Providing continuously variable feel forces for full authority flight control systems

Abstract

This application is entitled "Providing Continuous Variable Feel Forces for Full Authority Flight Control Systems." In one or more embodiments, a method for providing continuous variable feel forces for an aircraft includes sensing a force sensor value by each of at least one sensor associated with at least one aircraft control. The method further includes determining a net force value by using the force sensor value of each of the at least one sensor. In addition, the method includes comparing the net force value to a desired breakout force. Further, the method includes determining whether the net force value exceeds the desired breakout force. In addition, the method includes determining an adjusted force value by using the desired breakout force and the net force value when the net force value exceeds the desired breakout force. Further, the method includes determining an actuator torque command based on the adjusted force value. In addition, the method includes commanding an autopilot actuator to apply a torque with the actuator torque command.

Description

Technical Field

[0001] This disclosure relates to sensing power for flight control systems. Specifically, this disclosure relates to providing continuously variable sensing power for all-powered flight control systems. Background Technology

[0002] During certain phases of flight (e.g., takeoff), pilots desire reduced feel force on aircraft controls such as cockpit sticks and / or rudder pedals. Providing pilots with variable feel force in aircraft with fully powered flight controls has historically involved considerable complexity. For example, in older applications, variable feel force was achieved through mechanical means (e.g., pneumatic and / or hydraulic means) that altered spring geometry. In contemporary applications, variable feel force has been achieved by electronically mimicking the behavior of springs through motor control. Both solutions require complexity accompanied by undesirable costs.

[0003] In view of the above, an improved design is needed to provide variable tactile force for aircraft controls. Summary of the Invention

[0004] This disclosure relates to methods, systems, and apparatus for providing continuously variable sensed force for a fully powered flight control system. In one or more embodiments, a method for providing continuously variable sensed force for an aircraft includes sensing force sensor values ​​by each of at least one sensor associated with at least one aircraft control. The method further includes determining a net force value by at least one processor using the force sensor values ​​of each of the at least one sensor. Furthermore, the method includes comparing the net force value with a desired start-up force by at least one processor. Furthermore, the method includes determining whether the net force value exceeds the desired start-up force by at least one processor. Furthermore, the method includes determining an adjusted force value by at least one processor using the desired start-up force and the net force value when the net force value exceeds the desired start-up force. Furthermore, the method includes determining an actuator torque command by at least one processor based on the adjusted force value. Furthermore, the method includes commanding an autopilot actuator to apply torque by at least one processor using the actuator torque command.

[0005] In at least one embodiment, at least two sensors are present. In one or more embodiments, the sensors are associated with the aircraft's control stick. In some embodiments, at least one of the sensors is associated with the right rudder pedal, and another of at least one of the sensors is associated with the left rudder pedal.

[0006] In one or more embodiments, the method further includes determining, by at least one processor, whether an engine failure exists on the aircraft. In at least one embodiment, the method further includes determining a thrust asymmetry offset value when an engine failure exists. In some embodiments, the method further includes determining a compensated pedal force value by at least one processor based on an adjusted force value and the thrust asymmetry offset value.

[0007] In one or more embodiments, the starting force is expected to be predetermined.

[0008] In at least one embodiment, the method further includes determining, by at least one processor, whether it is desirable to reduce the sensed force of at least one aircraft control associated with each of at least one of the at least one sensors. In one or more embodiments, the method further includes engaging an autopilot actuator in a torque command mode when it is desirable to reduce the sensed force. In some embodiments, at least one processor determines whether it is desirable to reduce the sensed force based on the flight phase of the aircraft.

[0009] In one or more embodiments, the autopilot actuator is a pedal autopilot actuator or a lever autopilot actuator.

[0010] In at least one embodiment, a system for providing continuously variable sensed force to an aircraft includes at least one aircraft control. The system further includes at least one sensor associated with each of the at least one aircraft control. In one or more embodiments, each of the at least one sensor is configured to sense a force sensor value. Furthermore, the system includes at least one processor configured to: determine a net force value using the force sensor values ​​of each of the at least one sensor; compare the net force value with a desired starting force; determine whether the net force value exceeds the desired starting force; when the net force value exceeds the desired starting force, determine an adjusted force value using the desired starting force and the net force value; determine an actuator torque command based on the adjusted force value; and command an autopilot actuator to apply torque using the actuator torque command.

[0011] In one or more embodiments, at least one processor is further configured to determine whether an engine failure exists on the aircraft. In at least one embodiment, at least one processor is further configured to determine a thrust asymmetry offset value when an engine failure exists. In some embodiments, at least one processor is further configured to determine a compensated pedal force value based on an adjusted force value and the thrust asymmetry offset value.

[0012] These features, functions, and advantages may be implemented independently in the various embodiments of this disclosure or may be combined in other embodiments. Attached Figure Description

[0013] These and other features, aspects, and advantages of this disclosure will be better understood from the following description, the appended claims, and the accompanying drawings, wherein:

[0014] Figure 1 This is an illustration of an exemplary cockpit of an aircraft according to at least one embodiment of the present disclosure, which may employ the disclosed system for providing continuously variable sensing forces for an aircraft with a fully powered flight control system.

[0015] Figure 2 This is a schematic diagram illustrating a disclosed system for providing continuously variable sensing force for an aircraft with a fully powered flight control system, according to at least one embodiment of the present disclosure.

[0016] Figure 3 This is a flowchart illustrating a disclosed method for providing continuously variable feel force to a control stick of an aircraft having a fully powered flight control system, according to at least one embodiment of the present disclosure.

[0017] Figure 4 This is a flowchart illustrating a disclosed method for providing continuously variable feel force to the rudder pedal of an aircraft having a fully powered flight control system, according to at least one embodiment of the present disclosure.

[0018] Figure 5 This is a graph illustrating the exemplary rudder pedal force according to at least one embodiment of the present disclosure.

[0019] Figure 6 A block diagram of an illustrative computing system suitable for implementing embodiments of the present disclosure is shown, according to at least one embodiment of the present disclosure. Detailed Implementation

[0020] The methods and apparatus disclosed herein provide an operating system for providing continuously variable sense force to a full-powered flight control system. In one or more embodiments, the system of this disclosure provides variable sense force to aircraft controls by using existing conventional hardware (e.g., conventional autopilot reverse-drive actuators) currently present on many aircraft having full-powered flight control systems.

[0021] As mentioned above, during certain phases of flight (e.g., takeoff), pilots expect to reduce the perceived force of aircraft controls (e.g., control yokes and / or rudder pedals). Providing pilots with variable "feel" force to aircraft with fully powered flight controls has historically involved considerable complexity. For example, in older applications, variable feel force was achieved through mechanical means (e.g., pneumatic and / or hydraulic means) that altered spring geometry. In contemporary applications, variable feel force has been achieved by electronically simulating the behavior of springs through motor control. Both solutions require complexity accompanied by undesirable costs.

[0022] The disclosed system employs a flight control system comprising a conventional simple mechanical spring "feel" system and a conventional autopilot inverse actuator, with appropriately added force sensors and control logic to provide the same functionality (e.g., providing variable "feel" force) as the more complex "variable feel" solutions described above (e.g., mechanical and electronic "variable feel" solutions). During operation of the disclosed system, the autopilot inverse actuator operates in a new torque command mode and is commanded proportionally to the applied pilot force measured (sensed) by the force sensors, effectively transforming the existing autopilot inverse actuator into a boost servo that can be effectively subtracted from the "feel" force provided by the conventional simple mechanical spring "feel" system.

[0023] Implementing the disclosed system to provide pilots with variable "feel" forces in aircraft with full-powered flight control does not require a significant increase in cost compared to conventional, simple mechanical spring "feel" systems. Instead of using a more complex "feel" solution, providing pilots with variable "feel" forces by repurposed existing hardware (e.g., autopilot reverse actuators) reduces the complexity, cost, and weight of the aircraft. The disclosed system allows for reduced complexity (and therefore cost) in terms of mechanical moving parts and complex closed-loop electronic control systems.

[0024] In the following description, numerous details are set forth to provide a more comprehensive description of the system. However, it will be apparent to those skilled in the art that the disclosed system can be practiced without these specific details. In other instances, well-known features have not been described in detail to avoid unnecessarily obscuring the system.

[0025] Embodiments of this disclosure are described herein based on functional and / or logical components and various processing steps. It should be understood that such components can be implemented by any number of hardware, software, and / or firmware components configured to perform specified functions. For example, embodiments of this disclosure may employ various integrated circuit components (e.g., memory elements, digital signal processing elements, logic elements, lookup tables, etc.) that can perform multiple functions under the control of one or more processors, microprocessors, or other control devices. Furthermore, those skilled in the art will understand that embodiments of this disclosure can be implemented in combination with other components, and the systems described herein are merely exemplary embodiments of this disclosure.

[0026] For the sake of brevity, conventional technologies and components related to other functional aspects of the aircraft controls and systems (as well as the various operational components of the systems) may not be described in detail herein. Furthermore, the connecting lines shown in the various figures included herein are intended to illustrate exemplary functional relationships and / or physical couplings between the various elements. It should be noted that many alternative or additional functional relationships or physical connections may exist in one or more embodiments of this disclosure.

[0027] Figure 1 This is a figure 100 illustrating an exemplary cockpit 120 of an aircraft 110 according to at least one embodiment of the present disclosure, which may employ a disclosed system for providing continuously variable feel forces for an aircraft with a fully powered flight control system. In this figure, the cockpit 120 of the aircraft 110 is shown as including a plurality of aircraft controls operated by two pilots. The aircraft controls of the cockpit 120 include sticks 130a, 130b and rudder pedals 140a, 140b, 140c, 140d. Specifically, the aircraft controls for the first pilot include stick 130a, left rudder pedal 140a, and right rudder pedal 140b. Furthermore, the aircraft controls for the second pilot include stick 130b, left rudder pedal 140c, and right rudder pedal 140d.

[0028] It should be noted that, in one or more embodiments, the cockpit 120 of the aircraft 110 may include a larger... Figure 1 Showing more or fewer aircraft controls, and may include with Figure 1 The aircraft controls shown (e.g., joysticks 130a, 130b and rudder pedals 140a, 140b, 140c, 140d) are of different types. Furthermore, in one or more embodiments, the aircraft 110 can be any type of aircraft with a fully powered flight control system, including but not limited to civil and military aircraft.

[0029] During flight of aircraft 110, the pilot uses control sticks 130a and 130b to control the attitude of aircraft 110, typically both pitch and roll. When control sticks 130a and 130b are pulled back by the pilot (i.e., pulled towards the pilot's chest), the nose of aircraft 110 rises. Conversely, when control sticks 130a and 130b are pushed forward by the pilot (i.e., pushed away from the pilot's chest), the nose of aircraft 110 descends.

[0030] When the pilot pulls back control sticks 130a and 130b to raise the nose of the aircraft 110, a conventional simple mechanical spring mechanism connected to control sticks 130a and 130b will provide the pilot with an artificial drag "feel" force. The farther the pilot pulls back control sticks 130a and 130b, the greater the drag "feel" force the spring mechanism will provide to the pilot.

[0031] Similarly, during flight of aircraft 110, the pilot uses rudder pedals 140a, 140b, 140c, and 140d to control the rudder on the vertical stabilizer of aircraft 110 to control the yaw of aircraft 110. Pressing down (e.g., pushed down by the pilot) the left rudder pedals 140a and 140c will move the rudder to the left, causing the nose of aircraft 110 to turn left. Similarly, pressing down (e.g., pushed down by the pilot) the right rudder pedals 140b and 140d will move the rudder to the right, causing the nose of aircraft 110 to turn right.

[0032] When the pilot depresses rudder pedals 140a, 140b, 140c, and 140d (i.e., pushes rudder pedals 140a, 140b, 140c, and 140d downwards), a conventional, simple mechanical spring mechanism connected to rudder pedals 140a, 140c, and 140d provides the pilot with a sense of drag. The more the pilot depresses rudder pedals 140a, 140b, 140c, and 140d, the greater the sense of drag provided by the spring mechanism.

[0033] Figure 2 This is a schematic diagram illustrating a disclosed system 200 for providing continuously variable feel force to an aircraft having a full-power flight control system, according to at least one embodiment of the present disclosure. In this figure, system 200 is shown as an aircraft 110 (reference 110) including two pilots (e.g., a first pilot and a second pilot). Figure 1 The aircraft controls include a joystick 130a, a left rudder pedal 140a, and a right rudder pedal 140b for the first pilot. For the second pilot, the aircraft controls include a joystick 130b, a left rudder pedal 140c, and a right rudder pedal 140d. It should be noted that in one or more embodiments, system 200 may include... Figure 2 Showing more or fewer aircraft controls.

[0034] Also in this figure, system 200 is shown as including sensors (e.g., force sensors) 230a, 230b, 230c, 230d, 235a, 235b respectively connected (e.g., communicatively coupled) to aircraft controls (e.g., sticks 130a, 130b; left rudder pedals 140a, 140c; and right rudder pedals 140b, 140d). Specifically, sensor 230a is connected to the left rudder pedal 140a of the first pilot, sensor 230b is connected to the right rudder pedal 140b of the first pilot, sensor 235a is connected to the stick 130a of the first pilot, sensor 230c is connected to the left rudder pedal 140c of the second pilot, sensor 230d is connected to the right rudder pedal 140d of the second pilot, and sensor 235b is connected to the stick 130b of the second pilot. It should be noted that, compared to... Figure 2 In one or more embodiments, system 200 may include more than one sensor connected to each aircraft control.

[0035] also, Figure 2 The system 200 is shown as including one or more processors 210 connected (e.g., communicatively coupled) to sensors 230a, 230b, 230c, 230d, 235a, 235b. The processors 210 are also connected (e.g., communicatively coupled) to the main flight controls 240 of the aircraft 110, as well as pedal autopilot actuators (e.g., conventional autopilot reverse actuators) 220a and stick autopilot actuators 220b (e.g., conventional autopilot reverse actuators).

[0036] An autopilot actuator (e.g., pedal autopilot actuator 220a and stick autopilot actuator 220b) is a device that physically moves aircraft controls (e.g., sticks 130a, 130b or rudder pedals 140a, 140b, 140c, 140d) when the aircraft 110 is operating in autopilot mode. In older applications, when the aircraft 110 is operating in autopilot mode, the autopilot actuator (e.g., pedal autopilot actuator 220a and stick autopilot actuator 220b) physically moves the aircraft controls to control the flight of the aircraft 110. In newer applications, in autopilot mode, the flight of the aircraft 110 is not controlled by the movement of the aircraft controls, but rather by the flight software. For these newer applications, autopilot actuators (e.g., pedal autopilot actuator 220a and stick autopilot actuator 220b) move aircraft controls based on the flight movement of aircraft 110 to provide the pilot with a reference to the movement of aircraft 110. Many aircraft with fully powered flight control systems include conventional autopilot actuators (e.g., pedal autopilot actuator 220a and stick autopilot actuator 220b). In one or more embodiments, compared to Figure 2 The system 200 may include more than two autopilot actuators (e.g., pedal autopilot actuator 220a and lever autopilot actuator 220b).

[0037] Currently, when the aircraft 110 operates in autopilot mode (e.g., the aircraft 110 is controlled by autopilot software), the conventional autopilot actuators installed on the aircraft 110 (e.g., pedal autopilot actuator 220a and stick autopilot actuator 220b) are configured to operate only in one mode (which is autopilot mode). However, for the disclosed system 200, during pilot-controlled flight of the aircraft 110 (i.e., not in autopilot mode), when the processor 210 commands this, the autopilot actuators (i.e., pedal autopilot actuator 220a and stick autopilot actuator 220b) are also configured to operate in a second mode (which is torque command mode).

[0038] Figure 2 System 200 operation in Figure 3 and Figure 4 This is discussed in the description. In particular, Figure 3 This illustrates at least one embodiment of a method for providing a fully powered flight control system for an aircraft (e.g., Figure 1 The control stick of the aircraft 110 (e.g., Figure 1The flowchart of the disclosed method 300 shows that the control sticks 130a and 130b provide continuously variable tactile feedback. It should be noted that during the disclosed method 300, the aircraft 110 is controlled (flying) by a pilot (i.e., the aircraft 110 is not flying in autopilot mode (e.g., by software flight)).

[0039] In this figure, at the beginning 305 of the operation method 300 of system 200, at least one processor (e.g., Figure 2 The processor 210) determines whether it is desired to reduce the lever (e.g., Figure 1 The sensor force (e.g., commanded by control rules) 310 of the control sticks 130a and 130b. Based on the flight phase of the aircraft 110 (e.g., during takeoff), it may be desirable to reduce the aircraft controls (e.g., Figure 1 The feel force of the control sticks 130a and 130b. For example, during takeoff of the aircraft 110, as the pilot pulls back the control sticks 130a and 130b (i.e., pulls the control sticks 130a and 130b toward the pilot's chest) to raise the nose of the aircraft 110, it may be desirable to provide a lower artificial drag "feel" force on the control sticks 130a and 130b.

[0040] If processor 210 determines that it desires to reduce the tactile force of control levers 130a and 130b, then processor 210 will send a signal to the autopilot actuator (e.g., Figure 2 The autopilot actuator 220b sends a command to engage the autopilot actuator 220b in a torque command mode, which is different from the autopilot mode in which the autopilot actuator 220b is operated. After receiving the command to engage in the torque command mode, the autopilot actuator 220b engages itself in the torque command mode 315.

[0041] Then, with the control lever (e.g., Figure 1 At least one sensor associated with each of the control levers 130a and 130b (e.g., Figure 2 Sensors 235a and 235b will sense (receive) the force applied by the pilot to the control sticks 130a and 130b (e.g., the magnitude of the force by which the pilot pulls back the control sticks 130a and 130b). Sensors 235a and 235b associated with each of the control sticks 130a and 130b will then send a force sensor value corresponding to the magnitude of the sensed force (e.g., a force of 10 pounds (lb)). For example, the force sensor value depends on how much force the pilot pulls back the control sticks 130a and 130b. Processor 210 will then send the force sensor value corresponding to the force applied by the pilot to the control sticks 130a and 130b (e.g., the magnitude of the force by which the pilot pulls back the control sticks 130a and 130b). Figure 1Each sensor 235a, 235b associated with the control levers 130a and 130b receives a force sensor value 320.

[0042] After receiving force sensor values ​​from each sensor 235a, 235b, the processor 210 determines a net force value 325 for the control sticks 130a, 130b using the force sensor values. This net force value is the force applied by the pilot to the control sticks (e.g., Figure 1 The total amount of force on control levers 130a and 130b.

[0043] After determining the net force value, processor 210 compares the net force value (e.g., 10 lb of force) with the desired breakout force (e.g., 7 lb of force) of the control levers (e.g., control levers 130a, 130b) 330. In one or more embodiments, the desired breakout force is a predetermined force applied to control levers 130b, 130b. Processor 210 then determines whether the net force value (e.g., 10 lb of force) exceeds the desired breakout force (e.g., 7 lb of force) 335.

[0044] If processor 210 determines that the net force value does not exceed the expected starting force, the method proceeds to step 320. However, if processor 210 determines that the net force value does exceed the expected starting force, processor 210 determines an adjusted force value by using the expected starting force and the net force value. In one or more embodiments, processor 210 determines the adjusted force value (e.g., 3 lb force) 340 by subtracting the expected starting force (e.g., 7 lb force) from the net force value (e.g., 10 lb force).

[0045] After processor 210 determines the adjusted force value (e.g., a force of 3 lb), processor 210 calculates and generates an actuator torque command 345 based on the adjusted force value (e.g., proportional to it). Processor 210 then sends the actuator torque command to the autopilot actuator (e.g., Figure 2 The autopilot actuator 220b commands the autopilot actuator 220b to apply a force proportional to an adjusted force value (e.g., 3 lb) to the control sticks 130a and 130b, thereby reducing the pilot's manual drag "feel" on the control sticks 130a and 130b. Once the autopilot actuator 220b receives an actuator torque command, the autopilot actuator 220b applies a force proportional to the adjusted force value (e.g., 3 lb) to the control sticks 130a and 130b.

[0046] The processor then determines whether autopilot actuator 220b should stop (end) applying a force proportional to the adjusted force value to the control sticks (e.g., control sticks 130a, 130b) 350. For example, processor 210 may determine whether autopilot actuator 220b should stop applying the force based on whether the flight phase has ended (e.g., the aircraft 110 is no longer in the takeoff phase of flight and has reached its cruising altitude). If processor 210 determines that autopilot actuator 220b should not stop applying the force, method 300 returns to step 310 to recalculate and confirm the force to be applied. However, if processor 210 determines that autopilot actuator 220b should stop applying the force, method 300 ends 355.

[0047] However, if processor 210 determines that a reduction in the feel force of control levers 130a and 130b is not desired, autopilot actuator 220b will remain disengaged, and the artificial "feel" force will be simply provided to control levers 130a and 130b via the unreinforced spring mechanism of aircraft 110 and 360. Method 300 then proceeds to step 350.

[0048] Figure 4 This illustrates at least one embodiment of a method for providing a fully powered flight control system for an aircraft (e.g., Figure 1 The rudder pedals (e.g., of the aircraft 110) Figure 1 The flowchart of the disclosed method 400 shows that the rudder pedals 140a, 140b, 140c, 140d provide continuously variable tactile feedback. It should be noted that during the disclosed method 400, the aircraft 110 is controlled (flying) by a pilot (i.e., the aircraft 110 is not flying in autopilot mode (e.g., by software flight)).

[0049] In this figure, at the start 405 of the operation method 400 of system 200, at least one processor (e.g., Figure 2 The processor 210) determines whether it is desired to reduce the rudder pedal (e.g., Figure 1 The tactile feedback of the rudder pedals 140a, 140b, 140c, 140d (e.g., commanded by control rules) 410. Based on the flight phase of the aircraft 110 (e.g., during landing), the aircraft controls (e.g., Figure 1 The feel force of the rudder pedals 140a, 140b, 140c, 140d may be desirable. For example, as the pilot lands the aircraft 110, it may be desirable to provide a lower artificial drag "feel" force to the rudder pedals 140a, 140b, 140c, 140d.

[0050] If processor 210 determines that it is desirable to reduce the rudder pedal (e.g., Figure 1 If the steering pedals 140a, 140b, 140c, and 140d are felt, the processor 210 will send a command to the autopilot actuator (e.g., Figure 2 The autopilot actuator 220a is commanded to engage in a torque command mode, which is different from the autopilot mode in which the autopilot actuator 220a is operated. After receiving the command to engage in the torque command mode, the autopilot actuator 220a engages itself in the torque command mode 415.

[0051] Then, with the rudder pedals (e.g., Figure 1 At least one sensor associated with the steering pedals 140a, 140b, 140c, 140d (e.g., Figure 2 Sensors 230a, 230b, 230c, and 230d will sense (receive) the force exerted by the pilot on each of the rudder pedals 140a, 140b, 140c, and 140d (e.g., the magnitude of the force by which the pilot presses down on each of the rudder pedals). Each of the sensors 230a, 230b, 230c, and 230d associated with the rudder pedals 140a, 140b, 140c, and 140d will then send a force sensor value corresponding to the magnitude of the sensed force (e.g., force in pounds) to processor 210. For example, the force sensor value depends on how much force the pilot exerts when pressing down on the rudder pedals 140a, 140b, 140c, and 140d. In one or more embodiments, the force sensor values ​​for the forces sensed on the left rudder pedal (e.g., rudder pedals 140a and 140c) are positive, and the force sensor values ​​for the forces sensed on the right rudder pedal (e.g., rudder pedals 140b and 140d) are negative. The processor 210 receives force sensor values ​​420 from each sensor 230a, 230b, 230c, 230d associated with each rudder pedal (e.g., rudder pedals 140a, 140b, 140c, 140d).

[0052] After receiving force sensor values ​​from each of the sensors 230a, 230b, 230c, and 230d, the processor 210 determines the net force value (e.g., 30 lb force) 425 of the rudder pedals (e.g., rudder pedals 140a, 140b, 140c, and 140d) using the force sensor values. The net force value is the total amount of force exerted by the pilot on the rudder pedals (e.g., rudder pedals 140a, 140b, 140c, and 140d) (e.g., force sensor values ​​are summed together).

[0053] After processor 210 determines the net force value, it compares the net force value (e.g., 30 lb of force) with the desired starting force (e.g., 22 lb of force) of the rudder pedals (e.g., rudder pedals 140a, 140b, 140c, 140d) 430. In one or more embodiments, the desired starting force is a predetermined force applied to the rudder pedals (e.g., rudder pedals 140a, 140b, 140c, 140d). Processor 210 then determines whether the net force value (e.g., 30 lb of force) exceeds the desired starting force (e.g., 22 lb of force) 435.

[0054] If processor 210 determines that the net force value does not exceed the expected starting force, the method proceeds to step 420. However, if processor 210 determines that the net force value does exceed the expected starting force, processor 210 determines an adjusted force value (e.g., an adjusted pedal force value) by using the expected starting force and the net force value. In one or more embodiments, processor 210 determines the adjusted force value (e.g., 8 lb force) 440 by subtracting the expected starting force (e.g., 22 lb force) from the net force value (e.g., 30 lb force).

[0055] After processor 210 determines the adjusted force value (e.g., 8 lb force), processor 210 optionally determines whether an engine malfunction (e.g., left engine malfunction or right engine malfunction) 445 exists on aircraft 110. If processor 210 determines that there is no engine malfunction on aircraft 110, processor 210 calculates and generates an actuator torque command 455 based on the adjusted force value (proportional to it). Processor 210 then sends the actuator torque command to autopilot actuator 220a to instruct autopilot actuator 220a to apply a force proportional to the adjusted force value (e.g., 8 lb force) to the rudder pedals (e.g., rudder pedals 140a, 140b, 140c, 140d), thereby reducing the pilot's artificial drag "feel" force on the rudder pedals 140a, 140b, 140c, 140d. Once the autopilot actuator 220a receives the actuator torque command, the autopilot actuator 220a applies a force proportional to the adjusted force value (e.g., 8 lb force) to the rudder pedals (e.g., rudder pedals 140a, 140b, 140c, 140d).

[0056] However, if processor 210 determines that there is an engine failure on aircraft 110, processor 210 will determine a compensated pedal force value 450 by adding a thrust asymmetry offset value to an adjusted force value (e.g., a force of 8 lb). Processor 210 determines the thrust asymmetry offset value based on the engine failure. For example, if there is a right-side engine failure, aircraft 110 will have more thrust on the left side of aircraft 110, and aircraft 110 will pull to the right during flight. Conversely, if there is a left-side engine failure, aircraft 110 will have greater thrust on the right side of aircraft 110, and aircraft 110 will pull to the left during flight. The thrust asymmetry offset value (e.g., the magnitude of a force in lbs) helps to counteract the asymmetry of aircraft 110 during flight (e.g., preventing aircraft 110 from being pulled to either side). For a right-side engine failure, the thrust asymmetry offset will be positive (e.g., 5 lb of force), while for a left-side engine failure, the thrust offset will be negative (e.g., -5 lb of force).

[0057] For this example, we will assume a right-side engine failure and that processor 210 determines a thrust offset asymmetry value of +5 lb. After processor 210 determines the thrust offset asymmetry value (e.g., +5 lb), processor 210 calculates and generates actuator torque command 455 based on a compensated pedal force value (proportionate to it). It should be noted that the compensated pedal force value (e.g., 13 lb) will be equal to the adjusted force value (e.g., 8 lb) plus the thrust offset asymmetry value (e.g., +5 lb). For this example, due to the right-side engine failure, a force proportional to the compensated pedal force value (e.g., 13 lb) will be applied to rudder pedals 140a, 140b, 140c, and 140d.

[0058] The processor 210 then sends an actuator torque command to the autopilot actuator 220a to instruct the autopilot actuator 220a to apply a force proportional to a compensated pedal force value (e.g., 13 lb) to the rudder pedals 140a, 140b, 140c, and 140d, thereby reducing the pilot's artificial drag "feel" on the left rudder pedals 140a and 140c. Once the autopilot actuator 220a receives the actuator torque command, the autopilot actuator 220 will apply a force proportional to the compensated pedal force value (e.g., 13 lb) to the rudder pedals 140a, 140b, 140c, and 140d.

[0059] The processor then determines whether autopilot actuator 220a should stop (end) applying a force proportional to the adjusted force value to the rudder pedals (e.g., rudder pedals 140a, 140b, 140c, 140d) 460. For example, processor 210 may determine whether autopilot actuator 220a should stop applying force based on whether the flight phase has ended (e.g., the aircraft 110 is no longer in the landing phase of flight and has landed on the runway). If processor 210 determines that autopilot actuator 220a should not stop applying force, method 400 returns to step 410 to recalculate and confirm the force to be applied. However, if processor 210 determines that autopilot actuator 220a should stop applying force, method 400 ends 465.

[0060] However, if processor 210 determines that it does not wish to reduce the feel force of the rudder pedals (e.g., rudder pedals 140a, 140b, 140c, 140d), then autopilot actuator 220a will remain disengaged, and the artificial "feel" force will be simply provided to the rudder pedals 140a, 140b, 140c, 140d, 470 via the unreinforced spring mechanism of aircraft 110. Method 400 then proceeds to step 460.

[0061] Figure 5 This is a graph 500 illustrating exemplary rudder pedal forces according to at least one embodiment of the present disclosure. The x-axis of graph 500 represents rudder pedal displacement in inches, and the y-axis of graph 500 represents rudder pedal force in pounds. Curve 510 on graph 500 represents the actual rudder pedal force (e.g., pedals 140a, 140b, 140c, or 140d) generated by a conventional spring mechanism of the rudder pedals (e.g., rudder pedals 140a, 140b, 140c, or 140d) attached to the aircraft 110. Curve 520 on graph 500 represents the force felt by the pilot assisted by autopilot actuator 220a (by subtracting 25%) when operating in torque command mode. And, curve 530 on graph 500 represents the force felt by the pilot assisted by autopilot actuator 220a (by subtracting 50%) when operating in torque command mode. It should be noted that although the graph 500 is for the rudder pedals (e.g., rudder pedals 140a, 140b, 140c or 140d), the graph for the control levers (e.g., control levers 130a or 130b) will have a similar curve.

[0062] Curves 510, 520, and 530 on graph 500 show that when the autopilot actuator 220a is operating in torque command mode, the rudder pedal will travel the same distance with less force required from the pilot. For example, curve 510 shows that the pilot needs to apply approximately 38 lb of force to move the rudder pedal by one inch, curve 520 shows that the pilot only needs to apply approximately 34 lb of force to move the rudder pedal by one inch, and curve 530 shows that the pilot only needs to apply approximately 30 lb of force to move the rudder pedal by one inch.

[0063] Furthermore, curves 510 and 530 on graph 500 show that when the autopilot actuator 220a operates in torque command mode, the rudder pedal will travel a greater distance when the same amount of force is applied. For example, curve 510 shows that when the pilot applies a force of 30 lb to the rudder pedal, the rudder pedal will only travel 0.4 inches. And curve 530 shows that when the pilot applies the same 30 lb force to the rudder pedal, the rudder pedal will travel 1 inch.

[0064] Figure 6 A block diagram illustrating an illustrative computing system 600 suitable for implementing embodiments of the present disclosure is shown, according to at least one embodiment of the present disclosure. For example, the disclosed system (e.g., Figure 2 At least one processor of the system 200) (e.g., Figure 2 The processor 210 may include and / or employ at least a portion of the disclosed computer system 600. The computing system 600 includes a bus 606 or other communication mechanism for transmitting information, interconnecting subsystems and devices such as a processor 607, system memory 608 (e.g., random access memory (RAM)), static storage device 609 (e.g., read-only memory (ROM)), disk drive 610 (e.g., magnetic or optical), communication interface 614 (e.g., modem or Ethernet card), display 611 (e.g., cathode ray tube (CRT) or liquid crystal display (LCD)), input device 612 (e.g., keyboard), and cursor control (not shown).

[0065] According to one embodiment of this disclosure, computing system 600 performs a specific operation by executing one or more sequences of one or more instructions contained in system memory 608 via processor 607. Such instructions may be read into system memory 608 from another computer-readable / usable medium, such as static storage device 609 or disk drive 610. In alternative embodiments, hardwired circuitry may be used in place of or in combination with software instructions to implement this disclosure. Therefore, embodiments of this disclosure are not limited to any particular combination of hardware circuitry and / or software. In one embodiment, the term "logic" should refer to any combination of software or hardware used to implement all or part of this disclosure.

[0066] As used herein, the terms "computer-readable medium" or "computer-usable medium" refer to any medium that participates in providing instructions to processor 607 for execution. Such media can take many forms, including but not limited to non-volatile media, volatile media, and transmission media. Non-volatile media include, for example, optical discs or magnetic disks, such as disk drive 610. Volatile media include dynamic memory, such as system memory 608.

[0067] Common forms of computer-readable media include, for example, floppy disks, floppy hard disks, magnetic tapes, any other magnetic media, optical disc read-only memory (CD-ROM), any other optical media, punched cards, paper tapes, any other physical media with a perforated pattern, RAM, programmable read-only memory (PROM), erasable programmable read-only memory (EPROM), FLASH-EPROM, any other memory chip or cassette tape, or any other media that a computer can read.

[0068] In embodiments of this disclosure, the execution of the instruction sequence of this disclosure is performed by a single computing system 600. According to other embodiments of this disclosure, two or more computing networks 600 coupled by a communication link 615 (e.g., a local area network (LAN), a public switched telephone network (PTSN), or a wireless network) can coordinate with each other to execute the instruction sequence required to practice this disclosure.

[0069] The computing system 600 can send and receive messages, data, and instructions, including programs (i.e., application code), via communication link 615 and communication interface 614. Once the program code is received, it can be executed by processor 607 and / or stored in disk drive 610 or other non-volatile memory for later execution. The computing system 600 can also interact with database 632 within database system 631 via data interface 633, whereby the computing system 600 can store electronic design information or data in database system 631 and retrieve electronic design information or data from database system 631.

[0070] Although specific embodiments have been shown and described, it should be understood that the above discussion is not intended to limit the scope of these embodiments. While embodiments and variations of many aspects of the invention have been disclosed and described herein, such disclosure is provided for illustrative purposes only. Therefore, various changes and modifications can be made without departing from the scope of the claims.

[0071] In cases where the above method indicates specific events occurring in a particular order, those skilled in the art who benefit from this disclosure will recognize that the order can be modified and such modifications are variations of this disclosure. Furthermore, some parts of the method can be executed simultaneously as a parallel process where possible, or they can be executed sequentially. Additionally, more or fewer steps of the method can be performed.

[0072] Therefore, these embodiments are intended to illustrate alternatives, modifications, and equivalents that may fall within the scope of the claims.

[0073] Furthermore, this disclosure includes embodiments as described in the following terms:

[0074] Clause 1. A method for providing continuously variable sensing force to an aircraft 110, the method comprising:

[0075] Force sensor values ​​are sensed by each of at least one ...

[0076] The net force value is determined by at least one processor 210 using the force sensor value of each of at least one of at least one of sensors 230a, 230b, 230c, 230d, 235a, 235b;

[0077] At least one processor 210 compares the net force value with the desired starting force.

[0078] At least one processor 210 determines whether the net force value exceeds the expected starting force;

[0079] When the net force value exceeds the expected starting force, at least one processor 210 determines the adjusted force value by using the expected starting force and the net force value;

[0080] The actuator torque command is determined by at least one processor 210 based on the adjusted force value; and

[0081] At least one processor 210 commands the autopilot actuators 220a, 220b to apply torque using actuator torque commands.

[0082] Clause 2. The method according to Clause 1, wherein at least two of at least one of sensors 230a, 230b, 230c, 230d, 235a, and 235b are present.

[0083] Clause 3. The method described in Clause 2, wherein at least two sensors 235a, 235b are associated with control sticks 130a, 130b of the aircraft 110.

[0084] Clause 4. The method according to Clause 2, wherein at least one of sensors 230b, 230d is associated with right rudder pedals 140b, 140d, and another sensor of at least one of sensors 230a, 230c is associated with left rudder pedals 140a, 140c.

[0085] Clause 5. The method according to Clause 4, wherein the method further includes determining by at least one processor 210 whether an engine malfunction exists on the aircraft 110.

[0086] Clause 6. The method according to Clause 5, wherein the method further includes determining a thrust asymmetry offset value by at least one processor 210 when an engine failure is present.

[0087] Clause 7. The method according to Clause 6, wherein the method further comprises determining a compensated pedal force value by at least one processor 210 based on an adjusted force value and a thrust asymmetry offset value.

[0088] Clause 8. The method described in Clause 1, wherein the expected starting force is predetermined.

[0089] Clause 9. The method according to Clause 1, wherein the method further comprises determining by at least one processor 210 whether it is desired to reduce the sensing force of at least one aircraft control 130a, 130b, 140a, 140b, 140c, 140d associated with each of at least one of at least one sensor 230a, 230b, 230c, 230d, 235a, 235b.

[0090] Clause 10. The method according to Clause 9, wherein the method further comprises engaging autopilot actuators 220a, 220b in torque command mode when a reduction in perceived force is desired.

[0091] Clause 11. The method according to Clause 9, wherein at least one processor 210 determines whether a reduction in perceived force is desired based on the flight phase of the aircraft 110.

[0092] Clause 12. The method according to Clause 1, wherein autopilot actuators 220a, 220b are one of pedal autopilot actuators or lever autopilot actuators.

[0093] Clause 13. A system for providing continuously variable sensing force to an aircraft 110, the system comprising:

[0094] At least one aircraft control 130a, 130b, 140a, 140b, 140c, 140d;

[0095] At least one sensor 230a, 230b, 230c, 230d, 235a, 235b associated with each of at least one aircraft control 130a, 130b, 140a, 140b, 140c, 140d, wherein each of at least one sensor 230a, 230b, 230c, 230d, 235a, 235b is configured to sense force sensor values; and

[0096] At least one processor 210 is configured to:

[0097] The net force value is determined by using the force sensor values ​​of at least one of the sensors 230a, 230b, 230c, 230d, 235a, and 235b.

[0098] Compare the net starting force with the expected starting force.

[0099] Determine if the net force value exceeds the expected starting force.

[0100] When the net force exceeds the expected starting force, the adjusted force is determined by using the expected starting force and the net force.

[0101] The actuator torque command is determined based on the adjusted force value, and

[0102] Use actuator torque commands to instruct autopilot actuators 220a and 220b to apply torque.

[0103] Clause 14. The system according to Clause 13, wherein at least two of at least one of sensors 230a, 230b, 230c, 230d, 235a, and 235b are present.

[0104] Clause 15. The system described in Clause 14, wherein at least two sensors 235a, 235b are associated with control sticks 130a, 130b of the aircraft 110.

[0105] Clause 16. The system according to Clause 14, wherein at least one of sensors 230b, 230d is associated with right rudder pedals 140b, 140d, and another of at least one of sensors 230a, 230c is associated with left rudder pedals 140a, 140c.

[0106] Clause 17. The system according to Clause 16, wherein at least one processor 210 is further configured to determine whether an engine failure exists on the aircraft 110.

[0107] Clause 18. The system according to Clause 17, wherein at least one processor 210 is further configured to determine a thrust asymmetry offset value in the presence of an engine failure.

[0108] Clause 19. The system according to Clause 18, wherein at least one processor 210 is further configured to determine a compensated pedal force value based on an adjusted force value and a thrust asymmetry offset value.

[0109] Clause 20. The system described in Clause 13, wherein the desired starting force is predetermined.

[0110] Although certain illustrative embodiments and methods have been disclosed herein, it will be apparent to those skilled in the art from the foregoing disclosure that variations and modifications can be made to these embodiments and methods without departing from the true spirit and scope of this disclosure. Many other examples exist, each differing from the others only in details. Therefore, this disclosure is intended to be limited only to the scope required by the appended claims and the rules and principles of applicable law.

Claims

1. A method for providing continuously variable sensing forces to an aircraft (110), the method comprising: The force sensor value is obtained from each of the at least one sensor (230a, 230b, 230c, 230d, 235a, 235b) associated with at least one aircraft control (130a, 130b, 140a, 140b, 140c, 140d). The net force value is determined by at least one processor (210) using the force sensor value of each of the at least one sensor (230a, 230b, 230c, 230d, 235a, 235b); The at least one processor (210) compares the net force value with the desired starting force; The at least one processor (210) determines whether the net force value exceeds the expected starting force; When the net force value exceeds the desired starting force, the at least one processor (210) determines an adjusted force value by using the desired starting force and the net force value; The at least one processor (210) determines the actuator torque command based on the adjusted force value; The at least one processor (210) commands the autopilot actuators (220a, 220b) to apply force to the at least one aircraft control using the actuator torque command, thereby reducing the artificial drag sensation on the aircraft control; and When it is desired to reduce the perceived force, the autopilot actuators (220a, 220b) engage in torque command mode.

2. The method according to claim 1, wherein at least two of the at least one sensor (230a, 230b, 230c, 230d, 235a, 235b) are present; and wherein the at least two sensors (235a, 235b) are associated with the control stick of the aircraft (110).

3. The method of claim 2, wherein at least one of the sensors (230b, 230d) is associated with the right rudder pedal, and another of at least one of the sensors (230a, 230c) is associated with the left rudder pedal.

4. The method of claim 3, wherein the method further comprises determining a thrust asymmetry offset value by the at least one processor (210) when an engine failure is present.

5. The method of claim 4, wherein the method further comprises determining a compensated pedal force value by the at least one processor (210) based on the adjusted force value and the thrust asymmetry offset value.

6. The method of claim 1, wherein the method further comprises determining by the at least one processor (210) whether it is desired to reduce the sensing power of the at least one aircraft control (130a, 130b, 140a, 140b, 140c, 140d) associated with each of the at least one sensor (230a, 230b, 230c, 230d, 235a, 235b).

7. The method of claim 6, wherein the at least one processor (210) determines whether a reduction in perceived force is desired based on the flight phase of the aircraft (110).

8. The method according to any one of claims 1-7, wherein the autopilot actuator (220a, 220b) is one of a pedal autopilot actuator or a lever autopilot actuator.

9. A system for providing continuously variable sensing force to an aircraft (110), the system comprising: At least one aircraft control (130a, 130b, 140a, 140b, 140c, 140d). At least one autopilot actuator (220a, 220b) configured to apply force to the at least one aircraft control (130a, 130b, 140a, 140b, 140c, 140d). At least one sensor (230a, 230b, 230c, 230d, 235a, 235b) associated with each of the at least one aircraft control (130a, 130b, 140a, 140b, 140c, 140d), wherein each of the at least one sensor (230a, 230b, 230c, 230d, 235a, 235b) is configured to sense force sensor values; and At least one processor (210) is configured to: The net force value is determined by using the force sensor value of each of the at least one sensor (230a, 230b, 230c, 230d, 235a, 235b). The net force value is compared with the expected starting force. Determine whether the net force value exceeds the expected starting force. When the net force value exceeds the desired starting force, an adjusted force value is determined by using the desired starting force and the net force value. The actuator torque command is determined based on the adjusted force value. The actuator torque command is used to command at least one autopilot actuator (220a, 220b) to apply force to at least one aircraft control, thereby reducing the artificial drag sensation on the aircraft control; When it is desired to reduce the perceived force, the at least one autopilot actuator (220a, 220b) engages in torque command mode.

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