Control system of tilt-rotor aircraft

Abstract

The invention, which relates to the technical field of aircraft control, discloses a control system for a tilt-rotor aircraft, comprising: a control instruction unit for sending an aircraft control instruction; the flight management unit is used for generating a guidance resolving instruction; the outer ring control unit is used for receiving a guidance resolving instruction and generating a flight control quantity and a pitching / rolling angle instruction; the course control unit is used for receiving a yaw angle rate and a track / course angle instruction for coordinating turning, and the inner ring attitude control unit is used for receiving a pitch angle instruction and a roll angle instruction and generating a corresponding channel control quantity; the control distribution unit is used for receiving all channel control quantities and distributing driving actuators; and the sensor feedback unit is used for monitoring attitude, speed and position information of the aircraft in real time. Through a highly integrated control system and an intelligent control algorithm, the operation process of the aircraft is simplified, the operation difficulty is reduced, and the user experience is improved.

Description

Technical Field

[0001] This invention relates to the field of aircraft control technology, specifically a control system for a tiltrotor aircraft. Background Technology

[0002] With the rapid development of the low-altitude economy, the research and development of vertical takeoff and landing (VTOL) manned aircraft has become increasingly diverse. Currently, most mainstream aircraft configurations adopt tiltrotor configurations to meet the requirements of vertical takeoff and landing and high-speed forward flight. For manned aircraft, ensuring the safety of the controller and its movement according to predetermined response characteristics is crucial. However, tiltrotor aircraft need to switch or transform between rotor mode, transition mode, and fixed-wing mode during flight, which poses a significant challenge to controller design. Traditional methods mainly employ cascade control strategies, that is, designing one rotor controller and one fixed-wing controller, and switching weights according to changes in airspeed. While this control method can meet the functional flight test requirements of unmanned aerial vehicles, its safety and airworthiness are difficult to guarantee in manned aircraft, and it cannot meet the flight quality and handling control quality requirements for airworthiness.

[0003] Therefore, how to rationally design a control architecture suitable for manned tilt-layout aircraft, making it safer and able to meet airworthiness requirements to achieve the goal of airworthiness certification, has become a pressing technical and practical problem to be solved in this field. Summary of the Invention

[0004] To achieve the above objectives, the present invention provides the following technical solution:

[0005] A control system for a tiltrotor aircraft includes an aircraft body, a tilting mechanism, a rotor drive device, and a control system, wherein the control system includes:

[0006] The control command unit is used to issue aircraft control commands; it generates control commands through the pilot's control stick or autopilot system, and the commands include, but are not limited to, takeoff, landing, hovering, forward movement, and turning.

[0007] The flight management unit is used to receive aircraft control commands and generate guidance calculation commands based on the aircraft control commands.

[0008] The outer ring control unit is used to receive guidance calculation commands sent by the flight management unit, and generate flight control quantities, pitch angle commands, and roll angle commands according to the guidance calculation commands.

[0009] The heading control unit is used to receive the yaw rate command for coordinated turning sent by the outer ring control unit and the track angle control command or heading angle control command sent by the flight management unit, and generate a heading control quantity according to the yaw rate command.

[0010] The inner ring attitude control unit is used to receive pitch angle commands and roll angle commands sent by the outer ring control unit, and generate pitch channel control quantities and roll channel control quantities according to the pitch angle commands and roll angle commands respectively.

[0011] The control allocation unit, which is connected to the outer loop control unit, the heading control unit, and the inner loop attitude control unit, is used to receive the forward control quantity, pitch channel control quantity, roll channel control quantity, altitude control quantity, and heading control quantity, and to complete the control quantity allocation according to the forward control quantity, pitch channel control quantity, roll channel control quantity, altitude control quantity, and heading control quantity, so as to drive the corresponding actuator to perform the corresponding action.

[0012] The sensor feedback unit, located on the aircraft body, is connected to the outer ring control unit, heading control unit, inner ring attitude control unit, and control distribution unit. It is used to monitor the aircraft's attitude, speed, and position information in real time and transmit the monitoring data to the flight management unit.

[0013] Compared with the prior art, the present invention has the following beneficial effects:

[0014] This invention simplifies the operation of aircraft through a highly integrated control system and intelligent control algorithms. Pilots only need to issue basic control commands, and the system can automatically complete complex guidance calculations and control variable allocation, reducing operational difficulty and improving the user experience.

[0015] Specifically, through the coordinated operation of multiple control units (flight management unit, outer loop control unit, heading control unit, and inner loop attitude control unit), this system can achieve precise control of the aircraft's attitude, speed, and position. The sensor feedback unit monitors the aircraft's status in real time and transmits the data to the flight management unit, ensuring the real-time nature and accuracy of control commands.

[0016] Meanwhile, this invention, through the diversification of guidance calculation commands (including forward ground speed commands, airspeed commands, altitude commands, etc.), can flexibly respond to different flight missions and operating conditions. The coordinated action of the outer loop control unit and the heading control unit enables the aircraft to maintain good attitude control and heading stability during coordinated turns. Furthermore, the control allocation unit in the system rationally allocates various control variables (forward control variables, altitude control variables, roll angle control variables, etc.) to drive the corresponding actuators to perform actions, thereby achieving a smooth transition between tiltrotor and fixed-wing modes, greatly enhancing the aircraft's maneuverability and adaptability. Attached Figure Description

[0017] Figure 1 A schematic diagram of the overall control architecture of a tiltrotor aircraft;

[0018] Figure 2 This is a schematic diagram of the angle loop control circuit for a tiltrotor aircraft.

[0019] Figure 3 A schematic diagram of the angle ring reference model for a tiltrotor aircraft;

[0020] Figure 4 A comparison diagram of angle loop commands and responses for a tiltrotor aircraft;

[0021] Figure 5 This is a flowchart of the rate control process for a tiltrotor aircraft.

[0022] In the diagram, 1 is the control command unit, 2 is the flight management unit, 3 is the outer loop control unit, 4 is the heading control unit, 5 is the inner loop attitude control unit, 6 is the control allocation unit, 7 is the sensor feedback unit, and 8 is the command filtering unit. Detailed Implementation

[0023] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0024] Example 1:

[0025] like Figure 1 As shown, this embodiment provides a control system for a tiltrotor aircraft, including an aircraft body, a tilting mechanism, a rotor drive device, and a control system. The control system includes:

[0026] The control command unit 1 is used to issue aircraft control commands; the control commands are generated through the pilot's control stick or autopilot system, and the commands include, but are not limited to, takeoff, landing, hovering, forward movement, and turning.

[0027] Flight management unit 2 is used to receive aircraft control commands, convert the control commands into specific guidance calculation commands through a built-in flight control algorithm, and generate guidance calculation commands based on the aircraft control commands. In this embodiment, the guidance calculation commands include one or more of the following: forward ground speed command, airspeed command, altitude command, side offset command, side velocity command, track angle command, heading angle command, and roll angle command.

[0028] The outer loop control unit 3 is used to receive guidance calculation commands sent by the flight management unit 2, and convert the guidance calculation commands into specific flight control quantities and pitch angle commands and roll angle commands through the outer loop control algorithm. In this embodiment, the flight control quantities include one or more of the following: forward control quantity, altitude control quantity, roll angle control quantity, pitch angle control quantity, and yaw rate command for coordinated turning.

[0029] The heading control unit 4 is used to receive the yaw rate command for coordinated turning sent by the outer ring control unit 3 and the track angle control command or heading angle control command sent by the flight management unit 2. Through the heading control algorithm, it integrates the yaw rate command and the track angle / heading angle control command to generate a precise heading control quantity.

[0030] The inner-loop attitude control unit 5 receives pitch and roll commands from the outer-loop control unit 3. Through the inner-loop attitude control algorithm, it converts these commands into specific pitch and roll channel control quantities, ensuring stable and precise attitude control of the aircraft. It serves as a crucial bridge between the outer-loop control unit and the inner-loop attitude control unit. The specific inner-loop attitude control algorithm includes a command smoothing system and a delay-based state feedback gain algorithm. The command smoothing system, acting as a key bridge between the outer-loop control unit 3 and the inner-loop attitude control unit 5, transforms the original angle commands (potentially containing step jumps) into physically feasible third-order reference commands. It receives the original attitude commands (including pitch and roll commands) from the outer-loop control unit 3, outputs reference angle commands, smoothed attitude angles for position tracking reference, outputs reference angular rate commands for the desired rotational speed to provide dynamic feedforward, and outputs reference angular acceleration commands for the desired rotational acceleration to compensate for system inertia. The inner-loop attitude control unit receives reference angle commands, reference angular rate commands (and possibly reference angular acceleration commands) from the command smoothing system, and actual angles and angular rates from the sensor feedback unit. Specifically, in tiltrotor aircraft, delays in signal transmission, computation, and actuator response can degrade the control system's performance (e.g., phase lag, increased overshoot). Therefore, when calculating control quantities in the inner-loop attitude control algorithm of the inner-loop attitude control unit 5, a state feedback control law is used to reduce or eliminate the impact of delays on system stability and performance. Delay-based state feedback gain design is one of the key technologies for achieving high-precision attitude control in the inner-loop attitude control unit 5, used to compensate for the impact of delays on angle commands. It directly affects the inner-loop control algorithm, ensuring that the system can still track commands quickly, stably, and accurately even with delays.

[0031] The control allocation unit 6, connected to the outer ring control unit 3, the heading control unit 4, and the inner ring attitude control unit 5, receives the forward control quantity, pitch channel control quantity, roll channel control quantity, altitude control quantity, and heading control quantity. It then allocates the control quantity according to these quantities to drive the corresponding actuators to perform corresponding actions. Through a control allocation algorithm, it rationally distributes each control quantity to the corresponding actuators, including one or more of the following: rotor drive motor speed control, fixed wing control surface servo control, and tilt rotor tilt mechanism control, thereby changing the overall flight state or position of the aircraft.

[0032] The sensor feedback unit 7, located on the aircraft body, connects to the outer loop control unit 3, the heading control unit 4, the inner loop attitude control unit 5, and the control distribution unit 6. It is used to monitor the aircraft's attitude, speed, and position information in real time and transmit the monitoring data to the flight management unit 2. In this embodiment, the sensor feedback unit 7 includes, but is not limited to, IMU (gyroscope, accelerometer), magnetometer, GPS / GNSS (including RTK), barometer, pitot tube, laser / ultrasonic altimeter (near ground), visual / LiDAR (SLAM, obstacle avoidance navigation), tilt mechanism angle sensor, etc. By using multiple sensors (such as gyroscopes, accelerometers, GPS, etc.) to collect the aircraft's attitude, speed, and position information in real time, the control system has a high-precision feedback mechanism, improving control performance.

[0033] The rotor drive device control module includes: a rotor drive motor for driving the rotation of the rotor; and a rotor speed sensor for real-time monitoring of the rotor speed and transmitting the monitoring data to the flight management unit 2. The rotor drive device control module controls the speed and direction of the rotor drive motor according to the control signal generated by the flight management unit 2 to achieve precise control of the rotor.

[0034] Furthermore, a command filtering unit 8 is provided between the control command unit 1 and the flight management unit 2 to filter the control commands and convert them into a series of specific control command signals. Because the operator's control actions may have a certain degree of randomness and instability, and external environmental factors may interfere with signal transmission, the original control command signals may contain noise or other unwanted frequency components. To ensure the accuracy and reliability of the control command signals received by the flight management unit 2, the command filtering unit 8 is introduced into the signal transmission link. This unit, through its built-in filtering algorithm and hardware circuitry, can effectively filter out noise components in the control command signals, smooth signal fluctuations, and thus improve signal purity and stability. The design of the command filtering unit 8 comprehensively considers real-time performance and filtering effect, employing advanced digital filtering technology to dynamically filter the control command signals in real time. This ensures both real-time signal transmission and that the filtered signal accurately reflects the operator's control intentions. The command filtering unit 8 also possesses a certain degree of self-adaptability, automatically adjusting filtering parameters according to actual operating conditions and signal characteristics to adapt to different control command inputs and changes in the flight environment, thereby further enhancing the flexibility and robustness of the entire flight control system.

[0035] Example 2

[0036] The only difference between this embodiment and Embodiment 1 is the design of the anti-delay control and the reference model. The yaw rate command for coordinated turning is calculated using the following formula:

[0037]

[0038] Where, r cmd To coordinate the yaw rate during turns; u is the forward velocity of the aircraft; v is the lateral velocity of the aircraft; V is the airspeed of the aircraft, and V > V_r_min, where V_r_min is the minimum yaw rate control speed, which is the lowest airspeed at which the aircraft's yaw rate control system can be triggered and begin to work effectively during flight. It is a technical term in the field of aeronautical engineering, applied in flight control systems, automatic flight systems, yaw damper systems, and other related fields. It is slightly greater than the minimum airspeed at which the rudder begins to have control effect (rudder effectiveness speed), and is quantitatively determined through theoretical calculations and simulations based on the aircraft characteristics; k is the sideslip control gain; β is the feedback sideslip angle; acc_y is the lateral acceleration of the aircraft; g is the gravitational acceleration; and θ is the pitch angle. For roll angle, r stick The yaw rate command given to the pilot for active control, where k_coff is the airspeed adaptation coefficient that varies with airspeed.

[0039] The airspeed adaptation coefficient k_coff is obtained by the following formula:

[0040]

[0041] That is, when V ≤ V_r_min, k_coff = 0;

[0042] When V_r_min < V < V_r_max, k_coff = smoothly increases from 0 to 1 with V 2 and increases smoothly from 0 to 1;

[0043] When V ≥ V_r_max, k_coff = 1.

[0044] Among them, V_r_min is the minimum yaw rate starting speed selected according to the aircraft characteristics, and V_r_max is the fixed-wing cruise speed selected according to the aircraft characteristics.

[0045] The pitch angle and roll angle enter the inner-loop attitude control unit 5, calculate the control amounts of the roll channel and pitch channel and enter the control allocation; the control allocation unit 6 can use one or several optimization algorithms such as the least squares method, quadratic programming method, and active set method to allocate the control amount requirements to the motor speed, fixed-wing rudder surface, or tilt angle / angle rate, while avoiding the problem of actuator anti-saturation, ensuring that the required control amount is maximally achieved. Among them, the least squares method is often used to solve the problems of sensor redundancy or control allocation in the flight control system; the quadratic programming method is used to solve the problem of multi-objective control amount allocation with constraints; the active set method is the core numerical algorithm for solving the quadratic programming method with linear constraints and is used in flight control to solve the optimal allocation problem of rudder surface commands with high precision and high efficiency.

[0046] Regarding the flight quality requirements for the fixed-wing and multi-rotor states, to meet the airworthiness requirements, reasonable feedforward commands can be provided by designing a feedforward command reference model. Finally, in combination with the feedback controller, the responses of different loops can perfectly track the commands and have anti-disturbance capabilities. The specific steps for obtaining the feedforward command reference model are as follows:

[0047] Step a: Construct a second-order transfer function model:

[0048] Equivalent the response characteristics of the attitude command attitude hold (ACAH) control response to a second-order link. The calculation formula is: In the formula, G ref (s) is the reference model, θ(s) is the Laplace transform of the actual attitude angle (pitch angle, roll angle or yaw angle), θ cmd (s) is the Laplace transform of the command attitude angle, ω act is the equivalent natural frequency, ∈ act is the equivalent damping ratio;

[0049] Step b: Determine model parameters:

[0050] Since the flight quality of attitude and altitude control is strongly correlated with the motor's rotational speed response and lift response under load, the initial values ​​of the attitude channel's ω are selected based on airworthiness standards, taking into account the response time constants of different motors and different flight quality standards. act Between 1.4 and 4.8 rad / s, the ω of the heading channel act Less than 1 rad / s, ω of the height channel act Between 1 and 5 rad / s;

[0051] Based on the time constant correction formula for the dynamic system:

[0052] In the formula ω' act The corrected natural frequency is given by Tm, which is the time constant of the power system. The damping ratio is selected according to airworthiness standards. act Generally between 0.8 and 1;

[0053] Step c: Generate feedforward instructions:

[0054] Generate feedforward instructions using an approximate inverse model:

[0055]

[0056] Where u ff This is a feedforward control command. It is the reference model G ref The approximate inverse model of (s), τ f The time constant of the low-pass filter is denoted as . A low-pass filter is connected in series after the ideal inverse model to transform the higher-order derivative into an actualable approximate inverse model. The feedforward command generated by this approximate inverse model can enable the system to respond quickly when tracking commands, thereby improving the system's robustness and disturbance rejection capability.

[0057] Figure 3 This is a command smoothing system based on a second-order reference model, used to generate physically feasible attitude command trajectories. Its core function is to improve tracking performance and suppress latency effects through feedforward compensation. Based on the second-order reference model, it smooths the original attitude commands and outputs reference angle commands, reference angular rate commands, and reference angular acceleration commands as feedback to improve the tracking capability of the desired signal. Specifically, the input angle command passes through two integrators to obtain the reference angle command, reference angular rate command, and reference angular acceleration command. K3 amplifies the angle error (the difference between the input command and the current reference angle output) and includes it as part of the angular acceleration command. K4 feeds back the reference angular rate and multiplies it by a damping coefficient to suppress oscillations. K3 represents the system stiffness gain, which determines the response speed. The larger the K3 value, the faster the response. K4 = 2∈ act ω act K4 represents the system damping gain, used to suppress oscillations, ω act The equivalent natural frequency is determined by the actuator bandwidth, ∈ act For the equivalent damping ratio, the damping ratio ∈ act The value range is 0.8-1.

[0058] Due to the transmission delay of actuators and sensors, the gain margin and phase margin of the system are reduced, which affects the stability of the system. Therefore, if the transmission delay is not considered, the designed controller may become unstable. Thus, this embodiment is based on the state feedback gain design of the delay to actively compensate for the delay effect. The state feedback control gain matrix [K1, K2] is calculated by combining the reference and feedback angle command and the feedback angular rate command, where K1 is the attitude angle feedback gain (compensating for the angle command delay) and K2 is the angular rate feedback gain (compensating for the angular rate command delay).

[0059] Figure 2This paper demonstrates a hierarchical closed-loop control structure for the angle loop of a tiltrotor aircraft. Its core is the generation of smoothing commands through a smoothing system reference model, combined with state feedback gain to achieve high-precision attitude tracking. Specifically, the operator or autopilot system generates raw angle commands (e.g., pitch angle commands) and inputs them into the smoothing system reference model. The smoothing system reference model smooths the raw commands, generating a physically achievable reference trajectory (avoiding actuator saturation or system oscillation caused by step commands). The smoothing system reference model outputs third-order command reference commands: angle reference commands, angular rate reference commands, and angular acceleration reference commands. The inner-loop attitude control unit 5 receives the reference trajectory output by the smoothing system reference model, as well as the actual angle and actual angular rate from the sensors. It obtains the control quantity through the state feedback control gain matrix, and then transforms it through the control allocation matrix scaling factor K5. The scaled control quantity is then input to the control allocation unit 6. The control allocation unit 6 allocates the generalized control quantity to each actuator (e.g., motors, control surfaces, tilt mechanism) based on the current aircraft configuration (e.g., rotor tilt angle) and the availability of actuators. For example, during hovering, the control allocation primarily distributes the pitch torque demand to the rotor differential; during forward flight, it is allocated to the elevator. The control allocation output commands are fed into the actuator model (typically including the actuator's dynamic response and saturation limits). The actuator model simulates the actual actuator response; its output forces and torques act on the aircraft body, generating linear and angular accelerations through rigid body dynamics equations, which in turn integrally change the aircraft's attitude (angle) and angular rate. The aircraft's actual angle and angular rate are measured by sensor feedback unit 7; sensor feedback unit 7 feeds back to the inner-loop attitude control unit 5 to form a closed loop. The actual angle and angular rate are fed back to the state feedback control gain matrix, compared with the reference values ​​generated by the smoothing system reference model, forming an error signal, which in turn generates the control input.

[0060] The state feedback control gain matrix is ​​obtained based on a method for calculating the state feedback gain with time delay, the calculation method including the following steps:

[0061] Step S1, System Modeling

[0062] First, consider the delays of the actuator and sensor. The actuator can be approximated as a first-order inertial system with a pure time delay, and the settling time of this first-order inertial system with a pure time delay is Ts. For ease of analysis, a second-order element approximation is used. The sensor filtering also uses a second-order element approximation. Therefore, a second-order transfer function with a damping ratio of 1 can be approximately equivalent to a first-order time delay.

[0063]

[0064] Where, ωn Let be the natural frequency of the second-order transfer function. For a real actuator and sensor with a damping ratio of 1, the two second-order delays can be approximated as:

[0065]

[0066] Where, ω actuator ω is the natural frequency of the second-order transfer function of the equivalent actuator's operating element (e.g., motor speed response, servo response). filter The natural frequency of the second-order transfer function of the equivalent sensor filter;

[0067] Step S2: Extended-dimensional state space modeling

[0068] If we assume the desired response of the entire tiltrotor aircraft's control system (order n=2) is... This is used to describe the attitude response (such as pitch angle and roll angle) of a tiltrotor aircraft. Here, b is the system gain, and the larger the value of b, the higher the tracking accuracy. s is the Laplace differential operator, and the desired pole of the system is P, which controls the response speed. By scheduling P and b, the airworthiness requirements of multi-rotor (stability) and fixed-wing (agility) can be uniformly met. The actuator is approximated as a first-order inertial system with pure time delay, and the approximate first-order time delay of the system is Ts.

[0069] Calculate the A and B matrices of the extended-dimensional system based on the system order n=2 and the first-order delay system denoted as Ts:

[0070]

[0071] Where A is the state feedback matrix and B is the mapping from control input to state. The full state feedback and pole placement are calculated, that is, the state space is expanded based on A, B and the desired poles of the system. The delay element is regarded as part of the system to realize state augmentation and obtain state feedback gain.

[0072] Step S3: Calculation of the controllable canonical matrix

[0073] Calculate matrix P - =[P;P] characteristic polynomial f(λ) = |λE-P - |=λ 2 +2*P*λ+P 2 And the remaining coefficient a after removing the coefficients of higher-order terms. d =[P 2 [2P]; Then, based on matrices A and B, the controllable standard form A is calculated and obtained. n 、B n and T n , where P - Let λ be the expected pole matrix, λ be the eigenvalue variables, E be the identity matrix, and a be the eigenvalue matrix. dLet A be the coefficient vector of the desired characteristic polynomial. n B is a system matrix in standard controllable form (an n*n matrix); n The input matrix is ​​in controllable standard form; T n Let A be a transformation matrix, and satisfy A n =T n -1 AT n ;

[0074] Step S4: Controllability Verification

[0075] Determine whether the system is controllable based on the following criteria:

[0076] Let the controllability matrix be s = [B AB…A n-1 [B] Matrix s is an n×(n×m) matrix (where n is the state dimension and m is the input dimension), used to determine whether the system is completely controllable. If the rank of matrix s is n, the system is controllable; if the rank of the matrix is ​​less than n, the system is uncontrollable. When the system is controllable, the controllable canonical form A is calculated. n Transformation matrix T n :

[0077]

[0078] in, This is to take the last row of the inverse of the controllability matrix.

[0079]

[0080] A n =T n -1 AT n

[0081] B n =T n -1 B

[0082] Step S5: Gain Calculation

[0083] For a second-order system After obtaining matrix A n After that, A can be calculated. n The corresponding characteristic polynomial is f(λ) = |λE-An| = λ 2 +a1λ+a0, where E is the identity matrix, and assuming the poles of the expectation are μ, then the characteristic polynomial of the expectation is f(λ)=λ 2 +2*μ*λ+μ 2 And take the coefficients of the last two, the actual characteristic polynomial coefficients a n =[μ 2 ;2μ];

[0084] Based on the feedback gain K of the controllable canonical space n =a d -a n Calculateable and controllable standard type A n Feedback gain K n And based on the feedback gain K n Calculate the state feedback control gain matrix [K1,K2]:

[0085] [K1,K2]=K n *T n =[P 2 -μ 2 ;2P-2μ]*T n n = 2;

[0086] The attitude angle of the aircraft is controlled based on the state feedback control gain matrix [K1,K2].

[0087] Calculate the characteristic polynomial of the system matrix under the controllable canonical form and extract the coefficients;

[0088] Calculate the feedback gain in the controllable canonical form based on the desired pole locations;

[0089] The state feedback control gain matrix [K1,K2] is transformed into the original state space by the transformation matrix, and the attitude angle of the aircraft is controlled according to the state feedback control gain matrix.

[0090] The control system of the tiltrotor aircraft in this embodiment also includes an airworthiness performance verification module, which is used to verify the tracking performance of the command smoothing system through a step response test; wherein the verification method includes the following steps:

[0091] Input a 1-degree step attitude command into the smoothing system reference model and compare the reference angle trajectory output by the smoothing system reference model with the actual angle trajectory fed back by the sensor.

[0092] The control system is deemed compliant when the overlap between the actual trajectory and the reference trajectory is ≥95% and the following airworthiness mode requirements are met.

[0093] Translational Rate Command (TRC) Response Mode: For airworthiness testing of the Translational Rate Command (TRC), its step response is generally required to exhibit first-order system characteristics, with a rise time t... r ∈[2.5s,5s], overshoot = 0%;

[0094] Vertical Rate Height Hold Mode: The Vertical Rate Command Height Hold (RCHH) mode requires its step response to exhibit first-order system characteristics, with a rise time t r ≤5s, pure delay <0.2s, overshoot = 0%;

[0095] Rate Command Direction Hold (RCDH) mode: This verification method is applicable to the Rate Command Direction Hold (RCDH) mode and verifies the step response characteristics according to customized parameters.

[0096] Figure 4 The comparison between the output angle of a 1-degree step command after processing by the reference model of the smoothing system and the actual feedback angle is shown. This is used to verify the preprocessing module of the command smoothing system. It serves as the airworthiness performance verification standard and basis for the outer loop control unit in the control system of a tiltrotor aircraft. It can be found that the response can track the expected reference model, which basically meets the airworthiness requirements.

[0097] Example 3

[0098] The only difference between this embodiment and Embodiments 1 and 2 is in the implementation of rate control and airworthiness. The control quantity output by the parallel controller is a full control quantity with dimensions. It is necessary to subtract the current control quantity to obtain the increment of the control quantity. The increment of the control quantity is then divided by the maximum angular acceleration or vertical acceleration of the aircraft, i.e., K5 = 1 / maximum angular acceleration or vertical acceleration, to obtain the normalized incremental virtual control quantity. Finally, the incremental control allocation of the incremental virtual control quantity is performed based on methods such as weighted least squares method and quadratic programming method.

[0099] The current control variables can be obtained by identifying the actuator transfer function and the sensor equivalent transfer function, or by using feedback speed and feedback control surfaces to calculate the current control variables using a pseudo-inverse method. The current control variables include vertical control variables and three-axis torque control variables (roll, pitch, and yaw).

[0100] Figure 5 This is a flowchart of the rate control process for a tiltrotor aircraft. It represents the core algorithm of the inner-loop attitude control unit 5 in the tiltrotor aircraft's control system. The inner-loop attitude control unit 5 implements rate control through a cascaded control architecture, which includes:

[0101] The outer loop control module receives angle / position reference commands and sensor feedback angles, generates angle errors, and processes these errors using a proportional controller and an integral controller. It performs first-order model shaping on the yaw rate / speed commands. The time constant for TRC mode ranges from [2.5, 5] seconds, while the time constant for RCHH mode ranges from [0, 5] seconds. The feedforward gain K8 ranges from [0, 1], and the speed command is output. A saturation switch is installed after the integral stage. When the speed integral exceeds a threshold, the feedforward gain K8 is set to 0, and the integral controller is reset.

[0102] The inner loop control module compares the rate command with the feedback rate to generate a rate error, and generates a control quantity through one or more of proportional, proportional-integral, or active disturbance rejection control.

[0103] The control distribution module inputs control quantities into the control distribution unit 6 and distributes them to the actuators (rotor motors / control surfaces / tilt mechanisms) to change the aircraft's state. The sensor feedback unit 7 collects attitude / rate data in real time in a closed loop.

[0104] The control architecture works as follows: Angle error is obtained by comparing the angle reference command with the feedback angle. This error is processed by a proportional controller (gain K8) and an integral controller (the integral time constant is set by the "1 / time constant" module), outputting a rate command. A saturation switch is installed after the integral stage to prevent integral saturation. The rate error is obtained by comparing the rate command output from the outer loop with the feedback rate. After passing through the inner loop attitude control unit 5, a control quantity is output, which then generates actuator commands through the control distribution unit 6. The commands output by the control distribution unit 6 drive the actuators to change the aircraft's state. The sensor feedback unit 7 measures the aircraft's angle / position and rate and feeds them back to the input. The time constant is no less than 2.5 seconds and no more than 5 seconds for TRC, and no more than 5 seconds for RCHH. K8 is the feedforward gain, which is generally 0-1. If the integral of the rate exceeds the set threshold, the feedforward gain is 0. Proportional control is sufficient for angle / position reference command control. The yaw rate / rate command control of the inner loop generally adopts proportional, proportional-integral, active disturbance rejection and other control methods. After obtaining the control quantity, the control distribution link is carried out, and finally it is applied to the aircraft body to control the normal movement of the aircraft.

[0105] In summary, this invention simplifies operation through a highly integrated multi-level control system (flight management, outer loop, heading, and inner loop units) and intelligent algorithms. Pilots only need basic commands to automatically complete guidance calculations and control allocation. The sensor feedback unit monitors the aircraft's status in real time, ensuring command accuracy. The system supports diverse guidance commands (forward ground speed, airspeed, altitude, etc.), flexibly adapting to different missions. The outer loop and heading units work together to ensure attitude stability during coordinated turns, while the control allocation unit rationally allocates forward / altitude / roll control quantities, driving actuators to achieve a smooth transition between tiltrotor and fixed-wing modes, significantly improving maneuverability and adaptability while reducing operational difficulty.

[0106] It should be noted that the technical features in embodiments 1-3 above can be combined arbitrarily, and the resulting technical solutions all fall within the protection scope of this application. In this document, terms such as "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0107] While embodiments of the present invention have been shown and described, it will be appreciated by those skilled in the art that various changes, modifications, substitutions, and variations may be made to these embodiments without departing from the principles and spirit of the invention, and that the scope of the invention is defined by the appended claims and their equivalents.

Claims

1. A control system for a tiltrotor aircraft, characterized in that, include: The control command unit (1) is used to issue aircraft control commands; Flight management unit (2) is used to receive aircraft control commands and generate guidance calculation commands based on the aircraft control commands; The outer ring control unit (3) is used to receive guidance calculation commands sent by the flight management unit (2); The heading control unit (4) is used to receive the yaw rate command for coordinated turning sent by the outer ring control unit (3) and the track angle control command or heading angle control command sent by the flight management unit (2); The inner ring attitude control unit (5) is used to receive pitch angle commands and roll angle commands sent by the outer ring control unit (3). Through the inner ring attitude control algorithm, the pitch angle and roll angle commands are converted into specific pitch channel and roll channel control quantities. The inner ring attitude control algorithm includes a command smoothing processing system and a delay-based state feedback gain algorithm. The state feedback control gain matrix is ​​used to achieve stable and precise control of the aircraft attitude. The control distribution unit (6) is connected to the outer ring control unit (3), the heading control unit (4), and the inner ring attitude control unit (5) and is used to receive flight control quantities; The sensor feedback unit (7) is arranged on the aircraft body and is connected to the outer ring control unit (3), the heading control unit (4), the inner ring attitude control unit (5) and the control distribution unit (6) for real-time monitoring of the aircraft's attitude, speed and position information.

2. The control system of the tiltrotor aircraft according to claim 1, characterized in that, The contents of the control command unit (1) include one or more of the following: takeoff, landing, hovering, forward movement, and turning.

3. The control system for the tiltrotor aircraft according to claim 2, characterized in that, An instruction filtering unit (8) is provided between the control instruction unit (1) and the flight management unit (2) for filtering the control instructions.

4. The control system of the tiltrotor aircraft according to claim 1, characterized in that, The flight control variables include one or more of the following: forward control variables, altitude control variables, roll angle control variables, pitch angle control variables, and yaw rate commands for coordinated turns.

5. The control system of the tiltrotor aircraft according to claim 4, characterized in that, The yaw rate command for coordinated turning is calculated using the following formula: Where, r cmd To coordinate the yaw rate during turns; u is the forward velocity of the aircraft; v is the lateral velocity of the aircraft; V is the airspeed of the aircraft, and V > V_r_min, where V_r_min is the minimum yaw rate control speed, which is quantitatively determined through theoretical calculations and simulations based on the aircraft characteristics; k is the sideslip control gain; β is the feedback sideslip angle; acc_y is the lateral acceleration of the aircraft; g is the gravitational acceleration; and θ is the pitch angle. For roll angle, r stick The yaw rate command given to the pilot for active control, where k_coff is the airspeed adaptation coefficient that varies with airspeed.

6. The control system of the tiltrotor aircraft according to claim 5, characterized in that, The airspeed adaptation coefficient k_coff is obtained by the following formula: Where V_r_min is the minimum yaw rate for takeoff control selected based on the aircraft characteristics, and V_r_max is the fixed-wing cruise speed selected based on the aircraft characteristics.

7. The control system of the tiltrotor aircraft according to claim 1, characterized in that, To meet airworthiness requirements, reasonable feedforward commands are provided by designing a feedforward command reference model. The feedforward command reference model is obtained through the following steps: The response characteristics of the attitude command attitude hold (ACAH) control response are equivalent to a second-order circuit. The calculation formula is: Where θ(s) is the Laplace transform of the actual attitude angle, θ cmd (s) represents the Laplace transform of the command attitude angle, ω act For the equivalent natural frequency, ∈ act For the equivalent damping ratio, the attitude ω act Between 1.4 and 4.8 rad / s, the heading ω act Less than 1 rad / s, the altitude ω act Between 1 and 5 rad / s; damping ratio ∈ act Between 0.8 and 1; The original attitude command is smoothed based on a second-order reference model and outputs reference angle command, reference angular rate command, and reference angular acceleration command as feedback. The angle error is amplified via K3, and the reference angular rate is fed back via K4 and multiplied by a damping coefficient to suppress oscillations. K4=2∈ act ω act K3 represents the system stiffness gain, K4 represents the system damping gain, and ω represents the system damping gain to suppress oscillations. act The equivalent natural frequency is determined by the actuator bandwidth, ∈ act For the equivalent damping ratio, the damping ratio ∈ act The value range is 0.8-1.

8. The control system of the tiltrotor aircraft according to claim 1, characterized in that, The state feedback control gain matrix is ​​obtained based on a method for calculating the state feedback gain with time delay, the calculation method including the following steps: Step S1, System Modeling The actuator can be approximated as a first-order inertial system with a pure time delay, and the settling time of this first-order inertial system with a pure time delay is Ts. For ease of analysis, a second-order element approximation is used. The sensor filtering also uses a second-order element approximation. Therefore, considering a second-order transfer function with a damping ratio of 1, it can be approximately equivalent to a first-order time delay: Where, ω n Let be the natural frequency of the second-order transfer function. For a real actuator and sensor with a damping ratio of 1, the two second-order delays can be approximated as: Where, ω actuator ω is the natural frequency of the second-order transfer function of the equivalent actuator's moving element. filter The natural frequency of the second-order transfer function of the equivalent sensor filter; Step S2: Extended-dimensional state space modeling Let the desired response of the entire tiltrotor aircraft's control system (order n=2) be: Used to describe the attitude response of a tiltrotor aircraft, where b is the system gain, s is the Laplace differential operator, P is the desired pole of the control system, and the actuator is approximated as a first-order inertial plus pure delay system, with the approximate first-order delay of the system being Ts. Calculate the A and B matrices of the extended-dimensional system based on the system order n=2 and the first-order delay system denoted as Ts: Where A is the state feedback matrix and B is the mapping from control input to state. The full state feedback and pole placement are calculated, that is, the state space is expanded based on A, B and the desired poles of the system. The delay element is regarded as part of the system to realize state augmentation and obtain state feedback gain. Step S3: Calculation of the controllable canonical matrix Calculate matrix P - =[P;P] characteristic polynomial f(λ) = |λE-P - |=λ 2 +2*P*λ+P 2 And the remaining coefficient a after removing the coefficients of higher-order terms. d =[P 2 [2P]; Then, based on matrices A and B, the controllable standard form A is obtained. n 、B n and T n , where P - Let λ be the expected pole matrix, λ be the eigenvalue variables, E be the identity matrix, and a be the eigenvalue matrix. d Let A be the coefficient vector of the desired characteristic polynomial. n A system matrix in a controllable standard form; B n The input matrix is ​​in controllable standard form; T n Let A be a transformation matrix, and satisfy A n =T n -1 AT n ; Step S4: Controllability Verification Determine whether the control system is controllable based on the following criteria: Let the controllability matrix be s = [B AB…A n-1 [B], matrix s has dimensions n×(n×m), where n is the state dimension and m is the input dimension. It is used to determine whether the control system is completely controllable. If the rank of matrix s is n, the control system is controllable; if the rank of matrix s is less than n, the control system is uncontrollable. When the control system is controllable, the controllable canonical form A is calculated. n Transformation matrix T n : in, To take the last row of the inverse of the controllability matrix; A n =T n -1 A T n B n =T n -1 B; Step S5: Gain Calculation For a second-order system After obtaining matrix A n Then, calculate A. n The corresponding characteristic polynomial is f(λ0=|λE-An|=λ 2 +a1λ+a0, where E is the identity matrix, and let μ be the pole of the expectation, then the characteristic polynomial of the expectation is f(λ)=λ 2 +2*μ*λ+μ 2 And take the coefficients of the last two, the actual characteristic polynomial coefficients a n =μ 2 ;2μ]; Based on the feedback gain K of the controllable canonical space n =a d -a n Calculate controllable standard type A n Feedback gain K n And based on the feedback gain K n Calculate the state feedback control gain matrix [K1,K2]: [K1,K2]=K n *T n =[P 2 -μ 2 ;2P-2μ]*T n ,n=2; The attitude angle of the aircraft is controlled based on the state feedback control gain matrix [K1,K2].

9. The control system of the tiltrotor aircraft according to claim 1, characterized in that, The control system also includes an airworthiness performance verification module, used to verify the tracking performance of the command smoothing system through a step response test; wherein the verification method includes the following steps: Input a 1-degree step attitude command into the smoothing system reference model and compare the reference angle trajectory output by the smoothing system reference model with the actual angle trajectory fed back by the sensor. The control system is deemed compliant when the overlap between the actual trajectory and the reference trajectory is ≥95% and the following airworthiness mode requirements are met. Translational rate response mode: The step response exhibits first-order system characteristics, with a rise time t r ∈[2.5s,5s], overshoot = 0%; Vertical rate height holding mode: The step response exhibits first-order system characteristics, with rise time tr ≤ 5s, pure delay < 0.2s, and overshoot = 0%; Rate command direction hold mode: Verify step response characteristics with customized parameters.

10. The control system of the tiltrotor aircraft according to claim 1, characterized in that, The inner-loop attitude control unit (5) implements rate control through a cascaded control architecture, which includes: The outer loop control module receives angle / position reference commands and sensor feedback angles, generates angle errors, processes the errors using a proportional controller and an integral controller, and performs first-order model shaping on the yaw rate / rate commands. The inner loop control module compares the rate command with the feedback rate to generate a rate error, and generates a control quantity through one or more of the proportional, proportional-integral, and active disturbance rejection control methods. The control allocation module inputs the control quantity into the control allocation unit (6) and allocates it to the actuator ( ) to change the aircraft state. The sensor feedback unit (7) collects attitude / rate data in real time in a closed loop.

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