CN110647160A - Flight control method and device for aircraft - Google Patents
Flight control method and device for aircraft Download PDFInfo
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- CN110647160A CN110647160A CN201910958089.5A CN201910958089A CN110647160A CN 110647160 A CN110647160 A CN 110647160A CN 201910958089 A CN201910958089 A CN 201910958089A CN 110647160 A CN110647160 A CN 110647160A
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/10—Simultaneous control of position or course in three dimensions
- G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
Abstract
A flight control method and apparatus for an aircraft is disclosed herein. A flight control device may include: an input component that generates flight maneuver instructions based on the received input; a multi-frequency trap that receives a feedback signal based on the sensor measurements, wherein the multi-frequency trap has a plurality of notch frequencies and traps the feedback signal at the plurality of notch frequencies to generate a trapped feedback signal; and a control law module that generates flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the notched feedback signals.
Description
Technical Field
The present invention relates generally to aircraft and, more particularly, to flight control methods and apparatus for aircraft.
Background
The Flight Control System ("FCS") of a modern aircraft is the core of the entire aircraft onboard System, and is one of the most complex systems of the entire aircraft. 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. Airplanes using fly-by-wire technology have a lower weight than airplanes using conventional mechanical control systems because fly-by-wire eliminates a large amount of complex, redundant mechanical equipment in the mechanical transmission control system, reduces the total weight of system components, and further frees up the interior space of the airplane. In addition, the introduction of fly-by-wire flight control means that the aircraft can adopt a static and unstable design for breaking the traditional aerodynamic layout, so that parts of the aircraft structure for controlling the flight stability can be omitted or the specific gravity can be reduced, for example, the horizontal stable surface and the vertical stable surface of the tail of the aircraft are reduced, and the range and the load of the reduced aircraft are improved.
Since the birth of an airplane, aeroelasticity problems such as divergence, flutter, gust response and the like become important factors influencing the stability and flight performance of the airplane. The wide application of the composite material in the design of the airplane enables the flexibility of the airplane to be larger, and the problem of rigid body motion and structure elastic coupling is outstanding, so that the telex control law design is particularly needed to be concerned, and the problem of pneumatic servo elasticity is solved.
One solution in the prior art is to use a single frequency trap to filter the signal (e.g., the rudder command) output by the teletype control law. The single frequency trap solves the aeroelastic servo problem by limiting the amplitude of the rudder instruction of a specific frequency. However, the single-frequency wave trap can only adapt to the worst signal, but cannot meet the requirement of stability margin of each signal loop, and is difficult to meet the requirement of pneumatic servo elasticity in a full frequency domain.
Accordingly, there is a need in the art for improved flight control methods and apparatus for aircraft.
Disclosure of Invention
The invention provides a flight control method and device for an aircraft. A flight control device may include: an input component that generates flight maneuver instructions based on the received input; a multi-frequency trap that receives a feedback signal based on the sensor measurements, wherein the multi-frequency trap has a plurality of (i.e., different) notch frequencies and traps the feedback signal at the plurality of notch frequencies to generate a trapped feedback signal; and a control law module that generates flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the notched feedback signals.
In one embodiment, a flight control device for an aircraft is provided, comprising: an input component that generates flight maneuver instructions based on the received input; a first multi-frequency trap that receives a first feedback signal based on a first sensor measurement, wherein the first multi-frequency trap has a plurality of notch frequencies and traps the first feedback signal at the plurality of notch frequencies to generate a trapped first feedback signal; and a control law module that generates flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the notched first feedback signal.
In one aspect, the flight control apparatus further comprises: a second multi-frequency trap that receives a second feedback signal based on second sensor measurements, wherein the second multi-frequency trap has a plurality of notch frequencies and traps the second feedback signal at the plurality of notch frequencies of the second multi-frequency trap to generate a trapped second feedback signal, wherein the control law module generates flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions, the trapped first feedback signal, and the trapped second feedback signal.
In one aspect, the control law modules comprise longitudinal control law modules, and wherein: the first feedback signal comprises an inertial angle of attack rate of the aircraft; the second feedback signal comprises a steady axis normal overload of the aircraft; and the flight commands include pitch commands for controlling an elevator of the aircraft.
In one aspect, the flight control apparatus further comprises: a parameter synthesis module that receives a notched first feedback signal from the first multi-frequency trap and a notched second feedback signal from the second multi-frequency trap and generates a synthetic feedback signal based on the notched first feedback signal and the notched second feedback signal, wherein the control law module generates flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the synthetic feedback signal.
In one aspect, the control law modules comprise lateral control law modules, and wherein: the first feedback signal comprises a body roll rate of the aircraft; the second feedback signal comprises a body axis yaw rate of the aircraft; the composite feedback signal includes a steady-axis roll rate of the aircraft; and the flight commands include roll commands for controlling ailerons and spoilers of the aircraft.
In one aspect, the flight control apparatus further comprises: a third multi-frequency trap that receives a third feedback signal based on a third sensor measurement, wherein the third multi-frequency trap has a plurality of notch frequencies and notches the third feedback signal at the plurality of notch frequencies of the third multi-frequency trap to generate a notched third feedback signal, the parameter synthesis module generates the synthetic feedback signal based on the notched first feedback signal from the first multi-frequency trap, the notched second feedback signal from the second multi-frequency trap, and the notched third feedback signal from the third multi-frequency trap, wherein the control law module generates flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the synthetic feedback signal.
In one aspect, the control law module comprises a heading control law module, and wherein: the first feedback signal comprises a body roll rate of the aircraft; the second feedback signal comprises a body axis yaw rate of the aircraft; the third feedback signal comprises a lateral overload of a body axis of the aircraft; the composite feedback signal includes a rate of change of a side-slip angle of inertia of the aircraft; and the flight command comprises a yaw command for controlling a rudder of the aircraft.
In one aspect, the desired frequency response of each multi-frequency trap is based on the aircraft's most severe structural modal frequency response under different conditions and stability margin requirements.
In an aspect, each multi-frequency trap comprises a plurality of single-frequency traps, wherein each single-frequency trap has a notch center frequency and a damping parameter, wherein the damping parameter is selected to meet the stability margin requirement over a range of frequencies that includes the notch center frequency.
In an aspect, the damping parameters are selected to minimize a phase lag of the single frequency trap.
In an aspect, the input component includes one or more of: side lever, pedal, handle, control panel.
In another embodiment, a flight control method for an aircraft is provided, comprising: generating flight maneuver instructions based on the received input; receiving a first feedback signal based on the first sensor measurement; notching the first feedback signal at a first plurality of notch frequencies to generate a notched first feedback signal; and generating flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the notched first feedback signal.
In one aspect, the flight control method further comprises: receiving a second feedback signal based on a second sensor measurement; notching the second feedback signal at a second plurality of notch frequencies to generate a notched second feedback signal; and generating flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions, the notched first feedback signal, and the notched second feedback signal.
In one aspect, the first feedback signal comprises an inertial angle of attack rate of the aircraft; the second feedback signal comprises a steady axis normal overload of the aircraft; and the flight commands include pitch commands for controlling an elevator of the aircraft.
In one aspect, the flight control method further comprises: generating a composite feedback signal based on the notched first feedback signal and the notched second feedback signal, and generating flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the composite feedback signal.
In one aspect, the first feedback signal comprises a body roll rate of the aircraft; the second feedback signal comprises a body axis yaw rate of the aircraft; the composite feedback signal includes a steady-axis roll rate of the aircraft; and the flight commands include roll commands for controlling ailerons and spoilers of the aircraft.
In one aspect, the flight control method further comprises: receiving a third feedback signal based on a third sensor measurement; notching the third feedback signal at a third plurality of notch frequencies to generate a notched third feedback signal; generating the composite feedback signal based on the notched first feedback signal, the notched second feedback signal, and the notched third feedback signal; and generating flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the synthetic feedback signals.
In one aspect, the first feedback signal comprises a body roll rate of the aircraft; the second feedback signal comprises a body axis yaw rate of the aircraft; the third feedback signal comprises a lateral overload of a body axis of the aircraft; the composite feedback signal includes a rate of change of a side-slip angle of inertia of the aircraft; and the flight command comprises a yaw command for controlling a rudder of the aircraft.
In one aspect, each feedback signal is separately notched using a multi-frequency trap, the desired frequency response of each multi-frequency trap being based on the most severe structural modal frequency response and stability margin requirements of the aircraft under different conditions.
In an aspect, each multi-frequency trap comprises a plurality of single-frequency traps, wherein each single-frequency trap has a notch center frequency and a damping parameter, wherein the damping parameter is selected to meet the stability margin requirement over a range of frequencies that includes the notch center frequency.
In an aspect, the damping parameters are selected to minimize a phase lag of the single frequency trap.
In an aspect, the input is received through one or more of: side lever, pedal, handle, control panel.
According to the invention, the multi-frequency filter is arranged in front of the control law module so as to trap the feedback signal measured by the sensor, so that the high-frequency vibration in the feedback signal can be prevented from entering the control law, the problem of high-frequency stability is solved, the requirements of stability margins of a plurality of signal loops can be met, and the influence on the low-frequency stability margin is reduced. Further optionally, by optimizing the damping parameters of each wave trap, the multi-frequency filter can meet the high-frequency pneumatic servo stability requirement and minimize the influence on the low-frequency stability margin.
Drawings
Fig. 1 is a block diagram of a flight control device for an aircraft according to an embodiment of the invention.
Fig. 2 is a block diagram of a flight control device for an aircraft according to another embodiment of the invention.
FIG. 3 is a flow chart of a flight control method for an aircraft according to one embodiment of the invention.
FIG. 4 is a schematic illustration of a longitudinal control portion of a flight control device according to one embodiment of the present invention.
FIG. 5 is a schematic view of a lateral control portion of a flight control apparatus according to one embodiment of the present invention.
FIG. 6 is a schematic view of a heading control portion of a flight control device according to one embodiment of the invention.
Figure 7 is a schematic diagram of a multi-frequency trap in accordance with one embodiment of the present invention.
Figure 8 is a schematic diagram of the frequency response of a multiple frequency trap according to one embodiment of the present invention.
Figure 9 is a flow diagram of a method for optimizing trap parameters according to one embodiment of the present 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.
Fly-by-wire flight control system (Fly by wire flight control system) is an advanced electronic flight control system, replaces the traditional mechanical control system, adopts the feedback control principle, converts the mechanical instruction generated when a pilot operates the airplane into an electric signal through a special sensor, and drives the control surface of the airplane to deflect through a control law model by a flight control computer receiving the electric signal, thereby controlling the flight state of the airplane. The introduction of fly-by-wire flight control breaks through the static and unstable design of the traditional pneumatic layout, improves the operability of the airplane and improves the driving precision and safety.
The flight control device according to the invention may comprise: an input component that generates flight maneuver instructions based on the received input; a multi-frequency trap that receives a feedback signal based on the sensor measurements, wherein the multi-frequency trap has a plurality of notch frequencies and traps the feedback signal at the plurality of notch frequencies to generate a trapped feedback signal; and a control law module that generates flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the notched feedback signals. The invention is applicable to various types of aircrafts, such as civil aircraft, military aircraft, unmanned aerial vehicles and the like.
Fig. 1 is a block diagram of a flight control device 100 for an aircraft according to an embodiment of the invention. Flight control device 100 can include an input assembly 110, a control law module 120, an actuator 130, a first sensor 150, a first parameter processing module 152, a first multi-frequency trap 154, and the like. The first sensor 150, the first parameter processing module 152, and the first multi-frequency trap 154 may form a feedback channel. Optionally, flight control apparatus 100 may also include another feedback channel (e.g., including second sensor 160, second parameter processing module 162, second multi-frequency trap 164, etc.).
The input component 110 may be an input device for maneuvering the aircraft, and may generate flight maneuver instructions based on the received input. For example, in a civilian aircraft, the input assembly 110 may include a sidebar or stick pan, a foot peg, a speed brake handle, a trim control panel, a slat handle, a flight mode control panel, and the like for operator manipulation and generation of flight manipulation instructions. In a drone, the input component 110 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 law module 120 may generate flight commands based on flight maneuver instructions, aerodynamic parameters, measurement feedback parameters, and the like. 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 tunnel acquisition or test flight. 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 control law module 120 may be implemented by a computer or other electronic device for converting flight maneuver instructions generated by the input assembly 110 into flight commands for controlling the actuators 130 according to various algorithms. The control laws module 120 can implement a longitudinal control law, a lateral control law, and a heading control law. The control law module 120 enables the pilot to fine-tune the flight attitude of the aircraft, simplifies the control process and improves the pilot accuracy. To further improve flight safety, multiple sets of control law systems can be used to execute the same command, which is called a redundant fly-by-wire system. When a certain part of the redundancy system fails, the flight control system can automatically isolate the failed part and use the rest of the intact systems to continue to execute corresponding instructions. For example, the four-redundancy fly-by-wire system can adopt four sets of independent channels and can prevent double-channel faults, thereby greatly improving the safety performance of the airplane.
The actuators 130 may include various drive components for controlling aircraft motion, such as slats, flaps, spoilers, rudders, elevators, horizontal stabilizers, spoilers, flaps, slats, and the like. The actuators 130 vary the forces and moments experienced by the aircraft based on the flight commands provided by the control law module 120 to achieve axial control (e.g., pitch, roll, yaw, etc.) and trim, lift and drag control, auto-flight, etc. functions for the aircraft.
The first parameter processing module 152 and the second parameter processing module 162 are optional and may each process a respective measured parameter to generate a feedback signal. The processing performed by the first parameter processing module 152 and the second parameter processing module 162 may include, for example, but is not limited to, filtering, modification, coordinate transformation, parameter combination, parameter calculation, and the like. For example, for the longitudinal control law, the first parameter processing module 152 may calculate the inertial rate of change of angle of attack based on the aircraft angle of attack and with reference to the parameters of airspeed, normal overload, pitch angle, and roll angle, and provide the inertial rate of change of angle of attack as a first feedback signal to the first multi-frequency trap 154. The second parameter processing module 162 may generate a stable normal overload based on the body axis normal overload of the aircraft and with reference to the pitch angle rate, the aircraft longitudinal overload, the angle of attack, etc. as a second feedback signal to be provided to the second multi-frequency trap 162.
First multi-frequency trap 154 and second multi-frequency trap 164 may receive feedback signals from first parameter processing module 152 and second parameter processing module 162, respectively. Each multi-frequency trap may have a plurality of notch frequencies and notch (i.e., attenuate/clip) the feedback signal at these notch frequencies to generate a notched feedback signal that is provided to control law module 120. The notch frequency of the first multi-frequency trap 154 and the notch frequency of the second multi-frequency trap 164 may be the same or different or partially the same, and the notch frequency and corresponding notch depth may be determined based on the desired frequency response characteristics of the corresponding feedback channel, respectively. Thus, control law module 120 may generate flight maneuver instructions to control actuators 130 based on the flight maneuver instructions generated by input assembly 110 and the trapped feedback signals provided by one or more feedback channels (e.g., by first multi-frequency trap 154 and second multi-frequency trap 164). The specific algorithm employed by the control law module 120 may be designed or adjusted as needed or empirically by one skilled in the art.
In alternative embodiments, first parameter processing module 152 and/or second parameter processing module 162 may not be included. Thus, measurement parameters generated by the sensors (e.g., first sensor 150 and/or second sensor 160) may be provided directly as feedback signals to the multi-frequency trap. Alternatively, the first parameter processing module 152 and/or the second parameter processing module 162 may not change the measured parameter generated by the sensor, thereby making the feedback signal the same as the measured parameter.
By arranging the multi-frequency filter in front of the control law module 120 to trap the feedback signal, the high-frequency vibration in the feedback signal can be prevented from entering the control law, the stability margin requirements of a plurality of signal loops can be met, the accuracy of the control law module 120 is improved, the problem of high-frequency stability is solved, and the airplane is prevented from fluttering. The multi-frequency filter can solve the problem of pneumatic servo elasticity by trapping the feedback signals at a plurality of frequencies, and the problem that the traditional single-frequency trap filter can only adapt to the worst signal and can not meet the stability margin of each signal loop is solved. The effect on the low frequency stability margin is reduced from the control law architecture by placing a multi-frequency filter before the control law module 120 to notch the feedback signal provided to the control law module 120. Therefore, the flight control method and the flight control device for the aircraft provided by the invention can enable the multi-frequency filter to meet the requirement of high-frequency pneumatic servo stability and reduce the influence on low-frequency stability margin.
Fig. 2 is a block diagram of a flight control device 200 for an aircraft according to another embodiment of the invention.
Flight control device 200 can include an input assembly 210, a control law module 220, an actuator 230, a first sensor 250, a first multi-frequency trap 254, a second sensor 260, a second multi-frequency trap 264, and the like. The components described above have similar functions to the similar components in fig. 1 and are not described in detail. For example, the first sensor 250 and the second sensor 260 may each be used to sense various measured parameters of the aircraft. First multi-frequency trap 254 may receive the first feedback signal from first sensor 250 and trap (i.e., attenuate/clip) the first feedback signal at a plurality of trap frequencies to generate a trapped first feedback signal. The second multi-frequency trap 264 may receive the second feedback signal from the second sensor 260 and trap the second feedback signal at a plurality of trap frequencies to generate a trapped second feedback signal. Although not shown, optional parameter processing modules may also be included between first sensor 250 and first multi-frequency trap 254 and/or between second sensor 260 and second multi-frequency trap 264 to perform appropriate processing on the sensor measurements, such as filtering, modification, coordinate transformation, parameter combination, parameter calculation, and the like.
As shown in fig. 2, flight control apparatus 200 also includes a parameter synthesis module 270 that receives a notched first feedback signal from first multi-frequency trap 254 and a notched second feedback signal from second multi-frequency trap 264, and generates and provides a synthetic feedback signal to control law module 220 based on the notched first feedback signal and the notched second feedback signal. Control law module 220 may generate flight commands for controlling the flight motions of the aircraft via actuators 230 based on the flight maneuver instructions provided by input assembly 210 and the synthetic feedback signals provided by parameter synthesis module 270.
Although not shown, flight control device 100 may include more feedback channels. Each feedback channel may provide a different feedback signal based on the sensor measurements. These feedback signals may be provided to control law module 220 for use in generating flight commands and/or to parameter synthesis module 270 for use with one or more other feedback signals to generate a synthetic feedback signal, which may then be provided to control law module 220 for use in generating flight commands.
FIG. 3 is a flow chart of a flight control method for an aircraft according to one embodiment of the invention.
At step 302, flight maneuver instructions may be generated based on the received input. For example, the driver may operate input components such as side bars, foot pedals, handles, control panels, and the like. The input assembly may convert the corresponding input into flight maneuver instructions.
At step 312, a first feedback signal based on the first sensor measurement may be received. For example, sensors distributed throughout the aircraft may measure a first measured parameter relating to the aircraft, the measurement of which may be converted into or as a first feedback signal.
At step 314, the first feedback signal may be notched at a first plurality of notch frequencies to generate a notched first feedback signal. For example, step 314 may be implemented using a first multi-frequency trap. The desired frequency response of the first multi-frequency trap may be based on the most severe structural modal frequency response and stability margin requirements of the aircraft under different conditions. The first multi-frequency trap comprises a plurality of single-frequency traps, wherein each single-frequency trap has a trap center frequency and a damping parameter, wherein the damping parameter is selected to meet a stability margin requirement over a frequency range encompassing the trap center frequency. In a further example, the damping parameters may be selected to minimize a phase lag of the single frequency trap.
In one embodiment, at step 330, a flight command may be generated for controlling a flight action of the aircraft based on the flight maneuver instruction and the notched first feedback signal.
In another embodiment, the method may further include optional steps 322 and 324. At step 322, a second feedback signal based on the second sensor measurement may be received. The second feedback signal may be different from the first feedback signal.
At step 324, the second feedback signal may be notched at a second plurality of notch frequencies to generate a notched second feedback signal. For example, step 324 may be implemented using a second multi-frequency trap. The first plurality of notch frequencies and the second plurality of notch frequencies may be the same or different, or partially the same. Thus, at step 330, a flight command may be generated for controlling a flight action of the aircraft based on the flight maneuver instruction, the notched first feedback signal, and the notched second feedback signal.
In another embodiment, a composite feedback signal may be generated based on the notched first feedback signal and the notched second feedback signal (and one or more other feedback signals) prior to step 330. Then at step 330, flight commands may be generated for controlling the flight actions of the aircraft based on the flight maneuver instructions and the synthetic feedback signals.
FIG. 4 is a schematic illustration of the longitudinal control law portion of a flight control device according to one embodiment of the present invention. The longitudinal control law of the aircraft can be implemented based on the architecture shown in fig. 1 and generates elevator commands based on the aircraft sidestick maneuvers. In the longitudinal control law part, a Multi-frequency trap Multi-Filter1 and a Multi-Filter2 can be respectively arranged on two feedback channels of the incidence angle change rate and the stable axial normal overload, as shown in fig. 4.
The angle of attack sensor 410 may measure the aircraft angle of attack αvaneAnd inertial angle of attack rate estimation module 412 may be based at least in part on aircraft angle of attack αvaneTo calculate the inertial angle of attack rate
Although not shown, inertial angle of attack rate estimation module 412 may also use additional parameters in calculating the inertial angle of attack rate, such as the aircraft corrected airspeed, the normal overload in the aircraft's airflow coordinate system, the pitch angle of the aircraft, the roll angle of the aircraft, etc., which may be measured or calculated with corresponding sensors, and the manner in which the inertial angle of attack rate is calculated is known in the art. The first Multi-frequency trap (Multi-Filter1)414 may be responsive to a rate of change of inertial angle of attack at a plurality of trap frequencies
The notch is performed to generate and provide a notched inertial angle of attack rate 416 to the longitudinal control law model 430. Longitudinal control law model 430 may be implemented as part of control law modules 120 or 220 shown in fig. 1 or 2.
The inertial sensor 420 may measure the aircraft body axial normal overload Nzb_gAnd the coordinate transformation module 422 may overload the normal of the aircraft body coordinate system by Nzb_gNormal overload N converted into airflow coordinatezb_s(i.e., steady axis normal overload). Although not shown, the coordinate transformation module 422 may also use additional parameters in performing the coordinate transformation, such as aircraft pitch rate, longitudinal overload of the aircraft body axis, angle of attack of the aircraft, etc., which may be measured with corresponding sensors, and the transformation between the aircraft body coordinate system and the air flow coordinate system is known in the art. The second Multi-frequency trap (Multi-Filter2)424 may be overloaded N normal to the stable axis at multiple trap frequencieszb_sThe notch is performed to generate a notched stable axial normal overload 426 and provided to the longitudinal control law model 430.
The airplane sidestick 440 may be manipulated by the pilot to generate sidestick manipulation inputs. The sidebar steering inputs may include steering inputs for pitch and/or roll. The sidebar input processing component 442 may generate a sidebar manipulation instruction (e.g., a pitch instruction 444 and/or a roll instruction) based on the sidebar manipulation input. Pitch instructions 444 may be provided to pitch control law model 430 and summer 450. Thus, longitudinal control law model 430 generates pitch correction instructions 432 based on at least longitudinal pitch instructions 444, notched inertial angle of attack rate 416, and notched steady-axis normal overload 426. Summer 450 superimposes pitch command 444 and pitch correction command 432 to generate pitch command 460(δ)ecnd) For controlling the pitching action of the aircraft elevator. Although a separate vertical control law model 430 and summer 450 are shown in fig. 4, the vertical control law model 430 and summer 450 may also be implemented in combination (e.g., as vertical control law modules). Similarly, while a separate airplane sidebar 440 and sidebar-input processing assembly 442 are shown in FIG. 4, the airplane sidebar 440 and the sidebar-input processing assembly 442 may also be implemented in combination (e.g., as an input assembly).
FIG. 5 is a schematic view of a lateral control portion of a flight control apparatus according to one embodiment of the present invention. The lateral control law for an aircraft may be implemented based on the architecture shown in fig. 2, and roll commands generated based on sidestick steering for controlling the ailerons and spoilers of the aircraft.
As described above, the sidestick manipulation input will be generated when the pilot manipulates the sidestick, and the sidestick input processing component 510 may generate a sidestick manipulation command (e.g., a pitch command and/or a roll command) based on the sidestick manipulation input. The sidebar input processing component 510 may provide the lateral roll command to the lateral control law model 530. Lateral control law model 530 may be implemented as part of control law modules 120 or 220 shown in FIG. 1 or FIG. 2.
Inertial sensors of the aircraft may provide a body roll rate. The first multi-frequency trap 512 may trap the body axis roll rate at a plurality of trap frequencies and provide the trapped body axis roll rate to the steady axis roll rate calculation module 520.
In addition, inertial sensors of the aircraft may also provide a body axis yaw rate. The second multi-frequency trap 514 may trap the body axis yaw rate at a plurality of trap frequencies and provide the trapped body axis yaw rate to the steady axis roll rate calculation module 520.
The steady-axis roll rate calculation module 520 may be implemented as the parameter synthesis module 270 as described with reference to FIG. 2 and calculate the steady-axis roll rate based on the notched body-axis roll rate and the notched body-axis yaw rate. The steady-axis roll rate calculation module 520 may also utilize other auxiliary parameters in calculating the steady-axis roll rate, such as angle of attack measurements provided by an angle of attack sensor, and the like. The steady-axis roll rate calculation module 520 may provide the steady-axis roll rate to the lateral control law model 530.
The lateral control law model 530 may generate roll commands for controlling the ailerons and spoilers of the aircraft based on lateral roll instructions entered by the sidebar and the steady-axis roll rate provided by the steady-axis roll rate calculation module 520.
FIG. 6 is a schematic view of a heading control portion of a flight control device according to one embodiment of the invention. The aircraft's yaw control law may be implemented based on the architecture shown in fig. 2 and generates yaw commands for controlling the rudder of the aircraft based on foot-pedal maneuvers.
The foot pedal manipulation input will be generated when the pilot manipulates the foot pedal, and the foot pedal input processing component 610 may generate a yaw command based on the foot pedal manipulation input and may provide the yaw command to the heading control law model 630. Heading control law model 630 may be implemented as part of control law module 120 or 220 shown in FIG. 1 or FIG. 2.
As described above with reference to fig. 5, inertial sensors of the aircraft may provide a body roll rate. The first multi-frequency trap 612 may trap the body roll rate at a plurality of trap frequencies and provide the trapped body roll rate to the side-slip angle rate of inertia estimation module 620. In addition, inertial sensors of the aircraft may also provide a body axis yaw rate. The second multi-frequency trap 614 may trap the body axis yaw rate at a plurality of trap frequencies and provide the trapped body axis yaw rate to the inertial side-slip angle change rate estimation module 620.
In addition, inertial sensors of the aircraft may also provide lateral overloading of the body axis. The third multi-frequency trap 616 may trap the body axis lateral overload at a plurality of trap frequencies and provide the trapped body axis lateral overload to the inertial side-slip angle change rate estimation module 620.
The inertial side-slip angle rate estimation module 620 may be implemented as the parameter synthesis module 270 as described with reference to FIG. 2 and estimates the inertial side-slip angle rate based on the notched body axis roll angle rate, the notched body axis yaw angle rate, and the notched body axis lateral overload. The inertial side-slip angle change rate estimation module 620 may also utilize other sensor measurements in determining the inertial side-slip angle change rate, such as roll and pitch angles provided by inertial sensors, angle of attack provided by an angle of attack sensor, airspeed provided by a pitot tube sensor, and so forth. The sideslip angle change rate estimation module 620 may provide the sideslip angle change rate to the heading control law model 630.
The yaw control law model 630 may generate yaw commands for controlling the rudder of the aircraft based on the yaw commands input by the foot pedals and the rate of change of the angle of sideslip provided by the angle of inertia change rate estimation module 620.
In some embodiments, lateral and heading laws may be implemented in combination. For example, steady-axis roll rate calculation module 520 and inertial slip angle rate estimation module 620 may be implemented in combination and provide a steady-axis roll rate and an inertial slip angle rate. Lateral control law model 530 and heading control law model 630 may be implemented in combination and provide commands for controlling the ailerons, spoilers, and rudder of the aircraft, respectively, based on the control signals input by the sidesticks and foot pedals.
Figure 7 is a schematic diagram of a multi-frequency trap 700 according to one embodiment of the present invention. The multi-frequency trap 700 may include a plurality of single frequency traps 710, 720, 730, etc., each of which may have a different notch frequency and corresponding notch amplitude. Thus, multi-frequency trap 700 may achieve trapping at a plurality of different trapping frequencies. The number of single frequency traps can be determined according to the number of frequency points of the desired trap. By way of example and not limitation, figure 7 illustrates a single frequency trap as a second order trap. In particular implementations, however, higher order traps may be employed.
In fig. 7, k denotes the number of single frequency traps, s denotes the input signal, ωnkIs the trapped wave center frequency of the kth single frequency trap, damping parameter xi1、ξ2Affecting notch bandwidth and depth (e.g., maximum amplitude attenuation). By way of example, and not limitation, the ratio ξ1/ξ2The magnitude of the notch can be determined, with the smaller the ratio the deeper the notch and the more severe the corresponding phase lag. Xi1、ξ2The larger the absolute value, the wider the notch width and the more severe the corresponding low band phase lag. Number k and frequency ω of single frequency trapsnkMay be determined based on structural modal characteristics of the aircraft body, as described with reference to fig. 5.
Figure 8 is a schematic diagram of the frequency response of a multiple frequency trap according to one embodiment of the present invention. According to the elastic aircraft dynamic model, the control law model, the sensor model and the actuator model, different configurations, weights, flight speeds and the like are considered, and the frequency response characteristics gamma (j omega) of the most severe structural mode can be determined. According to the stability margin requirement (e.g., ASE stability margin 6dB), by first making Γ (j ω) symmetric about 0dB and then shifting the margin downward (e.g., 6dB), the desired frequency response of the multi-frequency trap Φ (j ω) can be obtained. Accordingly, a notch frequency and a corresponding notch amplitude to be clipped may be determined, wherein each notch frequency may be implemented as a single frequency trap, and a plurality of single frequency traps may be combined to form a multi-frequency trap.
According to one embodiment, for an flight control system having a plurality of multi-frequency traps, the transfer function of each multi-frequency trap may be determined in turn. For example, in fig. 1, the transfer function of the first multi-frequency trap 154 may be first determined, when the first multi-frequency trap 154 is switched into the first feedback path and the second multi-frequency trap 164 is not switched into the second feedback path, i.e., the feedback signal of the second parameter processing module 162 is directly provided to the control law module 120. If there are other feedback channels, the other feedback channels similarly do not have multi-frequency traps coupled thereto.
The closed loop is opened at the input of the first multi-frequency trap 154, the transfer function of the first multi-frequency trap 154 is calculated and the frequency response Γ (j ω) of the most severe structural mode is determined, the desired multi-frequency trap frequency response Φ (j ω) is generated and the parameters of the first multi-frequency trap 154 are solved.
The designed first multi-frequency trap 154 is then coupled into a first feedback path and the second multi-frequency trap 164 is coupled into a second feedback path to calculate the parameters of the second multi-frequency trap 164. The closed loop is broken at the input of the second multi-frequency trap 164 and the transfer function of the second multi-frequency trap 164 is calculated and the parameters of the second multi-frequency trap 164 are solved in a similar manner.
The flight control apparatus described herein with reference to figures 2 and 4-6 may similarly solve for the parameters of each multi-frequency trap.
Figure 9 is a flow diagram of a method for optimizing trap parameters according to one embodiment of the present invention. By way of example and not limitation, figure 9 illustrates how the parameters of a second order single frequency trap therein may be optimized with reference to the multiple frequency trap of figure 7. In other implementations, the single frequency trap may have different forms and the damping parameters may be determined in a similar process. As described above, the center frequency ω of the single frequency trapnCan be determined according to the structural modal characteristics of the airplane body.
In step 902, a notch frequency range (ω) is determined1,ω2). For example, the frequency range may be a notch center frequency ωnIn a specified frequency range of the vicinity, e.g. 2Hz, 3Hz, 4Hz, etc. The size of the frequency range may be selected as desired and is not limited to the above examples.
In step 904, a damping parameter ξ is initialized1、ξ2. As will be appreciated by those skilled in the art, the damping parameter ξ1、ξ2May be determined empirically or experimentally, or may be a value in a specified set containing candidate values.
In step 906, it is determined that the frequency response of the designed single frequency trap is in the notch frequency range (ω)1,ω2) Whether the desired notch amplitude is met.
For example, it may be in the notch frequency range (ω)1,ω2) And determining whether the cost function I ═ G (j ω) | - | Φ (j ω) is less than a threshold value, wherein Φ (j ω) is the desired trap frequency response, and G (j ω) represents the designed trap frequency response. If the cost function is in the notch frequency range (ω)1,ω2) Above the threshold, it can indicate that the frequency response of the designed single-frequency trap is in the trap frequency range (ω)1,ω2) To meet the desired notch amplitude. By way of example, and not limitation, the threshold may be 0. In other examples, the threshold may be selected as desired, and in some cases, the error may be accounted for.
If the frequency response of the single frequency trap is in the notch frequency range (omega)1,ω2) Above the desired notch amplitude, ξ are saved in step 9081、ξ2The current value of (d) is taken as the effective damping parameter and may proceed to step 912. Conversely, if the frequency response of the single frequency trap is in the notch frequency range (ω)1,ω2) If it does not meet the desired notch amplitude, then ξ is discarded1、ξ2And may likewise proceed to step 912.
In step 912, it may be determined whether ξ is present1、ξ2The next set of candidate values. By way of example, and not limitation, ξ1、ξ2The candidate values of (a) may be empirically or experimentally determined values in a given set, or may be values in a given range of values adjusted by a given step size.
If xi is present1、ξ2Then ξ are processed in step 9101、ξ2Adjust to the next set of candidate values and proceed to step 906 to determine that the next set of candidate values is in the notch frequency range (ω)1,ω2) Whether the desired notch amplitude is met.
If xi is not present1、ξ2Then each set of valid ξ is determined in step 9141、ξ2Causing a phase lag. For example, each set of effective ξ may be determined1、ξ2Cost function ofWherein the content of the first and second substances,
for traps at cross-over frequency omegac(i.e., the frequency at which the amplitude-frequency curve crosses 0 dB), λ is a weighting factor and may be determined experimentally or empirically, and the second term in the above equation is the notch frequency range (ω)1,ω2) The phase error accumulation.
In step 916, an optimization parameter ξ may be determined based on phase lag1、ξ2. For example, the parameter ξ may be selected such that the phase lag is minimized1、ξ2. In other implementations, the parameter ξ may be selected for phase lag below a threshold1、ξ2. One skilled in the art can also select from the effective ξ according to other criteria1、ξ2Medium selection optimization parameter xi1、ξ2。
By utilizing the optimization algorithm, a single wave trap G in the multi-frequency wave traps can be solved in sequencekDamping parameter xi in1k、ξ2kFor example from high to low in accordance with the notch frequency. Each time a single trap G is completedkAfter the parameters are solved, the formula can be adoptedTo update the frequency response of the multi-frequency trap, and thenSuccessive solving of a single trap Gk+1Until all the damping parameters are solved, the multi-frequency wave trap can be obtained.
Compared with the prior art, the multi-frequency trap is used, and the problem that the traditional single-frequency trap cannot meet the stability margin of each signal loop is solved. The multi-frequency filter is arranged in front of the control law module, so that the requirements of stability margins of a plurality of signal loops can be met, and the influence on the low-frequency stability margin is reduced from the control law framework. Optionally, by optimizing the damping parameters of each wave trap, the multi-frequency filter can meet the high-frequency pneumatic servo stability requirement and minimize the influence on the low-frequency stability margin.
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 (22)
1. A flight control device for an aircraft, comprising:
an input component that generates flight maneuver instructions based on the received input;
a first multi-frequency trap that receives a first feedback signal based on a first sensor measurement, wherein the first multi-frequency trap has a plurality of notch frequencies and traps the first feedback signal at the plurality of notch frequencies to generate a trapped first feedback signal; and
a control law module that generates flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the notched first feedback signal.
2. The flight control apparatus of claim 1, further comprising:
a second multi-frequency trap that receives a second feedback signal based on a second sensor measurement, wherein the second multi-frequency trap has a plurality of notch frequencies and traps the second feedback signal at the plurality of notch frequencies of the second multi-frequency trap to generate a trapped second feedback signal,
wherein the control law module generates flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions, the notched first feedback signal, and the notched second feedback signal.
3. The flight control apparatus of claim 2, wherein the control law module comprises a longitudinal control law module, and wherein:
the first feedback signal comprises an inertial angle of attack rate of the aircraft;
the second feedback signal comprises a steady axis normal overload of the aircraft; and is
The flight commands include pitch commands for controlling an elevator of the aircraft.
4. The flight control apparatus of claim 2, further comprising:
a parameter synthesis module that receives a notched first feedback signal from the first multi-frequency trap and a notched second feedback signal from the second multi-frequency trap and generates a synthetic feedback signal based on the notched first feedback signal and the notched second feedback signal, wherein the control law module generates flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the synthetic feedback signal.
5. The flight control apparatus of claim 4, wherein the control law module comprises a lateral control law module, and wherein:
the first feedback signal comprises a body roll rate of the aircraft;
the second feedback signal comprises a body axis yaw rate of the aircraft;
the composite feedback signal includes a steady-axis roll rate of the aircraft; and is
The flight commands include roll commands for controlling ailerons and spoilers of the aircraft.
6. The flight control apparatus of claim 4, further comprising:
a third multi-frequency trap that receives a third feedback signal based on a third sensor measurement, wherein the third multi-frequency trap has a plurality of notch frequencies and traps the third feedback signal at the plurality of notch frequencies of the third multi-frequency trap to generate a trapped third feedback signal,
the parameter synthesis module generates the synthetic feedback signal based on a notched first feedback signal from the first multi-frequency trap, a notched second feedback signal from the second multi-frequency trap, and a notched third feedback signal from the third multi-frequency trap, wherein the control law module generates flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the synthetic feedback signal.
7. The flight control device of claim 6, wherein the control law module comprises a heading control law module, and wherein:
the first feedback signal comprises a body roll rate of the aircraft;
the second feedback signal comprises a body axis yaw rate of the aircraft;
the third feedback signal comprises a lateral overload of a body axis of the aircraft;
the composite feedback signal includes a rate of change of a side-slip angle of inertia of the aircraft; and is
The flight commands include a yaw command for controlling a rudder of the aircraft.
8. A flight control apparatus according to any one of claims 1 to 7, wherein the desired frequency response of each multi-frequency trap is based on the most severe structural modal response of the aircraft under different conditions and stability margin requirements.
9. The flight control apparatus of claim 8, wherein each multi-frequency trap comprises a plurality of single-frequency traps, wherein each single-frequency trap has a trap center frequency and a damping parameter, wherein the damping parameter is selected to meet the stability margin requirement over a range of frequencies including the trap center frequency.
10. The flight control apparatus of claim 9, wherein the damping parameters are selected to minimize phase lag of the single frequency trap.
11. The flight control apparatus of claim 1, wherein the input component comprises one or more of: side lever, pedal, handle, control panel.
12. A flight control method for an aircraft, comprising:
generating flight maneuver instructions based on the received input;
receiving a first feedback signal based on the first sensor measurement;
notching the first feedback signal at a first plurality of notch frequencies to generate a notched first feedback signal; and
generating flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the notched first feedback signal.
13. The flight control method of claim 12, further comprising:
receiving a second feedback signal based on a second sensor measurement;
notching the second feedback signal at a second plurality of notch frequencies to generate a notched second feedback signal; and
generating flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions, the notched first feedback signal, and the notched second feedback signal.
14. The flight control method according to claim 13, wherein:
the first feedback signal comprises an inertial angle of attack rate of the aircraft;
the second feedback signal comprises a steady axis normal overload of the aircraft; and is
The flight commands include pitch commands for controlling an elevator of the aircraft.
15. The flight control method of claim 13, further comprising:
generating a composite feedback signal based on the notched first feedback signal and the notched second feedback signal, and generating flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the composite feedback signal.
16. The flight control method according to claim 15, wherein:
the first feedback signal comprises a body roll rate of the aircraft;
the second feedback signal comprises a body axis yaw rate of the aircraft;
the composite feedback signal includes a steady-axis roll rate of the aircraft; and is
The flight commands include roll commands for controlling ailerons and spoilers of the aircraft.
17. The flight control method of claim 15, further comprising:
receiving a third feedback signal based on a third sensor measurement;
notching the third feedback signal at a third plurality of notch frequencies to generate a notched third feedback signal;
generating the composite feedback signal based on the notched first feedback signal, the notched second feedback signal, and the notched third feedback signal; and
generating flight commands for controlling flight actions of the aircraft based on the flight maneuver instructions and the synthetic feedback signals.
18. The flight control method according to claim 17, wherein:
the first feedback signal comprises a body roll rate of the aircraft;
the second feedback signal comprises a body axis yaw rate of the aircraft;
the third feedback signal comprises a lateral overload of a body axis of the aircraft;
the composite feedback signal includes a rate of change of a side-slip angle of inertia of the aircraft; and is
The flight commands include a yaw command for controlling a rudder of the aircraft.
19. A flight control method according to any one of claims 12 to 18, wherein each feedback signal is notched using a respective multi-frequency trap, the desired frequency response of each multi-frequency trap being based on the most severe structural modal response and stability margin requirements of the aircraft under different conditions.
20. The method of flight control of claim 19, wherein each multi-frequency trap comprises a plurality of single-frequency traps, wherein each single-frequency trap has a trap center frequency and a damping parameter, wherein the damping parameter is selected to meet the stability margin requirement over a range of frequencies including the trap center frequency.
21. The flight control method of claim 20, wherein the damping parameters are selected to minimize phase lag of the single frequency trap.
22. The flight control method of claim 12, wherein the input is received through one or more of: side lever, pedal, handle, control panel.
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