Method and apparatus for aircraft force contention monitoring

Abstract

A method and apparatus for monitoring force conflicts in an aircraft are disclosed. The apparatus may include a first actuator and a second actuator driving a control surface. The first actuator drives the control surface through a first actuator rod, a first force sensor is configured to detect a first output force applied by the first actuator to the first actuator rod, and a first position sensor is configured to detect a first displacement of the first actuator rod. The second actuator drives the control surface through a second actuator rod, a second force sensor is configured to detect a second output force applied by the second actuator to the second actuator rod, and a second position sensor is configured to detect a second displacement of the second actuator rod. The force conflict monitor is configured to determine that a force conflict occurs on the control surface when the output force difference between the first output force and the second output force is greater than a first threshold value and the displacement difference between the first displacement and the second displacement is greater than a second threshold value.

Description

Technical Field

[0001] The present invention relates to the field of aircraft, and in particular to a method and device for monitoring force disputes in aircraft. Background Art

[0002] Redundant servo actuators are widely used in flight control systems, which adopts a strategy in which two or more actuators actively control the same control surface in parallel. The entire servo loop of the control surface actuation system mainly includes various sensors, electro-hydraulic servo valves and actuators.

[0003] When multiple actuators are used to control a control surface at the same time, the manufacturing tolerances of the components in the control circuit and their installation lead to differences in the responses of the various actuators, so that the actuators restrain each other and produce disputes. This is the phenomenon of "force dispute" between the actuators. Rudder surface force dispute is an inherent characteristic of the flight control system using redundant actuators. Excessive force dispute will impose additional loads on the aircraft's control surface, which has a greater impact on the static strength of the control surface.

[0004] Therefore, there is a need in the art for an effective method and apparatus for monitoring aircraft force conflicts. Summary of the invention

[0005] The present invention provides a force dispute monitoring method and device that is easy to implement, highly applicable, and highly robust. By using the output force difference and output displacement difference between multiple actuators as input signals of a force dispute monitor, force dispute monitoring is performed separately through dual channels. When the thresholds of both channels are triggered, it is determined that a force dispute occurs on the rudder surface, thereby improving the robustness of the force dispute monitor.

[0006] According to one embodiment of the present invention, there is provided a device for monitoring force conflicts of an aircraft, comprising: a first actuator and a first force sensor and a first position sensor associated with the first actuator, wherein the first actuator drives the control surface of the aircraft through a first actuator rod, the first force sensor is configured to detect a first output force applied by the first actuator to the first actuator rod, and the first position sensor is configured to detect a first displacement of the first actuator rod; a second actuator and a second force sensor and a second position sensor associated with the second actuator, wherein the second actuator drives the control surface through a second actuator rod, the second force sensor is configured to detect a second output force applied by the second actuator to the second actuator rod, and the second position sensor is configured to detect a second displacement of the second actuator rod; and a force conflict monitor, wherein the force conflict monitor is configured to determine that a force conflict occurs on the control surface when the output force difference between the first output force and the second output force is greater than a first threshold value and the displacement difference between the first displacement and the second displacement is greater than a second threshold value.

[0007] On the one hand, each of the first actuator and the second actuator includes a cavity and a piston installed in the cavity, the first actuator rod and the second actuator rod are respectively connected to the pistons of the first actuator and the second actuator, wherein the piston divides the cavity into two chambers, and each of the first output force and the second output force includes a pressure difference between the two chambers separated by the corresponding piston.

[0008] In one aspect, the control surface includes a spoiler, an aileron, a rudder, or an elevator.

[0009] In one aspect, the force conflict monitor is further configured to multiply the displacement difference by a coefficient to generate a conversion force value, compare the conversion force value with the output force difference value, and issue a force sensor fault signal if the conversion force value differs from the output force difference value by more than a third threshold.

[0010] In one aspect, the force conflict monitor inhibits force conflict monitoring of the control surface in response to the force sensor fault signal.

[0011] In one aspect, in response to a force conflict occurring with the control surface, one or both of the first actuator or the second actuator is inhibited from driving the control surface.

[0012] According to one embodiment of the present invention, a method for monitoring force conflicts on an aircraft is provided, comprising: receiving a first output force and a first displacement of a first actuator, wherein the first actuator drives a rudder surface of the aircraft through a first actuator rod, the first output force represents a first output force applied by the first actuator to the first actuator rod and the first displacement represents a first displacement of the first actuator rod; receiving a second output force and a second displacement of a second actuator, wherein the second actuator drives the rudder surface through a second actuator rod, the second output force represents a second output force applied by the second actuator to the second actuator rod and the second displacement represents a second displacement of the second actuator rod; and determining that a force conflict occurs on the rudder surface when the output force difference between the first output force and the second output force is greater than a first threshold value and the displacement difference between the first displacement and the second displacement is greater than a second threshold value.

[0013] On the one hand, each of the first actuator and the second actuator includes a cavity and a piston installed in the cavity, the first actuator rod and the second actuator rod are respectively connected to the pistons of the first actuator and the second actuator, wherein the piston divides the cavity into two chambers, and each of the first output force and the second output force includes a pressure difference between the two chambers separated by the corresponding piston.

[0014] In one aspect, the control surface includes a spoiler, an aileron, a rudder, or an elevator.

[0015] On the one hand, the method for aircraft force conflict monitoring also includes: multiplying the displacement difference by a coefficient to generate a conversion force value, comparing the conversion force value with the output force difference value, and issuing a force sensor fault signal if the conversion force value differs from the output force difference value by more than a third threshold.

[0016] In one aspect, the method for aircraft force conflict monitoring further includes inhibiting force conflict monitoring of the control surface in response to the force sensor fault signal.

[0017] In one aspect, the method for aircraft force conflict monitoring further includes: in response to the force conflict occurring on the control surface, causing one or both of the first actuator or the second actuator to inhibit driving the control surface.

[0018] According to one embodiment of the present invention, there is provided a flight control system, which includes the device for monitoring aircraft force conflicts as described in any one of the above items. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] Figure 1 FIG. 1 is a schematic diagram of a control surface force conflict according to an embodiment of the present invention.

[0020] Figure 2 FIG. 4 is a schematic diagram of the architecture of a device for monitoring force contention according to an embodiment of the present invention.

[0021] Figure 3 is a logical diagram of a force contention monitor according to one embodiment of the present invention.

[0022] Figure 4 is a flow chart of a method for force contention monitoring according to an embodiment of the present invention. DETAILED DESCRIPTION

[0023] The present invention will be further described below in conjunction with specific embodiments and drawings, but the protection scope of the present invention shall not be limited thereto.

[0024] The device for monitoring force conflicts of an aircraft according to the present invention may include a first actuator and a second actuator for driving a control surface. The first actuator drives the control surface through a first actuator rod, a first force sensor is configured to detect a first output force applied by the first actuator to the first actuator rod, and a first position sensor is configured to detect a first displacement of the first actuator rod. The second actuator drives the control surface through a second actuator rod, a second force sensor is configured to detect a second output force applied by the second actuator to the second actuator rod, and a second position sensor is configured to detect a second displacement of the second actuator rod. The force conflict monitor is configured to determine that a force conflict occurs on the control surface when the output force difference between the first output force and the second output force is greater than a first threshold value and the displacement difference between the first displacement and the second displacement is greater than a second threshold value.

[0025] Figure 1 FIG. 1 is a schematic diagram of a control surface force conflict according to an embodiment of the present invention. Figure 1 The figure shows a control surface 101 of an aircraft and two actuators 110 and 120 driving the control surface. The two actuators 110 and 120 generate output forces F1 and F2 respectively in response to the control surface position command from the flight control computer, thereby driving the control surface 101 to move. Due to the response difference, installation clearance and control surface torsional stiffness of the two actuators 110 and 120 used to control the same control surface 101, the forces F1 and F2 output by the actuators 110 and 120 associated with the control surface 101 are inconsistent, resulting in a "force dispute" phenomenon.

[0026] Rudder surface force dispute is an inherent characteristic of a flight control system using redundant actuators. Excessive force dispute will impose additional loads on the aircraft rudder surface, which has a greater impact on the static strength of the rudder surface. According to an embodiment of the present invention, the flight control system can monitor static force disputes and isolate faults by setting a force dispute monitor.

[0027] Figure 2 2 is a schematic diagram of a device 200 for monitoring force contention according to an embodiment of the present invention. The device 200 for monitoring force contention may include a first actuator 250 for driving a control surface 201, a second actuator 260, and a force contention monitor 222. As an example and not a limitation, Figure 2 The force dispute monitor 222 is shown as part of the flight control computer (FCM) 220. In other embodiments, the force dispute monitor 222 may exist as a separate device, or may be integrated with suitable electronic equipment on the aircraft. The flight control computer 220 and the force dispute monitor 222 may each be implemented using a computer, a processor, an integrated circuit, a programmable logic device, a microprocessor, a controller, a microcontroller, or a state machine.

[0028] The flight control computer 220 transmits the control surface position instruction to the first actuator control unit 210 and the second actuator control unit 230. The first actuator control unit 210 sends an actuator position instruction signal to the first actuator 250 based on the control surface position instruction to drive the control surface 201, and the second actuator control unit 230 sends an actuator position instruction signal to the second actuator 260 based on the control surface position instruction to drive the control surface 201. The first actuator 250 and the second actuator 260 can drive different parts of the same control surface 201. The control surface 201 can be, for example, a spoiler, an aileron, a rudder, an elevator, etc.

[0029] The first actuator 250 and the second actuator 260 may each be an electro-hydraulic servo actuator. As an example and not a limitation, Figure 2 The first actuator 250 is shown to be an electro-hydraulic servo actuator, which may include a piston 251 installed in a cavity, the piston 251 dividing the cavity into two cavities 252 and 253, and the two cavities 252 and 253 may be filled with liquid. The piston 251 is connected to a first actuating rod 254, and the output end of the first actuating rod 254 is connected to the rudder surface 201. When the first actuator 250 receives a control command from the first actuator control unit 210, the liquid content in the two cavities 252 and 253 is adjusted according to the control command, thereby causing a pressure difference between the two cavities 252 and 253, and the pressure difference causes the piston 251 to move, thereby driving the first actuating rod 254 to move and correspondingly driving the rudder surface 201 to move.

[0030] The second actuator 260 may have a cavity and piston structure similar to that of the first actuator 250 , and drive the control surface 201 to move via a second actuating rod 264 .

[0031] According to an embodiment of the present invention, the first actuator 250 has an associated first force sensor 256 and a first position sensor 258, wherein the first force sensor 256 is configured to detect a first output force applied by the first actuator 250 to the first actuator rod 254, and the first position sensor 258 is configured to detect a first displacement of the first actuator rod 254.

[0032] The second actuator 260 has an associated second force sensor 266 and a second position sensor 268 , wherein the second force sensor 266 is configured to detect a second output force applied by the second actuator 260 to the second actuating rod 264 , and the second position sensor 268 is configured to detect a second displacement of the second actuating rod 264 .

[0033] The information detected by the first force sensor 256 and the first position sensor 258 may be provided to the force dispute monitor 222 via the first actuator control unit 210, and the information detected by the second force sensor 266 and the second position sensor 268 may be provided to the force dispute monitor 222 via the second actuator control unit 230. According to one embodiment of the present invention, the force dispute monitor 222 may be configured to determine that a force dispute occurs on the control surface 201 when the output force difference between the first output force and the second output force is greater than a first threshold value and the displacement difference between the first displacement and the second displacement is greater than a second threshold value. For example, the force dispute monitor 222 may provide a signal indicating that a force dispute occurs on the control surface to the flight control computer 220. In response to the signal that a force dispute occurs on the control surface, the flight control computer 220 may issue a control surface inhibition signal, for example, causing one or both of the first actuator 250 or the second actuator 260 to inhibit driving the control surface 201.

[0034] In the embodiment using the electro-hydraulic servo actuator, the force sensors 256 and 266 can be implemented as pressure sensors, which can detect the pressure of two chambers separated by the piston, and the detected pressure of the two chambers can be transmitted to the flight control computer 220 or the force dispute monitor 222 through the actuator control units 210 and 230. In the flight control computer 220 or the force dispute monitor 222, the pressure difference (Delta Pressure, DP) of the two chambers of a single actuator can be calculated according to the data of the pressure sensor, that is, the output force of the corresponding actuator can be obtained. Further, the flight control computer 220 or the force dispute monitor 222 can calculate the output force difference (Difference Delta Pressure, DDP) between the output forces of the two actuators.

[0035] In one embodiment, the force dispute monitor 222 may also be configured to multiply the displacement difference between the two actuators by a coefficient to generate a conversion force value, and compare the conversion force value with the output force difference between the two actuators. If the conversion force value differs from the output force difference by more than a third threshold, the force dispute monitor 222 may issue a force sensor failure signal. In one embodiment, in response to the force sensor failure signal, the flight control computer 220 or the force dispute monitor 222 may suppress force dispute monitoring for the control surface 201.

[0036] As described above, the present invention provides a design of a multi-source input and highly robust force dispute monitor, which monitors the force dispute of the rudder surface through the actuator output force and the actuator displacement respectively. The rudder surface force dispute monitor will only be triggered when the thresholds of the two monitoring channels are triggered. This dual-channel monitoring form reduces the failure of a single rudder surface due to the false triggering of the force dispute monitor, and improves the robustness of the monitor and the aircraft control performance. At the same time, the design of the force dispute monitor according to the present invention can prevent the rudder surface from being fatigued and causing force disputes due to excessive deformation.

[0037] Figure 3 1 is a logic diagram of a force contention monitor according to one embodiment of the present invention. The force contention monitor may be implemented as part of a flight control computer or may be implemented separately. For example, the force contention monitor may be implemented using a computer, a processor, an integrated circuit, a programmable logic device, a microprocessor, a controller, a microcontroller, or a state machine.

[0038] The force dispute monitor receives the two-chamber pressure signals 301 and 302 of the inner actuator (IB), the two-chamber pressure signals 303 and 304 of the outer actuator (OB), the inner actuator displacement signal 305, and the outer actuator displacement signal 306 transmitted by the actuator control unit (ACE). Then the force dispute monitor judges the actuator output force and displacement signals respectively:

[0039] 1) The force dispute monitor takes the difference of the two-chamber pressure signals 301 and 302 of the inner actuator and takes the absolute value to obtain the two-chamber pressure difference of the inner actuator, that is, the output force 308 of the inner actuator; takes the difference of the two-chamber pressure signals 303 and 304 of the outer actuator and takes the absolute value to obtain the two-chamber pressure difference of the outer actuator, that is, the output force 309 of the outer actuator; then the output forces 308 and 309 are taken the difference to obtain the output force difference (DDP) 310. When DDP 310 is greater than the set threshold value (constant1) 311 and lasts for a specified time, it is judged to be true, and the judgment signal 312 is 1;

[0040] 2) The force contention monitor calculates the absolute value of the displacement signals 305 and 306 of the two actuators by subtracting them to obtain a displacement difference 313. When the displacement difference 313 is greater than a threshold value (constant2) 314 and lasts for a certain period of time, the judgment is true and the judgment signal 315 is 1;

[0041] 3) The system monitor of the flight control system suppresses the signal 307 and performs a negation operation to obtain a signal 316, and performs an AND gate calculation on the judgment signals 312, 315 and the signal 316 of the force dispute monitor. Therefore, only when the output force difference and the displacement difference reach the corresponding thresholds and the system monitor is not suppressed, the force dispute monitor will trigger a fault signal and output a signal 317 indicating that a force dispute occurs on the control surface (the fault flag Fault Flag is set);

[0042] 4) If the displacement difference signal 313 is greater than the threshold value (constant2) 314, the displacement difference signal 313 is multiplied by the coefficient k to generate a conversion force value, and the conversion force value is compared with the output force difference value 310. If the difference between the conversion force value and the output force difference value 310 is greater than the threshold value (for example, for a certain period of time), a force sensor fault signal 318 (Pressure FaultFlag is set) is issued. The difference between the conversion force value and the output force difference value 310 greater than the threshold value may indicate that the force sensor may be faulty. In response to the force sensor fault signal 318, force dispute monitoring may be suppressed. The coefficient k may be a preconfigured constant, for example, it may be determined based on the relationship between the output force and the output displacement of the actuator. In different embodiments, different coefficients k may be used. Therefore, since the signal integrity of the position sensor (such as LVDT) is higher, the data of the force sensor can be cross-validated from different sources to detect force sensor failures.

[0043] Figure 4 4 is a flow chart of a method 400 for force contention monitoring according to one embodiment of the present invention. The method may be implemented using a flight control system or a force contention monitor as described above, or a processor, an integrated circuit, a programmable logic device, a microprocessor, a controller, a microcontroller, or a state machine, etc. The method 400 for force contention monitoring may be performed automatically or in response to an activation operation.

[0044] At step 402, a first output force and a first displacement of a first actuator may be received, wherein the first actuator drives a control surface of the aircraft via a first actuator rod, the first output force represents a first output force applied by the first actuator to the first actuator rod and the first displacement represents a first displacement of the first actuator rod. As described above, the first force sensor may be configured to detect the first output force applied by the first actuator to the first actuator rod and the first position sensor may be configured to detect the first displacement of the first actuator rod. For example, the control surface may include a spoiler, an aileron, a rudder, or an elevator.

[0045] In step 404, a second output force and a second displacement of a second actuator are received, wherein the second actuator drives the control surface through a second actuator rod, the second output force represents a second output force applied by the second actuator to the second actuator rod and the second displacement represents a second displacement of the second actuator rod. As described above, the second force sensor may be configured to detect a second output force applied by the second actuator to the second actuator rod and the second position sensor may be configured to detect a second displacement of the second actuator rod. For example, each of the first actuator rod and the second actuator rod may be connected to a piston installed in a cavity, the piston dividing the cavity into two cavities, and each of the first output force and the second output force includes a pressure difference between the two cavities separated by the corresponding piston.

[0046] In step 406, when it is determined that the output force difference between the first output force and the second output force is greater than the first threshold value and the displacement difference between the first displacement and the second displacement is greater than the second threshold value, it can be determined in step 408 that the control surface has a force dispute. In one embodiment, in response to determining that the control surface has a force dispute, an alarm signal indicating that the control surface has a force dispute can be provided. The alarm signal can be transmitted to the flight control system or other onboard electronic components, or provided to the user. For example, the flight control system or the force dispute monitor can provide an audio and / or visual indication. In an optional embodiment, in response to the control surface having a force dispute, one or both of the first actuator or the second actuator can be inhibited from driving the control surface.

[0047] At optional step 410 , the displacement difference may be multiplied by a coefficient to generate a converted force value, which is compared to the output force difference value. If the converted force value differs from the output force difference value by more than a threshold, a force sensor fault signal may be issued at step 412 .

[0048] In optional step 412, in response to the force sensor fault signal, force dispute monitoring for the control surface may be inhibited. For example, the force dispute monitor or other components that perform force dispute monitoring may stop performing force dispute monitoring.

[0049] Although Figure 4 Each step is described in a certain order, but it should be understood that these steps can be performed in different orders or some steps can be performed concurrently. For example, steps 402 and 404 can be performed in different orders or concurrently, steps 406 and 410 can be performed in different orders or concurrently, and so on.

[0050] The present invention has at least the following advantages and benefits:

[0051] 1. Provide dual-channel monitoring, which makes the monitoring more robust, reduces the probability of false triggering, and improves the efficiency of route operation;

[0052] 2. Strong adaptability. The force contention monitor can use the existing sensors or sensor signals in the flight control system without modifying the equipment. The force contention monitor algorithm can be implemented in the FCM of most aircraft.

[0053] 3. The position sensor is used to perform heterogeneous cross-verification on the force sensor data, thereby improving the integrity of the pressure signal.

[0054] It should be noted that the numerical values, threshold values, etc. given in each embodiment are only examples and are not intended to limit the scope of the present invention. In addition, as an overall technical solution, there are other components or steps that are not listed in the claims or description of the present invention. Moreover, a single name of a component does not exclude other names of the component.

[0055] In addition, it should be noted that the use of serial words such as "first" and "second" to limit components is only for the convenience of distinguishing the corresponding components. If not otherwise stated, the above words have no special meaning and therefore cannot be understood as limiting the scope of protection of this application.

[0056] The various steps and modules of the methods and devices described above can be implemented with hardware, software, or a combination thereof. If implemented in hardware, the various illustrative steps, modules, and circuits described in conjunction with the present disclosure can be implemented or executed with a general-purpose processor, a digital signal processor (DSP), an application-specific integrated circuit (ASIC), a field programmable gate array (FPGA), or other programmable logic components, hardware components, or any combination thereof. The general-purpose processor can be a processor, a microprocessor, a controller, a microcontroller, or a state machine, etc. If implemented in software, the various illustrative steps and modules described in conjunction with the present disclosure can be stored on a computer-readable medium or transmitted as one or more instructions or codes. The software modules that implement the various operations of the present disclosure can reside in a storage medium, such as RAM, flash memory, ROM, EPROM, EEPROM, registers, hard disks, removable disks, CD-ROMs, 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 corresponding program modules to implement the various steps of the present disclosure. Moreover, the software-based embodiments can be uploaded, downloaded, or remotely accessed by appropriate communication means. Such appropriate communications means include, for example, the Internet, the World Wide Web, an intranet, software applications, cable (including fiber optic cables), magnetic communications, electromagnetic communications (including RF, microwave and infrared communications), electronic communications or other such communications means.

[0057] It should also be noted that these embodiments may be described as a process depicted as a flow chart, flow diagram, structure diagram, or block diagram. Although the flow chart may describe the operations as sequential processes, many of these operations can be performed in parallel or concurrently. In addition, the order of these operations can be rearranged.

[0058] The disclosed methods, devices, and systems should not be limited in any way. On the contrary, the present disclosure covers all novel and non-obvious features and aspects of the various disclosed embodiments (alone and in various combinations and sub-combinations with each other). The disclosed methods, devices, and systems are not limited to any specific aspects or features or combinations thereof, nor do any disclosed embodiments require the existence of any one or more specific advantages or the resolution of specific or all technical problems.

[0059] The present invention is not limited to the above-mentioned specific embodiments, which are merely illustrative and not restrictive. Under the guidance of the present invention, ordinary technicians in this field can make many forms without departing from the scope of protection of the present invention and the claims, all of which belong to the protection scope of the present invention.

Claims

1. A device for monitoring aircraft force conflicts, It is characterized in that include: a first actuator and a first force sensor and a first position sensor associated with the first actuator, wherein the first actuator drives the control surface of the aircraft through a first actuator rod, the first force sensor being configured to detect a first output force applied by the first actuator to the first actuator rod and the first position sensor being configured to detect a first displacement of the first actuator rod; a second actuator and a second force sensor and a second position sensor associated with the second actuator, the second actuator driving the control surface via a second actuator rod, the second force sensor being configured to detect a second output force applied by the second actuator to the second actuator rod and the second position sensor being configured to detect a second displacement of the second actuator rod; as well as A force conflict monitor is configured to determine that a force conflict occurs on the control surface when an output force difference between the first output force and the second output force is greater than a first threshold and a displacement difference between the first displacement and the second displacement is greater than a second threshold.

2. The device for monitoring aircraft force conflicts as claimed in claim 1, Features: Each of the first actuator and the second actuator includes a cavity and a piston installed in the cavity, the first actuator rod and the second actuator rod are respectively connected to the pistons of the first actuator and the second actuator, wherein the piston divides the cavity into two chambers, and each of the first output force and the second output force includes a pressure difference between the two chambers separated by the corresponding piston.

3. The device for monitoring aircraft force conflicts as claimed in claim 1, It is characterized in that The control surfaces include spoilers, ailerons, rudders, or elevators.

4. The device for monitoring aircraft force conflicts according to claim 1, Features: The force conflict monitor is further configured to multiply the displacement difference by a coefficient to generate a conversion force value, compare the conversion force value with the output force difference value, and issue a force sensor fault signal if the conversion force value differs from the output force difference value by more than a third threshold.

5. The device for monitoring aircraft force conflicts as claimed in claim 4, Features: The force conflict monitor inhibits force conflict monitoring of the control surface in response to the force sensor fault signal.

6. The device for monitoring aircraft force conflicts as claimed in claim 1, Features: In response to a force conflict occurring on the control surface, one or both of the first actuator or the second actuator is restrained from driving the control surface.

7. A method for monitoring aircraft force contention, It is characterized in that include: Receiving a first output force and a first displacement of a first actuator, wherein the first actuator drives a control surface of the aircraft through a first actuator rod, the first output force represents a first output force applied by the first actuator to the first actuator rod and the first displacement represents a first displacement of the first actuator rod; receiving a second output force and a second displacement of a second actuator, wherein the second actuator drives the control surface through a second actuator rod, the second output force represents a second output force applied by the second actuator to the second actuator rod and the second displacement represents a second displacement of the second actuator rod; as well as When the output force difference between the first output force and the second output force is greater than a first threshold and the displacement difference between the first displacement and the second displacement is greater than a second threshold, it is determined that a force conflict occurs on the control surface.

8. The method for monitoring aircraft force conflicts according to claim 7, Features: Each of the first actuator and the second actuator includes a cavity and a piston installed in the cavity, the first actuator rod and the second actuator rod are respectively connected to the pistons of the first actuator and the second actuator, wherein the piston divides the cavity into two chambers, and each of the first output force and the second output force includes a pressure difference between the two chambers separated by the corresponding piston.

9. The method for monitoring aircraft force conflicts according to claim 7, It is characterized in that The control surfaces include spoilers, ailerons, rudders, or elevators.

10. The method for monitoring aircraft force conflicts according to claim 7, It is characterized in that Also includes: The displacement difference is multiplied by a coefficient to generate a conversion force value, the conversion force value is compared with the output force difference value, and if the conversion force value differs from the output force difference value by more than a third threshold, a force sensor fault signal is issued.

11. The method for aircraft force contention monitoring according to claim 10, It is characterized in that Also includes: Force conflict monitoring of the control surface is inhibited in response to the force sensor fault signal.

12. The method for monitoring aircraft force contention according to claim 7, It is characterized in that Also includes: In response to a force conflict occurring on the control surface, one or both of the first actuator or the second actuator is restrained from driving the control surface.

13. A flight control system, It is characterized in that include: A device for monitoring aircraft force conflicts as described in any one of claims 1 to 6.

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