maneuver, the formation error can be quickly eliminated, and the forward formation can be quickly formed and maintained.

[0192] Those skilled in the art will appreciate that the embodiments described herein are intended to help readers understand the principles of the present invention, and should be understood that the protection scope of the present invention is not limited to such specific statements and embodiments. Those skilled in the art can make various other specific variations and combinations that do not deviate from the essence of the present invention based on the technical revelations disclosed by the present invention, and these variations and combinations are still within the protection scope of the invention.

Claims

1. A method for controlling rapid formation flight of aircraft based on direct force, characterized in that: The following steps are involved: S1: Design the aircraft's torque control channel based on dynamic inversion, direct force control channel based on dynamic inversion and long aircraft track tracking guidance law; The design of the direct force control channel based on dynamic inversion in S1 includes the following steps: B1: Select the aircraft state track inclination angle and state track azimuth as the control state variables, and the control surfaces are the trailing edge flaps and vertical canards. The formula is: in, To control the state variables, is the state track inclination angle, is the state track azimuth, with superscript is the transpose of the matrix, To control the rudder, For the trailing edge flap, It is a vertical canard; Control state variables The differential equation is expressed as: in, is the control state function, is the system track angle output function, is the nonlinear coupling force affecting the track, To directly control the control force generated by the rudder surface, To control the rudder function, For time, superscript represents the first-order derivative, is the aircraft state variable; in, is the nonlinear coupling force affecting the vertical track, is the nonlinear coupling force affecting the lateral track, is the aircraft mass, is the aircraft speed, For thrust, is the track roll angle, is the angle of attack, is the sideslip angle, is the lateral force, For lift, To remove the residual lift of the aircraft after the flaps generate lift, is the acceleration due to gravity, In order to remove the residual side force of the aircraft caused by the side force generated by the canard, is the control force generated by the flaps, is the control force generated by the vertical canard, is the flap lift coefficient, is the canard side force coefficient; B2: Use nonlinear dynamic inversion theory to calculate the control command signal of the direct force control channel and complete the design of the direct force control channel based on dynamic inversion. The formula is: in, is the expected track tilt angle response, is the expected track azimuth response, is the track tilt angle response bandwidth, is the track azimuth response bandwidth, is the track tilt angle command, is the track azimuth instruction; S2: Design aircraft formation constraints; S3: Based on the aircraft formation constraints, design the wingman consistency formation collaborative control law to complete the aircraft rapid formation flight control based on direct force.

2. The method for controlling rapid formation flight of aircraft based on direct force according to claim 1, characterized in that: The design of the torque control channel based on dynamic inversion in S1 includes the following steps: A1: The aircraft state variables are divided into angular rate loop and angle loop according to the singular perturbation theory; A2: Take the angular rate loop as the inner loop and the angle loop as the outer loop, design dynamic inverse control laws for the angular rate loop and the angle loop respectively, and complete the design of the torque control channel based on dynamic inversion.

3. The method for controlling rapid formation flight of aircraft based on direct force according to claim 2, characterized in that: In A1, the aircraft state variables are divided into an angular rate loop and an angle loop according to the singular perturbation theory, and the formula is: in, is the aircraft state variable, and are the angular rate state variable and the angle state variable, respectively. , and are the roll angular rate, pitch angular rate and yaw angular rate, respectively. , and are pitch angle, sideslip angle and roll angle respectively, with superscript is the transpose of the matrix; The dynamic inverse control law for the diagonal rate loop design in A2 is specifically: The differential equation for angular rate is expressed as: in, is the angular rate state function, is the system angular rate state output function, is the nonlinear coupling torque, is the control torque generated by the rudder surface, is the input vector function, For time, superscript represents the first-order derivative; in, is the nonlinear coupled rolling moment, is the nonlinear coupled pitching moment, is the nonlinear coupled yaw moment, is the forward moment of inertia of the aircraft, is the lateral moment of inertia of the aircraft, is the vertical moment of inertia of the aircraft, is the product of inertia, , and They are respectively the rolling moment, pitching moment and yaw moment in addition to the moment caused by the deflection of the rudder surface; in, is the rolling control moment generated by the aileron rudder, is the rolling control torque generated by the rudder, is the pitch control moment produced by the elevator, is the yaw control torque generated by the aileron rudder, is the yaw control torque generated by the rudder; Then the dynamic inverse control law of the angular rate loop is: in, is the input vector, is the expected dynamic characteristic of the roll angular rate, is the expected dynamic characteristics of the pitch rate, is the expected dynamic characteristics of the yaw rate, Roll rate bandwidth, is the pitch angle rate bandwidth, is the yaw rate bandwidth, The roll rate control command generated by the dynamic reverse angle loop control law, The pitch rate control command generated by the dynamic reverse angle loop control law, The yaw rate control command generated by the dynamic reverse angle loop control law; The dynamic inverse control law for the diagonal loop design in A2 is specifically: The angular differential equation is expressed as: in, is the angular state function, is the system angular state output function, is the nonlinear coupling force on the sideslip angle The impact of is the angular rate state variable and angle state variables The kinematic relationship between is the angular rate state variable changes; in, is the effect of nonlinear coupling force on pitch, is the effect of nonlinear coupling force on sideslip, is the effect of nonlinear coupling force on roll, is the aircraft mass, is the aircraft speed, The dynamic pressure of the aircraft, is the wing area, is the side force coefficient affected by the sideslip angle, is the acceleration due to gravity, is the track inclination angle, is the track roll angle, is the angle of attack, For thrust; Then the dynamic inverse control law of the corner loop is: in, is the expected dynamic characteristic of the pitch angle, is the expected dynamic characteristic of the sideslip angle, is the expected dynamic characteristic of the roll angle, Elevation bandwidth, Sideslip angle bandwidth, is the roll angle bandwidth, is the pitch angle control command of the dynamic reverse angle loop control law, The sideslip angle control command is based on the dynamic reverse angle loop control law. is the roll angle control command of the dynamic reverse angle loop control law.

4. The method for controlling rapid formation flight of aircraft based on direct force according to claim 3, characterized in that: The design of the S1 mid- and long-range aircraft trajectory tracking guidance law includes the following steps: C1: Select a track point through the L1 guidance algorithm, calculate the L1 distance between the current position of the aircraft and the track point, and obtain the lateral acceleration instruction based on the speed and position of the aircraft. , the formula is: in, is the aircraft speed and The angle between the distances; C2: According to the lateral acceleration command , calculate the roll angle command required for the aircraft to track the desired trajectory , and the roll angle command Convert to track azimuth command , the formula is: in, and is the proportionality coefficient, is the roll angle command signal of the previous beat, is the Laplace variable; C3: Use proportional conversion to obtain altitude instructions , complete the design of the lead aircraft's track tracking guidance law, the formula is: in, is the proportionality coefficient, is the height instruction, It is the aircraft altitude signal.

5. The method for controlling rapid formation flight of aircraft based on direct force according to claim 4, characterized in that: The aircraft formation constraints designed in S2 are specifically: S21: Taking the lead aircraft as the reference point, convert the position information of all aircraft in the ground coordinate system into the position information in the lead aircraft body coordinate system. The formula is: in, For the The forward position of the aircraft in the ground coordinate system, For the The lateral position of the aircraft in the ground coordinate system, No. The forward position of the aircraft in the lead aircraft system, For the The lateral position of the aircraft in the lead aircraft system, is the yaw angle; S22: According to the expected formation, position constraint information relative to the lead aircraft is given to obtain Constraint matrix for a formation of aircraft , the formula is: in, For the The position constraint matrix of the aircraft, , For the The first aircraft The forward position constraint of the aircraft, For the The first aircraft The lateral position constraints of the aircraft, For the The first aircraft The vertical position constraints of the aircraft, ; S23: Based on Constraint matrix for a formation of aircraft , design the cooperative control law to realize the control of aircraft formation, the formula is: in, For the The first aircraft The forward relative position distance of the aircraft, For the The first aircraft The lateral relative position distance of the two aircraft, For the The first aircraft The vertical relative distance between the two aircraft.

6. The method for controlling rapid formation flight of aircraft based on direct force according to claim 5, characterized in that: The wingman consistency formation cooperative control law in S3 includes a height consistency formation control law, a lateral consistency formation control law and a forward consistency formation control law.

7. The method for controlling rapid formation flight of aircraft based on direct force according to claim 6, characterized in that: The highly consistent formation control law is: in, For the The track angle command of the aircraft, and is the proportionality coefficient, For the The first aircraft The communication relationship between aircraft, For the The aircraft's flight path inclination angle, For the The flight path inclination angle of the aircraft.

8. The method for controlling rapid formation flight of aircraft based on direct force according to claim 7, characterized in that: The lateral consistency formation control law is: in, For the The track azimuth command of the aircraft, and is the proportionality coefficient, For the The aircraft's track azimuth, For the The flight path angle of the aircraft.

9. The method for controlling rapid formation flight of aircraft based on direct force according to claim 8, characterized in that: The forward consistency formation control law is: in, For the The speed command of the aircraft, and is the proportionality coefficient, For the The speed of the aircraft, For the The speed of the aircraft.

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