Test method for compliance of hydraulic source failure effect indicator for fly-by-wire flight control system

Abstract

The present application is a test method for demonstrating compliance of the impact of hydraulic source failure of an electric flight control system. The method includes: the SOV valve of the flight control system actuator corresponding to the No. 2 hydraulic source is cut off by cutting off the SOV valve of the flight control system actuator corresponding to the No. 1 hydraulic source and injecting a hydraulic source failure signal, so that the control command cannot perform closed-loop control in the servo circuit of its actuator, and the main characteristics of the flight control system when the hydraulic source dual source failure on the aircraft is actually failed are equivalent. The present application enables the flight control system to reasonably enter the hydraulic source dual source failure scenario, fully reflects the working characteristics of the main equipment and system of the flight control system when the dual source fails, and ensures 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. At the same time, it has been approved by the authority to demonstrate the compliance of the flight control system with the CCAR25.672 (c) clause when the failure occurs, and has strong promotion.

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 show / demonstrate to the authority that the safety and reliability of the aircraft meet the requirements of airworthiness regulations. CCAR25.672(c) stipulates:

[0003] It must be shown that, following any single failure of the stability augmentation system or any other automatic or power-controlled system, the following provisions are complied with:

[0004] (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 that is critical for the type of malfunction;

[0005] (2) The controllability and maneuverability requirements of 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.

[0006] 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 onboard hydraulic source is the focus of airworthiness attention. The hydraulic source is a key energy system on board. 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 both 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.

[0007] At present, 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 the engineering simulator can simulate a variety of flight states and fault conditions, it cannot fully reproduce all factors in real flight. Verification only through engineering simulators has certain limitations.

[0008] Therefore, it is necessary to propose a solution to improve one or more problems existing in the above-mentioned related technical solutions.

[0009] 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 the present application, and therefore may include information that does not constitute the prior art known to ordinary technicians in the field. Summary of the invention

[0010] The present application provides a method for testing the compliance of a hydraulic source failure impact indicator for an electric-by-wire flight control system, wherein the hydraulic source of the electric-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:

[0011] 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;

[0012] 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;

[0013] 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.

[0014] In an exemplary embodiment of the present application, 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.

[0015] In an exemplary embodiment of the present application, 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:

[0016] 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;

[0017] 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.

[0018] In an exemplary embodiment of the present application, in the scenario of 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 the dual hydraulic source failure in the real scenario includes:

[0019] 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.

[0020] 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 remaining two hydraulic sources after the hydraulic source No. 1 is selected.

[0021] In an exemplary embodiment of the present application, 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.

[0022] In an exemplary embodiment of the present application, before the step of trimming the aircraft in the air and making 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 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:

[0023] The aircraft control surface actuators include elevator control surface actuators, aileron control surface actuators and rudder control surface actuators;

[0024] 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;

[0025] 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;

[0026] 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.

[0027] 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, 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 center hydraulic source.

[0028] In an exemplary embodiment of the present application, 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, 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.

[0029] 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.

[0030] The present application proposes a compliance test method for the impact of hydraulic source failure on the fly-by-wire flight control system. On the one hand, through the real failure of hydraulic source No. 1 and the injection of hydraulic source failure, 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 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, indicating 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

[0031] The drawings herein are incorporated into the specification 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 for ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0032] Figure 1 A schematic diagram showing the steps of a test method for the hydraulic source failure impact indication compliance test of a fly-by-wire flight control system in an exemplary embodiment of the present application is shown;

[0033] Figure 2 A schematic diagram showing the configuration of the elevator surface actuator and three-way hydraulic sources in an exemplary embodiment of the present application is shown;

[0034] Figure 3 A schematic diagram showing the configuration of the aileron rudder actuator and the three-way hydraulic source in the exemplary embodiment of the present application is shown;

[0035] Figure 4 A schematic diagram showing the configuration of a rudder surface actuator and three-way hydraulic sources in an exemplary embodiment of the present application is shown;

[0036] Figure 5 A schematic diagram showing longitudinal maneuverability in an exemplary embodiment of the present application is shown;

[0037] Figure 6 A schematic diagram showing the lateral maneuverability in an exemplary embodiment of the present application is shown;

[0038] Figure 7 A schematic diagram of the heading maneuvering capability in an exemplary embodiment of the present application is shown. DETAILED DESCRIPTION

[0039] Example embodiments will now be described more fully with reference to the accompanying drawings. However, example embodiments can be implemented in a variety of forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided so that this application will be more comprehensive and complete and 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.

[0040] In addition, the accompanying drawings are only schematic illustrations of the present application and are not necessarily drawn to scale. The same reference numerals in the figures represent the same or similar parts, and thus their repeated description will be omitted. Some of the block diagrams shown in the accompanying drawings are functional entities and do not necessarily correspond to physically or logically independent entities. These functional entities can be implemented in software form, or implemented in one or more hardware modules or integrated circuits, or implemented in different networks and / or processor devices and / or microcontroller devices.

[0041] The fly-by-wire control system is a key control system for aircraft. It generates aircraft control surface control instructions through the solution of the control loop according to the current motion state of the aircraft and the operating instructions of the pilot in the cockpit, drives the deflection of the aircraft control surface, and realizes the attitude / trajectory control desired by the pilot. Among them, the servo actuator system is the actuator of the flight control system. It receives the control instructions of the system and drives the control surface (such as ailerons, rudders, elevators, etc.) to move according to the size of the instructions to achieve the angular movement or trajectory movement of the aircraft. After receiving the command signal from the control system, the servo actuator system converts the electrical signal into a hydraulic signal (flow), and then outputs force and speed through the power method to drive the hydraulic steering gear to move and drive the control surface to move. SOV valve (Solenoid Operated Valve, electromagnetic control valve).

[0042] This example embodiment provides a method for testing the hydraulic source failure effect of a fly-by-wire flight control system for compliance. Figure 1 As shown, the following steps may be included:

[0043] 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;

[0044] Step S101: 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, so as to simulate a single hydraulic source failure in a real scenario.

[0045] Step S102: In a 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.

[0046] Step S103: the No. 3 hydraulic source is in a normal state, the hydraulic energy sources of the three hydraulic sources are all in a pre-positioned state, and the crew performs a test flight action.

[0047] The embodiment of the present application proposes a test method for demonstrating compliance of the impact of a hydraulic source failure on an electric flight control system. On the one hand, a real failure of hydraulic source No. 1 causes the SOV valve of the flight control system actuator corresponding to hydraulic source No. 1 to be cut off, and a hydraulic source failure is injected to cause the SOV valve of the flight control system actuator corresponding to hydraulic source No. 2 to be cut off, so that the corresponding control command cannot form a closed-loop control in the servo circuit of its actuator, 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 flight tests, and demonstrating that the flight control system complies with CCAR25.672(c) clause when a failure occurs, which has strong generalizability.

[0048] Next, a compliance test method for a hydraulic source failure impact indicator for an electric-by-wire flight control system proposed in this example embodiment will be described in more detail.

[0049] Step S101: 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, so as to simulate a single hydraulic source failure in a real scenario.

[0050] Specifically, step S101 includes the following sub-steps:

[0051] Step S1001: The aircraft is trimmed in the air according to the set altitude, speed, center of gravity, weight and flap position.

[0052] 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.

[0053] 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 actuator of the flight control system corresponding to hydraulic source No. 1 is still in the cut-off state. The control command cannot perform closed-loop control in the servo circuit of its actuator, simulating a single hydraulic source failure in a real scenario.

[0054] Step S102: In a 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.

[0055] Specifically, 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, 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 command cannot perform closed-loop control in the servo circuit of its actuator, so as to simulate the dual hydraulic source failure in the real scenario.

[0056] Step S103: the No. 3 hydraulic source is in a normal state, the hydraulic energy sources of the three hydraulic sources are all in a pre-positioned state, and the crew performs a test flight action.

[0057] Specifically, the hydraulic source No. 3 is in a normal state, the hydraulic energy of the hydraulic source No. 1 is restored, the hydraulic energy of the hydraulic source No. 2 is normal, the hydraulic energy of the three hydraulic sources are all in a pre-positioned state, and the crew performs a test flight action.

[0058] In the disclosed embodiment, the positions of the three hydraulic sources include: a left hydraulic source, a middle hydraulic source, and a right hydraulic source. 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 hydraulic sources in the fly-by-wire flight control system are also included:

[0059] The aircraft control surface actuators include elevator control surface actuators, aileron control surface actuators and rudder control surface actuators;

[0060] Step S2001: 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. The elevator rudder surface is divided into a left elevator rudder surface and a right elevator rudder surface, and each of the two actuators is configured with three hydraulic sources in a master-slave mode.

[0061] Specifically, Figure 2As shown, the left rudder surface of the elevator includes the left inner rudder surface of the elevator and the left outer rudder surface of the elevator, and the right rudder surface of the elevator includes the right inner rudder surface of the elevator and the right outer rudder surface of the elevator. The actuator of the left outer rudder surface of the elevator is controlled by the left hydraulic source, the actuator of the right outer rudder surface of the elevator is controlled by the right hydraulic source, and the actuator of the left inner rudder surface of the elevator and the actuator of the right inner rudder surface of the elevator are controlled by the center hydraulic source.

[0062] 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 three hydraulic sources by two actuators in a master-slave mode.

[0063] Specifically, Figure 3 As 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, 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.

[0064] Step S2003: configure the rudder surface actuators 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 actuators of the upper rudder, the middle rudder, and the lower rudder are configured with the three hydraulic sources in a main-standby-standby manner.

[0065] Specifically, 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.

[0066] In the embodiment of the present disclosure, the dual-source failure in the hydraulic energy of the three hydraulic sources includes: left hydraulic source failure and middle hydraulic source failure, left hydraulic source failure and right hydraulic source failure, right hydraulic source failure and middle hydraulic source failure.

[0067] Specifically, according to the configuration of the aircraft control surface actuators and the three-way hydraulic source in the flight control system, the dual-source failure of the hydraulic source is divided into the following three scenarios.

[0068] Scenario 1: Left hydraulic power source failure and middle hydraulic power source failure;

[0069] At this time, the left elevator rudder surface is lost and unusable, and the right elevator rudder surface can only be controlled by the right outer elevator rudder surface to achieve aircraft pitch control;

[0070] The left aileron rudder surface is lost and unusable, and the aircraft roll control is achieved by controlling the right aileron rudder surface only by the right outer aileron rudder surface;

[0071] The rudder relies only on the lower rudder control to achieve aircraft heading control.

[0072] Scenario 2: Left hydraulic source failure and right hydraulic source failure;

[0073] 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, so as to realize the pitch control of the aircraft;

[0074] The left inner aileron control surface controls the left aileron control surface, and the right inner aileron control surface controls the right aileron control surface to achieve aircraft roll control;

[0075] The rudder is controlled only by the Chinese rudder to achieve aircraft heading control.

[0076] Scenario 3: Right hydraulic power source failure and middle hydraulic power source failure;

[0077] At this time, the right elevator rudder surface is lost and unusable, and the aircraft pitch control is achieved by controlling the left elevator rudder surface only by the left outer elevator rudder surface;

[0078] The right aileron rudder surface is lost and unusable, and the aircraft roll control is achieved by controlling the left aileron rudder surface only by the left outer aileron rudder surface;

[0079] The rudder is controlled only by the upper rudder to achieve aircraft heading control.

[0080] For the above fault scenarios, dual-source failure will cause the flight control system to control with minimum rudder, and the aircraft is in a highly dangerous state at this time, and only one set of hydraulic energy is available on board. At the same time, in order to avoid bringing operational burden to the crew when conducting dual-source failure tests, the crew's unnecessary 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 is implemented for the three scenarios of dual-source failure, where hydraulic source No. 1 is any one of the left hydraulic source, the middle hydraulic source, and the right hydraulic source, and hydraulic source No. 2 is any one of the two remaining hydraulic sources after hydraulic source No. 1 is selected.

[0081] Specifically, when a dual-source failure of the hydraulic source occurs, a control valve is correspondingly provided for each hydraulic source on the onboard control panel. 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, and the operation 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, and 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 a cut-off state, and the control instruction cannot perform closed-loop control in the servo circuit 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, thereby 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, its servo circuit is disconnected and the fault is latched.

[0082] The aircraft is equipped with an onboard test device, which is used to inject a hydraulic source fault signal into the flight control system. In the scenario of a single hydraulic source fault in the flight control system, the hydraulic source fault signal is injected to simulate the scenario of a dual hydraulic source fault in the flight control system. The test device consists of a fault injection panel and a fault forwarding computer. The fault injection panel is equipped with a left hydraulic source fault button, a middle hydraulic source fault button, and a right hydraulic source fault button. After pressing the corresponding button, the fault injection panel sends the fault signal to the fault forwarding computer. It should be understood that when the control button of the left hydraulic source is turned off on the control panel, the right hydraulic source fault button or the middle hydraulic source fault button needs to be pressed on the fault injection panel to achieve a dual source fault. The fault forwarding computer sends the hydraulic source failure signal to the flight control system. After the flight control system recognizes the fault signal, it cuts off the SOV valve of the flight control system actuator corresponding to the No. 2 hydraulic source. The No. 2 hydraulic source does not actually cut off the hydraulic pipeline, and the hydraulic energy of the No. 2 hydraulic source is normal. The control failure of the SOV valve of the actuator corresponding to the No. 2 hydraulic source is used to simulate the control failure of the SOV valve of the actuator corresponding to the second hydraulic source when the second hydraulic source fails in the real scene, forming a dual hydraulic source failure. This dual hydraulic source failure is not a real dual hydraulic source energy loss failure. It reflects the working characteristics of the main equipment and system of the flight control system when the dual source fails, ensures that the real hydraulic source redundancy on the aircraft is not lost, and ensures the safety of the crew during the high-risk subject test flight. It shows that the flight control system complies with the CCAR25.672(c) clause when the failure occurs, and has strong promotion.

[0083] It should be understood that among the hydraulic energy of 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, the hydraulic energy of hydraulic source No. 3 is normal, and the hydraulic energy of the three hydraulic sources are all in the pre-position state, and the crew performs the test flight action.

[0084] It is also necessary to understand that when exiting the dual-source failure scenario, the hydraulic energy of the three hydraulic sources is in the pre-position state, restarting the control computer corresponding to hydraulic source No. 1, restoring the SOV valve of the flight control system actuator corresponding to hydraulic source No. 1, and restoring the control command to perform closed-loop control in the servo circuit of its actuator, powering off the test equipment that injects the hydraulic source failure signal, restoring the SOV valve of the flight control system actuator corresponding to hydraulic source No. 2, and restoring the control command to perform closed-loop control in the servo circuit of its actuator. That is, the exit from the dual-source failure scenario of the hydraulic source is completed.

[0085] Example 1

[0086] 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.

[0087] At this time, close the control valve of the right hydraulic source in the onboard 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.

[0088] The flight control system injects a fault signal of the central hydraulic source through the onboard test equipment of the flight control system. The flight control system receives the command and actively cuts off the SOV valve of the actuator of the right inner rudder surface of the elevator. At this time, the right rudder surface of the elevator fails. The SOV valve of the actuator of the right inner rudder surface of the aileron is actively cut off. At this time, the right rudder surface of the aileron fails. The SOV valve of the actuator of the upper rudder is actively cut off. At this time, the upper rudder surface fails, and the central rudder controls the heading.

[0089] Open the right hydraulic source control valve of the onboard hydraulic system, restore the hydraulic pipeline of the right hydraulic source, 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, 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.

[0090] Confirm the dual-source failure scenario and determine the flight control system and aircraft characteristics that meet the actual dual-source failure.

[0091] Observe the warning information report on board:

[0092] The hydraulic source fails;

[0093] The right elevator failed;

[0094] The right aileron failed;

[0095] The rudder margin is reduced;

[0096] The crew performed test flight maneuvers to verify the control functions of pitch, roll and yaw.

[0097] The details are as follows:

[0098] Execute double 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.

[0099] Execute step ailerons, establish a stable turning flight with a bank angle of ±30 degrees, and perform a step 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.

[0100] 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.

[0101] At this time, the aircraft's maneuverability is as follows:

[0102] Using a rod displacement of about -94 to 80 mm, an overload of 1.26 to 0.6 g is achieved, and the longitudinal maneuverability is normal. Figure 5 shown.

[0103] The flight maneuver of the compliance criterion action is completed within 11 seconds, and the lateral maneuverability is normal. Figure 6 shown.

[0104] 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.

[0105] Through the above embodiment, the control valve of one hydraulic source among the three hydraulic sources is closed, the hydraulic pipeline of the hydraulic source is cut off, the flight control system cannot obtain the hydraulic energy of the hydraulic source, and the SOV valve of the actuator of the hydraulic source is cut off. At this time, the corresponding flight control system actuator loses control ability, and the control valve of the hydraulic source is opened, the hydraulic energy of the hydraulic source is restored, the hydraulic pipeline of the hydraulic source is restored, and the SOV valve of the actuator is still in a cut-off state. The control command cannot perform closed-loop control in the servo circuit of its actuator, and the other actuator on the same side becomes the "master" to control the movement of the rudder surface, so that the control failure of the actuator is used to simulate the real scenario of a single hydraulic source failure.

[0106] In the scenario of a single hydraulic source failure, a hydraulic source failure signal is injected to cut off the SOV valve of the flight control system actuator corresponding to the other hydraulic source. The hydraulic pipeline of this hydraulic source is not actually cut off, and the hydraulic energy of this hydraulic source is always available. The control command cannot perform closed-loop control in the servo circuit of its actuator, and the control failure of the actuator is used to simulate the real scenario of another hydraulic source failure. In this way, the scenario of dual hydraulic source failure is realized.

[0107] The hydraulic source dual-source failure is equivalent to the main characteristics of the flight control system when the real hydraulic source dual-source failure occurs on the aircraft, ensuring that the redundancy of the real hydraulic source on the aircraft is not lost, ensuring the safety of the crew during high-risk flight tests, and indicating that the flight control system complies with CCAR25.672(c) in the event of a failure.

[0108] 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 technical features indicated. Thus, 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 the present application, the meaning of "multiple" is two or more, unless otherwise clearly and specifically defined.

[0109] In the description of this specification, the description with reference to the terms "one embodiment", "some embodiments", "example", "specific example" or "some examples" etc. means 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 representations of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described may be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art may combine and combine the different embodiments or examples described in this specification.

[0110] Those skilled in the art will readily appreciate other embodiments of the present application after considering the specification and practicing the invention disclosed herein. The present application is intended to cover any modification, use or adaptation of the present application, which follows the general principles of the present application and includes common knowledge or customary technical means in the art that are not disclosed in the present application.

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 the cut-off state to simulate a single hydraulic source failure in a real scenario, including: 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 the No. 1 hydraulic source in the hydraulic system, restore the hydraulic pipeline of the No. 1 hydraulic source in the hydraulic system, restore the hydraulic energy of the No. 1 hydraulic source, and make the SOV valve of the flight control system actuator corresponding to the No. 1 hydraulic source still in the cut-off state; 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 to simulate a dual hydraulic source failure in a real scenario, including: 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 the No. 2 hydraulic source to be cut off, while the hydraulic pipeline of the No. 2 hydraulic source is not cut off, and the hydraulic energy of the No. 2 hydraulic source is normal. 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 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.

4. The test method for the compliance of the hydraulic source failure effect indicator of the fly-by-wire flight control system according to claim 3 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.

5. The test method for the hydraulic source failure impact indication compliance of the fly-by-wire flight control system according to claim 4 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.

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 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.

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: 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.

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 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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