Airplane control surface jamming fault test method and test modification system thereof
By injecting faults into the fly-by-wire flight control system through flight control methods, control commands for the control surfaces are generated. This solves the shortcomings of the traditional mechanical jamming method, realizes fault testing and reset at multiple state points, and improves the safety and efficiency of aircraft flight testing.
Patent Information
- Application Number
- CN202511599144.8
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-11-04
- Publication Date
- 2026-02-27
AI Technical Summary
Traditional mechanical methods of jamming control surfaces cannot achieve continuous setting of multiple fault state points in fly-by-wire flight control systems, and cannot be reset, which increases the number of test flights and manpower consumption, and affects the safety of aircraft takeoff and landing.
The system injects faults into flight tests using a flight control method, generates control commands for the control surfaces through the flight test control panel and modules, enables testing of multiple fault state points, and provides a reset function to avoid mechanical jamming.
This technology enables testing at multiple fault states without altering the fly-by-wire aircraft configuration, improving the safety and efficiency of aircraft flight testing while reducing the number of test flights and manpower consumption.
Smart Images
Figure CN121573202A_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The present application belongs to the technical field of aircraft flight test, and relates to a method for simulating aircraft rudder surface blocking failure in aircraft flight test, in particular to a method for aircraft rudder surface blocking failure test and a test modification system thereof. BACKGROUND
[0002] At present, more and more large and medium-sized aircrafts in China adopt fly-by-wire control, which will become the mainstream form of flight control system development. The fly-by-wire flight control system is different from the traditional pure mechanical control system or mechanical assisted control system. For the flight test subject: rudder surface blocking, the traditional mechanical device is used to block the rudder surface, which is not conducive to the take-off and landing of the aircraft, affects the safety of take-off and landing of the aircraft, and this failure setting method can only set one state point for one flight, cannot be set continuously, and cannot be reset, which increases the consumption of flight test times and human resources. SUMMARY
[0003] Therefore, the present application provides a method for aircraft rudder surface blocking failure test and a test modification system thereof, which does not change the entity configuration of the fly-by-wire aircraft control system, and completes the aircraft rudder surface blocking failure test by using the flight control method, solves the test of multiple failure state points in one flight test, and can reset in real time, thereby improving the safety of aircraft flight test.
[0004] The technical scheme of the present application is as follows: A method for aircraft rudder surface blocking failure test, in the flight test, the rudder surface blocking function is turned on, then the failure is injected into the aircraft transmission matching actuator control instruction, the actuator instruction bias is formed by combining the original actuator control instruction; after setting the blocking bias position, the blocking bias position and the initial blocking position increment are input into the actuator instruction.
[0005] Further, the actuator instruction bias m and the control law y(t) generated by the actuator control instruction are summed to obtain the rudder surface control instruction m+y(t)=y after the rudder surface blocking function is turned on, and the control instruction y is sent to the rudder surface control structure, and the rudder surface is controlled to the rudder surface deflection y at the failure injection time as the initial blocking position.
[0006] Further, after setting the blocking bias position, the increment of the input blocking bias position relative to the initial blocking position is taken as the variable x of the control law to generate the rudder surface control instruction z(t)=y+x, and the control instruction y+x is sent to the rudder surface control structure to control the rudder surface to the expected deflection y+x.
[0007] Further, the reset method is also provided, and after the reset signal is injected, the bias signals m and x injected by the failure are removed from the control law.
[0008] The test modification system of the aircraft control surface jamming fault test method comprises a flight test control panel, a flight test interface unit and a flight control module, the flight test interface unit is connected and controls the flight control module, the flight control module is provided with flight software and issues control instructions to control the control structure of the control surface; the flight test control panel is connected with the flight test interface unit, and the flight test control panel is provided with an input and output interface.
[0009] Further, the flight test control panel is provided with keys corresponding to specific control surfaces on the aircraft and a confirmation key, after pressing the specific control surface key and then pressing the confirmation key, the flight control module starts the jamming function test of the corresponding control surface.
[0010] Further, the flight test control panel is further provided with a digital keyboard and a display box, after inputting the jamming bias position in the digital keyboard and pressing the confirmation key, the flight control module inputs the variable x into the control instruction of the control surface.
[0011] Further, the flight test control panel is further provided with a reset key, after pressing the reset key, the flight control module immediately removes the bias signals m and x from the control law.
[0012] Technical effects: 1. Without changing the entity configuration of the airborne equipment, the flight test control panel, the flight test interface unit and the flight control module are used to realize the data transmission link, the flight test control panel is used as the instruction input to select the control surface jamming and the jamming parameters, and the flight control module is used as the basis of the flight test software module version to issue the control instructions to realize the jamming of different control surfaces.
[0013] 2. By avoiding the use of mechanical jamming structure, the present application can verify multiple jamming fault injection state points in one test flight, and solves the problem of flight resources. DETAILED DESCRIPTION
[0014] Figure 1 is a flight test control panel.
[0015] Figure 2 is a control surface jamming data transmission link diagram.
[0016] Figure 3 is a control surface jamming data processing flowchart. DETAILED DESCRIPTION
[0017] The features and exemplary embodiments of the various aspects of the present application will be described in detail below. In the following detailed description, numerous specific design details are set forth in order to provide a more thorough understanding of the present application. However, it will be apparent to one of ordinary skill in the art that the present application can be practiced without some or all of these specific details. The description of the embodiments is merely intended to provide a more detailed understanding of the present application. The present application is not limited to any particular setting or method described below, but covers any modifications, equivalents, and alternatives falling within the spirit and scope of the present application. In the drawings and the following description, well-known functions or constructions are not described in detail because they would obscure the application. Embodiments of the present application will be described with reference to the accompanying drawings, in which:
[0018] It should be noted that the embodiments of the present application and the features in the embodiments can be combined with each other without conflict, and each embodiment can be referred to and cited by each other. The present application will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.
[0019] Embodiment 1: An aircraft control surface jam fault test method, in a flight test, the control surface jam function is turned on, then the fault is injected into the aircraft transmission matching actuator control command, and the original actuator control command is combined to form an actuator command bias; after setting the jam bias position, the jam bias position and the initial jam position increment are input into the actuator command.
[0020] The actuator command bias m and the control law y(t) generated by the actuator control command are summed to obtain the control surface control command m+y(t)=y after the control surface jam function is turned on, and the control command y is sent to the control surface control structure. The control surface deflection y at the fault injection time is taken as the initial jam position.
[0021] After setting the jam bias position, the increment of the input jam bias position relative to the initial jam position is taken as the variable x of the control law, and the control surface control command z(t)=y+x is generated. The control command y+x is sent to the control surface control structure, and the control surface is controlled to the expected deflection y+x.
[0022] There is also a reset method, after the reset signal is injected, the bias signals m and x injected by the fault are removed from the control law.
[0023] A test modification system of an aircraft control surface jam fault test method, comprising a flight test control panel, a flight test interface unit and a flight control module, the flight test interface unit is connected and controls the flight control module, the flight control module is provided with flight software and sends control commands to control the control surface control structure; the flight test control panel is connected with the flight test interface unit, and the flight test control panel is provided with an input and output interface.
[0024] The flight test control panel is provided with keys corresponding to specific rudders on the aircraft and a confirmation key. After pressing the specific rudder key and then pressing the confirmation key, the flight control module starts the test of the corresponding rudder blocking function.
[0025] The flight test control panel is also provided with a digital keyboard and a display box. After inputting the blocking bias position in the digital keyboard and pressing the confirmation key, the flight control module inputs the variable x into the rudder control instruction.
[0026] The flight test control panel is also provided with a reset key. After pressing the reset key, the flight control module immediately removes the bias signals m and x from the control law.
[0027] Embodiment 2: The specific implementation of the technical solution of the application includes components such as Figure 2 As shown in the figure, the flight test control panel FTCP, the flight test interface unit FTIU, the test interface unit TIU, the flight control module FCM1, the flight control module FCM2, the flight control module FCM3, the cable 1, the cable 2, the cable 3, the cable 4, and the cable 5.
[0028] Data link definition: FTCP sets the fault mode and transmits it to FTIU through the 422 bus. FTIU connects with the main flight control FCM through the network cable and injects the fault signal into the system control loop.a) FTCP and FTIU communicate through the 422 bus. FTIU sends data frames to FTIU at a fixed period of 40 Hz. Each frame of data is fixed at 10 bytes.b) TIU and FCM communicate through the UDP protocol. FTIU sends data frames to FCM at a fixed period of 40 Hz. The specific link is shown in the attached Figure 2 .
[0029] Rudder blocking fault test flight software module version selection: through the test interface unit (TIU) linked to FCM2, different versions are selected, for example: 1, 2, 3…; representing different rudder blocks, for example: left elevator blocking, right elevator blocking, left aileron blocking, etc.
[0030] The surface jamming data transmission: the required surface jamming mode can be set through the "FTCP", different surface jamming mode buttons have the function of mutual exclusion, that is, only a single surface can be set to enter the jamming mode at a time. Taking the "left aileron jamming" as an example, after pressing the "left aileron jamming" button through the "FTCP", the button state quantity is checked and transmitted to the FCM through the data link, the FCM software records the current surface / actuator control instruction y, the actuator instruction bias (fault injection) m=y-y(t)+x is combined with the control law generated control law y(t) to obtain the final surface control instruction z(t)=m+y(t)=y (at this time there is no input digital keyboard x=0), the FCM sends the control instruction y to the ACE to control the surface to the surface deflection y at the jamming injection time; and the jamming bias position is set through the digital keyboard, which is the increment of the initial jamming position, the digital keyboard setting value is checked and transmitted to the FCM through the data link, the FCM recognizes the signal x to generate the surface control instruction z(t)=m+y(t)=y+x, and the FCM sends the control instruction y+x to the ACE to control the surface to the expected deflection y+x.
[0031] When the flight verification action is performed, the "reset" button of the "FTCP" is pressed, the FCM recognizes that the fault injection signal is eliminated, the fault injection function is cancelled, the bias signal m=y-y(t)+x of the fault injection is removed from the control law, and the FCM normally issues the control law calculation instruction y(t). The specific process is shown in the attached Figure 3 According to the above data link definition, surface jamming fault flight software module version selection and surface jamming data transmission method, while ensuring the installation and crew operation requirements, the installation test has been completed on a certain type of aircraft, the surface jamming fault injection flight test requirements have been realized, the surface jamming flight test subjects have been completed, and the expected effect has been achieved.
[0032] The above only describes the preferred embodiments of the present application, and it should be noted that those skilled in the art can make several improvements without departing from the principles of the present application, and these improvements should also be considered as the protection scope of the present application.
Claims
1. An aircraft control surface jam fault test method, characterized by, In the flight test, the stuck function of the control surface is opened, then the fault is injected into the aircraft transmission matching actuator control command, and the original actuator control command is combined to form the actuator command bias; after setting the stuck bias position, the stuck bias position and the initial stuck position increment are input into the actuator command.
2. The method of claim 1, wherein, The actuator command bias m and the control law y(t) generated by the actuator control command are summed to obtain the control command m+y(t)=y of the control surface after the stuck function of the control surface is opened, and the control command y is sent to the control surface control structure. The deflection y of the control surface at the fault injection time is used as the initial stuck position.
3. The method of claim 1, wherein, After setting the stuck bias position, the increment of the input stuck bias position relative to the initial stuck position is used as the variable x of the control law, and the control command z(t)=y+x of the control surface is generated. The control command y+x is sent to the control surface control structure, and the control surface is controlled to the expected deflection y+x.
4. The method of claim 1, wherein, There is also a reset method, after the reset signal is injected, the bias signals m and x injected by the fault are removed from the control law.
5. A test modification system for a method of testing a flight control surface jam fault, according to any one of claims 1 to 4, wherein the method is performed. The flight test control panel, the flight test interface unit and the flight control module are included, the flight test interface unit is connected and controls the flight control module, the flight software is arranged in the flight control module and the control command is issued to control the control surface control structure; the flight test control panel is connected with the flight test interface unit, and the flight test control panel is provided with an input-output interface. 6. The test retrofit system for a flight control surface jam fault test method of claim 5, wherein, The flight test control panel is provided with keys corresponding to specific control surfaces on the aircraft and a confirmation key. After pressing the specific control surface key and then pressing the confirmation key, the flight control module opens the stuck function test of the corresponding control surface.
7. The test retrofit system for a flight control surface jam fault test method of claim 6, wherein, The flight test control panel is also provided with a digital keyboard and a display box. After inputting the stuck bias position in the digital keyboard and pressing the confirmation key, the flight control module inputs the variable x into the control command of the control surface.
8. The test retrofit system for a flight control surface jam fault test method of claim 6, wherein, The flight test control panel is also provided with a reset key. After pressing the reset key, the flight control module immediately removes the bias signals m and x from the control law.