A method and apparatus for monitoring for aileron buzz

Abstract

A method for monitoring control surface force rivalry oscillation for an aircraft is disclosed. The method can include determining a difference between a control surface command and an actual actuator displacement. If the difference exceeds a first threshold, it is determined that the control surface enters an oscillation period and the difference is continuously monitored for an oscillation amplitude of the oscillation period. The method further includes accumulating a fatigue damage amount corresponding to the oscillation amplitude of each oscillation period, and when the accumulated fatigue damage amount exceeds a second threshold, it can be determined that the control surface force rivalry exceeds a limit and that there is a control surface force rivalry oscillation fault. A remote controller for an aircraft implementing the control surface force rivalry oscillation monitoring is also disclosed.

Description

Technical Field

[0001] This invention relates to the field of aircraft, and more particularly to a method and apparatus for monitoring oscillations caused by control surface force conflicts. Background Technology

[0002] During their service life, aircraft structures may experience expected or unexpected oscillations due to control loop failures or the coupling of various external forces. These oscillations can be categorized into different levels based on their impact range. Among these, the inconsistent alternating output forces among multiple actuators on the same control surface caused by uncommanded oscillations of a single actuation loop (generally measured by DDP, the difference in actuator pressure) is called control surface force conflict oscillation. The alternating stress caused by DDP will cause fatigue damage to the control surface actuation loop structure, and in severe cases, lead to localized damage to the control surface structure, affecting aircraft safety.

[0003] Control surface force conflicts are classified into long-period and short-period oscillations based on the severity of the oscillation (i.e., the amount of fatigue damage to the structure over its entire life cycle caused by the oscillation at that frequency and amplitude). Generally, after a long-period oscillation occurs, the faulty actuator is cut off but the current flight can continue. After a short-period oscillation is triggered, the faulty actuator is cut off but the nearest landing is required.

[0004] To protect the structure of the control surface actuation circuit from damage caused by force conflict oscillations, the flight control system needs to be designed with a force conflict oscillation monitor to cut off the operation of the actuation circuit that has failed due to oscillations before fatigue damage exceeds the allowable threshold. Realizing DDP monitoring requires the ability to simultaneously read the pressure signals from both chambers of different actuators on the same control surface; therefore, this function is usually housed in the Flight Controller Digital Computer (FCM). However, receiving signals from each actuator and performing control surface force conflict oscillation monitoring by the FCM may have a delay.

[0005] Therefore, there is a need in the art for an improved method and apparatus for monitoring rudder surface force conflict oscillations. Summary of the Invention

[0006] This invention proposes an improved method and device for monitoring rudder surface force conflict oscillations, which can reside in a remote controller located near the rudder surface and used to control the rudder surface actuators, thereby enabling timely and accurate identification of rudder surface force conflict oscillation failures.

[0007] In one embodiment of the present invention, a method for monitoring control surface force conflict oscillations of an aircraft is provided, comprising: receiving a control surface control command from an actuator control unit; causing an actuator associated with the control surface to generate an actuator displacement to drive the control surface based on the control surface control command; determining a difference between the control surface control command and the actuator displacement; if the difference exceeds a first threshold, determining that the control surface has entered an oscillation cycle and monitoring the oscillation amplitude of the difference in the oscillation cycle; accumulating fatigue damage corresponding to the oscillation amplitude of each oscillation cycle; and indicating the presence of a control surface force conflict oscillation fault when the accumulated fatigue damage exceeds a second threshold.

[0008] In one aspect, the method for monitoring the oscillation of control surface forces for an aircraft also includes, for each oscillation cycle: triggering the accumulation of fatigue damage for that oscillation cycle when the difference changes sign, or after the difference reaches the minimum amplitude of that oscillation cycle if the difference does not change sign.

[0009] On the one hand, for each oscillation cycle: the oscillation amplitude of the difference in that oscillation cycle includes the maximum amplitude of the difference in that oscillation cycle.

[0010] On one hand, the amount of fatigue damage is proportional to the amplitude of the oscillation.

[0011] On one hand, the control surface is driven by the actuator and at least one additional actuator.

[0012] On one hand, the actuator displacement is detected by a displacement sensor associated with the actuator.

[0013] In one embodiment of the present invention, a control surface force conflict oscillation monitoring device for an aircraft is provided, comprising: a command receiving module configured to receive control surface control commands from an actuator control unit; a drive module configured to cause an actuator associated with the control surface to generate actuator displacement based on the control surface control commands to drive the control surface; and a monitoring module configured to: determine a difference between the control surface control commands and the actuator displacement; if the difference exceeds a first threshold, determine that the control surface has entered an oscillation cycle and monitor the oscillation amplitude of the difference in the oscillation cycle; accumulate fatigue damage corresponding to the oscillation amplitude of each oscillation cycle; and indicate the presence of a control surface force conflict oscillation fault when the accumulated fatigue damage exceeds a second threshold.

[0014] On one hand, the monitoring module is also configured to, for each oscillation cycle, trigger the accumulation of fatigue damage for that oscillation cycle when the difference changes sign, or after the difference reaches the minimum amplitude of that oscillation cycle if the difference does not change sign.

[0015] On the one hand, for each oscillation cycle: the oscillation amplitude of the difference in that oscillation cycle includes the maximum amplitude of the difference in that oscillation cycle.

[0016] On one hand, the amount of fatigue damage is proportional to the amplitude of the oscillation.

[0017] On one hand, the control surface is driven by the actuator and at least one additional actuator.

[0018] On one hand, the actuator displacement is detected by a displacement sensor associated with the actuator.

[0019] In one embodiment of the present invention, an aircraft is provided, including a control surface force conflict oscillation monitoring device for an aircraft as described in any of the preceding claims. Attached Figure Description

[0020] Figure 1 This is an architecture diagram of a fly-by-wire flight control system according to one embodiment.

[0021] Figure 2 This is a flowchart of a method for monitoring rudder force conflict oscillations according to an embodiment of the present invention.

[0022] Figure 3 This is a schematic diagram of a force dispute oscillation monitoring logic according to an embodiment of the present invention.

[0023] Figure 4 This is an example of an oscillation waveform according to an embodiment of the present invention.

[0024] Figure 5 This is a flowchart of the control surface force conflict oscillation monitoring logic according to an embodiment of the present invention.

[0025] Figure 6 This is a block diagram of a control surface force conflict oscillation monitoring device for an aircraft according to an embodiment of the present invention. Detailed Implementation

[0026] The present invention will be further described below with reference to specific embodiments and accompanying drawings, but this should not be construed as limiting the scope of protection of the present invention.

[0027] To monitor and identify rudder surface force conflict oscillation faults in a timely and accurate manner, and to protect the structure on the rudder surface actuation circuit from damage due to force conflict oscillations, this invention proposes a rudder surface force conflict oscillation fault monitoring method residing in a remote controller (REU). By comparing the command position with the actual position of a single actuation circuit, force conflict oscillation faults are identified, thus improving the timeliness and accuracy of identifying rudder surface force conflict oscillation faults.

[0028] Figure 1This is an architectural diagram of a fly-by-wire flight control system 100 according to one embodiment. The fly-by-wire flight control system may include one or more flight control computers 101 (e.g., FCMs) and one or more actuator control units 111, 112 (e.g., ACEs). In other embodiments, the fly-by-wire flight control system may include more or fewer flight control electronics.

[0029] The flight control computer 101 can receive control inputs from cockpit control equipment, external system interfaces, and aircraft attitude data. Based on these input signals, it can calculate control commands for the control surfaces according to control laws to control aircraft control surfaces such as elevators, rudders, ailerons, and spoilers. For example, current civil aircraft can use fly-by-wire flight control systems, with side sticks, pedals, and speed brake levers as control input devices, a redundant flight control digital computer (FCM), actuator control unit (ACE), and remote controller (REU) as core processing components, and electromechanical and electrohydraulic servo actuators as execution components. Through control law calculation, it controls the corresponding control surfaces, achieving full-time, full-authority fly-by-wire control.

[0030] After calculating the control surface commands, the flight control computer 101 sends them to each actuator control unit. Each actuator control unit then transmits the commands to its controlled REU via a separate bus and can receive feedback data from the REU. Ultimately, the local position closed-loop control of the actuators is achieved through the REU. Based on redundancy allocation, different actuators on the same control surface can be controlled by different actuator control units, with each actuator controlled by a different REU. In the event of a flight control computer failure, the actuator control unit can still calculate the control surface drive signals based on cockpit control equipment commands, ensuring safe flight and landing of the aircraft.

[0031] refer to Figure 1 The flight control computer 101 generates control surface commands and provides them to the first actuator control unit 111 and the second actuator control unit 112, respectively. The first actuator control unit 111 transmits the received (or generated) control surface commands to the first REU 113, which causes the first actuator 114 to drive the control surface 120 according to the control surface commands. The second actuator control unit 112 transmits the received (or generated) control surface commands to the second REU 115, which causes the second actuator 116 to drive the control surface 120 according to the control surface commands.

[0032] As mentioned earlier, when the output forces of multiple actuators 114 and 116 on the same control surface 120 are inconsistent, it causes control surface force conflict oscillation. Control surface force conflict oscillation will cause fatigue damage to the control surface actuation circuit structure, and in severe cases, it will lead to local structural damage to the control surface, affecting aircraft safety.

[0033] This invention proposes an improved method and device for monitoring rudder surface force conflict oscillations, which can reside in a remote controller (REU) close to the rudder surface, thereby enabling timely and accurate identification of rudder surface force conflict oscillation failures.

[0034] Figure 2 This is a flowchart of a control surface force conflict oscillation monitoring method 200 according to an embodiment of the present invention. The method can be implemented within a remote controller (REU), for example, referring to... Figure 1 The first REU 113 and / or the second REU 115 are implemented, for example, by a device such as a computer, processor, controller, integrated circuit, etc.

[0035] In step 202, control surface commands may be received from the actuator control unit. For example, the remote controller (REU) may receive control surface commands from the actuator control unit, which may be calculated by the flight control computer or the actuator control unit based on commands from the cockpit control equipment.

[0036] In step 203, the actuators associated with the control surface can be displaced based on control surface commands to drive the control surface. For example, refer to Figure 1 The REU can be a first REU 113, which causes the first actuator 114 to drive the control surface 120 based on the control surface control command provided by the first actuator control unit 111. The control surface 120 can also be driven by at least one additional actuator (e.g., a second actuator 116), which can drive the same control surface 120 based on the control surface control command received by the second REU 115.

[0037] In step 204, the difference between the control surface command and the actuator displacement can be determined. The actuator displacement can be detected by a displacement sensor associated with the actuator. For example, an actuator position sensor can detect the actuator piston position as the actuator displacement. The difference between the control surface command and the actuator displacement can be obtained by subtracting the actuator displacement from the control surface command, or by subtracting the control surface command from the actuator displacement.

[0038] In one embodiment, the timeline can be divided into multiple frames (e.g., 1 millisecond), with each frame representing the delay from receiving a surface control command to completing an actuator displacement. The REU can receive a surface control command in the first frame and generate or detect an actuator displacement in a subsequent second frame. To account for the delay in actuator displacement, the difference between the surface control command in the previous frame and the actuator displacement detected in the current frame can be used.

[0039] In step 205, it is determined whether the difference exceeds a first threshold. If the difference exceeds the first threshold, proceed to step 206; otherwise, return to step 202. By way of example and not limitation, if the difference is greater than zero, exceeding the first threshold may include the difference being greater than a positive first threshold. Conversely, if the difference is less than zero, exceeding the first threshold may include the difference being less than a negative first threshold.

[0040] In step 206, it can be determined that the control surface has entered an oscillation cycle and the oscillation amplitude of the difference during that oscillation cycle can be monitored. By way of example and not limitation, the difference can be monitored continuously or periodically at a certain interval (or sampling rate).

[0041] In step 207, the fatigue damage amount corresponding to the oscillation amplitude of each oscillation cycle can be accumulated. In one embodiment, the fatigue damage amount of each oscillation cycle can be accumulated (e.g., summed) at the end of each oscillation cycle. For example, for each oscillation cycle, the oscillation cycle is considered to have ended when the difference changes sign (crosses 0), or when the difference reaches the minimum amplitude of the oscillation cycle if the difference does not change sign, and the accumulation of the fatigue damage amount of the oscillation cycle can be triggered.

[0042] For each oscillation cycle, the oscillation amplitude of the difference in that cycle includes the maximum amplitude of the difference in that cycle. The amount of fatigue damage can be proportional to the oscillation amplitude. The larger the oscillation amplitude, the greater the corresponding amount of fatigue damage. Fatigue damage can accumulate over a predetermined time period. As an example, and not a limitation, fatigue damage can be accumulated separately for each segment of the voyage, or it can be accumulated over multiple segments of the voyage.

[0043] Force-induced oscillations are constrained by the corresponding DDP amplitude at different frequency points f. The DDP limit at different frequencies can typically be expressed as a*f^(b) (Formula 1) based on the aircraft structural characteristics. Here, a and b are constant coefficients, typically taken as 10% (or another ratio) of the structural fatigue life damage as the allowable damage consumption boundary condition. An allowable oscillation time limit t is set for the limit line; exceeding this time is considered to have unacceptable fatigue effects on the aircraft structure. Usually, two time limits t are set, for example, 3 hours (long period) and 3 minutes (short period). This time limit indicates that if the actuator oscillates at the corresponding frequency and amplitude, exceeding time t will consume 10% of the fatigue life damage. After capturing one oscillation cycle, the fatigue damage generated by that oscillation cycle is converted into the fatigue damage amount x% of that cycle using the following formula:

[0044]

[0045] Where Y_peak is the peak value of the current period's oscillation (i.e., the oscillation amplitude), t is the allowable oscillation time limit constant (which can correspond to the time limit of a short period or a long period), and a and b are the coefficient constants of the force conflict oscillation limit line function (Formula 1), respectively. The specific values ​​of a, b, and t can be set according to the aircraft structure and airworthiness configuration, and are not limited in this invention.

[0046] It should be understood that the above formula is merely an example for determining fatigue damage, not a limitation. In practice, different relationships between oscillation amplitude and fatigue damage can be used. For example, various coefficient values ​​can be set, coefficients can be increased or decreased, or other parameters can be modified as needed. Besides using formulas to determine the fatigue damage corresponding to the oscillation amplitude for each oscillation cycle, other methods can also be used to determine fatigue damage. For example, for a given aircraft structure and airworthiness configuration, a mapping table or graph between oscillation amplitude and fatigue damage may exist. After obtaining the oscillation amplitude, the corresponding fatigue damage can be obtained.

[0047] In step 208, it can be determined whether the accumulated fatigue damage exceeds a second threshold. If so, proceed to step 209; otherwise, return to step 202.

[0048] In step 209, when the accumulated fatigue damage exceeds a second threshold, a control surface force conflict oscillation fault may be indicated. In one embodiment, the REU reports the fault, and the FCM provides corresponding crew alarms upon receiving the fault signal, such as audible alarms, visual alarms, etc. In a further embodiment, when the accumulated fatigue damage exceeds a second or third threshold, the FCM may selectively disconnect the associated actuators or take appropriate remedial measures.

[0049] In one embodiment, the accumulated fatigue damage can be reset (e.g., cleared to zero) if predetermined conditions are met. For example, the accumulated fatigue damage can be reset if it does not exceed a second threshold within a specified time, when the aircraft completes maintenance, or after a fault alarm is triggered due to the accumulated fatigue damage exceeding the second threshold. This process can then be repeated. Figure 2 The process described herein restarts the accumulation of fatigue damage. By resetting the accumulated fatigue damage, the continuous accumulation of fatigue damage can be prevented, or accidental triggering caused by fatigue accumulated under normal command actions can be avoided.

[0050] It should be understood that the various thresholds used in this invention may be the same or different, and their specific values ​​may be determined as needed for specific implementation, without limiting the invention.

[0051] The force conflict oscillation fault monitoring function proposed in this invention can reside within a remote controller (REU). The REU has a high data processing and refresh rate, thus this method can accurately capture oscillation peaks with high reliability. According to this invention, no additional signals or equipment are required; the oscillation monitoring function can be achieved by comparing the existing position feedback signal of the actuation loop with the command. Furthermore, one REU is responsible for controlling only one servo actuator. This method can accurately identify failed actuation loops, while other actuators on the same control surface can still retain active operation capabilities, ensuring the availability of the control surface and the remaining configurable redundancy of the flight control system.

[0052] Figure 3 This is a schematic diagram of the force conflict oscillation monitoring logic according to an embodiment of the present invention. The force conflict oscillation monitoring logic flow of the control surface can be implemented by a REU (e.g., the control surface force conflict oscillation monitoring module or monitor therein).

[0053] Figure 4 This is an example of an oscillation waveform according to an embodiment of the present invention, wherein the solid line represents the oscillation waveform and the dashed line represents one recognition cycle.

[0054] according to Figure 3 As shown in box 301, the command-position difference (ΔY) is obtained by subtracting the position command signal (command (t-1)) from the previous frame's position command signal and the current servo actuator position signal. When the difference exceeds a set threshold A... start At this point, if a potential force conflict oscillation is suspected, in box 302, the monitor begins searching for waveform peaks (amplitudes): comparing the current ΔY with ΔY. max (Initial value or ΔY obtained from previous periods) max ), update ΔY based on the larger of the two values. max until ΔY max If the value stops increasing (e.g., it stops increasing after a predetermined number of cycles, or ΔY drops from its highest value beyond a specified threshold), update Y_peak (oscillation peak) to ΔY for the current cycle. max . Figure 4 This shows what happens when ΔY exceeds the set threshold A. start Find and obtain ΔY max The waveform.

[0055] refer to Figure 3 If in ΔY max If ΔY then decreases, the difference between the current ΔY and Y_peak is obtained in box 311. When this difference exceeds the threshold A... stop At that time, starting from box 312, search for and record the minimum difference ΔY within the period. min For example, comparing the current ΔY with ΔY min (Initial value or ΔY obtained from previous periods) min ), update ΔY based on the smaller of the two values.min until ΔY min If the value stops decreasing (e.g., it stops decreasing after a predetermined number of periods, or ΔY increases from its minimum value to exceed a specified threshold), update Y_valley (oscillation valley value) to ΔY for the current period. min .

[0056] If we are looking for ΔY min If ΔY does not change sign during the process, then in box 322, when ΔY changes from Y_valley (or ΔY min The value increases and rises with the current ΔY and Y_valley (or ΔY). min When the difference between ΔY and ΔY exceeds the threshold Y', the recording of the current cycle is completed, capturing half an oscillation cycle (ΔY). max ~ΔY min At this point, a trigger signal can be generated to calculate the fatigue damage amount for the current cycle in block 330. See also Figure 4 The figure in (1).

[0057] In another embodiment, when the difference between the current ΔY and Y_peak exceeds a threshold A... stop Start searching for and recording the minimum difference ΔY within the period. min If ΔY changes sign (box 321), then when the sign of ΔY changes, the recording of the current cycle is completed, capturing half an oscillation cycle (ΔY). max ~0), at which point a trigger signal can be generated to calculate the fatigue damage amount for the current cycle in block 330. See also Figure 4 Figure (2) in the figure. In this case, ΔY can be considered as min (or Y_valley) is 0.

[0058] In box 330, the oscillation amplitude (Y_peak or ΔY) captured in the current period is displayed. max The calculated fatigue damage percentage (x%) is converted to an equivalent fatigue damage amount (x%). In box 331, the fatigue damage percentage (x%) calculated in the current cycle is added to the previously accumulated fatigue damage amount that has been stored. When the accumulated value exceeds a set allowable value (e.g., a second threshold), the monitor is triggered to indicate the presence of a rudder surface force conflict oscillation fault.

[0059] Figure 5 This is a flowchart of the control surface force conflict oscillation monitoring logic according to an embodiment of the present invention. Figure 5 It can be Figure 2 This illustrates one implementation of the control surface force conflict oscillation monitoring method 200. The control surface force conflict oscillation monitoring logic flow can be implemented by a REU (e.g., a control surface force conflict oscillation monitoring module or monitor therein).

[0060] In step 501, the difference ΔY between the actuator position signal and the command signal can be calculated. If the difference ΔY exceeds a first threshold (e.g., A...), start In step 502, the oscillation peak value ΔY can be found and recorded. max If the current ΔY and ΔY max If the difference exceeds the second threshold, then in step 503, find and record the minimum value ΔY. min If the current ΔY and ΔY min If the difference exceeds the third threshold, the current oscillation cycle is captured in step 504.

[0061] In another implementation, if ΔY is recorded... max Then, in step 505, if ΔY decreases and changes sign, it is also considered that the capture of the current oscillation period is complete, where ΔY min It is 0.

[0062] In step 506, it can be based on ΔY max Calculate the fatigue damage amount x% for the current cycle.

[0063] In step 507, the fatigue damage amount x% of the current cycle can be added to the previously stored (previous cycles') fatigue accumulation amount to obtain a new fatigue accumulation amount.

[0064] If the accumulated fatigue exceeds a threshold, the monitor can be triggered in step 508. For example, the monitor can issue a fault signal, causing the REU to report a surface force conflict oscillation fault.

[0065] Figure 6 This is a block diagram of a control surface force conflict oscillation monitoring device 600 for an aircraft according to an embodiment of the present invention. The control surface force conflict oscillation monitoring device 600 may be implemented by a remote controller, or implemented in a remote controller.

[0066] The rudder force conflict oscillation monitoring device 600 may include a command receiving module 601, which is configured to receive rudder control commands from the actuator control unit 610.

[0067] The control surface force conflict oscillation monitoring device 600 may include a drive module 602 configured to cause an actuator 620 associated with the control surface to generate actuator displacement to drive the control surface based on control surface commands. The control surface may also be driven by at least one additional actuator.

[0068] The control surface force conflict oscillation monitoring device 600 may further include a monitoring module 603 configured to determine the difference between the control surface command and the actuator displacement. The actuator displacement may be detected by a displacement sensor associated with the actuator.

[0069] If the difference exceeds the first threshold, the monitoring module 603 determines that the control surface has entered an oscillation cycle and monitors the oscillation amplitude of the difference during that oscillation cycle. For each oscillation cycle, the oscillation amplitude of the difference during that oscillation cycle includes the maximum amplitude of the difference during that oscillation cycle.

[0070] The monitoring module 603 can accumulate fatigue damage corresponding to the oscillation amplitude of each oscillation cycle. The fatigue damage is proportional to the oscillation amplitude. When the accumulated fatigue damage exceeds a second threshold, the monitoring module 603 can indicate the presence of a control surface force conflict oscillation fault. For each oscillation cycle, the calculation / accumulation of fatigue damage for that oscillation cycle is triggered when the difference changes sign, or if the difference does not change sign, after the difference reaches the minimum amplitude of that oscillation cycle.

[0071] The control surface force conflict oscillation monitoring device 600 may also include other components, such as a memory 607. The memory 607 can store various received or generated data, such as control surface commands, actuator displacements, and various thresholds.

[0072] This invention proposes a method for monitoring control surface force conflict oscillation faults residing in a remote controller. It compares the actual actuator displacement with the commanded position (provided by the flight control computer) using a monitor. When the difference exceeds a set threshold, a force conflict fault is identified, and the system begins searching for the oscillation peak (amplitude). The amplitude of the oscillation is then converted into fatigue damage. When the accumulated fatigue damage exceeds a threshold, a control surface force conflict oscillation fault can be identified, improving the timeliness and accuracy of fault identification.

[0073] The various steps and modules of the methods and apparatus described above can be implemented in hardware, software, or a combination thereof. If implemented in hardware, the various illustrative steps, modules, and circuits described in connection with this disclosure can be implemented or executed using a general-purpose processor, digital signal processor (DSP), application-specific integrated circuit (ASIC), field-programmable gate array (FPGA), or other programmable logic components, hardware components, or any combination thereof. A general-purpose processor can be a processor, microprocessor, controller, microcontroller, or state machine, etc. If implemented in software, the various illustrative steps and modules described in connection with this disclosure can be stored as one or more instructions or codes on a computer-readable medium or transmitted. Software modules implementing the various operations of this disclosure can reside in a storage medium, such as RAM, flash memory, ROM, EPROM, EEPROM, registers, hard disk, removable disk, CD-ROM, cloud storage, etc. The storage medium can be coupled to a processor so that the processor can read and write information from / to the storage medium and execute the corresponding program modules to implement the various steps of this disclosure.

[0074] The numerical values ​​given in the various embodiments are merely examples and are not intended to limit the scope of the invention. In practice, the specific parameters of each component and various thresholds can be appropriately set as needed, and are not limited to the specific values ​​given as examples herein. Furthermore, as a whole technical solution, there are other components or steps not listed in the claims or specification of this invention. Moreover, a single name for a component does not preclude other names for that component.

[0075] It should also be noted that these embodiments may be described as processes depicted as flowcharts, flow diagrams, structure diagrams, or block diagrams. Although a flowchart may describe the operations as a sequential process, many of these operations can be executed in parallel or concurrently. Furthermore, the order of these operations can be rearranged.

[0076] The disclosed methods, apparatuses, and systems should not be limited in any way. Rather, this disclosure covers all novel and non-obvious features and aspects of the various disclosed embodiments (individually and in various combinations and sub-combinations of each other). The disclosed methods, apparatuses, and systems are not limited to any particular aspect or feature or combination thereof, and no disclosed embodiment is required to have any one or more specific advantages or to solve any particular or all technical problems.

[0077] This invention is not limited to the specific embodiments described above. The specific embodiments described above are merely illustrative and not restrictive. Those skilled in the art can make many other modifications based on the teachings of this invention without departing from the spirit and scope of the claims. All of these modifications are within the scope of protection of this invention.

Claims

1. A method for monitoring control surface force rivalry oscillations for an aircraft, characterized in that, comprising: receiving surface control commands from an actuator control unit; causing an actuator associated with a surface to produce an actuator displacement to drive the surface based on the surface control commands; determining a difference between the surface control commands and the actuator displacement; if the difference exceeds a first threshold, determining that the surface enters an oscillation cycle and monitoring the difference for an oscillation amplitude of the oscillation cycle; accumulating a fatigue damage quantity corresponding to the oscillation amplitude of each oscillation cycle, wherein for each oscillation cycle, the fatigue damage quantity of the oscillation cycle is triggered to be accumulated when the difference changes sign or after the difference reaches a minimum amplitude of the oscillation cycle if the difference does not change sign; and when the accumulated fatigue damage quantity exceeds a second threshold, indicating that a surface force rivalry oscillation fault exists.

2. The method of control surface force rivalry oscillation monitoring for an aircraft as in claim 1, wherein, for each oscillation cycle: the oscillation amplitude of the difference for the oscillation cycle includes a maximum amplitude of the difference for the oscillation cycle.

3. The surface force rivalry oscillation monitoring method for an aircraft of claim 1, wherein: the fatigue damage quantity is proportional to the oscillation amplitude.

4. The surface force rivalry oscillation monitoring method for an aircraft of claim 1, wherein: the surface is driven by the actuator and at least one additional actuator.

5. The surface force rivalry oscillation monitoring method for an aircraft of claim 1, wherein: the actuator displacement is detected by a displacement sensor associated with the actuator.

6. An oscillating control surface force disagreement monitoring device for an aircraft, characterized by, comprising: an instruction receiving module configured to receive surface control commands from an actuator control unit; a driving module configured to cause an actuator associated with a surface to produce an actuator displacement to drive the surface based on the surface control commands; a monitoring module configured to: determine a difference between the surface control commands and the actuator displacement, if the difference exceeds a first threshold, determine that the surface enters an oscillation cycle and monitor the difference for an oscillation amplitude of the oscillation cycle; accumulate a fatigue damage quantity corresponding to the oscillation amplitude of each oscillation cycle, wherein for each oscillation cycle, the fatigue damage quantity of the oscillation cycle is triggered to be accumulated when the difference changes sign or after the difference reaches a minimum amplitude of the oscillation cycle if the difference does not change sign; and when the accumulated fatigue damage quantity exceeds a second threshold, indicate that a surface force rivalry oscillation fault exists.

7. The control surface force disagreement oscillation monitoring apparatus for an aircraft as defined in claim 6, wherein, for each oscillation cycle: the oscillation amplitude of the difference for the oscillation cycle includes a maximum amplitude of the difference for the oscillation cycle.

8. The surface force rivalry oscillation monitoring apparatus for an aircraft of claim 6, wherein: the fatigue damage quantity is proportional to the oscillation amplitude.

9. The surface force rivalry oscillation monitoring apparatus for an aircraft of claim 6, wherein: the surface is driven by the actuator and at least one additional actuator.

10. The surface force rivalry oscillation monitoring apparatus for an aircraft of claim 6, wherein: the actuator displacement is detected by a displacement sensor associated with the actuator.

11. An aircraft comprising the control surface force disagreement oscillation monitoring apparatus for an aircraft as claimed in any one of claims 6 to 10.

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