CN112947527A - Flight control method and device for airplane - Google Patents
Flight control method and device for airplane Download PDFInfo
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Abstract
A flight control method and apparatus for an aircraft is disclosed herein. A flight control apparatus for an aircraft may include: a vertical tail load processing module for determining an upper limit of a sideslip angle according to the vertical tail load of the airplane; a sideslip command module that receives a sideslip manipulation input and an upper sideslip angle limit provided by the vertical tail load processing module, and generates a sideslip angle command that does not exceed the upper sideslip angle limit based on the sideslip manipulation input; a feedback signal processing module that generates sideslip angle feedback for the aircraft based on sensor measurements; and a controller that generates a yaw control command to control a rudder of the aircraft based on the sideslip angle command and the sideslip angle feedback. Corresponding flight control methods and systems for an aircraft are also disclosed.
Description
Technical Field
The present invention relates generally to the field of aircraft, and more particularly to a flight control method and apparatus for an aircraft.
Background
The Flight Control System ("FCS", for short, Flight Control System) of a modern aircraft is the core of the whole aircraft onboard System. The flight control system controls the flight attitude and the flight track of the airplane by utilizing the movement of the pneumatic control surface. Flight control systems have evolved to the present time, and have completed the transition from mechanical maneuvering to Fly-by-Wire. Fly-by-wire flight control systems have become the standard configuration for newly developed aircraft. In order to further improve the economy of civil aircrafts, the main stream two-channel wide-body airliner teleplane transverse direction control law adopts P-Beta control (namely, a side lever/steering wheel instruction roll angle rate and a foot pedal instruction sideslip angle).
25.353 'Rudder control reverts conditions' requirements are added to CS25 attribute 22 newly released by EASA at 11/5/2018, and pedal repeated operation is used as the restriction of maneuvering load; the FAA issued "Yaw manager Conditions-Rudder reviews" on 16.7.2018 and proposed that the terms solicit public opinion, and soon was expected to supplement the FAR25.353 "Rudder control reviews" requirements.
Accordingly, there is a need in the art for improved flight control methods and apparatus.
Disclosure of Invention
A flight control technique is provided that generates a sideslip angle command based on a sideslip maneuver input, the sideslip angle command not exceeding a sideslip angle upper limit determined from a vertical tail load of an aircraft, thereby providing an aircraft lateral heading (P-Beta) control law that satisfies a load limiting requirement.
In one embodiment of the present invention, a flight control apparatus for an aircraft may comprise: a vertical tail load processing module for determining an upper limit of a sideslip angle according to the vertical tail load of the airplane; a sideslip command module that receives a sideslip manipulation input and an upper sideslip angle limit provided by the vertical tail load processing module, and generates a sideslip angle command that does not exceed the upper sideslip angle limit based on the sideslip manipulation input; a feedback signal processing module that generates sideslip angle feedback for the aircraft based on sensor measurements; and a controller that generates a yaw control command to control a rudder of the aircraft based on the sideslip angle command and the sideslip angle feedback.
In an aspect, the sideslip instruction module generates a sideslip angle change rate instruction based on the sideslip manipulation input, the feedback signal processing module generates inertial sideslip angle change rate feedback of the aircraft based on sensor measurements, and the controller generates a sideslip angle rate command as part of the yaw control command based on the sideslip angle change rate instruction and the inertial sideslip angle change rate feedback.
In one aspect, the flight control apparatus further comprises: a roll command module that receives a roll maneuver input and generates a roll rate command based on the roll maneuver input, wherein the feedback signal processing module generates roll rate feedback for the aircraft based on sensor measurements, and the controller generates roll control commands based on the roll rate command and the roll rate feedback to control ailerons and spoilers of the aircraft.
In one aspect, the feedback signal processing module comprises: an inertial sideslip angle change rate estimation module that determines an inertial sideslip angle change rate of the aircraft based on an aircraft yaw rate, a roll angle, and an airframe lateral overload measured by sensors; and a complementary filter that complementarily filters the rate of change of the inertial sideslip angle and the sideslip angle measured by a sideslip angle sensor on the aircraft to generate the sideslip angle feedback.
In an aspect, the vertical tail load of the aircraft is a current vertical tail load estimated by the vertical tail load processing module based on a speed, a side slip angle, a yaw rate, and a rudder deflection of the aircraft.
In an aspect, the side-slip command module adjusts a coefficient that converts the side-slip maneuver input to the side-slip angle command based on the current estimated vertical tail load.
On one hand, the upper limit of the side slip angle corresponding to the smaller vertical tail load is larger than the upper limit of the side slip angle corresponding to the larger vertical tail load, and on the basis of meeting the airplane manipulation capability, the vertical tail load is guaranteed not to exceed the limit.
In an aspect, the vertical tail load of the aircraft is a rated vertical tail load, wherein the vertical tail load processing module determines the upper sideslip angle limit based on the rated vertical tail load of the aircraft, and a speed, a yaw rate, and a rudder deflection of the aircraft.
In one aspect, the controller includes a linear quadratic regulation with integral (LQR) controller that performs proportional-integral control of a difference of the sideslip angle command and the sideslip angle feedback to generate the yaw control command.
In another embodiment of the invention, a flight control method for an aircraft may include: receiving a sideslip maneuver input of the aircraft; determining a sideslip angle upper limit based on a vertical tail load of the aircraft; generating a sideslip angle command not exceeding the upper sideslip angle limit based on the sideslip manipulation input; generating sideslip angle feedback for the aircraft based on sensor measurements; and generating a yaw control command to control a rudder of the aircraft based on the sideslip angle command and the sideslip angle feedback.
In one aspect, the flight control method further comprises: generating a side-slip angle change rate command based on the side-slip manipulation input; generating inertial sideslip angle change rate feedback for the aircraft based on sensor measurements; and generating a side-slip angle rate command as part of the yaw control command based on the side-slip angle change rate command and the inertial side-slip angle change rate feedback.
In one aspect, the flight control method further comprises: receiving a roll maneuver input and generating a roll angle rate instruction based on the roll maneuver input; generating roll angle change rate feedback for the aircraft based on sensor measurements; and generating roll control commands based on the roll rate command and the roll rate of change feedback to control ailerons and spoilers of the aircraft.
In one aspect, the flight control method further comprises: determining a rate of inertial sideslip angle change of the aircraft based on an aircraft yaw rate, a roll angle, and an aircraft body lateral overload measured by sensors; and complementarily filtering the inertial sideslip angle rate of change and a sideslip angle measured by a sideslip angle sensor on the aircraft to generate the sideslip angle feedback.
In one aspect, the vertical tail load of the aircraft is a current vertical tail load estimated based on a speed, a side slip angle, a yaw rate, and a rudder deflection of the aircraft.
In one aspect, the flight control method further comprises: adjusting coefficients that convert the side-slip steering input to the side-slip angle command based on the current estimated vertical tail load.
In one aspect, a lower vertical tail load corresponds to a greater upper side slip angle limit than a higher vertical tail load.
In an aspect, the vertical tail load of the aircraft is a rated vertical tail load, wherein the upper sideslip angle limit is determined based at least in part on the rated vertical tail load of the aircraft, and a speed, a yaw rate, and a rudder deflection of the aircraft.
In one aspect, the flight control method further comprises: further comprising performing proportional-integral control of a difference of the sideslip angle command and the sideslip angle feedback to generate the yaw control command.
In another embodiment of the present invention, a flight control system for an aircraft may comprise: a processor; and a memory for storing processor-executable instructions, wherein the processor is configured to execute the processor-executable instructions to implement a flight control method as described above.
The invention is applicable to various types of airplanes, such as civil airplanes, military airplanes, unmanned planes and the like. For example, the flight control techniques of the present invention may provide an aircraft lateral heading (P-Beta) control law that satisfies load limiting requirements. Therefore, the vertical tail load is not over-limited on the basis of meeting the airplane control capacity. According to the flight control method and the flight control device, the course static stability of the large airplane can be enhanced, the characteristics of Dutch rolling and coordinated turning are improved, and the riding comfort level is improved.
Drawings
FIG. 1 is a block diagram of a control architecture for an aircraft according to one embodiment of the invention.
FIG. 2 is a schematic diagram of a roll command module according to one embodiment of the invention.
FIG. 3 is a schematic diagram of a sideslip command module according to one embodiment of the present disclosure.
Fig. 4 is a schematic diagram of a feedback signal processing module according to an embodiment of the invention.
FIG. 5 is a schematic diagram of a coordinated turn time history curve according to one embodiment of the invention.
FIG. 6 is a schematic illustration of a yaw maneuver time history curve according to an embodiment of the present invention.
Figure 7 is a schematic diagram of a sideslip maneuver input time history, in accordance with one embodiment of the present invention.
FIG. 8 is a flow chart of a flight control method for an aircraft according to one embodiment of the invention.
Detailed Description
The present invention will be further described with reference to the following specific examples and drawings, but the scope of the present invention should not be limited thereto.
A flight control technique is provided that generates a sideslip angle command based on a sideslip maneuver input, the sideslip angle command not exceeding a sideslip angle upper limit determined from a vertical tail load of an aircraft, thereby providing an aircraft lateral heading (P-Beta) control law that satisfies a load limiting requirement. According to the flight control method and the flight control device, the course static stability of the large airplane can be enhanced, the characteristics of Dutch rolling and coordinated turning are improved, and the riding comfort level is improved.
Fig. 1 is an architectural block diagram of a control system 100 for an aircraft according to one embodiment of the invention. The aircraft may include various input components (e.g., a side bar 101, foot pedals 102, a speed brake handle, trim switches, a slat handle, a flight mode control panel, etc.). These input assemblies receive pilot steering inputs and may generate flight steering commands accordingly for controlling the actuators 106 of the aircraft, thereby controlling the movement (e.g., longitudinal, lateral, heading) of the aircraft. In a drone, the input component may be a remote control device and may include various remote control components for manipulation by a user and generating flight manipulation instructions accordingly.
The control system 100 drives the control surface deflection of the airplane through a control law module (or model) so as to control the flight state of the airplane. Such control laws modules can implement longitudinal, lateral, and heading control laws. The control law module may generate flight commands based on flight maneuver instructions, aerodynamic parameters, measurement feedback parameters, and the like. FIG. 1 shows that the control system 100 may include a P-Beta (lateral and heading) control law module 110 that may control roll rate (i.e., lateral, e.g., control ailerons and spoilers) based on input commands from the sidebar 101 (or the steering wheel, or other input component) and sideslip angle (i.e., heading, e.g., control rudder) based on input commands from the foothold 102 (or other input component). Although not shown, the vertical control law module may be implemented in combination with or separately from the P-Beta control law module 110. In other embodiments, the lateral and heading control laws in the P-Beta control law module 110 may also be implemented separately.
The control law (control law) represents the functional relation between the deflection of the control surface and the flight state and the flight control input. The aerodynamic parameters of the aircraft can be obtained through wind tunnels or test flights. The measurement feedback parameters may be obtained using measurements from sensors mounted on the aircraft, including, for example, true airspeed, angle of attack, pitch angle, pitch rate, overload, rudder deflection, etc. The signals detected by the sensors may be pre-processed, such as, but not limited to, filtering, correction, coordinate transformation, parameter combination, parameter calculation, and the like. The control law modules (e.g., P-Beta control law module 110) may be implemented by a computer or other electronic device for converting flight maneuver instructions generated by input components (e.g., side poles 101, foot pedals 102, control panels, etc.) into flight commands for controlling the actuators 106 according to various algorithms. To further improve flight safety, multiple sets of control systems may be employed to execute the same command, which is referred to as a redundant fly-by-wire system.
As shown in FIG. 1, as the pilot manipulates the sidestick 101 to generate a roll maneuver input (e.g., sidestick lateral displacement), the roll command module 120 may generate a roll rate command p based on the roll maneuver inputm. Controller 140 may be based on roll rate command pmAnd roll angle rate of change feedback from the feedback signal processing module 170 to generate the roll control command δa_cmdTo control the ailerons and spoilers of the aircraft to cause the aircraft to roll laterally. For example, the feedback signal processing module 170 may generate roll angle rate of change feedback for the aircraft based on the signals measured by the sensors 160. In one embodiment, the controller 140 may be configured to control the roll rate based on the roll rate command pmAnd roll angle change rate feedback from the feedback signal processing module 170 to perform feedback control to obtain a roll control command deltaa_cmd. For example, the roll rate command p may be adjustedmProportional plus integral control with the roll angle rate of change feedback difference to generate a roll control command deltaa_cmdSo that the roll rate of change of the aircraft tracks the roll rate command pm. In another embodiment, the roll control command δa_cmdMay be superimposed at summer 105 with sidebar-steering-based sidebar feedforward compensation 103 to generate a compensated roll control command δ'a_cmdFor controlling aileron actuators and spoiler actuators of an aircraft. By way of example and not limitation, controller 140 may be a Linear-Quadratic-Regulator (LQR) controller with integration or a Linear-Quadratic Regulator (LSRC) controller thereofHis type of controller.
Additionally or alternatively, a sideslip maneuver input (e.g., a pedal displacement) will be generated when the pilot manipulates the pedals 102, and the sideslip command module 130 may generate a sideslip angle command β based on the sideslip maneuver inputmAnd sideslip angle rate of change command
Controller 140 may be based on a sideslip angle command βmSideslip angle rate of change command
And sideslip angle feedback, inertial sideslip angle change rate feedback from feedback signal processing module 170 to generate yaw control command δr_cmdTo control the rudder of the aircraft to yaw the aircraft. For example, the feedback signal processing module 170 may generate sideslip angle feedback and inertial sideslip angle change rate feedback for the aircraft based on the signals measured by the sensors 160.
In one embodiment, controller 140 may command β according to the sideslip anglemAnd sideslip angle feedback from feedback signal processing module 170 to obtain the sideslip angle command (e.g., as yaw control command δ)r_cmdA portion of (a). For example, controller 140 may command a slip angle βmAnd performing proportion + integral control on the difference value fed back by the sideslip angle to generate a sideslip angle command so that the sideslip angle of the airplane tracks the sideslip angle command betam。
Additionally or alternatively, in one embodiment, the controller 140 is based on a side-slip angle change rate command
And inertial sideslip angle change rate feedback from feedback signal processing module 170 to obtain a sideslip angle rate command (e.g., as yaw control command δ)r_cmdA portion of (a). For example, the controller 140 may command the rate of change of slip angle
Proportional control is carried out on the difference value fed back by the sideslip angle change rate to generate a sideslip angle speed command, so that the sideslip angle change rate of the airplane tracks the sideslip angle change rate command
In another embodiment, the controller 140 generates a yaw control command δr_cmdMay be superimposed at summer 105 with pedal feedforward compensation 104 based on pedal manipulation to generate compensated yaw control command δ'r_cmdFor controlling rudder actuators of an aircraft.
While fig. 1 illustrates the roll manipulation input being generated by the sidebar 101 and the yaw manipulation input being generated by the foot peg 102, in other implementations, other input components may be used to generate the roll manipulation input and the yaw manipulation input, such as a remote control, a control panel, and so forth.
The aircraft body dynamics model 107 can calculate the motion parameters of the aircraft, such as the rolling rate, the yaw rate, the sideslip angle, the lateral overload, and the like, according to the control surface control commands output by the control law module 110 and/or the flight state of the aircraft. These motion parameters may then be used for flight modeling or calculation.
Onboard aircraft mounted sensors 160 may measure aircraft flight parameters such as true airspeed, angle of attack, pitch angle, pitch rate, overload, rudder deflection angle, and the like. According to one embodiment of the invention, the feedback signal processing module 170 may estimate the rate of change of the sideslip angle of inertia based on the angular rate, roll angle, and body side overload measured by the sensors 160
This rate of inertial slip angle change may be fed back to the controller 140. In another embodiment, an AOS (Angle of sideslip) sensor on the aircraft may measure the sideslip angle of the aircraft, which may be fed back to the controller 140 via the feedback signal processing module 170. In one embodiment according to the invention, the feedback signal processing module 170 may correlate the estimated inertial sideslip angle rate of change with the sideslip measured by AOS (sideslip angle) sensors on the aircraftThe angles are complementarily filtered to generate a new sideslip angle betaestAs a side slip angle feedback signal to the controller 140. Accordingly, the controller 140 may command β according to the sideslip anglemAnd sideslip angle rate of change command
And the sideslip angle β provided by the feedback signal processing module 170estAnd rate of change of inertial side slip angle
Performing feedback control to generate a yaw control command δr_cmd。
As described above, repeated actuation of the foot peg by the pilot may cause the droop loads to exceed limits. According to one embodiment of the invention, the vertical tail load processing module 150 may determine the upper limit of the side slip angle according to the vertical tail load of the aircraft, and may provide the vertical tail load of the aircraft and/or the upper limit of the side slip angle corresponding to the vertical tail load to the control law module 110 (specifically, to an input of the sideslip command module 130), whereby the sideslip command module 130 may generate the side slip angle command β not exceeding the upper limit of the side slip angle based on the sideslip maneuver inputm. Generating a side-slip angle command beta based on a side-slip steering input and a vertical-tail loadmThe vertical tail load can be prevented from exceeding the limit even when the pilot repeatedly operates the pedals.
In one aspect, the vertical tail load processing module 150 may estimate the current vertical tail load of the aircraft based on aircraft aerodynamic data and flight parameters. For example, the vertical tail load processing module 150 may calculate the current estimated vertical tail load based on the speed, the side slip angle, the yaw rate, and the rudder deflection of the aircraft. Based on the current estimated droop load, a corresponding upper side slip angle limit may be determined. By way of example and not limitation, there may be a predetermined mapping between aircraft vertical tail loads and upper side slip angles, and the corresponding upper side slip angle limits may be determined based on the current vertical tail loads (as well as aircraft speed, yaw rate, and rudder deflection) using functions, look-up tables, interpolation methods, and the like. For example, a smaller vertical tail load may correspond to a greater upper side slip angle limit than a larger vertical tail load.
In one example, the sideslip command module 130 may generate the sideslip angle command β using a function or algorithm based on the sideslip maneuver input and the vertical tail loadm. In another example, conversion of the side-slip steering input to the side-slip angle rate command β may be adjusted based on the estimated vertical-tail loadmThe coefficient of (a). In another example, estimating the droop load may limit a side-slip angle command β generated based on the side-slip maneuver inputmMaximum value of (beta)cmd_max. For example, when the estimated vertical tail load is large, the maximum value βcmd_maxMay be small (not exceeding the upper limit of side slip angle for maximum load) whereby a large side slip maneuver input results in a limited side slip angle command βmSo that the actual vertical tail load of the aircraft does not exceed the threshold value.
In another aspect, the droop load processing module 150 may determine the upper side-slip angle limit based on the rated droop load of the aircraft, and the speed, yaw rate, and rudder deflection of the aircraft. By way of example and not limitation, there may be a predetermined mapping between the rated vertical tail load (as well as aircraft speed, yaw rate, and rudder deflection) and the upper limit of the side slip angle, and the corresponding upper limit of the side slip angle may be determined using a function, a look-up table, interpolation, or the like. That is, under different flight conditions, the upper sideslip angle limit may be different, and the sideslip angle command βmThe actual vertical fin loading of the aircraft is not caused to exceed the threshold.
By way of example and not limitation, aircraft configuration data, current flight conditions, side slip angle, yaw rate, lateral overload, rudder deflection, and aircraft related aerodynamic data versus aircraft vertical tail loading may be represented as follows:
FV＝[-(CYβ-CYβto)β-(CYβ-CYβto)rbLV/VT+CYδrδr]QeSW
wherein:
1)FVrepresents the vertical tail load;
2)CYβrepresenting the derivative of the coefficient of total machine side force against the sideslip angle, CYβtoRepresenting the derivative of the coefficient of tailless side force of the entire machine to the sideslip angle, CYδrThe derivative of the full-aircraft lateral force coefficient to the rudder is represented, and the derivative belongs to the aerodynamic data of the aircraft;
3) beta represents the true sideslip angle of the aircraft, rbRepresenting the aircraft yaw rate, deltarRepresenting rudder deflection;
4)VTrepresents the vacuum velocity, QeRepresents the dynamic pressure of flight;
5)LVindicating the length of the vertical tail arm, SWThe area of the vertical tail is shown as a constant.
For a fixed flight state (e.g., 300 knots, height 30000ft), the aerodynamic data in 2) above is fixed, and the true airspeed, dynamic flight pressure of the aircraft in 4) above is determined. The vertical tail loading is therefore directly related to only the side slip angle, yaw rate, rudder deflection in 3) above. If the maximum load (e.g., rated load), yaw rate, rudder deflection, and aerodynamic data for the droop are known, the maximum allowable sideslip angle (i.e., upper sideslip angle limit) may be calculated.
The true sideslip angle beta of the aircraft is directly related to the sideslip angle command betamAnd (4) correlating. In the control law, the estimated sideslip angle β of the aircraft is fed backestWith sideslip angle command betamIn comparison, the rudder deflection is then controlled so that the sideslip angle of the aircraft tracks the sideslip angle commanded by the control law. The control law thus adjusts the maximum sideslip angle command value β based on the foot pedal maneuver input and the aircraft vertical tail loadingcmd,maxThe airplane vertical tail control device can meet the requirement of airplane control capacity and can control vertical tail load not to exceed the limit.
Although fig. 1 shows a separate vertical tail load processing module 150, in other embodiments, the vertical tail load processing module 150 may also be implemented in the feedback signal processing module 170 or in the side-slip command module 130. The various components shown in fig. 1, such as the control law module 110 or sub-components thereof, the vertical tail load processing module 150, the feedback signal processing module 170, the aircraft body dynamics model 107, etc., may be implemented separately or collectively using hardware, software, hardware, firmware, etc., such as using a processor, computer, controller, application specific integrated circuit, etc., without departing from the scope of the present invention.
FIG. 2 is a schematic diagram of a roll command module according to one embodiment of the invention. The roll command module 120 (see FIG. 1) may generate a roll rate command p based on the roll maneuver inputmFor controlling the ailerons and spoilers of an aircraft.
As described above, roll maneuver inputs, such as lateral sidestick lateral displacement, will be generated when the pilot maneuvers the sidestick 101. The roll signal conversion module 202 may generate a roll signal p based on a roll manipulation input (e.g., lateral rod lateral displacement)cmd. The roll signal conversion module 202 may convert the side-bar lateral displacement into a roll signal p based on a predetermined function, algorithm, or the likecmd. Roll signal pcmdMay indicate roll direction, roll rate, etc. For example, for a 50% sidebar lateral input, the roll signal pcmdMay be 10 degrees/second.
The adder 204 adds the roll signal pcmdSubtracting feedback (e.g., p) from the output of roll command module 120m) The difference is passed through a gain module 206 (e.g., multiplied by the frequency ω of the roll command model)pm) Then passes through an integrator 208 to output a roll rate command pm. Roll rate command pmCan be used for controlling the ailerons and spoilers of the aircraft so that the aircraft performs a rolling action.
Referring to FIG. 1, the feedback signal processing module 170 may generate a roll rate p for the stable axis of the aircraft based on the measurement signals of the sensors 160s. For example, based on the body roll rate and yaw rate provided by the inertial sensors, and the angle of attack measurements provided by the angle of attack sensors, etc., the roll rate p for the stable axis of the aircraft may be calculateds. The controller 140 may be configured to control the roll rate based on the roll rate command pmAnd the roll rate p of the aircraft's stable axis provided by the feedback signal processing module 170sPerforming feedback control to generate a roll control command δa_cmd(e.g., roll rate commands) for controlling aileron actuators and spoiler actuators of the aircraft. By way of example and not limitation, the controller 140 may be based on roll angleRate instruction pmAnd roll rate p of the aircraft stabilizer shaftsThe difference between them is subjected to proportional plus integral control to generate a roll control command deltaa_cmd. For example, roll control command δa_cmdThe roll rate increment can be indicated for controlling the aileron actuators and spoiler actuators of the aircraft such that the roll rate p of the aircraftsTracking roll rate command pm。
FIG. 3 is a schematic diagram of a sideslip command module according to one embodiment of the present disclosure. The side-slip command module 130 (see FIG. 1) may generate a side-slip angle command β based on the side-slip manipulation inputmFor controlling the rudder of an aircraft. Alternatively, the sideslip command module 130 may further generate a sideslip angle change rate command based on the sideslip maneuver input
As well as for controlling the rudder of an aircraft.
Sideslip maneuver inputs, such as pedal displacement, are generated when the pilot maneuvers the pedals 102. The side-slip signal conversion module 302 may generate the side-slip signal β based on a side-slip manipulation input (e.g., pedal displacement)cmd. The side-slip signal conversion module 302 may convert the pedal displacement into the side-slip signal β based on a predetermined function, algorithm, or the likecmd. Side-slip signal betacmdThe direction of sideslip, angle of sideslip, etc. may be indicated.
As described above, in one embodiment according to the present invention, the sideslip command module 130 may use aircraft vertical tail loading to limit the sideslip angle command βm. For example, estimating the vertical-tail load may adjust the generation of the side-slip signal β based on the side-slip steering inputcmdLimiting the sideslip signal betacmd(or sideslip Angle Command βm) Maximum value of (beta)cmd_maxAnd the like.
Referring to fig. 1, the controller 140 respectively controls the slip angles according to the slip angle commands βmAnd sideslip angle rate of change command
Modifying the sideslip angle feedback and inertial sideslip angle change rate feedback from the feedback signal processing module 170 to generate a yaw control command δr_cmdThe rudder of the aircraft is controlled so that the aircraft is yawed. The feedback signal processing module 170 may estimate the rate of change of the side-slip angle of inertia based on the angular rate, roll angle, and body side-loading measured by the sensors 160
And fed back as the rate of change of side slip angle. In one embodiment, the feedback signal processing module 170 may generate sideslip angle feedback based on the actual sideslip angle of the aircraft measured by the sensor 160 (e.g., AOS sensor). In another embodiment, the feedback signal processing module 170 may estimate the rate of change of the inertial slip angle
Complementarily filtering the sideslip angle measured by the AOS sensor to generate a composite sideslip angle betaestAs a side slip angle feedback.
As described above, controller 140 may command β based on the sideslip anglemAnd the sideslip angle feedback provided by the feedback signal processing module 170 (e.g., measured sideslip angle, or synthetic sideslip angle βest) The difference between the two is subjected to proportional plus integral control to generate a yaw control command deltar_cmd. For example, yaw control command δr_cmdThe sideslip angle increment may be indicated for controlling a rudder actuator of the aircraft such that an actual sideslip angle of the aircraft tracks the sideslip angle command βm. Additionally, controller 140 may command the rate of change of slip angle
And sideslip angle rate of change feedback
Is proportionally controlled to generate a sideslip angular rate command (as a yaw control command delta)r_cmdPart of) such that the aircraft sideslip angle change rate tracks the sideslip angle change rate command
The structures and parameters shown in fig. 2 and 3 are exemplary only and not limiting. In particular practice, other structures or algorithms (e.g., control law models) may be employed to convert the roll maneuver input into the roll rate command pmAnd converting the sideslip manipulation input into a sideslip angle command betamAnd sideslip angle rate of change command
By way of example and not limitation, according to the requirement of the grade 1 flight quality, the time constant of the rolling mode is 0.3-0.8 seconds, the amplitude doubling time of the spiral mode is at least 12 seconds, the damping ratio of the Dutch rolling mode is at least 0.4, and the frequency of the Dutch rolling mode is up toThe specific requirements of 0.4rad/s less and the Dutch roll damping not lower than 0.35 and the like can be determined, and the parameters omega of the roll instruction module 120 and the sideslip instruction module 130 can be determinedpm(frequency of roll instruction model), ωβm(frequency of sideslip command model), ξm(damping of sideslip command model), and the like.
Fig. 4 is a schematic diagram of a feedback signal processing module according to an embodiment of the invention. As described above, the feedback signal processing module 170 may complementarily filter the estimated rate of change of the inertial side-slip angle and the measured side-slip angle of the AOS sensor using a complementary filter to generate the synthetic side-slip angle βest。
In one aspect, as shown in FIG. 4, the sideslip angle change rate estimation module 410 may be based on a yaw rate r of the aircraft's stable axissRolling angle phi of airplane body and lateral overload n of airplane bodyybTo estimate the rate of change of the inertial side slip angle
Yaw rate r of aircraft stabilizer shaftsRolling angle phi of airplane body and lateral overload n of airplane bodyybMay be measured by sensors on the aircraft or calculated from sensor measurement data. Estimating inertial side slip angle rate of change
May employ existing algorithms or algorithms developed in the future without limiting the scope of the invention. In one example, the feedback signal processing module 170 may estimate the rate of change of the inertial slip angle
To the controller 140.
On the other hand, AOS sensor 401 may be mounted on the aircraft (e.g., nose position) for directly measuring the sideslip angle of the aircraft. The sideslip angle signal measured by AOS sensor 401 may be filtered by low pass filter 402 (where τ is the filter time constant) to generate a measured sideslip angle βADS. The feedback signal processing module 170 may estimate the rate of change of the inertial slip angle
Sideslip angle beta measured with AOS sensorADSComplementary filtering to generate a composite sideslip angle betaest。
The block 403-408 is shown in FIG. 4 as a complementary filter (e.g., a second order complementary filter with a frequency ωnDamping xi). Measuring the sideslip angle betaADSThe feedback signal (e.g., beta) from the output of the feedback signal processing module 170 is subtracted at adder 403est) The difference reaches the adder 407 via the gain module 404 and the integrator 405 (e.g., signal)
) While the difference output by adder 403 also reaches adder 407 via gain module 406 (e.g., signal)
). The adder 407 adds the estimated rate of change of the side-slip angle of inertia
And signal
And
the result of the addition is provided to an integrator 408 to generate a composite slip angle βest。
The block 403-408 shown in fig. 4 is only one specific implementation of a complementary filter. Other forms of complementary filters may be used by those skilled in the art to estimate the rate of change of the side-slip angle of inertia
Sideslip angle beta measured with AOS sensorADSPerforming fusion to generate a synthetic sideslip angle βestSuch alternative implementations fall within the scope of the inventionWithin.
Accordingly, referring to FIG. 1, controller 140 responds to sideslip angle command βmAnd sideslip angle rate of change command
Estimated rate of change of side-slip angle of inertia provided to feedback signal processing module 170
And resultant sideslip angle betaestMaking a correction to generate a yaw control command deltar_cmd. In one aspect, the controller 140 can be based on
And
and the difference value between the two is subjected to proportional control to generate a sideslip angular rate command. In another aspect, the controller 140 can be based on βmAnd betaestThe difference between the two is subjected to proportion and integral control to generate a sideslip angle command, and the sideslip angle of the airplane is controlled to track the sideslip angle command betam。
Synthetic sideslip angle beta of the present inventionestCapable of estimating inertial sideslip angle rate of change from single
Or angle of sideslip beta measured by AOS sensorADSProviding greater accuracy and stability. For example, it may reduce errors in the inertial sideslip angle change rate estimation caused by aircraft aerodynamic data, reduce the effects of gusts on the sideslip angle sensor, and the like.
After the P-Beta control law parameter design as in fig. 1-4 is completed, it is necessary to check whether the performance indexes of the control law frequency domain and the control law time domain meet the requirements, for example, the method includes:
(1) the stability margin of the control law needs to meet the condition that the amplitude margin is more than or equal to 6dB and the phase margin is more than or equal to 45 degrees;
(2) the open loop crossing frequency is required to be between 1.0rad/s and 5.0 rad/s;
(3) the rolling mode, the spiral mode and the Dutch rolling mode of the closed-loop control law meet the requirement of grade 1 flight quality;
(4) the transient side slip angle is less than or equal to 2 degrees and the transient side overload is less than or equal to 0.1g in the coordinated turning process;
(5) the steady state side slip angle is less than or equal to 0.5 degrees and the steady state side overload is less than or equal to 0.02g in the coordinated turning process;
(6) the steady-state control precision of the course sideslip angle is 0.25 degrees;
(7) the P-Beta control law at least has the anti-crosswind capability of 35 kn;
(8) under the condition of pedal full stroke operation, the operation allowance of at least 20-30% is transversely provided.
After the control law check is confirmed, the yaw maneuvering load is checked according to the latest CS25.353 clause of EASA (for example, the foot pedal performs the maneuvering input according to the requirement of fig. 7, and the vertical tail load is checked to meet the requirement).
FIG. 5 is a schematic diagram of a coordinated turn time history curve according to one embodiment of the invention. FIG. 5 shows the result of comparing the P-Beta control law according to the present invention (dashed line) with the conventional yaw damping control law (solid line) at an airspeed of 180kn and a height of 2000 ft. During a coordinated turn, the roll angles produced by the P-Beta control law and the yaw damping control law are similar for the same flight maneuver input (e.g., sidestick input), but the sideslip angle produced by the P-Beta control law of the present invention is significantly less than the sideslip angle produced by the yaw damping control law, i.e., the sideslip angle during a coordinated turn is reduced.
FIG. 6 is a schematic illustration of a yaw maneuver time history curve according to an embodiment of the present invention. FIG. 6 shows the results of a comparison of the P-Beta control law according to the present invention (dashed line) with the conventional yaw damping control law (solid line) at an airspeed of 180kn and a height of 2000 ft. In yawing maneuvers, the maximum magnitude of sideslip angle of the P-Beta control law of the present invention may be less than the maximum magnitude of sideslip angle of the yaw damping control law at the same flight maneuver input (e.g., foot pedal input), thereby effectively limiting aircraft droop loads (e.g., without exceeding the droop load limit for airworthiness requirements).
Specifically, when the foot pedal is not operated (pedal stroke is 0), the aircraft sideslip angle is 0. When the foot pedal is manipulated (e.g., reaches and remains at 20 of travel), the actual sideslip angle of the aircraft will gradually increase based on the output command of the sideslip command module 130. Since the P-Beta control law of the present invention may limit the slip angle command Beta based on the estimated vertical tail loadmTherefore, the maximum amplitude of the sideslip angle of the P-Beta control law of the invention can be smaller than that of the yaw damping control law. For example, as shown, the foot peg is operated at 20 ° for a period of 10-15 s, and the P-Beta control law of the present invention may have the maximum amplitude of the droop load less than the maximum amplitude of the droop load of the yaw damping control law. During the time period 20-25 s, the foot pedal is operated at-20 ° continuously, and the P-Beta control law of the present invention can reduce the vertical tail load maximum amplitude to 38.3kN (about 18.5% reduction) compared to the vertical tail load maximum amplitude of 47kN of the yaw damping control law. Analysis shows that VAThe speed droop tail load gain is more significant. Therefore, the flight control technology can more effectively provide the aircraft lateral heading P-Beta control law system meeting the load limitation requirement.
Figure 7 is a schematic diagram of a sideslip maneuver input time history, in accordance with one embodiment of the present invention. When not operating, the pedal stroke is 0 degree. When the Pedal continues to be positively forced for a period of time (e.g., Δ t), the Pedal may be positively displaced until a maximum positive stroke Pedal _ max is reached. After releasing the foot peg, the foot peg gradually returns to the original position. When force is applied to the pedal in the reverse direction, the pedal can generate reverse displacement until the maximum reverse stroke is reached. After releasing the foot peg, the foot peg gradually returns to the original position. The magnitude of the maximum forward stroke may be the same as or different from the magnitude of the maximum reverse stroke.
FIG. 8 is a flow chart of a flight control method for an aircraft according to one embodiment of the invention. The flight control method 800 may be implemented by a computer, a processor, a flight control system as described above with reference to FIG. 1, or the like.
At step 802, a sideslip maneuver input for an aircraft may be received. For example, sideslip maneuver inputs may be received from foot pedals or other control components (e.g., control panels).
At step 804, an upper sideslip angle limit may be determined based on the vertical tail loading of the aircraft. As described above, the current estimated vertical tail load may be calculated based on aerodynamic data of the aircraft and flight parameters including, for example, the aircraft's speed, side slip angle, yaw rate, and rudder deflection. For a fixed flight state, the aerodynamic data of the aircraft is fixed. Thus, a corresponding upper sideslip angle limit may be determined based on the current estimated vertical tail load of the aircraft. In another example, the vertical tail load of the aircraft may be a rated vertical tail load, such that the upper sideslip angle limit may be determined based on the rated vertical tail load of the aircraft, and the speed, yaw rate, and rudder deflection of the aircraft.
At step 806, a sideslip angle command not exceeding a sideslip angle upper limit may be generated based on the sideslip maneuver input. Thus, the maximum value of the sideslip angle command may be limited based on the vertical tail load. For example, a smaller vertical tail load may correspond to a greater upper side slip angle limit than a larger vertical tail load. In another example, the coefficient that converts the side-slip steering input to the side-slip angle command may be adjusted based on the current estimated vertical tail load such that the side-slip angle command does not exceed the side-slip angle upper limit. Therefore, the vertical tail load is not over-limited on the basis of meeting the airplane control capacity.
At step 808, sideslip angle feedback for the aircraft may be generated based on the sensor measurements. In one aspect, the sideslip angle measured by a sideslip angle sensor on the aircraft may be used as sideslip angle feedback. In another aspect, a rate of change of an inertial side-slip angle of the aircraft may be estimated based on the aircraft yaw rate, roll angle, and body side-loading measured by the sensors, and the rate of change of the inertial side-slip angle and a side-slip angle measured by a side-slip angle sensor on the aircraft may be complementarily filtered to generate the side-slip angle feedback.
At step 810, a yaw control command may be generated based on the sideslip angle command and the sideslip angle feedback to control a rudder of the aircraft. For example, the difference of the sideslip angle command and the sideslip angle feedback may be proportional-integral controlled to generate a yaw control command.
In another example, a sideslip angle change rate command may also be generated based on the sideslip maneuver input (e.g., at step 806), inertial sideslip angle change rate feedback for the aircraft may be generated based on the sensor measurement (e.g., at step 808), and a sideslip angle rate command may be generated as part of the yaw control command based on the sideslip angle change rate command and the inertial sideslip angle change rate feedback (e.g., at step 810).
At step 812, a roll maneuver input may be received. For example, sideslip maneuver inputs may be received from a sidebar, a steering wheel, or other control components (e.g., a control panel).
At step 814, a roll angle rate instruction may be generated based on the roll maneuver input.
At step 816, roll angle change rate feedback for the aircraft may be generated based on the sensor measurements.
At step 818, roll control commands may be generated based on the roll rate command and the roll rate of change feedback to control the ailerons and spoilers of the aircraft. For example, the roll rate command may be proportional-integral controlled from the difference in roll angle rate feedback detected by the sensor to generate a roll control command.
It should be noted that the operations shown in fig. 8 may be performed in a different order, and some operations may be performed in parallel. For example, steps 802 and 804 may be performed in a different order or in parallel, step 802 and 810 may be performed before or after step 812 and 818, or step 802 and 810 may be performed in parallel with step 812 and 818.
In the invention, a structure that the side lever is adopted to control the roll angle rate and the pedal is adopted to control the sideslip angle is adopted, and meanwhile, yaw maneuvering load limiting constraint is introduced into a control law structure, so that a control law-load integrated design is provided. The invention can estimate the change rate of the inertial sideslip angle of the airplane according to the angular rate, overload and attitude signals measured by inertial navigation, and carries out complementary filtering on the change rate and sideslip angle signals measured by a sideslip angle sensor, and the change rate and the sideslip angle signals are combined to be used as input signals of a P-Beta control law. In the aspect of control law sideslip signal acquisition, a sideslip angle sensor and inertial navigation data fusion algorithm is adopted, and the atmospheric disturbance resistance of the sideslip signal is improved. Compared with the prior art, the P-Beta control law framework not only enhances the course stability, improves the damping of the Dutch roll and the coordinated turning characteristic, but also meets the latest CS25.353 yaw maneuvering load limiting requirement of the EASA. According to the flight control method and the flight control device, the course static stability of the large airplane is enhanced, the characteristics of Dutch rolling and coordinated turning can be improved, and the riding comfort level is improved.
The various steps and modules of the methods and apparatus described above may be implemented in hardware, software, or a combination thereof. If implemented in hardware, the various illustrative steps, modules, and circuits described in connection with the disclosure may be implemented or performed with a general purpose processor, a Digital Signal Processor (DSP), an Application Specific Integrated Circuit (ASIC), a Field Programmable Gate Array (FPGA), or other programmable logic component, hardware component, or any combination thereof. A general purpose processor may be a processor, microprocessor, controller, microcontroller, or state machine, among others. If implemented in software, the various illustrative steps, modules, etc. described in connection with the disclosure may be stored on or transmitted over as one or more instructions or code on a computer-readable medium. A software module implementing various operations of the present disclosure may reside in a storage medium such as RAM, flash memory, ROM, EPROM, EEPROM, registers, hard disk, a removable disk, a CD-ROM, cloud storage, and the like. A storage medium may be coupled to the processor such that the processor can read information from, and write information to, the storage medium, and execute the corresponding program modules to perform the various steps of the present disclosure. Furthermore, software-based embodiments may be uploaded, downloaded, or accessed remotely through suitable communication means. Such suitable communication means include, for example, the internet, the world wide web, an intranet, software applications, cable (including fiber optic cable), magnetic communication, electromagnetic communication (including RF, microwave, and infrared communication), electronic communication, or other such communication means.
It is also noted that the embodiments may be described as a process which is depicted as a flowchart, a flow diagram, a structure diagram, or a block diagram. Although a flowchart may describe the operations as a sequential process, many of the operations can be performed in parallel or concurrently. In addition, the order of the operations may be rearranged.
The disclosed methods, apparatus, and systems should not be limited in any way. Rather, the present disclosure encompasses all novel and non-obvious features and aspects of the various disclosed embodiments, both individually and in various combinations and sub-combinations with each other. The disclosed methods, apparatus, and systems are not limited to any specific aspect or feature or combination thereof, nor do any of the disclosed embodiments require that any one or more specific advantages be present or that a particular or all technical problem be solved.
While the present invention has been described with reference to the embodiments shown in the drawings, the present invention is not limited to the embodiments, which are illustrative and not restrictive, and it will be apparent to those skilled in the art that various changes and modifications can be made therein without departing from the spirit and scope of the invention as defined in the appended claims.
Claims (19)
1. A flight control apparatus for an aircraft, comprising:
a vertical tail load processing module for determining an upper limit of a sideslip angle according to the vertical tail load of the airplane;
a sideslip command module that receives a sideslip manipulation input and an upper sideslip angle limit provided by the vertical tail load processing module, and generates a sideslip angle command that does not exceed the upper sideslip angle limit based on the sideslip manipulation input;
a feedback signal processing module that generates sideslip angle feedback for the aircraft based on sensor measurements; and
a controller that generates a yaw control command to control a rudder of the aircraft based on the sideslip angle command and the sideslip angle feedback.
2. The flight control apparatus of claim 1, wherein:
the side-slip command module generates a side-slip angle change rate command based on the side-slip manipulation input,
the feedback signal processing module generates inertial side-slip angle change rate feedback for the aircraft based on sensor measurements,
and the controller generates a side-slip angle rate command as part of the yaw control command based on the side-slip angle change rate command and the inertial side-slip angle change rate feedback.
3. The flight control apparatus of claim 1, further comprising:
a roll command module that receives a roll maneuver input and generates a roll angle rate command based on the roll maneuver input,
wherein the feedback signal processing module generates roll angle rate of change feedback for the aircraft based on sensor measurements, and the controller generates roll control commands based on the roll angle rate commands and the roll angle rate of change feedback to control ailerons and spoilers of the aircraft.
4. The flight control apparatus of claim 1, wherein the feedback signal processing module comprises:
an inertial sideslip angle change rate estimation module that determines an inertial sideslip angle change rate of the aircraft based on an aircraft yaw rate, a roll angle, and an airframe lateral overload measured by sensors; and
a complementary filter that complementarily filters the rate of change of inertial sideslip angle and the sideslip angle measured by a sideslip angle sensor on the aircraft to generate the sideslip angle feedback.
5. The flight control apparatus of claim 1, wherein the vertical tail load of the aircraft is a current vertical tail load estimated by the vertical tail load processing module based on a speed, a sideslip angle, a yaw rate, and a rudder deflection of the aircraft.
6. The flight control apparatus of claim 5, wherein the sideslip command module adjusts a coefficient that converts the sideslip maneuver input into the sideslip angle command based on the current vertical tail load.
7. The flight control apparatus of claim 5, wherein the upper limit for side slip angle for smaller vertical tail loads is greater than the upper limit for side slip angle for larger vertical tail loads.
8. The flight control apparatus of claim 1, wherein the vertical tail load of the aircraft is a rated vertical tail load, wherein the vertical tail load processing module determines the upper sideslip angle limit based on the rated vertical tail load of the aircraft, and a speed, a yaw rate, and a rudder deflection of the aircraft.
9. The flight control device of claim 1, wherein the controller comprises a linear quadratic regulation with integral (LQR) controller that performs proportional integral control of a difference of the sideslip angle command and the sideslip angle feedback to generate the yaw control command.
10. A flight control method for an aircraft, comprising:
receiving a sideslip maneuver input of the aircraft;
determining a sideslip angle upper limit based on a vertical tail load of the aircraft;
generating a sideslip angle command not exceeding the upper sideslip angle limit based on the sideslip manipulation input;
generating sideslip angle feedback for the aircraft based on sensor measurements; and
generating a yaw control command based on the sideslip angle command and the sideslip angle feedback to control a rudder of the aircraft.
11. The flight control method of claim 10, further comprising:
generating a side-slip angle change rate command based on the side-slip manipulation input;
generating inertial sideslip angle change rate feedback for the aircraft based on sensor measurements; and
generating a sideslip angle rate command as part of the yaw control command based on the sideslip angle change rate command and the inertial sideslip angle change rate feedback.
12. The flight control method of claim 10, further comprising:
receiving a roll maneuver input and generating a roll angle rate instruction based on the roll maneuver input;
generating roll angle change rate feedback for the aircraft based on sensor measurements; and
generating roll control commands based on the roll rate command and the roll rate of change feedback to control ailerons and spoilers of the aircraft.
13. The flight control method of claim 10, further comprising:
determining a rate of inertial sideslip angle change of the aircraft based on an aircraft yaw rate, a roll angle, and an aircraft body lateral overload measured by sensors; and
complementarily filtering the inertial sideslip angle rate of change and a sideslip angle measured by a sideslip angle sensor on the aircraft to generate the sideslip angle feedback.
14. The flight control method of claim 10, wherein the vertical tail load of the aircraft is a current vertical tail load estimated based on a speed, a side slip angle, a yaw rate, and a rudder deflection of the aircraft.
15. The flight control method of claim 14, further comprising adjusting a coefficient that converts the sideslip maneuver input to the sideslip angle command based on the current vertical tail load.
16. The flight control method of claim 14, wherein the upper limit for side slip angle for smaller vertical tail loads is greater than the upper limit for side slip angle for larger vertical tail loads.
17. The flight control method of claim 10, wherein the vertical tail load of the aircraft is a rated vertical tail load, and wherein the upper sideslip angle limit is determined based at least in part on the rated vertical tail load of the aircraft, and a speed, a yaw rate, and a rudder deflection of the aircraft.
18. The flight control method of claim 10, further comprising proportional-integral controlling a difference of the sideslip angle command and the sideslip angle feedback to generate the yaw control command.
19. A flight control system for an aircraft, comprising:
a processor; and
a memory for storing processor-executable instructions,
wherein the processor is configured to execute the processor-executable instructions to implement the flight control method of any one of claims 10-18.
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| CN117784833B (en) * | 2024-02-23 | 2024-06-11 | 四川腾盾科技有限公司 | System, method, equipment and medium for controlling speed of plane-symmetric aircraft |
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