Flight-by-wire flight control system hydraulic source fault influence indication conformity test method

Abstract

The invention relates to a fly-by-wire flight control system hydraulic source fault influence indication conformity test method, which comprises the following steps of: cutting off an SOV valve of a flight control system actuator corresponding to a No.1 hydraulic source and injecting a hydraulic source fault signal to cut off an SOV valve of a flight control system actuator corresponding to a No.2 hydraulic source; and the control instruction cannot be subjected to closed-loop control in an actuator servo loop, and is equivalent to the main characteristic of the flight control system when the double-source fault of the hydraulic source on the aircraft really fails. The flight control system can reasonably enter a hydraulic source double-source fault scene, main equipment of the flight control system and working characteristics of the system are fully embodied when the double-source fault occurs, it can be guaranteed that the redundancy of a real hydraulic source on a machine is not lost, the safety of a unit during test flight of a high-risk subject is guaranteed, and the reliability of the flight control system is improved. And meanwhile, the method is accepted by a bureau to indicate the conformity of the flight control system to the CCAR25.672 (c) clause when the flight control system fails, so that the method has relatively high generalization performance.

Description

Technical Field

[0001] The present application relates to the technical field of aviation engineering flight control systems, and in particular to a compliance test method for indicating the impact of hydraulic source failures in fly-by-wire flight control systems. Background Art

[0002] Civil aircraft have high safety and reliability requirements for fly-by-wire flight control systems. If a civil aircraft wants to be put into operation, it must prove / demonstrate to the authority that the safety and reliability of the aircraft meet the requirements of airworthiness regulations. CCAR25.672(c) stipulates: It must be shown that, following any single failure of the stability augmentation system or any other automatic or power-operated system, the following provisions are complied with: (1) The airplane can be safely controlled when the malfunction or malfunction occurs at any speed or altitude that is within the approved operating limitations and is critical for the type of malfunction; (2) The controllability and maneuverability requirements specified in this part can be met within the flight envelope (e.g., speed, altitude, normal acceleration, and airplane configuration) specified in the Airplane Flight Manual for actual use.

[0003] Therefore, it is necessary to carry out flight control system fault tests to demonstrate the compliance of the system under a single fault state. Among them, the failure of the hydraulic source on the aircraft is the focus of airworthiness attention. The hydraulic source is a key energy system on the aircraft. When a hydraulic source fails or fails, the performance of the flight control system is the key to the authority's review. If a single set of hydraulic source fails, whether the flight control system has the hydraulic energy required for normal operation; when two sets of hydraulic sources fail at the same time, whether the aircraft is controllable and usable, and whether the aircraft can still fly and land safely.

[0004] Currently, most aircraft hydraulic energy configurations are three independent oil source pipelines, which supply oil to the actuators of the servo actuation system respectively. Failure verification tests are divided into single-source failure and dual-source failure. Single-source failure is mostly demonstrated through flight test verification, and dual-source failure is mostly demonstrated through engineering simulator tests. Although engineering simulators can simulate a variety of flight states and fault conditions, they cannot fully reproduce all factors in real flight. Verification only through engineering simulators has certain limitations.

[0005] Therefore, it is necessary to propose a solution to improve one or more problems existing in the above-mentioned related technical solutions.

[0006] It should be noted that the information disclosed in the above background technology section is only used to enhance the understanding of the background of this application, and therefore may include information that does not constitute prior art known to ordinary technicians in this field. Summary of the Invention

[0007] The present invention provides a method for testing the compliance of a hydraulic source failure impact indicator for a fly-by-wire flight control system, wherein the hydraulic source of the fly-by-wire flight control system is a three-way hydraulic source consisting of a No. 1 hydraulic source, a No. 2 hydraulic source, and a No. 3 hydraulic source, including the following steps: The aircraft is trimmed in the air so that the SOV valve of the actuator corresponding to the No. 1 hydraulic source is in the cut-off state to simulate a single hydraulic source failure in a real scenario; In the scenario of a single hydraulic source failure, a hydraulic source failure signal is injected to cut off the SOV valve of the actuator corresponding to the No. 2 hydraulic source, simulating a dual hydraulic source failure in a real scenario. The No. 3 hydraulic source is in a normal state, so that the hydraulic energy of the three hydraulic sources are all in a pre-positioned state, and the crew performs a test flight action.

[0008] In an exemplary embodiment of the present application, when the dual hydraulic sources simulating a real scenario fail, control instructions cannot perform closed-loop control on the actuator servo circuits corresponding to the No. 1 hydraulic source and the No. 2 hydraulic source.

[0009] In an exemplary embodiment of the present application, the step of trimming the aircraft in the air and placing the SOV valve of the actuator corresponding to the No. 1 hydraulic source in a cut-off state to simulate a single hydraulic source failure in a real scenario includes: Closing the control valve of the No. 1 hydraulic source in the hydraulic system, thereby cutting off the hydraulic pipeline of the No. 1 hydraulic source in the hydraulic system and cutting off the SOV valve of the flight control system actuator corresponding to the No. 1 hydraulic source; Open the control valve of hydraulic source No. 1 in the hydraulic system, restore the hydraulic pipeline of hydraulic source No. 1 in the hydraulic system, and restore the hydraulic energy of hydraulic source No. 1, so that the SOV valve of the flight control system actuator corresponding to hydraulic source No. 1 remains in the cut-off state.

[0010] In an exemplary embodiment of the present application, in a scenario where a single hydraulic source fails, the step of injecting a hydraulic source failure signal to cut off the SOV valve of the actuator corresponding to the second hydraulic source to simulate a dual hydraulic source failure in a real scenario includes: In the scenario of a single hydraulic source failure, a hydraulic source failure signal is injected. The hydraulic source failure signal causes the SOV valve of the flight control system actuator corresponding to hydraulic source No. 2 to be cut off. The hydraulic pipeline of hydraulic source No. 2 is not cut off, and the hydraulic energy of hydraulic source No. 2 is normal.

[0011] In an exemplary embodiment of the present application, the positions of the three hydraulic sources include: a left hydraulic source, a middle hydraulic source and a right hydraulic source, the hydraulic source No. 1 is any one of the left hydraulic source, the middle hydraulic source and the right hydraulic source, and the hydraulic source No. 2 is any one of the two remaining hydraulic sources after the hydraulic source No. 1 is selected.

[0012] In an exemplary embodiment of the present application, the dual hydraulic source failure includes: left hydraulic source failure and center hydraulic source failure, left hydraulic source failure and right hydraulic source failure, and right hydraulic source failure and center hydraulic source failure.

[0013] In an exemplary embodiment of the present application, before the step of trimming the aircraft in mid-air and placing the SOV valve of the actuator corresponding to the No. 1 hydraulic source in a cut-off state to simulate a single hydraulic source failure in a real scenario, the following steps are further included: configuring the aircraft control surface actuators and the three hydraulic sources in the fly-by-wire flight control system: The aircraft control surface actuators include elevator control surface actuators, aileron control surface actuators and rudder control surface actuators; The elevator rudder surface actuators are configured with three hydraulic sources. The elevator rudder surfaces are divided into a left elevator rudder surface and a right elevator rudder surface. The left elevator rudder surface and the right elevator rudder surface are each configured with two actuators in a master-standby mode and three hydraulic sources. The aileron rudder surface actuators are configured with three hydraulic sources. The aileron rudder surfaces are divided into a left aileron rudder surface and a right aileron rudder surface. The left aileron rudder surface and the right aileron rudder surface are each configured with two actuators and three hydraulic sources in a master-standby manner. The rudder surface actuator is configured with three hydraulic sources. The rudder surface is composed of an upper rudder, a middle rudder, and a lower rudder vertically distributed to form a rudder surface. The actuator of the upper rudder, the actuator of the middle rudder, and the actuator of the lower rudder are configured with the three hydraulic sources in a main-backup-backup manner.

[0014] In an exemplary embodiment of the present application, the left elevator rudder surface includes a left inner elevator rudder surface and a left outer elevator rudder surface, and the right elevator rudder surface includes a right inner elevator rudder surface and a right outer elevator rudder surface. The actuator of the left outer elevator rudder surface is controlled by a left hydraulic source, the actuator of the right outer elevator rudder surface is controlled by a right hydraulic source, and the actuators of the left inner elevator rudder surface and the right inner elevator rudder surface are controlled by a center hydraulic source.

[0015] In an exemplary embodiment of the present application, the left rudder surface of the aileron rudder includes the left inner rudder surface of the aileron rudder and the left outer rudder surface of the aileron rudder, and the right rudder surface of the aileron rudder includes the right inner rudder surface of the aileron rudder and the right outer rudder surface of the aileron rudder. The actuator of the left outer rudder surface of the aileron rudder is controlled by the left hydraulic source, the actuator of the right outer rudder surface of the aileron rudder is controlled by the right hydraulic source, and the actuator of the left inner rudder surface of the aileron rudder and the actuator of the right inner rudder surface of the aileron rudder are controlled by the middle hydraulic source.

[0016] In an exemplary embodiment of the present application, the actuator of the upper rudder is controlled by a left hydraulic source, the actuator of the middle rudder is controlled by a middle hydraulic source, and the actuator of the lower rudder is controlled by a right hydraulic source.

[0017] The present application proposes a compliance test method for demonstrating the impact of a hydraulic source failure on an electric flight control system. On the one hand, through a real failure of hydraulic source No. 1 and a failure of an injected hydraulic source, the SOV valves of the flight control system actuators corresponding to hydraulic source No. 1 and hydraulic source No. 2 are cut off, so that the corresponding control instructions cannot form a closed-loop control in the servo circuit of their actuators, which is equivalent to the main characteristics of the flight control system when a real hydraulic source dual-source failure occurs on the aircraft. On the other hand, the flight control system can reasonably enter a hydraulic source dual-source failure scenario, reflecting the working characteristics of the main equipment and system of the flight control system when a dual-source failure occurs, ensuring that the redundancy of the real hydraulic source on the aircraft is not lost, ensuring the safety of the crew during high-risk test flights, and demonstrating that the flight control system complies with CCAR25.672(c) clause in the event of a failure, and has strong generalizability. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] The accompanying drawings are incorporated into and constitute a part of the specification, illustrate embodiments consistent with the present application, and together with the specification, are used to explain the principles of the present application. Obviously, the drawings described below are only some embodiments of the present application, and those skilled in the art can derive other drawings based on these drawings without inventive effort.

[0019] Figure 1 A schematic diagram showing the steps of a compliance test method for a hydraulic source failure impact indicator for a fly-by-wire flight control system in an exemplary embodiment of the present application is shown; Figure 2 A schematic diagram showing the configuration of an elevator surface actuator and three hydraulic sources in an exemplary embodiment of the present application is shown; Figure 3 A schematic diagram showing the configuration of the aileron and rudder actuators and three hydraulic sources in an exemplary embodiment of the present application is shown; Figure 4 A schematic diagram showing the configuration of a rudder surface actuator and three hydraulic sources in an exemplary embodiment of the present application is shown; Figure 5 A schematic diagram showing longitudinal maneuverability in an exemplary embodiment of the present application is shown; Figure 6 A schematic diagram showing lateral maneuverability in an exemplary embodiment of the present application is shown; Figure 7 A schematic diagram of the heading maneuvering capability in an exemplary embodiment of the present application is shown. DETAILED DESCRIPTION

[0020] Example embodiments will now be described more fully with reference to the accompanying drawings. However, example embodiments can be embodied in many forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided so that this disclosure will be thorough and complete and will fully convey the concepts of the example embodiments to those skilled in the art. The described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments.

[0021] In addition, the accompanying drawings are merely schematic illustrations of the present application and are not necessarily drawn to scale. Identical reference numerals in the figures denote identical or similar parts, and thus repetitive descriptions thereof will be omitted. Some of the blocks shown in the accompanying drawings are functional entities that do not necessarily correspond to physically or logically separate entities. These functional entities may be implemented in software, in one or more hardware modules or integrated circuits, or in different networks and / or processor devices and / or microcontroller devices.

[0022] The fly-by-wire control system is a critical aircraft control system. Based on the aircraft's current motion and the pilot's commands from the cockpit, it generates control surface commands through control loop calculations, driving the deflection of the control surfaces to achieve the pilot's desired attitude / trajectory control. The servo actuator system is the flight control system's actuator. It receives control commands from the system and drives the control surfaces (such as ailerons, rudder, and elevator) in accordance with the commands, achieving angular or trajectory motion. After receiving the control system's command signals, the servo actuator system converts the electrical signals into hydraulic signals (flow rate). Using power methods, it outputs force and speed to drive the hydraulic servos and, in turn, the control surfaces. SOV valves (solenoid operated valves) are electromagnetically controlled valves.

[0023] This example embodiment provides a method for testing the hydraulic source failure impact of a fly-by-wire flight control system. Figure 1 As shown, the following steps may be included: The hydraulic source of the fly-by-wire flight control system is a three-way hydraulic source consisting of hydraulic source No. 1, hydraulic source No. 2 and hydraulic source No. 3; Step S101: The aircraft is trimmed in the air so that the SOV valve of the actuator corresponding to the No. 1 hydraulic source is in a cut-off state, thereby simulating a single hydraulic source failure in a real scenario.

[0024] Step S102: In a scenario where a single hydraulic source fails, a hydraulic source failure signal is injected to cut off the SOV valve of the actuator corresponding to the No. 2 hydraulic source, thereby simulating a dual hydraulic source failure in a real scenario.

[0025] Step S103: The No. 3 hydraulic source is in a normal state, the hydraulic energy of the three hydraulic sources are all in a pre-positioned state, and the crew performs a test flight action.

[0026] The embodiment of the present application proposes a compliance test method for demonstrating the impact of a hydraulic source failure on an electric flight control system. On the one hand, through a real failure of hydraulic source No. 1, the SOV valve of the flight control system actuator corresponding to hydraulic source No. 1 is cut off, and a hydraulic source failure is injected to cut off the SOV valve of the flight control system actuator corresponding to hydraulic source No. 2, so that the corresponding control instruction cannot form a closed-loop control in the servo loop of its actuator, which is equivalent to the main characteristics of the flight control system when the real hydraulic source dual-source failure on the aircraft occurs. On the other hand, the flight control system can reasonably enter the hydraulic source dual-source failure scenario, reflecting the working characteristics of the main equipment and system of the flight control system when the dual-source failure occurs, ensuring that the redundancy of the real hydraulic source on the aircraft is not lost, ensuring the safety of the crew during high-risk subject test flights, and demonstrating the compliance of the flight control system with CCAR25.672(c) clause in the event of a failure, which has strong generalizability.

[0027] Below, a more detailed description will be given of a compliance test method for a hydraulic source failure impact indicator for a fly-by-wire flight control system proposed in this example embodiment.

[0028] Step S101: The aircraft is trimmed in the air so that the SOV valve of the actuator corresponding to the No. 1 hydraulic source is in a cut-off state, thereby simulating a single hydraulic source failure in a real scenario.

[0029] Specifically, step S101 includes the following sub-steps: Step S1001: The aircraft is trimmed in the air according to the set altitude, speed, center of gravity, weight and flap position.

[0030] Step S1002: Close the control valve of hydraulic source No. 1 in the hydraulic system, so that the hydraulic pipeline of hydraulic source No. 1 in the hydraulic system is cut off, and the SOV valve of the flight control system actuator corresponding to hydraulic source No. 1 is cut off.

[0031] Step S1003: Open the control valve of hydraulic source No. 1 in the hydraulic system, restore the hydraulic pipeline of hydraulic source No. 1 in the hydraulic system, restore the hydraulic energy of hydraulic source No. 1, and the SOV valve of the flight control system actuator corresponding to hydraulic source No. 1 is still in the cut-off state. The control instruction cannot perform closed-loop control in the servo circuit of its actuator, simulating a single hydraulic source failure in a real scenario.

[0032] Step S102: In a scenario where a single hydraulic source fails, a hydraulic source failure signal is injected to cut off the SOV valve of the actuator corresponding to the No. 2 hydraulic source, thereby simulating a dual hydraulic source failure in a real scenario.

[0033] Specifically, in the scenario of a single hydraulic source failure, a hydraulic source failure signal is injected. The hydraulic source failure signal causes the SOV valve of the flight control system actuator corresponding to hydraulic source No. 2 to be cut off. The hydraulic pipeline of hydraulic source No. 2 is not cut off, the hydraulic energy of hydraulic source No. 2 is normal, and the control instruction cannot perform closed-loop control in its actuator servo circuit, thereby simulating a dual hydraulic source failure in a real scenario.

[0034] Step S103: The No. 3 hydraulic source is in a normal state, the hydraulic energy of the three hydraulic sources are all in a pre-positioned state, and the crew performs a test flight action.

[0035] Specifically, the No. 3 hydraulic source is in a normal state, the hydraulic energy of the No. 1 hydraulic source is restored, the hydraulic energy of the No. 2 hydraulic source is normal, the hydraulic energy of the three hydraulic sources are all in the pre-position state, and the crew performs a test flight action.

[0036] In the disclosed embodiment, the positions of the three hydraulic sources include: a left hydraulic source, a center hydraulic source, and a right hydraulic source. Prior to the step of trimming the aircraft in mid-air to cut off the SOV valve of the actuator corresponding to the No. 1 hydraulic source to simulate a single hydraulic source failure in a real-world scenario, the following steps are also included for configuring the aircraft control surface actuators and the three hydraulic sources in the fly-by-wire flight control system: The aircraft control surface actuators include elevator control surface actuators, aileron control surface actuators and rudder control surface actuators; Step S2001: The elevator rudder surface actuators are configured with three hydraulic sources. The elevator rudder surfaces are divided into a left elevator rudder surface and a right elevator rudder surface. The elevator rudder surfaces are divided into a left elevator rudder surface and a right elevator rudder surface, and are respectively configured with two actuators and three hydraulic sources in a master-slave mode.

[0037] Specifically, such as Figure 2 As shown, the left elevator rudder surface includes the left inner elevator rudder surface and the left outer elevator rudder surface, and the right elevator rudder surface includes the right inner elevator rudder surface and the right outer elevator rudder surface. The actuator of the left outer elevator rudder surface is controlled by the left hydraulic source, the actuator of the right outer elevator rudder surface is controlled by the right hydraulic source, and the actuator of the left inner elevator rudder surface and the actuator of the right inner elevator rudder surface are controlled by the center hydraulic source.

[0038] Step S2002: The aileron rudder surface actuators are configured with three hydraulic sources. The aileron rudder surfaces are divided into a left aileron rudder surface and a right aileron rudder surface. The left aileron rudder surface and the right aileron rudder surface are respectively configured with two actuators and three hydraulic sources in a master-slave mode.

[0039] Specifically, such as Figure 3As shown, the left rudder surface of the aileron rudder includes the left inner rudder surface of the aileron rudder and the left outer rudder surface of the aileron rudder, and the right rudder surface of the aileron rudder includes the right inner rudder surface of the aileron rudder and the right outer rudder surface of the aileron rudder. The actuator of the left outer rudder surface of the aileron rudder is controlled by the left hydraulic source, the actuator of the right outer rudder surface of the aileron rudder is controlled by the right hydraulic source, and the actuator of the left inner rudder surface of the aileron rudder and the actuator of the right inner rudder surface of the aileron rudder are controlled by the center hydraulic source.

[0040] Step S2003: The rudder surface actuators are configured with three hydraulic sources. The rudder surface is composed of an upper rudder, a middle rudder, and a lower rudder vertically distributed to form a rudder surface. The upper rudder actuator, the middle rudder actuator, and the lower rudder actuator are configured with the three hydraulic sources in a main-backup-backup manner.

[0041] Specifically, such as Figure 4 As shown, the actuator of the upper rudder is controlled by the left hydraulic source, the actuator of the middle rudder is controlled by the middle hydraulic source, and the actuator of the lower rudder is controlled by the right hydraulic source.

[0042] In the embodiment of the present disclosure, dual-source failures in the hydraulic energy of the three hydraulic sources include: left hydraulic source failure and middle hydraulic source failure, left hydraulic source failure and right hydraulic source failure, and right hydraulic source failure and middle hydraulic source failure.

[0043] Specifically, based on the configuration of the aircraft control surface actuators and the three hydraulic sources in the flight control system, dual-source failure of the hydraulic source is divided into the following three scenarios.

[0044] Scenario 1: Left hydraulic power source failure and center hydraulic power source failure; At this time, the left elevator rudder surface is lost and unusable, and the aircraft pitch control is achieved by relying only on the right outer elevator rudder surface to control the right elevator rudder surface; If the left aileron rudder is lost and unusable, the aircraft roll control can only be achieved by controlling the right aileron rudder with the right outer aileron rudder. The rudder is controlled only by the lower rudder to achieve aircraft heading control.

[0045] Scenario 2: Left and right hydraulic power source failures; At this time, the left inner rudder surface of the elevator controls the left rudder surface, and the right inner rudder surface of the elevator controls the right rudder surface to achieve aircraft pitch control; The left inner aileron rudder controls the left aileron rudder, and the right inner aileron rudder controls the right aileron rudder to achieve aircraft roll control; The rudder is controlled only by the Chinese rudder to achieve aircraft heading control.

[0046] Scenario 3: Right hydraulic power source failure and center hydraulic power source failure; At this time, the right elevator rudder surface is lost and unusable, and the aircraft pitch control is achieved by relying only on the left outer elevator rudder surface to control the left elevator rudder surface; If the right aileron rudder is lost and unusable, the aircraft roll control is achieved by controlling the left aileron rudder with the left outer aileron rudder. The rudder is controlled only by the upper rudder to achieve aircraft heading control.

[0047] For the above failure scenarios, dual-source failure will cause the flight control system to control with minimum rudder, which puts the aircraft in a highly dangerous state, and only one set of hydraulic energy is available on board. At the same time, in order to avoid placing an operational burden on the crew during dual-source failure tests, unnecessary crew actions are minimized to ensure that during high-risk test flights, the crew can concentrate on executing the flight actions specified in the terms. A normalized strategy was implemented for the three dual-source failure scenarios, where hydraulic source No. 1 is any one of the left, middle, and right hydraulic sources, and hydraulic source No. 2 is any one of the two remaining hydraulic sources after hydraulic source No. 1 is selected.

[0048] Specifically, when a dual-source failure of the hydraulic source occurs, each hydraulic source on the onboard control panel is provided with a corresponding control valve. The left hydraulic source corresponds to a left hydraulic source control valve, the middle hydraulic source corresponds to a middle hydraulic source control valve, and the right hydraulic source corresponds to a right hydraulic source control valve. The control valve of one of the hydraulic sources is closed by operation, which is to close the control valve of hydraulic source No. 1. At this time, the hydraulic pipeline of hydraulic source No. 1 in the hydraulic system is cut off, and the SOV valve of the flight control system actuator corresponding to hydraulic source No. 1 is cut off. The control valve of hydraulic source No. 1 in the hydraulic system is opened, the hydraulic pipeline of hydraulic source No. 1 in the hydraulic system is restored, the hydraulic energy of hydraulic source No. 1 is restored, and the SOV valve of the flight control system actuator corresponding to hydraulic source No. 1 is still in the cut-off state, and the control instruction cannot perform closed-loop control in the servo loop of its actuator. The control failure of the SOV valve of the actuator corresponding to hydraulic source No. 1 is used to simulate the control failure of the SOV valve of the actuator corresponding to the first hydraulic source when the first hydraulic source fails in a real scenario, generating a single hydraulic source failure. It should be understood that when the actuator recognizes that the hydraulic source to which it is connected has a fault, it cuts off its servo circuit and latches the fault.

[0049] The aircraft was equipped with an onboard test device that injected a hydraulic source fault signal into the flight control system. This signal simulated a dual-source failure scenario in the flight control system, even in a single-source failure scenario. The test device consisted of a fault injection panel and a fault forwarding computer. The panel contained buttons for the left, center, and right hydraulic sources. Pressing a corresponding button sent the fault signal to the fault forwarding computer. It's important to understand that if the left hydraulic source control button was turned off on the control panel, pressing either the right or center hydraulic source fault button on the fault injection panel would trigger a dual-source failure. The fault forwarding computer sent the hydraulic source failure signal to the flight control system. Upon recognizing the fault signal, the flight control system shut off the SOV valve on the flight control system actuator corresponding to hydraulic source No. 2. However, the hydraulic line to hydraulic source No. 2 was not actually shut off, and the hydraulic power supply to hydraulic source No. 2 remained normal. The failure of the SOV valve on the actuator corresponding to hydraulic source No. 2 simulates a real-world scenario in which the SOV valve on the actuator corresponding to the second hydraulic source fails, resulting in a dual-hydraulic source failure. This dual-hydraulic source failure does not represent a true loss of power from both hydraulic sources. It demonstrates the operating characteristics of the flight control system's key equipment and systems during a dual-source failure, ensuring that the aircraft's actual hydraulic source redundancy is maintained and crew safety is guaranteed during high-risk flight tests. This demonstrates the flight control system's compliance with CCAR 25.672(c) during failures, demonstrating its broad applicability.

[0050] It should be understood that among the hydraulic energy in the three hydraulic sources, the hydraulic energy of hydraulic source No. 1 has been restored, the hydraulic energy of hydraulic source No. 2 is normal, and the hydraulic energy of hydraulic source No. 3 is normal. The hydraulic energy of the three hydraulic sources are all in the pre-position state, and the crew performs the test flight action.

[0051] It's also important to understand that when exiting the dual-source failure scenario, all three hydraulic sources are armed, the control computer corresponding to hydraulic source No. 1 is restarted, the SOV valve of the flight control system actuator corresponding to hydraulic source No. 1 is restored, and the restoration control command is executed in closed-loop control of the actuator's servo circuit. The test equipment that injects the hydraulic source failure signal is powered off, the SOV valve of the flight control system actuator corresponding to hydraulic source No. 2 is restored, and the restoration control command is executed in closed-loop control of the actuator's servo circuit. This completes the exit from the dual-source failure scenario.

[0052] Example 1 The aircraft was trimmed in the air at an altitude of 3,000 meters, at a speed of 260 km / h, with a low front center of gravity and flaps at 0 degrees, keeping the wings level.

[0053] At this time, close the control valve of the right hydraulic source in the aircraft's hydraulic system and cut off the hydraulic pipeline of the right hydraulic source. At this time, the flight control system receives a right hydraulic source pressure of 0MPa, the SOV valve of the actuator of the right outer rudder surface of the elevator is cut off, the right inner rudder surface of the elevator controls the right rudder surface of the elevator, the SOV valve of the actuator of the right outer rudder surface of the aileron is cut off, the right inner rudder surface of the aileron controls the right rudder surface of the aileron, the SOV valve of the actuator of the lower rudder is cut off, and the upper rudder and the middle rudder control the heading.

[0054] A central hydraulic pressure source fault signal was injected through the flight control system's onboard test equipment. The flight control system received the command and proactively cut off the SOV valve of the actuator of the right inner elevator surface. At this time, the right elevator surface failed, and proactively cut off the SOV valve of the actuator of the right inner aileron surface. At this time, the right aileron surface failed, and proactively cut off the SOV valve of the actuator of the upper rudder. At this time, the upper rudder surface failed, and the central rudder controlled the heading.

[0055] Open the right hydraulic source control valve of the onboard hydraulic system, restore the hydraulic pipeline of the right hydraulic source, and restore the hydraulic energy of the right hydraulic source. The SOV valve of the actuator of the right outer rudder surface of the elevator is still in the cut-off state, the SOV valve of the actuator of the right outer rudder surface of the aileron is still in the cut-off state, and the SOV valve of the actuator of the lower rudder is still in the cut-off state. The actuator of the right outer rudder surface of the elevator, the actuator of the right outer rudder surface of the aileron, and the actuator of the lower rudder are still in a control failure state.

[0056] Confirm the dual-source failure scenario and determine the flight control system and aircraft characteristics that meet the requirements of a real dual-source failure.

[0057] Observe the onboard warning information reporting: Failure of hydraulic pressure source; The right elevator failed; The right aileron failed; reduced rudder margin; The crew performed test flight maneuvers to verify the control functions of pitch, roll and yaw.

[0058] The details are as follows: Execute pulse elevator, and the compliance criterion is that the aircraft has the ability to maneuver with a normal overload of 0.8~1.3g under the condition of losing one side of the control surface.

[0059] Execute stepped ailerons, establish a stable turning flight with a bank angle of ±30 degrees, and perform a stepped aileron maneuver with a roll angle change of 60 degrees in the opposite direction of the turn. The compliance criterion action can be established within 11 seconds.

[0060] The double pulse rudder is implemented, and the compliance criterion is that the pedal amount has no Dutch roll tendency within ±2 / 3 of the full stroke.

[0061] The aircraft's maneuverability at this point is as follows: Using a rod displacement of approximately -94 to 80 mm, we achieved an overload of 1.26 to 0.6 g, and the longitudinal maneuverability was normal. Figure 5 shown.

[0062] The flight maneuver of the compliance criterion action is completed within 11 seconds, and the lateral maneuverability is normal. Figure 6 shown.

[0063] The heading control did not trigger the angle of attack protection, the Dutch roll converged quickly, and the heading maneuverability was normal. Figure 7 shown.

[0064] In the above embodiment, the control valve of one of the three hydraulic sources is closed, cutting off the hydraulic pipeline of that hydraulic source. The flight control system cannot obtain the hydraulic energy of that source, and the SOV valve of that actuator is cut off. At this time, the corresponding flight control system actuator loses control capability. The control valve of that hydraulic source is opened, and the hydraulic energy and hydraulic pipeline of that hydraulic source are restored. However, the SOV valve of that actuator remains in a cut-off state, and control commands cannot be implemented in the closed-loop control of its actuator servo circuit. The other actuator on the same side becomes the "master" to control the movement of the control surface. The control failure of the actuator is used to simulate the real scenario of a single hydraulic source failure.

[0065] In the event of a single hydraulic source failure, a hydraulic source failure signal is injected, shutting off the SOV valve of the flight control system actuator corresponding to the other hydraulic source. This does not actually disconnect the hydraulic line of this hydraulic source, and hydraulic energy from this source remains. Control commands cannot be executed in the actuator's servo circuit for closed-loop control, simulating a real-world scenario of a hydraulic source failure by facilitating actuator control failure. This allows for a dual hydraulic source failure scenario.

[0066] The dual-source failure of the hydraulic source is equivalent to the main characteristics of the flight control system when the real dual-source failure of the hydraulic source on the aircraft occurs, ensuring that the redundancy of the real hydraulic source on the aircraft is not lost, ensuring the safety of the crew during high-risk test flights, and indicating that the flight control system complies with CCAR25.672(c) in the event of a failure.

[0067] In addition, the terms "first" and "second" are used for descriptive purposes only and should not be understood as indicating or implying relative importance or implicitly indicating the number of the technical features indicated. Therefore, a feature defined as "first" or "second" may explicitly or implicitly include one or more of the features. In the description of the embodiments of this application, the meaning of "plurality" is two or more, unless otherwise clearly specified.

[0068] In the description of this specification, the reference terms "one embodiment", "some embodiments", "example", "specific example" or "some examples" mean that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present application. In this specification, the schematic expressions of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art can combine and combine different embodiments or examples described in this specification.

[0069] Those skilled in the art will readily appreciate other embodiments of the present invention after considering the specification and practicing the invention disclosed herein. This application is intended to cover any variations, uses, or adaptations of the present invention that follow the general principles of this application and include common knowledge or customary techniques in the art that are not disclosed herein.

Claims

1. A test method for the compliance of the hydraulic source failure effect of the fly-by-wire flight control system, wherein the hydraulic source of the fly-by-wire flight control system is a three-way hydraulic source consisting of a No. 1 hydraulic source, a No. 2 hydraulic source and a No. 3 hydraulic source, characterized in that: The following steps are involved: The aircraft is trimmed in the air so that the SOV valve of the actuator corresponding to the No. 1 hydraulic source is in a cut-off state to simulate a single hydraulic source failure in a real scenario; In the scenario of a single hydraulic source failure, a hydraulic source failure signal is injected to cut off the SOV valve of the actuator corresponding to the No. 2 hydraulic source, so as to simulate a dual hydraulic source failure in a real scenario; The No. 3 hydraulic source is in a normal state, so that the hydraulic energy of the three hydraulic sources are all in a pre-positioned state, and the crew performs a test flight action.

2. The test method for the hydraulic source failure impact indication compliance of the fly-by-wire flight control system according to claim 1 is characterized in that: When the dual hydraulic sources simulating the real scene fail, the control instructions cannot perform closed-loop control on the actuator servo circuits corresponding to the No. 1 hydraulic source and the No. 2 hydraulic source.

3. The test method for the hydraulic source failure impact indication compliance of the fly-by-wire flight control system according to claim 1 is characterized in that: The step of trimming the aircraft in the air so that the SOV valve of the actuator corresponding to the No. 1 hydraulic source is in a cut-off state to simulate a single hydraulic source failure in a real scenario includes: Closing the control valve of the No. 1 hydraulic source in the hydraulic system, so that the hydraulic pipeline of the No. 1 hydraulic source in the hydraulic system is cut off, and the SOV valve of the flight control system actuator corresponding to the No. 1 hydraulic source is cut off; Open the control valve of hydraulic source No. 1 in the hydraulic system, restore the hydraulic pipeline of hydraulic source No. 1 in the hydraulic system, restore the hydraulic energy of hydraulic source No. 1, so that the SOV valve of the actuator of the flight control system corresponding to hydraulic source No. 1 is still in the cut-off state.

4. The test method for the hydraulic source failure impact indication compliance of the fly-by-wire flight control system according to claim 2 is characterized in that: In the scenario of a single hydraulic source failure, injecting a hydraulic source failure signal to cut off the SOV valve of the actuator corresponding to the No. 2 hydraulic source to simulate a dual hydraulic source failure in a real scenario includes: In the scenario of a single hydraulic source failure, a hydraulic source fault signal is injected. The hydraulic source fault signal causes the SOV valve of the flight control system actuator corresponding to hydraulic source No. 2 to be cut off, while the hydraulic pipeline of hydraulic source No. 2 is not cut off and the hydraulic energy of hydraulic source No. 2 is normal.

5. The test method for the hydraulic source failure impact indication compliance of the fly-by-wire flight control system according to claim 1 is characterized in that: The positions of the three hydraulic sources include: a left hydraulic source, a middle hydraulic source and a right hydraulic source, the No. 1 hydraulic source is any one of the left hydraulic source, the middle hydraulic source and the right hydraulic source, and the No. 2 hydraulic source is any one of the two remaining hydraulic sources after the No. 1 hydraulic source is selected.

6. The test method for the hydraulic source failure impact indicator compliance of the fly-by-wire flight control system according to claim 5 is characterized in that: The dual hydraulic source failure includes: a left hydraulic source failure and a middle hydraulic source failure, a left hydraulic source failure and a right hydraulic source failure, and a right hydraulic source failure and a middle hydraulic source failure.

7. The test method for the compliance of the hydraulic source failure effect indicator of the fly-by-wire flight control system according to claim 6 is characterized in that: Before the step of trimming the aircraft in the air so that the SOV valve of the actuator corresponding to the No. 1 hydraulic source is in a cut-off state to simulate a single hydraulic source failure in a real scenario, the configuration steps of the aircraft control surface actuator and the three-way hydraulic source in the fly-by-wire flight control system are also included: The aircraft control surface actuators include elevator control surface actuators, aileron control surface actuators and rudder control surface actuators; The elevator rudder surface actuator is configured with three hydraulic sources, the elevator rudder surface is divided into a left elevator rudder surface and a right elevator rudder surface, and the left elevator rudder surface and the right elevator rudder surface are respectively configured with two actuators and three hydraulic sources in a master-standby mode; The aileron rudder surface actuator is configured with three hydraulic sources, the aileron rudder surface is divided into a left aileron rudder surface and a right aileron rudder surface, and the left aileron rudder surface and the right aileron rudder surface are respectively configured with two actuators and three hydraulic sources in a master-slave mode; The rudder surface actuator is configured with three hydraulic sources. The rudder surface is vertically distributed with an upper rudder, a middle rudder and a lower rudder to form a rudder surface. The actuator of the upper rudder, the actuator of the middle rudder and the actuator of the lower rudder are configured with the three hydraulic sources in a main-standby-standby manner.

8. The test method for the hydraulic source failure impact indicator compliance of the fly-by-wire flight control system according to claim 7 is characterized in that: The left elevator rudder surface includes a left inner elevator rudder surface and a left outer elevator rudder surface, and the right elevator rudder surface includes a right inner elevator rudder surface and a right outer elevator rudder surface. The actuator of the left outer elevator rudder surface is controlled by a left hydraulic source, the actuator of the right outer elevator rudder surface is controlled by a right hydraulic source, and the actuator of the left inner elevator rudder surface and the actuator of the right inner elevator rudder surface are controlled by a central hydraulic source.

9. The test method for the hydraulic source failure impact indication compliance of the fly-by-wire flight control system according to claim 8, characterized in that: The left rudder surface of the aileron rudder includes a left inner rudder surface of the aileron rudder and a left outer rudder surface of the aileron rudder, and the right rudder surface of the aileron rudder includes a right inner rudder surface of the aileron rudder and a right outer rudder surface of the aileron rudder. The actuator of the left outer rudder surface of the aileron rudder is controlled by a left hydraulic source, the actuator of the right outer rudder surface of the aileron rudder is controlled by a right hydraulic source, and the actuator of the left inner rudder surface of the aileron rudder and the actuator of the right inner rudder surface of the aileron rudder are controlled by a middle hydraulic source.

10. The test method for the compliance of the hydraulic source failure effect indicator of the fly-by-wire flight control system according to claim 9, characterized in that: The actuator of the upper rudder is controlled by the left hydraulic source, the actuator of the middle rudder is controlled by the middle hydraulic source, and the actuator of the lower rudder is controlled by the right hydraulic source.

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