Flight control method and flight device

Abstract

The application provides a flight control method and a flight device. The method comprises the following steps: acquiring an aerodynamic lift component of the flight device in a current control step and gravity of the flight device; determining a first control weight group for an aerodynamic surface and a second control weight group for a rotor system according to a ratio relationship between the aerodynamic lift component and the gravity; acquiring first control efficiency information representing a dominant state of the aerodynamic surface and second control efficiency information representing a dominant state of the rotor system; fusing the first control efficiency information and the second control efficiency information according to the first control weight group and the second control weight group to obtain fused control efficiency information; and generating an actuator incremental instruction according to the fused control efficiency information, and driving the rotor system and the aerodynamic surface according to the actuator incremental instruction in the current control step.

Description

Technical Field

[0001] This application relates to the field of automation control technology, and in particular to a flight control method and flight equipment. Background Technology

[0002] During the full-envelope flight of a tiltrotor fixed-wing UAV from hovering to cruise, the generation mechanisms of dominant forces and moments continuously migrate from the rotor to the wings and aerodynamic control surfaces, requiring a smooth transfer of control authority from the rotor to the aerodynamic control surfaces. Traditional methods employ gain scheduling, segmented switching, or local linearization, relying on predefined mode-switching logic, which easily leads to control discontinuities and performance degradation at the boundaries. Linear interpolation or threshold switching based on airspeed lacks physical meaning and cannot allocate control authority according to the current actual force and moment generation state, easily causing a disconnect between control and physical mechanisms during the handover phase. Therefore, there is an urgent need for a method that can adaptively adjust the allocation of control authority based on real-time flight status to achieve a smooth transition during the full-envelope flight process. Summary of the Invention

[0003] This application provides a flight control method and flight equipment to at least solve the above-mentioned technical problems existing in the prior art.

[0004] According to a first aspect of this application, a flight control method is provided, the method comprising: Acquire the aerodynamic lift component of the flight equipment in the current control step and the gravity of the flight equipment; Based on the ratio of the aerodynamic lift component to the gravity, a first control reconfiguration for the aerodynamic control surfaces and a second control reconfiguration for the rotor system are determined. Acquire first control effectiveness information characterizing the dominant state of the aerodynamic control surface and second control effectiveness information characterizing the dominant state of the rotor system; Based on the first control reorganization and the second control reorganization, the first control effectiveness information and the second control effectiveness information are fused to obtain fused control effectiveness information; Based on the fused control performance information, an actuator incremental command is generated, and in the current control step, the rotor system and the aerodynamic control surfaces are driven according to the actuator incremental command.

[0005] In one possible implementation, acquiring the aerodynamic lift component of the flight device in the current control step includes: Obtain the aerodynamic model of the flight equipment; In response to the aerodynamic model meeting preset conditions, the angle of attack, airspeed, and atmospheric density of the flight equipment in the current control step are obtained; The dynamic pressure is determined based on the airspeed and the atmospheric density. The lift coefficient is determined based on the angle of attack. The aerodynamic lift component is determined based on the lift coefficient, the dynamic pressure, and the wing area of ​​the flight equipment.

[0006] In one possible implementation, acquiring the aerodynamic lift component of the flight device in the current control step includes: Obtain the aerodynamic model of the flight equipment; In response to the aerodynamic model not meeting the preset conditions, the thrust and tilt angle of each rotor in the rotor system in the current control step are obtained; Based on the thrust and tilt angle of each rotor, determine the resultant force of the rotor system thrust in the Z-axis direction of the fuselage; The aerodynamic lift component is determined based on the gravity and the resultant force.

[0007] In one possible implementation, obtaining the thrust of each rotor in the rotor system at the current control step includes: Obtain the rotational speed of each rotor in the current control step; The thrust of each rotor is determined based on the rotational speed of each rotor and the preset rotational speed-thrust mapping relationship.

[0008] In one possible implementation, the first control reorganization includes a first torque control weight and a first force control weight; the second control reorganization includes a second torque control weight and a second force control weight. The step of determining the first control reconfiguration for the aerodynamic control surfaces and the second control reconfiguration for the rotor system based on the ratio of the aerodynamic lift component to the gravity includes: Based on the ratio of the aerodynamic lift component to the gravity, the original aerodynamic weights of the force channel and the torque channel are determined respectively; wherein, each of the original aerodynamic weights is positively correlated with a power function of the ratio, and the exponent of the power function is a shape parameter; The original aerodynamic weights are saturated to obtain the original aerodynamic weights of the force channel and the original aerodynamic weights of the torque channel after saturation. The original aerodynamic weights of the force channel and the torque channel after saturation are subjected to first-order low-pass filtering with the filtering time constant to obtain the filtered aerodynamic weights of the force channel and the torque channel. The filtered aerodynamic weights of the force channel are subjected to singularity avoidance processing to obtain the first force control weights; the filtered aerodynamic weights of the torque channel are subjected to singularity avoidance processing to obtain the first torque control weights. The second force control weight is determined based on the first force control weight; the second torque control weight is determined based on the first torque control weight.

[0009] In one possible implementation, acquiring first control performance information characterizing the dominant state of the aerodynamic control surface and second control performance information characterizing the dominant state of the rotor system includes: Acquire aerodynamic data and the dynamic pressure and angle of attack of the flight equipment in the current control step, interpolate the aerodynamic data based on the dynamic pressure and the angle of attack, and determine the first control effectiveness information; the first control effectiveness information is used to characterize the sensitivity relationship between the deflection increment of each aerodynamic control surface and the force channel and torque channel. The operating point information of the flight equipment in the current control step is obtained, and the operating point information is analyzed and calculated by the rotor dynamics model, or by hovering state identification, to determine the second control performance information; the second control performance information is used to characterize the sensitivity relationship between the increment of each rotor actuator and the force channel and torque channel.

[0010] In one possible implementation, both the first control performance information and the second control performance information include force channel performance sub-information and torque channel performance sub-information; The step of fusing the first control performance information and the second control performance information based on the first control reorganization and the second control reorganization to obtain fused control performance information includes: Based on the first force control weight in the first control reorganization and the second force control weight in the second control reorganization, the force channel performance sub-information of the first control performance information and the force channel performance sub-information of the second control performance information are weighted and fused to obtain force channel fusion performance information; Based on the first torque control weight in the first control reorganization and the second torque control weight in the second control reorganization, the torque channel performance sub-information of the first control performance information and the torque channel performance sub-information of the second control performance information are weighted and fused to obtain torque channel fusion performance information; The force channel fusion performance information and the torque channel fusion performance information are fused to obtain the fused control performance information.

[0011] In one possible implementation, generating actuator incremental instructions based on the fused control performance information includes: Obtain the virtual control increment and the historical actuator increment of the previous control step; Based on the fused control performance information, the virtual control increment, and the historical actuator increment, the actuator incremental command is calculated using an incremental nonlinear dynamic inverse control law.

[0012] In one possible implementation, the method further includes: Obtain the reference control performance information of the flight equipment; Based on the reference control performance information and the fused control performance information, performance matching information is determined; Determine the minimum eigenvalue of the symmetric portion of the performance matching information; In response to the minimum eigenvalue being greater than zero, the fused control performance information is determined to have passed the stability verification.

[0013] According to a second aspect of this application, a flight device is provided, comprising: body; A rotor system, connected to the fuselage, includes multiple rotors; Aerodynamic control surfaces are connected to the fuselage; The flight controller is communicatively connected to both the rotor system and the aerodynamic control surfaces, and is used in the method described in this application.

[0014] The flight control method and flight equipment of this application directly reflect the current actual force and torque generation state of the flight equipment by the ratio of aerodynamic lift component to gravity. Based on the ratio, the control authority is reorganized, so that the incremental commands of the actuator are aligned with the actual physical mechanism. This avoids the problems of discontinuous control and disconnect between the transfer of authority and the physical mechanism caused by the reliance on airspeed for indirect judgment or preset switching logic in traditional methods, and realizes a smooth transition in the full envelope flight process.

[0015] It should be understood that the description in this section is not intended to identify key or essential features of the embodiments of this application, nor is it intended to limit the scope of this application. Other features of this application will become readily apparent from the following description. Attached Figure Description

[0016] The above and other objects, features, and advantages of exemplary embodiments of this application will become readily apparent from the following detailed description taken in conjunction with the accompanying drawings. Several embodiments of this application are illustrated in the drawings by way of example and not limitation, in which: In the accompanying drawings, the same or corresponding reference numerals indicate the same or corresponding parts.

[0017] Figure 1 A schematic diagram illustrating the implementation flow of the flight control method provided in an embodiment of this application is shown; Figure 2 A schematic diagram illustrating the implementation flow of the weight determination operation in the flight control method provided in this application embodiment is shown. Figure 3 A schematic diagram illustrating the implementation flow of the performance information acquisition operation of the flight control method provided in this application embodiment is shown; Figure 4 A schematic diagram illustrating the implementation flow of the fusion operation of the flight control method provided in this application embodiment is shown. Detailed Implementation

[0018] To make the objectives, features, and advantages of this application more apparent and understandable, the technical solutions in the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this application, and not all embodiments. Based on the embodiments of this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0019] The flight control method provided in the embodiments of this application will be described in detail below with reference to the accompanying drawings. This method can be executed by the flight controller of the flight equipment to achieve adaptive transfer of control authority between the rotor system and the aerodynamic control surfaces.

[0020] Figure 1 A schematic diagram illustrating the implementation flow of the flight control method provided in this application embodiment is shown.

[0021] refer to Figure 1 This application provides a flight control method, which includes: Operation 101: Obtain the aerodynamic lift component and gravity of the flight equipment in the current control step.

[0022] The flight control method provided in this application is applicable to flight equipment that includes a rotor system and aerodynamic control surfaces and needs to undergo a hovering to cruise transition process. Typically, this method can be applied to tiltrotor vertical takeoff and landing (VTOL) fixed-wing unmanned aerial vehicles (UAVs), which possess vertical takeoff and landing, hovering, and high-speed cruise capabilities, with control authority continuously shifting between the rotor system and aerodynamic control surfaces during flight. This application primarily uses a tiltrotor VTOL fixed-wing UAV as an example for illustration, but those skilled in the art will understand that this method is equally applicable to other flight equipment with similar control requirements. The current control step is the real-time control cycle of the flight controller, for example, executed at a frequency of 250Hz, and the duration of each control step corresponds to one control cycle.

[0023] The aerodynamic lift component refers to the aerodynamic lift component generated by the wing along the negative Z-axis of the fuselage. Its physical meaning is the magnitude of the lift actually carried by the wing under the current flight conditions. This aerodynamic lift component can be obtained in various ways, such as calculations using an airborne aerodynamic model or estimations using the force balance equations of the flight equipment; this application does not specifically limit this. Gravity is the gravitational force exerted by the Earth on the flight equipment, which can be determined by the product of the total mass of the flight equipment and the acceleration due to gravity, i.e. ,in The total mass of the UAV (known design parameters); Let gravitational acceleration be (take) ).

[0024] Operation 102: Based on the ratio of aerodynamic lift component to gravity, determine the first control reconfiguration for aerodynamic control surfaces and the second control reconfiguration for rotor system.

[0025] The ratio of aerodynamic lift to gravity characterizes the proportion of lift currently carried by the wing. A ratio close to 0 indicates that the wing generates almost no lift, and the flight equipment is in a rotor-dominated state; a ratio close to 1 indicates that the wing bears all the lift, and the flight equipment is in an aerodynamic control surface-dominated state; a ratio between 0 and 1 indicates a transitional state. The two sets of control authority reconfigurations determined based on this ratio correspond to the control authority allocation of the aerodynamic control surfaces and the rotor system, respectively. This ratio directly reflects the current actual force and torque generation state of the flight equipment, and has a clear physical meaning compared to indirect judgments based on airspeed.

[0026] Operation 103: Obtain first control effectiveness information characterizing the dominant state of the aerodynamic control surface and second control effectiveness information characterizing the dominant state of the rotor system.

[0027] Control performance information describes the mapping relationship between actuator increments and the forces and moments in each channel of the flight equipment. The first set of control performance information corresponds to the degree of influence of the deflection increment of each aerodynamic control surface on the forces and moments in each channel under aerodynamic control surface-dominated conditions; the second set of control performance information corresponds to the degree of influence of the increment of each rotor actuator on the forces and moments in each channel under rotor system-dominated conditions. This control performance information can be pre-established based on the dynamic model of the flight equipment and updated during flight according to the current flight state.

[0028] Operation 104: Based on the first control reorganization and the second control reorganization, the first control effectiveness information and the second control effectiveness information are fused to obtain fused control effectiveness information.

[0029] The two sets of control rights obtained in operation 102 are recombined and applied to the two sets of control performance information obtained in operation 103, respectively. A weighted fusion method is then used to obtain the fused control performance information for the current control step. This fused control performance information changes continuously throughout the full-envelope flight. During the hovering phase, it approximates the control performance information dominated by the rotor system; during the cruise phase, it approximates the control performance information dominated by the aerodynamic control surfaces; and during the transition phase, it represents a smooth transition between the two. Therefore, this fused control performance information accurately reflects the actual control performance of the hybrid actuator at the current operating point of the flight equipment.

[0030] Operation 105 generates actuator incremental commands based on fused control performance information, and drives the rotor system and aerodynamic control surfaces in the current control step according to the actuator incremental commands.

[0031] Based on the fused control performance information, the incremental commands for each actuator in the current control step are calculated. These actuator incremental commands include the rotor system's rotational speed increment, the tilt mechanism's tilt angle increment, and the aerodynamic control surface's deflection angle increment. The calculated actuator incremental commands are sent to the corresponding actuators, driving the rotor system and aerodynamic control surfaces to produce the corresponding physical actions. After the current control step is completed, the next control step is entered, and operations 101 to 105 are repeated to achieve real-time closed-loop control for full-envelope flight.

[0032] In this context, an actuator refers to a device in flight equipment that receives control commands and generates physical actions to change the flight state. In the embodiments of this application, actuators include, but are not limited to: the power motor and propeller in the rotor system (which generate thrust changes by changing the rotation speed), the tilt servo in the tilt mechanism (which changes the thrust direction by changing the angle), and the aileron servo, elevator servo, and rudder servo in the aerodynamic control surfaces (which generate aerodynamic torque changes by deflection).

[0033] Thus, this embodiment of the application obtains the aerodynamic lift component and gravity of the flight equipment in the current control step, and determines the first control reorganization for the aerodynamic control surfaces and the second control reorganization for the rotor system based on their ratio. Then, it fuses the two sets of control effectiveness information—one dominated by the aerodynamic control surfaces and the other by the rotor system—using the two sets of control reorganizations. This allows the fused control effectiveness information to continuously change with the flight state: rotor effectiveness dominates during hovering, aerodynamic control surface effectiveness dominates during cruise, and a smooth transition occurs between the two during transition phases. Finally, it generates actuator incremental commands based on the fused control effectiveness information and drives the actuators. Since the ratio of the aerodynamic lift component to gravity directly reflects the actual force and torque generation state of the flight equipment, determining the control reorganization based on this ratio ensures that the actuator incremental commands are aligned with the actual physical mechanism. This avoids the control discontinuity and disconnect between authority transfer and the physical mechanism caused by relying on airspeed indirect judgment or preset switching logic in traditional methods, achieving a smooth transition throughout the full envelope flight process.

[0034] In one embodiment of this application, obtaining the aerodynamic lift component of a flight device in the current control step includes: obtaining an aerodynamic model of the flight device; in response to the aerodynamic model satisfying preset conditions, obtaining the angle of attack, airspeed, and atmospheric density of the flight device in the current control step; determining the dynamic pressure based on the airspeed and atmospheric density; determining the lift coefficient based on the angle of attack; and determining the aerodynamic lift component based on the lift coefficient, dynamic pressure, and wing area of ​​the flight device.

[0035] The aerodynamic model is a pre-stored wing lift characteristic model used by the flight equipment to describe the correspondence between the lift coefficient and flight state. This aerodynamic model can be pre-established through airborne aerodynamic data lookup tables or updated in real time through online identification; this application does not specifically limit this. Preset conditions are the criteria for determining whether the aerodynamic model's accuracy meets the usage requirements, such as the model being successfully loaded and in an effective state, the angle-of-attack sensor signal being valid, and the pitot tube signal being valid. When these conditions are met, it indicates that the conditions for accurately calculating lift based on the aerodynamic model are currently available. The angle of attack can be directly measured by the angle-of-attack sensor or indirectly obtained through state estimation by an integrated navigation system (such as GPS / INS integrated navigation); airspeed is obtained through pitot tube measurement; atmospheric density is estimated from atmospheric pressure and temperature obtained by barometric altimeter measurements. All of the above sensor data are collected in real time by the flight controller at the beginning of each control step.

[0036] Dynamic pressure is the pressure exerted by an airflow; its physical meaning is the kinetic energy of a unit volume of airflow. The calculation method for dynamic pressure is as follows: Where V is the current airspeed and ρ is the current atmospheric density. Dynamic pressure is a key intermediate quantity in aerodynamic lift calculation, reflecting the impact intensity of airflow on the wing. The higher the airspeed or the greater the atmospheric density, the greater the dynamic pressure, and the greater the aerodynamic lift generated by the wing.

[0037] The lift coefficient is a dimensionless aerodynamic parameter that characterizes the efficiency of an airfoil in converting dynamic pressure into lift; its value depends on the angle of attack. The lift coefficient corresponding to the current angle of attack is obtained by querying a pre-stored lift coefficient curve or lookup table in the aerodynamic model. It should be noted that the relationship between lift coefficient and angle of attack is non-linear. Within a small angle of attack range, the lift coefficient increases approximately linearly with the increase of the angle of attack. However, when the angle of attack exceeds the stall angle of attack, the lift coefficient drops sharply.

[0038] Multiplying the lift coefficient, dynamic pressure, and wing area yields the aerodynamic lift components: .in, Let be the wing area of ​​the flight equipment, and be the known design parameters. This represents the aerodynamic lift component generated by the wing along the negative Z-axis of the fuselage (positive upwards).

[0039] The above process is executed at the beginning of each control step (e.g., 250Hz, i.e., once every 4ms). The flight controller collects data from each sensor in real time and performs calculations to ensure that the aerodynamic lift component is updated in a timely manner in each control step as the flight state changes.

[0040] In one embodiment of this application, obtaining the aerodynamic lift component of the flight equipment in the current control step includes: obtaining an aerodynamic model of the flight equipment; in response to the aerodynamic model not meeting preset conditions, obtaining the thrust and tilt angle of each rotor in the rotor system in the current control step; determining the resultant force of the rotor system thrust in the Z-axis direction of the aircraft based on the thrust and tilt angle of each rotor; and determining the aerodynamic lift component based on gravity and the resultant force.

[0041] In one embodiment of this application, obtaining the thrust of each rotor in the rotor system at the current control step includes: obtaining the rotational speed of each rotor at the current control step; and determining the thrust of each rotor based on the rotational speed of each rotor and a preset rotational speed-thrust mapping relationship.

[0042] At the start of each control step, it is first confirmed that the aerodynamic model does not meet the preset conditions, such as the aerodynamic model not being loaded, insufficient model accuracy, invalid angle of attack sensor signal, invalid pitot tube signal, or insufficient dynamic pressure signal-to-noise ratio due to low current airspeed. When any of the above conditions are met, the system switches to this embodiment to obtain the aerodynamic lift component.

[0043] The thrust of each rotor can be determined by the rotational speed-thrust mapping relationship of each rotor: obtain the rotational speed of each rotor in the current control step (feedback from the ESC or encoder), and calculate the thrust of each rotor according to the preset rotational speed-thrust mapping relationship. The tilt angle of each rotor was measured by the tilt servo encoder and denoted as . .

[0044] The preset speed-thrust mapping relationship refers to a pre-established data model describing the correspondence between rotor speed and thrust. This mapping relationship characterizes the quantitative relationship between motor speed (input) and propeller-generated thrust (output). This mapping relationship can be established in various ways: for example, by fixing the motor-propeller system on a thrust test bench, measuring the thrust values ​​at different speeds, and fitting the collected data points to obtain a speed-thrust curve; or by performing theoretical calculations based on propeller aerodynamic theories (such as blade element theory or momentum theory) combined with blade geometric parameters; or by performing real-time estimation through an approximate linear relationship between motor current and thrust. This application does not impose specific limitations on this, and an appropriate establishment method can be selected according to the actual configuration of the flight equipment.

[0045] For each rotor, its thrust Multiply by the rotor tilt angle The cosine value is used to obtain the component of the rotor thrust along the Z-axis of the fuselage. Then, sum the components of each rotor along the Z-axis of the fuselage to obtain the resultant force of the rotor system thrust along the Z-axis of the fuselage. Then, based on the difference between gravity and the net force, the aerodynamic lift component can be determined: .in For the gravity of the flight equipment, This is the resultant force of the rotor system thrust along the Z-axis of the fuselage. The physical meaning of the above formula is: in the Z-axis direction of the fuselage, the aerodynamic lift generated by the wing is equal to the total weight minus the portion of the rotor thrust borne in that direction. It should be noted that this approximation has good accuracy during steady-state flight, but deviations may be introduced during maneuvering flight due to inertial forces. These deviations can be corrected through inertial compensation terms, which are not specifically limited in this application.

[0046] Thus, the embodiments of this application achieve reliable aerodynamic lift components even when the aerodynamic model is not accurate enough or the sensor conditions are limited, through the above-described method.

[0047] Figure 2 A schematic diagram illustrating the implementation flow of the weight determination operation of the flight control method provided in this application embodiment is shown.

[0048] refer to Figure 2 In one embodiment of this application, the first control reconfiguration includes a first torque control weight and a first force control weight; the second control reconfiguration includes a second torque control weight and a second force control weight; the above operation 102, based on the ratio of aerodynamic lift component to gravity, determines the first control reconfiguration for aerodynamic control surfaces and the second control reconfiguration for the rotor system, including: Operation 201: Determine the original aerodynamic weights of the force channel and the torque channel based on the ratio of the aerodynamic lift component to gravity. The original aerodynamic weights are positively correlated with the power function of the ratio, and the exponent of the power function is the shape parameter.

[0049] First, calculate the aerodynamic lift components. With gravity ratio This ratio is the only physical quantity driving all subsequent weight calculations. This ratio directly reflects the proportion of lift currently carried by the wing. A ratio close to 0 indicates that the wing generates almost no lift, and the flight equipment is in a rotor-dominated state; a ratio close to 1 indicates that the wing bears all the lift, and the flight equipment is in an aerodynamic control surface-dominated state; a ratio between 0 and 1 indicates a transitional state.

[0050] Based on this ratio, the original aerodynamic weights for the force channel and torque channel are calculated separately. Taking the force channel as an example, the flight controller uses the absolute value of the lift ratio as the base and the shape parameters... The original aerodynamic weights of the force channel are obtained by exponentiation:

[0051] Accordingly, the original aerodynamic weights of the torque channel are:

[0052] in, For the force channel shape parameters, The shape parameters for the torque channel and the torque channel can be configured independently. The shape parameter determines the non-linearity of the weight change with the lift ratio: when the shape parameter equals 1, the weight is linearly proportional to the lift ratio; when the shape parameter is greater than 1, the curve bulges downwards, the weight increases more rapidly in the middle region, and the transition is more aggressive; when the shape parameter is less than 1, the curve bulges upwards, the weight changes faster at both ends, and the transition in the middle region is smoother. The shape parameter can be determined by simulation calibration to select a value that minimizes the tracking error during the transition phase. In this embodiment, different shape parameters are allowed for the two channels to adapt to actual working conditions where the switching behavior of the force channel and the torque channel is not completely synchronized during the transition phase.

[0053] Operation 202 involves saturating each original aerodynamic weight to obtain the original aerodynamic weights of the force channel and the torque channel after saturation.

[0054] Apply a saturation function to each original aerodynamic weight. Each original aerodynamic weight is limited to the range [0, 1]. When the weight is less than 0, it is taken as 0; when it is greater than 1, it is taken as 1; otherwise, the weight itself is taken.

[0055] Operation 203 involves performing first-order low-pass filtering on the original aerodynamic weights of the force channel and the torque channel after saturation processing, respectively, using the filtering time constant, to obtain the filtered aerodynamic weights of the force channel and the torque channel.

[0056] To prevent high-frequency jitter in the weights caused by sensor noise or aerodynamic disturbances, a first-order low-pass filter needs to be applied to the original aerodynamic weights after saturation processing obtained in operation 202.

[0057] The discrete form of the first-order low-pass filter for the force channel is:

[0058] in, Indicates the current control step. This indicates the previous control step. This is the filtered aerodynamic weight for the current control force channel. This is the filtered aerodynamic weight of the previous control force channel. This represents the original aerodynamic weights of the force channel after saturation processing for the current control step. The force channel filter coefficient determines the weighting of historical and current values ​​in the filtering result. The closer it is to 1, the more the filtering result depends on historical values, the stronger the smoothing effect, and the slower the response; The closer the coefficient is to 0, the more dependent the filtering result is on the original value; the response is faster, but the smoothing effect is weaker. Filter coefficients Force channel filtering time constant Sure:

[0059] To control the step size. Similarly, the filtered aerodynamic weights of the torque channel. With torque channel filtering time constant Perform the same filtering process. and It can be configured independently to determine the response speed of the force and torque channel weights to aerodynamic changes. The larger the filtering time constant, the smoother the weight changes and the stronger the noise suppression, but the more lagging the tracking of flight state changes. The smaller the filtering time constant, the faster the response, but the noise immunity decreases. The filtering time constant can be determined by simulation verification to ensure that the weight changes during the transition phase are smooth and free of chattering.

[0060] Operation 204: Singularity avoidance processing is performed on the filtered aerodynamic weights of the force channel to obtain the first force control weights; singularity avoidance processing is performed on the filtered aerodynamic weights of the torque channel to obtain the first torque control weights.

[0061] Singularity avoidance processing is applied to the filtered aerodynamic weights to prevent numerical singularities from being introduced into the aerodynamic performance matrix at low speeds or during hovering. The singularity avoidance processing works as follows: if the filtered aerodynamic weight is less than or equal to a preset threshold, the weight is reset to zero; otherwise, the weight remains unchanged. That is:

[0062] in, Preset a threshold for the force channel. A preset threshold is set for the torque channel; both can be configured independently. To control the weight of the first force, The weights are set as the first torque control weights. The physical meaning of this processing is that at low speeds or when hovering, aerodynamic lift is extremely small, and the original weights are approximately zero. However, sensor noise may cause small positive fluctuations. If these small values ​​are directly used in subsequent performance matrix fusion, the slight influx of these aerodynamic performance matrix components will introduce near-singular terms (at low speeds, the aerodynamic control surface performance is extremely weak, and its corresponding column is close to a zero vector), leading to a deterioration of the control assignment matrix condition number or even numerical singularity. By forcibly setting these small values ​​to zero through singularity avoidance processing, the fused control performance matrix is ​​completely dominated by the rotor in hovering, treating the small aerodynamic forces generated by the wing as bounded disturbances, thus ensuring numerical stability.

[0063] Operation 205: Determine the second force control weight based on the first force control weight; determine the second torque control weight based on the first torque control weight.

[0064] The first force control weight obtained from operation 204 Determine the second force control weight The relationship is as follows: Similarly, the weights are controlled according to the first torque. Determine the second torque control weight : The physical meaning of this complementary relationship is that the sum of the control authority of the rotor system and the aerodynamic control surfaces is 1; if one is strengthened, the other is weakened, and the sum of the two is always 1. Through this relationship, only the aerodynamic control surface weights need to be determined in a set of control weights, while the rotor system weights are directly obtained from the complementary relationship, ensuring the continuity and consistency of authority allocation.

[0065] This completes the first control reorganization for the aerodynamic control surfaces (including the first force control weights). and first torque control weight ) and the reconfiguration of the second control authority for the rotor system (including the second force control weights) Second torque control weight The determination of the above two sets of control authority reorganizations will serve as weighting coefficients for subsequent performance matrix fusion, used to achieve adaptive allocation of control authority between aerodynamic control surfaces and the rotor system.

[0066] Figure 3 This paper illustrates a schematic diagram of the implementation process of the performance information acquisition operation of the flight control method provided in the embodiments of this application.

[0067] refer to Figure 3 In one embodiment of this application, the above-mentioned operation 103, which acquires first control performance information characterizing the dominant state of the aerodynamic control surface and second control performance information characterizing the dominant state of the rotor system, includes: Operation 301: Acquire aerodynamic data and the dynamic pressure and angle of attack of the flight equipment in the current control step; interpolate the aerodynamic data based on the dynamic pressure and angle of attack to determine the first control effectiveness information; the first control effectiveness information is used to characterize the sensitivity relationship between the deflection increment of each aerodynamic control surface and the force channel and torque channel.

[0068] First control performance information This study describes the control effectiveness matrix for flight equipment during the high-speed / cruise phase (when the wings and aerodynamic control surfaces provide dominant lift and torque), depicting the sensitivity of each aerodynamic control surface deflection increment to the generalized force / torque in each channel. Specifically, pre-stored aerodynamic data is first acquired through aerodynamic analysis methods such as vortex lattice (VLM), computational fluid dynamics (CFD) simulation, or wind tunnel experiments, and stored in the form of a lookup table. The aerodynamic data records the aerodynamic derivatives of the flight equipment under different dynamic pressure and angle of attack conditions, i.e., the change in force / torque in each channel caused by a unit deflection of each aerodynamic control surface (such as aileron, elevator, and rudder).

[0069] Next, the dynamic pressure and angle of attack of the current control step are obtained. Using the current dynamic pressure and angle of attack as indices, aerodynamic data is interpolated. Since the aerodynamic data is discrete grid data (each taking a finite number of discrete points in the dynamic pressure and angle of attack dimensions), the currently measured dynamic pressure and angle of attack often do not fall exactly on the grid points. Therefore, two-dimensional interpolation is performed in both the dynamic pressure and angle of attack dimensions to obtain the aerodynamic derivatives under the current operating condition. Arranging these aerodynamic derivatives into a matrix according to the corresponding control surfaces and channels yields the first control effectiveness information. . Each column corresponds to the deflection increment of an aerodynamic control surface (e.g., the deflection angle of an aileron, elevator, rudder, etc.), and each row of the matrix corresponds to a generalized force or moment channel (e.g., the first row is the force channel, and subsequent rows are the roll moment channel, pitch moment channel, and yaw moment channel). Each element in the matrix represents the change in force or moment generated in the corresponding channel when the corresponding aerodynamic control surface deflects by a unit angle.

[0070] Among them, the aforementioned first control effectiveness information The system updates online based on the current operating point (dynamic pressure and angle of attack) of the flight equipment. When changes in flight conditions cause changes in dynamic pressure or angle of attack, the interpolation is re-performed to ensure accuracy. It always reflects the aerodynamic control surface performance under the current operating conditions.

[0071] Operation 302: Obtain the operating point information of the flight equipment in the current control step, and perform analytical calculations on the operating point information through the rotor dynamics model, or determine the second control effectiveness information through hovering state identification; the second control effectiveness information is used to characterize the sensitivity relationship between the increment of each rotor actuator and the force channel and torque channel.

[0072] Second control performance information This section describes the control effectiveness matrix for the flight equipment during the low-speed / hovering phase (when the rotors provide all lift and torque), depicting the sensitivity of each rotor actuator increment to the generalized force / torque of each channel. Specifically, the operating point information for the current control step is first acquired. This operating point information includes state variables such as the current rotational speed and tilt angle of each rotor, used to determine the rotor dynamics characteristics under the current operating condition. Then, the second control effectiveness information is determined using one of the following two methods: 1) Analytical Calculation Based on Rotor Dynamics Model. The current operating point information (rotor speeds, tilt angles, etc.) is substituted into a pre-established analytical rotor dynamics model for analytical calculation, yielding the sensitivity of each rotor actuator increment to the generalized force / torque of each channel. The analytical rotor dynamics model describes the quantitative relationship between rotor actuator increments (speed increments, tilt angle increments) and the force / torque changes in each channel of the flight equipment. This model can be established based on blade element theory or momentum theory, analytically expressing the influence of rotor speed changes on thrust and anti-torque, and the influence of tilt angle changes on thrust direction, as changes in force / torque, thus obtaining the sensitivity of each actuator increment to the force and torque channels.

[0073] 2) Identification through hovering. While hovering, a small excitation signal is applied to the flight equipment (e.g., a small change in rotational speed is applied to each rotor sequentially), and the force / torque response data of the flight equipment is collected synchronously. The sensitivity between the increment of each rotor actuator and the change in force / torque is then identified using a system identification algorithm (e.g., least squares method). This method is suitable for scenarios where the rotor dynamics model is inaccurate or the flight equipment configuration changes (e.g., changes in load leading to changes in rotational inertia).

[0074] Arrange the sensitivity values ​​obtained by method 1) or 2) above according to the row and column structure of a matrix to obtain the second control effectiveness information. The complete matrix form. Each column of the matrix corresponds to the increment of a rotor actuator, and each row corresponds to a generalized force or torque channel. Each element in the matrix represents the change in force or torque on the corresponding channel when the corresponding rotor actuator produces a unit increment. Here, rotor actuator refers to each independently controllable actuator unit belonging to the rotor system in the flight equipment, including the motors of each rotor and the servo motors of the tilt mechanism.

[0075] In this embodiment of the application, the second control performance information The system updates online along with the current operating point of the flight equipment. When the operating point changes (such as a change in rotor speed or tilt angle), it recalculates or updates its identification to ensure accuracy. It always reflects the control efficiency of the rotor system under the current operating conditions.

[0076] Figure 4A schematic diagram illustrating the implementation flow of the fusion operation of the flight control method provided in this application embodiment is shown.

[0077] refer to Figure 4 In one embodiment of this application, both the first control performance information and the second control performance information include force channel performance sub-information and torque channel performance sub-information; the above operation 104, based on the first control right reorganization and the second control right reorganization, fuses the first control performance information and the second control performance information to obtain fused control performance information, including: Operation 401: Based on the first force control weight in the first control reorganization and the second force control weight in the second control reorganization, the force channel performance sub-information of the first control performance information and the force channel performance sub-information of the second control performance information are weighted and fused to obtain the force channel fusion performance information.

[0078] First control performance information It contains row vectors for stress channels, denoted as force channel effectiveness sub-information. This row vector describes the sensitivity of each aerodynamic control surface deflection increment to the force channel. Second control performance information. It also includes row vectors for stress channels, denoted as force channel effectiveness sub-information. This row vector describes the sensitivity of each rotor actuator increment to the force channel generation.

[0079] Based on the first force control weight Second force control weight The above two force channel efficiency sub-information are weighted and fused:

[0080] in, This refers to the force channel fusion efficiency information. The physical meaning of the above formula is: the force channel fusion efficiency equals the rotor system force channel efficiency weighted by rotor force weights, plus the aerodynamic control surface force channel efficiency weighted by aerodynamic force weights. Because... (The sum of the two is always 1), the force channel fusion efficiency information is close to the rotor system force channel efficiency during hovering. During cruise, it approaches the efficiency of aerodynamic control surface force channels. ( ), which is a smooth transition between the two during the transition phase.

[0081] Operation 402: Based on the first torque control weight in the first control reorganization and the second torque control weight in the second control reorganization, the torque channel performance sub-information of the first control performance information and the torque channel performance sub-information of the second control performance information are weighted and fused to obtain torque channel fusion performance information.

[0082] Similarly, the first control effectiveness information It contains row vectors for stress moment channels (including roll moment channel, pitch moment channel, and yaw moment channel), denoted as moment channel effectiveness sub-information. Second control performance information It also includes row vectors for the stress torque channels, denoted as the torque channel effectiveness sub-information. .

[0083] Based on the first torque control weight Second torque control weight The above two torque channel efficiency sub-information are then weighted and fused:

[0084] in, This is for torque channel fusion performance information. The physical meaning of the above formula is similarly: due to... The torque channel fusion efficiency information is close to the rotor system torque channel efficiency during hovering. During cruise, it approaches the efficiency of aerodynamic control surface moment channels. ( ), which is a smooth transition between the two during the transition phase.

[0085] Operation 403: Fusion of force channel fusion performance information and torque channel fusion performance information to obtain fusion control performance information.

[0086] Integrate power channel efficiency information Integrating torque channel performance information By combining the matrix according to its row structure, with the force channel fusion performance information as the force channel row and the torque channel fusion performance information as the torque channel row, the complete fusion control performance information can be obtained. :

[0087] This integrated control performance information The control effectiveness matrix during full-envelope flight changes continuously, and during the hovering phase, it approximates the control effectiveness matrix dominated by the rotor system. During the cruise phase, the control effectiveness matrix is ​​close to that dominated by aerodynamic control surfaces. The transition phase involves a smooth, weighted fusion of the two. Therefore, It can accurately reflect the actual control performance of the hybrid actuator (rotor system and aerodynamic control surface acting simultaneously) of the flight equipment at the current operating point, and provide accurate performance matrix input for the calculation of subsequent actuator incremental commands.

[0088] In one embodiment of this application, the above operation 105 generates an actuator incremental command based on the fused control performance information. The actuator incremental command is used to drive the rotor system and aerodynamic control surfaces in the current control step. This includes: obtaining the virtual control increment and the historical actuator increment of the previous control step; and calculating the actuator incremental command using an incremental nonlinear dynamic inverse control law based on the fused control performance information, the virtual control increment, and the historical actuator increment.

[0089] Specifically, the virtual control increment given by the upper-level incremental nonlinear dynamic inverse (INDI) control law is first obtained. and the historical actuator increment of the previous control step .in, The expected increment of generalized force / torque is calculated by the upper-level guidance or INDI control law in the flight controller based on the deviation between the current flight state and the desired state. This refers to the incremental commands actually issued to each actuator in the previous control step. The flight controller will then integrate the control performance information. Virtual control increment and historical actuator increment Substituting into the INDI control law formula, calculate the actuator incremental instruction for the current control step:

[0090] The physical meaning of this formula is: first use the fusion control efficiency matrix Multiply by the historical actuator increment of the previous control step The generalized force / torque change actually generated by the actuator increment in the previous control step is obtained; then the desired virtual control increment is used. Subtracting the actual change yields the force / torque deviation that needs to be compensated for in the current control step; finally, multiply by the inverse of the fused control performance matrix. The force / torque deviation is then back-calculated into incremental commands for each actuator in the current control step. . This can be viewed as a column vector, where each element corresponds to an incremental command for an actuator, including the speed increment of each rotor in the rotor system, the tilt angle increment of the tilt mechanism, and the deflection angle increment of each control surface in the aerodynamic control surfaces. The calculated... Once the control is sent to the corresponding actuator, it can drive the rotor system and aerodynamic control surfaces to produce the corresponding physical actions in the current control step.

[0091] After each control step ends, the current control step's... Store as the historical actuator increment for the next control step Then, the next control step is executed repeatedly, thus forming a real-time closed-loop control loop for full-envelope flight.

[0092] In one embodiment of this application, the following verification operation is also performed: obtaining reference control performance information of the flight equipment; determining performance matching information based on the reference control performance information and the fused control performance information; determining the minimum eigenvalue of the symmetrical part of the performance matching information; and determining that the fused control performance information passes the stability verification in response to the minimum eigenvalue being greater than zero.

[0093] Specifically, in order to verify whether the fused control performance information generated by this method meets the stability conditions of the closed-loop system, thereby ensuring the theoretical reliability of the flight control process, this application also conducted Lyapunov stability verification.

[0094] First, define the virtual control tracking error. This is the difference between the expected generalized force / torque and the actual value. The INDI control law is designed as follows: ,in Let the feedback gain be a scalar. Substitute it into the dynamics of the real system. ( Let be a bounded perturbation vector, satisfying , (As the upper bound of the perturbation), the closed-loop error dynamics are obtained. ,in For the equivalent disturbance that includes performance estimation error and external disturbance, satisfying .

[0095] The specific verification steps are as follows: Obtain reference control performance information This reference control performance information is the true control performance matrix of the flight equipment at the current operating point, describing the true sensitivity of each actuator increment to the generalized force / torque of each channel in the current state of the flight equipment. These are theoretical values ​​used for offline analysis or simulation verification during actual flight. They are obtained through high-precision system identification or full-state aerodynamic models and serve as a reference benchmark for stability verification.

[0096] Based on reference control performance information and integrated control performance information Determine performance matching information : .in, It is the inverse matrix (or generalized inverse) of the integrated control effectiveness matrix. Describes the fusion control performance matrix With the real control performance matrix The degree of matching between them, if and Completely identical, then It is the identity matrix; if there is a discrepancy between the two, then The degree of deviation from the identity matrix reflects the magnitude of the performance estimation error.

[0097] Determine performance matching information Symmetrical part : .in, for The transpose of the matrix. Symmetric part. Used for subsequent stability determination, for any non-zero vector Quadratic Reflects the performance matching matrix in direction Positive definiteness on.

[0098] Choosing Lyapunov functions Differentiate along the closed-loop error dynamic trajectory: When the symmetric part of the matching matrix satisfies (Right now (During the definite period) .

[0099] when hour, The tracking error is uniformly ultimately bounded (UUB), and all trajectories converge to the sphere:

[0100] In response to The fusion control performance information was verified to pass stability testing. At this point, the tracking error is uniformly ultimately bounded (UUB), meaning the tracking error... Ultimately, it converges to a bounded spherical domain, and the radius of the sphere is determined by the upper bound of the perturbation, the minimum eigenvalue, and the feedback gain. This is the upper bound of the disturbance (including the maximum amplitude of bounded disturbances such as external wind field disturbances and model errors). Let be the feedback gain (a scalar, given by the control law design). The above equation shows that when a bounded disturbance exists, the tracking error will not diverge, but will eventually converge to a spherical domain that is proportional to the disturbance intensity and inversely proportional to the minimum eigenvalue and the feedback gain. The larger the disturbance, the larger the convergence domain; the larger the feedback gain and the minimum eigenvalue, the smaller the convergence domain.

[0101] It should be noted that, The establishment of this depends on the combined effect of the preceding operations in this application: on the one hand, the first-order low-pass filtering of operation 203 ensures the force channel weights. It changes continuously over time, while exist The upper continuum, therefore The function is continuous on this compact interval. By the extremum theorem for continuous functions on closed intervals, There exists a consistent lower bound within the total envelope. ,if only , Design guarantee Within the full-coverage line and The deviation does not exceed the limit that makes The degree to which positive definiteness is lost. On the other hand, the singularity avoidance of operation 204 resets the filtered aerodynamic weights to zero during low-speed flight, ensuring... By avoiding the introduction of near-singular aerodynamic efficiency matrices in the low-speed phase, further protection is achieved. Established.

[0102] Therefore, through the above verification, it is proven that the degree of matching between the fused control performance information generated in this application and the actual control performance information within the full envelope flight range satisfies the stability condition under the Lyapunov meaning. That is, when At this time, regardless of how the external disturbance changes (within a bounded range), the tracking error will not diverge, but will eventually converge to a definite bounded range, thus providing a theoretical stability guarantee for the flight control process.

[0103] To further illustrate the technical method of this application, a specific application example is given below.

[0104] In this specific application example of this application, the flight equipment is a tiltrotor vertical takeoff and landing fixed-wing unmanned aerial vehicle (UAV) with a total mass of [missing information]. ,gravity Wing area The flight controller operates at a frequency of 250Hz, which is the duration of each control step. The aerodynamic model selected is the VLM (vortex lattice method) lookup table model, with the angle of attack... and dynamic pressure Input: Lift coefficient The low-precision backup solution uses the thrust compensation method.

[0105] Before detailing the complete execution flow of each control step, the physical meaning of the weighting factors in this application will first be explained. Aerodynamic lift components. With gravity ratio This directly reflects the proportion of lift currently being carried by the wing. When (In hovering state) the flight equipment mainly relies on the rotor for lift, the lift ratio is approximately 0, the original aerodynamic weight is approximately 0, after filtering and saturation the aerodynamic weight is 0, and control authority is completely vested in the rotor system (integrated control performance information). ).when During cruise, the wings become the primary source of lift, with a lift ratio of approximately 1 and an aerodynamic weight of approximately 1. Control authority is then transferred to the aerodynamic control surfaces (integrating control performance information). During the transition phase, the weight changes smoothly as the aerodynamic lift increases, achieving a natural transfer of control authority.

[0106] In this specific application example, the relevant parameter configurations for weight calculation are as follows: Shape parameters Force control weighted filtering time constant Torque control weighted filtering time constant Singularity avoidance preset threshold .

[0107] Regarding the acquisition method of aerodynamic lift components, the lift acquisition step is triggered at the beginning of each control step (250Hz, i.e., every 4ms). This occurs when the sensors (angle of attack sensor, pitot tube) return valid data and the airspeed is relatively high ( When the sensor malfunctions or the signal-to-noise ratio is too low in the low-speed range, the aerodynamic lift component is calculated using the VLM aerodynamic lookup table method; when the sensor malfunctions or the signal-to-noise ratio is too low in the low-speed range, the aerodynamic lift component is calculated using the VLM aerodynamic lookup table method. When the aerodynamic lift component is obtained, the system automatically switches to the thrust compensation method.

[0108] Since the core function of this application is mainly reflected in the transition phase, the control authority in the hovering and cruise phases is clearly defined, with the rotor system and aerodynamic control surfaces each taking the lead independently. However, the transition phase is a critical interval where the control authority of the rotor and aerodynamic control surfaces fluctuates, and it is also the phase where traditional methods are most prone to control discontinuity and performance degradation. Therefore, the following uses the transition phase as an example to illustrate the complete execution flow of a typical control step.

[0109] Specifically, taking a certain moment in the transition phase as an example, the complete execution process is as follows: Step 1: Read the air speed Barometer height estimates density Calculate dynamic pressure Read the angle of attack From the table, we can find Wing area ,calculate Flight equipment quality , .

[0110] Step 2: Lift Ratio ; , ; .

[0111] Step 3: , , ; If the previous step was approximately 0.85, then the updated value in this step will be approximately 0.851, reflecting a smooth transition.

[0112] Step Four: It does not trigger the zeroing process; , .

[0113] Step 5: Calculate similarly , ( , ),have to , .

[0114] Step Six: Check for updates at work points (rotor efficiency) and (Aerodynamic efficiency), calculation: The torque channel is similarly applied. , ; Send to INDI control law. INDI control law: with For the performance matrix, calculate the actuator increment. It is then sent to the propulsion system and control surfaces, closing the control loop.

[0115] Step 7: Low-pass filtering ensures continuous, continuous, UUB holds true, and the error is bounded.

[0116] Therefore, in this specific application example, the aerodynamic weights are set to zero during hovering, and control authority belongs entirely to the rotor system; during cruise, the aerodynamic weights approach 1, and control authority is transferred to the aerodynamic control surfaces; during the transition state, the weights change continuously and smoothly with the lift ratio, and the rotor and control surfaces work together. The first-order low-pass filter ensures the smoothness of the weights, and singularity avoidance ensures numerical stability, achieving an adaptive and smooth transition of control authority across the entire control envelope.

[0117] To verify the actual control effect and robustness of the proposed method, a full-envelope flight simulation was also conducted on the MATLAB / Simulink platform. The simulated flight mission covered four stages: hovering, low-speed transition, high-speed cruise, and deceleration transition. Simulation results show that during the hovering stage, the first force control weight... and first torque control weight Maintaining a value of zero indicates that the singularity avoidance mechanism is in effect; at low speeds, the aerodynamic weights are set to zero, and control authority belongs entirely to the rotor system; during the transition phase... The curve rises monotonically along a smooth curve without chattering or oscillation, indicating that the first-order low-pass filter is effective and the weights transition smoothly with the increase of aerodynamic lift; during the cruise phase, and The system stabilized at approximately 0.9, confirming a complete transition to wing-dominated mode, with aerodynamic control surfaces assuming primary control authority. These simulation results validate the effectiveness of the proposed method for a smooth transition of control authority during full-envelope flight.

[0118] In addition, this application also applies to four wind conditions (wind speed) , , , The full envelope flight mission was repeated to verify the perturbation robustness of the proposed method. Table 1 shows the comparison results of the tracking performance RMSE of each channel under different wind conditions. As can be seen from Table 1, the maximum transient error occurs in the transition phase, and airspeed fluctuations caused by gusts will temporarily perturb the signal. However, the first-order low-pass filter effectively attenuated the disturbance. As the wind speed decreased... Increase to The tracking error of each channel increased accordingly, but remained within a controllable range. (Forward velocity channel) RMSE increased from 0.6582 to 2.8004. (Lateral velocity channel) increased from 0.4837 to 1.3820. (Vertical velocity channel) increased from 0.1783 to 2.8557, yaw angle channel Increasing the tracking error from 0.0556 rad to 0.0655 rad consistently satisfies the ultimate boundedness (UUB) constraint. These results verify the robustness of the proposed method under bounded perturbations, consistent with the theoretical conclusions of the aforementioned Lyapunov stability verification.

[0119] Table 1: Comparison of Tracking Performance RMSE under Different Wind Conditions

[0120] The simulation results above verify the effectiveness of the proposed method in smoothly transitioning control authority during full envelope flight, as well as its robustness under bounded perturbations, which is consistent with the theoretical conclusions of the aforementioned Lyapunov stability verification.

[0121] According to the above flight control method, this application also provides a flight device, which includes: a fuselage; a rotor system connected to the fuselage and including multiple rotors; aerodynamic control surfaces connected to the fuselage; and a flight controller that is communicatively connected to the rotor system and the aerodynamic control surfaces, respectively, for executing the flight control method of this application.

[0122] It should be noted that the description of the device in this application embodiment is similar to the description of the method embodiment above, and has similar beneficial effects as the method embodiment, therefore it will not be repeated. For any technical details not covered in the flight equipment provided in this application embodiment, please refer to... Figures 1 to 4 The meaning is understood in accordance with the description of any of the accompanying drawings.

[0123] It should be understood that the various forms of processes shown above can be used to rearrange, add, or delete steps. For example, the steps described in this application can be executed in parallel, sequentially, or in different orders, as long as the desired result of the technical solution disclosed in this application can be achieved, and this is not limited herein.

[0124] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this application, "a plurality of" means two or more, unless otherwise explicitly specified.

[0125] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A flight control method, characterized in that, The method includes: Acquire the aerodynamic lift component of the flight equipment in the current control step and the gravity of the flight equipment; Based on the ratio of the aerodynamic lift component to the gravity, a first control reconfiguration for the aerodynamic control surfaces and a second control reconfiguration for the rotor system are determined. Acquire first control effectiveness information characterizing the dominant state of the aerodynamic control surface and second control effectiveness information characterizing the dominant state of the rotor system; Based on the first control reorganization and the second control reorganization, the first control effectiveness information and the second control effectiveness information are fused to obtain fused control effectiveness information; Based on the fused control performance information, an actuator incremental command is generated, and in the current control step, the rotor system and the aerodynamic control surfaces are driven according to the actuator incremental command.

2. The method according to claim 1, characterized in that, The acquisition of the aerodynamic lift component of the flight equipment in the current control step includes: Obtain the aerodynamic model of the flight equipment; In response to the aerodynamic model meeting preset conditions, the angle of attack, airspeed, and atmospheric density of the flight equipment in the current control step are obtained; The dynamic pressure is determined based on the airspeed and the atmospheric density. The lift coefficient is determined based on the angle of attack. The aerodynamic lift component is determined based on the lift coefficient, the dynamic pressure, and the wing area of ​​the flight equipment.

3. The method according to claim 1, characterized in that, The acquisition of the aerodynamic lift component of the flight equipment in the current control step includes: Obtain the aerodynamic model of the flight equipment; In response to the aerodynamic model not meeting the preset conditions, the thrust and tilt angle of each rotor in the rotor system in the current control step are obtained; Based on the thrust and tilt angle of each rotor, determine the resultant force of the rotor system thrust in the Z-axis direction of the fuselage; The aerodynamic lift component is determined based on the gravity and the resultant force.

4. The method according to claim 3, characterized in that, The step of obtaining the thrust of each rotor in the rotor system in the current control step includes: Obtain the rotational speed of each rotor in the current control step; The thrust of each rotor is determined based on the rotational speed of each rotor and the preset rotational speed-thrust mapping relationship.

5. The method according to claim 1, characterized in that, The first control reorganization includes a first torque control weight and a first force control weight; the second control reorganization includes a second torque control weight and a second force control weight. The step of determining the first control reconfiguration for the aerodynamic control surfaces and the second control reconfiguration for the rotor system based on the ratio of the aerodynamic lift component to the gravity includes: Based on the ratio of the aerodynamic lift component to the gravity, the original aerodynamic weights of the force channel and the torque channel are determined respectively; wherein, each of the original aerodynamic weights is positively correlated with a power function of the ratio, and the exponent of the power function is a shape parameter; The original aerodynamic weights are saturated to obtain the original aerodynamic weights of the force channel and the original aerodynamic weights of the torque channel after saturation. The original aerodynamic weights of the force channel and the torque channel after saturation are subjected to first-order low-pass filtering with the filtering time constant to obtain the filtered aerodynamic weights of the force channel and the torque channel. The filtered aerodynamic weights of the force channel are subjected to singularity avoidance processing to obtain the first force control weights; the filtered aerodynamic weights of the torque channel are subjected to singularity avoidance processing to obtain the first torque control weights. The second force control weight is determined based on the first force control weight; the second torque control weight is determined based on the first torque control weight.

6. The method according to claim 1, characterized in that, The acquisition of first control effectiveness information characterizing the dominant state of the aerodynamic control surface and second control effectiveness information characterizing the dominant state of the rotor system includes: Acquire aerodynamic data and the dynamic pressure and angle of attack of the flight equipment in the current control step, interpolate the aerodynamic data based on the dynamic pressure and the angle of attack, and determine the first control effectiveness information; the first control effectiveness information is used to characterize the sensitivity relationship between the deflection increment of each aerodynamic control surface and the force channel and torque channel. The operating point information of the flight equipment in the current control step is obtained, and the operating point information is analyzed and calculated by the rotor dynamics model, or by hovering state identification, to determine the second control performance information; the second control performance information is used to characterize the sensitivity relationship between the increment of each rotor actuator and the force channel and torque channel.

7. The method according to claim 1, characterized in that, Both the first control performance information and the second control performance information include force channel performance sub-information and torque channel performance sub-information; The step of fusing the first control performance information and the second control performance information based on the first control reorganization and the second control reorganization to obtain fused control performance information includes: Based on the first force control weight in the first control reorganization and the second force control weight in the second control reorganization, the force channel performance sub-information of the first control performance information and the force channel performance sub-information of the second control performance information are weighted and fused to obtain force channel fusion performance information; Based on the first torque control weight in the first control reorganization and the second torque control weight in the second control reorganization, the torque channel performance sub-information of the first control performance information and the torque channel performance sub-information of the second control performance information are weighted and fused to obtain torque channel fusion performance information; The force channel fusion performance information and the torque channel fusion performance information are fused to obtain the fused control performance information.

8. The method according to claim 1, characterized in that, The step of generating actuator incremental instructions based on the fused control performance information includes: Obtain the virtual control increment and the historical actuator increment of the previous control step; Based on the fused control performance information, the virtual control increment, and the historical actuator increment, the actuator incremental command is calculated using an incremental nonlinear dynamic inverse control law.

9. The method according to claim 1, characterized in that, The method further includes: Obtain the reference control performance information of the flight equipment; Based on the reference control performance information and the fused control performance information, performance matching information is determined; Determine the minimum eigenvalue of the symmetric portion of the performance matching information; In response to the minimum eigenvalue being greater than zero, the fused control performance information is determined to have passed the stability verification.

10. A flight device, characterized in that, include: body; A rotor system, connected to the fuselage, includes multiple rotors; Aerodynamic control surfaces are connected to the fuselage; The flight controller is communicatively connected to both the rotor system and the aerodynamic control surfaces, and is used to execute the flight control method according to any one of claims 1-9.

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