A system for monitoring structural stress changes in aircraft and its application method

Abstract

This invention relates to the field of aircraft structural stress change detection technology, and particularly to a system for monitoring structural stress changes in aircraft and its usage method. It solves the problem that existing aircraft protective supports cannot automatically control and detect the aircraft's stress status in real time. The technical solution is as follows: A system for monitoring structural stress changes in aircraft includes: a ground operation module, multiple protection and support adjustment modules, and multiple attitude detection modules; the ground operation module, multiple protection and support adjustment modules, and attitude detection modules are connected wirelessly; a usage method for monitoring structural stress changes in aircraft includes the following steps: S1: manual mode operation; S2: adjusting the aircraft state to its original gravity state; S3: adjusting the aircraft state to its original attitude state; S4: automatic mode and system fault operation state. The beneficial effects of this invention are: this invention can monitor the stress changes of the aircraft in real time, effectively protecting the structural safety of the aircraft.

Description

Technical Field

[0001] This invention relates to the field of aircraft structural stress change detection technology, and in particular to a system for monitoring structural stress changes in aircraft and its usage method. Background Technology

[0002] As international trade gradually recovers, air transport, due to its speed and safety, has become one of the preferred long-distance transportation options for freighters. However, with the approaching retirement of the previous generation of cargo aircraft and the completion of previous passenger-to-freighter conversion projects, some cargo airlines and leasing companies believe that a portion of their cargo fleets may face retirement, potentially leading to a shortage of cargo fleet capacity. All-cargo aircraft currently account for approximately 61% of the market share and are gradually increasing, far exceeding earlier trends.

[0003] Cargo aircraft are mainly converted from retired passenger planes, commonly known as "passenger-to-cargo conversion". During the conversion process, the aircraft structure undergoes a lot of disassembly, replacement, removal and cutting. The requirements for structural changes during the construction process are very strict.

[0004] Current protection methods involve placing a fixed support on the bottom of the aircraft and fixing several observation points around it to conduct interval checks using instruments. However, instrument observation alone cannot detect changes in the stress on the aircraft structure in a timely manner. In the event of system failures (such as sudden hydraulic pressure loss in the protection device or sudden rapid movement of the actuator) or emergencies (including system failures and external factors, such as the sudden falling of aircraft parts), the aircraft structure may have already deformed or its stress may have changed. Measuring at this point could lead to permanent damage to the aircraft structure. Furthermore, existing protection supports cannot automatically control and monitor the aircraft's stress status in real time, nor can they record and analyze data during the protection process to determine the location of the problem. The existing protection devices also have slow response times and cannot effectively protect the aircraft in the event of system failures or emergencies.

[0005] How to solve the above problems is the challenge faced by this invention. Summary of the Invention

[0006] The purpose of this invention is to provide a system and method for monitoring structural stress changes in aircraft.

[0007] To achieve the above-mentioned objectives, the present invention adopts the following technical solution:

[0008] A system for monitoring structural stress changes in an aircraft includes: a ground operation module, multiple protection and support adjustment modules, and multiple attitude detection modules; the ground operation module, multiple protection and support adjustment modules, and attitude detection modules are connected wirelessly; the number of protection and support adjustment modules and attitude detection modules is specifically set according to the aircraft model.

[0009] The ground operation module includes a manual operation panel, a mode conversion knob, a status display screen, an audible and visual alarm, and a core controller;

[0010] The mode switching knob is electrically connected to the core controller and is used to select the operation mode of the system in manual, automatic and emergency modes;

[0011] The manual control panel is electrically connected to the core controller. The protection and support adjustment module can only be operated via the manual control panel when the mode switch knob is selected in manual mode.

[0012] When the mode switch knob is manually selected to automatic mode, the system will control the protection and support adjustment module.

[0013] In an emergency, the manual operation mode switch knob is turned to emergency mode. The system will stop the protection and support adjustment modules from working according to the pre-set program and maintain them in the current state. The system will then stop operating.

[0014] The actuator is activated and the lifting actuator cylinder is raised; the actuator support is locked in the current position.

[0015] The status display screen is electrically connected to the core controller and is used to display various operating parameters of the protection and support adjustment module and the attitude detection module in real time.

[0016] Specifically, this includes the aircraft's pitch value, roll value, pressure value of the pressure detection module, actuator 8, lifting actuator cylinder, fixed support, and position parameters of the actuator support.

[0017] The audible and visual alarm is electrically connected to the core controller. When system parameters exceed a certain threshold or a system fault occurs, it will emit sound and light to the outside world. The threshold is set as Gb = Gs - Gy > 200 or pitch value > 0.02 or tilt value > 0.02. The light is set to red, and the alarm sound is: "Attitude change, please monitor".

[0018] The protection and support adjustment module includes a pressure detector, an actuator, an actuator feedback detector, and a lifting actuator.

[0019] The pressure detector is fixedly installed inside the actuator. The upper end of the actuator cylinder feedback detector is fixedly connected to the actuator. The upper end of the lifting actuator cylinder is connected to the actuator.

[0020] The actuator includes a fixed support and an actuator support. The pressure detection module detects and reads the pressure data of the fixed support and sends it to the core controller for the operation control of the actuator.

[0021] The attitude detection module is mounted on the upper part of the fuselage via a suction cup and includes a tilt sensor and a pitch sensor. The two sensors detect the aircraft's tilt and pitch parameters respectively and transmit them to the core controller for system control.

[0022] A method for monitoring structural stress changes in an aircraft includes the following steps:

[0023] S1: Manual mode operation;

[0024] S2; Adjust the aircraft's state to its original gravity state;

[0025] S3: Adjust the aircraft status back to its original attitude state;

[0026] S4: Automatic mode and system fault operation status.

[0027] Step S1 specifically includes:

[0028] Select the mode switch knob to manual mode, move the lifting actuator cylinder from the maintenance warehouse to the bottom of the aircraft, and manually control the lifting actuator cylinder to rise through the manual operation panel, thereby driving the actuator to fit against the aircraft fuselage. The actuator feedback detector collects the displacement of the lifting actuator cylinder and transmits it to the core controller, which displays it on the status display screen in real time.

[0029] Step S2 specifically includes:

[0030] The aircraft is lifted by the manual remote control lifting actuator until the tires are off the ground. At this time, the fixed support bears the weight of the aircraft, and the pressure detector reads the current data value and transmits it to the core controller. The core controller records various parameters of the current protection and support adjustment module, including the pressure value of the pressure detector, the position parameters of the actuator, the lifting actuator, and the actuator support, and the aircraft gravity recorded by the fixed support. At this time, the parameter values ​​are referred to as the original gravity values ​​of the aircraft in this system. At this time, the aircraft is under balanced forces, which is equivalent to the state when the aircraft is on the ground, and is called the original gravity state.

[0031] Step S3 specifically includes:

[0032] When the aircraft is in its initial gravity state, if the pitch and roll values ​​in the aircraft's internal systems are not at zero, the elevator actuator is manually controlled to lift the aircraft. When the pitch and roll values ​​in the aircraft's internal systems are both at zero, control is stopped. At this time, the roll and pitch sensors in the attitude detection module detect and record the current aircraft roll and pitch values ​​and transmit them to the core controller. The parameter values ​​at this time are called the aircraft's initial attitude values ​​in this system, or the initial attitude state.

[0033] The pitch and roll values ​​in the aircraft's internal systems may be 0, but the sensor readings may not be 0. However, if the change in the original attitude value reaches a set threshold, the system will activate.

[0034] Step S4 specifically includes:

[0035] When the aircraft is in its original gravity and attitude states, the original gravity and attitude values ​​are transmitted to the core controller. The operator manually selects the automatic mode using the mode switching knob, and the system switches to automatic control. The elevator actuator and actuator support begin to work. That is, the elevator actuator and actuator support work according to the different values ​​of Gb, the original attitude values, and the ratio of Gs to Gy.

[0036] Where: Gb = Gs - Gy;

[0037] In the formula, Gb is the change in structural force, Gs is the real-time detection data, and Gy is the original value of gravity in KG. When the value of Gb is greater than 200 and the original value of attitude remains unchanged, the actuator support (12) will start to work, so that the value of Gb is ≤200. After the aircraft is supported, there are staff members walking and working on it. Under normal circumstances, the weight of the three people is about 200KG.

[0038] When the change in the original attitude value reaches 0.02, regardless of the Gb value, the lifting actuator and the actuator support will start working until the attitude value is consistent with the original attitude value and then stop working.

[0039] Once the aircraft has climbed to a position where its tilt and pitch values ​​are both 0, they generally will not change. However, when the initial attitude values ​​begin to change until they reach ≥0.02, it indicates that the aircraft's attitude has changed and will affect the structure. At this point, the system needs to work to correct the attitude.

[0040] The system automatically compares the Gs value with the Gy value. When the ratio of Gs to Gy exceeds a certain value (i.e., Gs:Gy = 1.2), the system alarms. At the same time as the alarm, the actuator supports work again until the Gs value reaches the Gy value and then stops.

[0041] When the actuator feedback detection module detects a fault in the lifting actuator system, the system automatically stops and runs and issues an alarm. At this time, the manual operation mode switch knob is selected to manual mode.

[0042] Compared with the prior art, the beneficial effects of the present invention are as follows:

[0043] 1. The number of protection and support adjustment modules of this invention can be freely grouped according to the size and type of the machine, and can be flexibly adjusted according to work needs without the need for interval settings.

[0044] 2. This invention can monitor and display the stress changes and stress trends of the aircraft structure in real time, solving the problem that the aircraft structure is prone to change and cannot be monitored and analyzed during the "passenger-to-cargo" conversion process.

[0045] 3. This invention can determine the location of sudden changes in the stress on the aircraft structure and effectively protect the aircraft by using protection and support adjustment modules and audible and visual warning devices.

[0046] 4. The lifting actuator cylinder and actuator support of the present invention can work according to the different values ​​of Gb, original attitude values, and the ratio of Gs to Gy, so as to effectively protect the aircraft.

[0047] 5. During the "passenger-to-freighter" conversion process, system malfunctions or emergencies may occur. In such cases, the manual operating system should be shut down to prevent damage to the aircraft caused by system malfunctions. Attached Figure Description

[0048] The accompanying drawings are provided to further illustrate the invention and form part of the specification. They are used together with the embodiments of the invention to explain the invention and do not constitute a limitation thereof.

[0049] Figure 1 This is a schematic diagram illustrating the system principle of the aircraft structure monitoring stress changes of the present invention.

[0050] Figure 2 This is a schematic diagram showing the structure and location of the system for monitoring structural stress changes in an aircraft according to the present invention.

[0051] Figure 3 This is a schematic diagram of the method for monitoring structural stress changes in an aircraft according to the present invention.

[0052] The attached diagram is labeled as follows: 1-Manual operation panel; 2-Mode conversion knob; 3-Status display screen; 4-Audio and visual alarm; 5-Core controller; 6-Attitude detection module; 7-Pressure detector; 8-Actuation mechanism; 9-Actuation mechanism feedback detector; 10-Lifting actuator cylinder; 11-Fixed support; 12-Actuation mechanism support; 13-Tilt sensor; 14-Pitch sensor. Detailed Implementation

[0053] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and embodiments. Of course, the specific embodiments described herein are merely illustrative and not intended to limit the invention.

[0054] Example 1

[0055] See Figure 1 and Figure 2The present invention provides a technical solution as follows: a system for monitoring structural stress changes in an aircraft, comprising: a ground operation module, multiple protection and support adjustment modules, and multiple attitude detection modules 6; the ground operation module, multiple protection and support adjustment modules, and attitude detection modules 6 are connected wirelessly; the number of protection and support adjustment modules and attitude detection modules 6 is specifically set according to the aircraft model; for example, the Airbus A320 aircraft uses 18 protection and support adjustment modules and 4 attitude detection modules 6.

[0056] The ground operation module includes a manual operation panel 1, a mode conversion knob 2, a status display screen 3, an audible and visual alarm 4, and a core controller 5;

[0057] The mode switching knob 2 is electrically connected to the core controller 5 and is used to select the operation mode of the system in three modes: manual, automatic and emergency.

[0058] The manual operation panel 1 is electrically connected to the core controller 5. The protection and support adjustment module 6 can only be operated through the manual operation panel 1 when the mode conversion knob 2 is selected in manual mode.

[0059] When the mode switch knob 2 is manually selected to automatic mode, the system will control the protection and support adjustment module.

[0060] In an emergency, manually switch the operation mode selector knob 2 to emergency mode. The system will stop the protection and support adjustment modules from working according to the pre-set program and maintain them in the current state. The system will then stop operating.

[0061] Actuating mechanism 8, lifting actuator cylinder 10; actuator support 12 locked in the current position.

[0062] The status display screen 3 is electrically connected to the core controller 5 and is used to display various working parameters of the protection and support adjustment module and the attitude detection module 6 in real time.

[0063] Specifically, this includes the aircraft's pitch value, roll value, pressure value of pressure detection module 7, and position parameters of actuator 8, lifting actuator cylinder 10, fixed support 11, and actuator support 12.

[0064] The audible and visual alarm 4 is electrically connected to the core controller 5. When the system parameters exceed a certain threshold or a system fault occurs, it will emit sound and light to the outside world. The threshold is set as Gb = Gs - Gy > 200 or pitch value > 0.02 or tilt value > 0.02. The light is set to red, and the alarm sound is: "Attitude change, please monitor".

[0065] The protection and support adjustment module includes a pressure detector 7, an actuation mechanism 8, an actuation cylinder feedback detector 9, and a lifting actuation cylinder 10;

[0066] Pressure detector 7 is fixedly installed inside actuator 8. The upper end of actuator cylinder feedback detector 9 is fixedly connected to actuator 8. The upper end of lifting actuator cylinder 10 is connected to actuator 8.

[0067] The actuator 8 includes a fixed support 11 and an actuator support 12. The pressure detection module 7 detects and reads the pressure data of the fixed support 11 and sends it to the core controller 5 for the operation control of the actuator 8.

[0068] The attitude detection module 6 is mounted on the upper part of the fuselage via a suction cup and includes a tilt sensor 13 and a pitch sensor 14. The two sensors detect the tilt and pitch parameters of the aircraft respectively and transmit them to the core controller 5 for system control.

[0069] refer to Figure 3 A method for monitoring structural stress changes in aircraft, specifically including the following steps:

[0070] S1: Manual mode operation;

[0071] S2; Adjust the aircraft's state to its original gravity state;

[0072] S3: Adjust the aircraft status back to its original attitude state;

[0073] S4: Automatic mode and system fault operation status.

[0074] Step S1 specifically includes:

[0075] Select the mode conversion knob 2 to manual mode, move the lifting actuator 10 from the maintenance warehouse to the bottom of the aircraft, and manually control the lifting actuator 10 to rise through the manual operation panel 1, thereby driving the actuator 8 to fit against the aircraft fuselage. The actuator feedback detector 9 collects the displacement of the lifting actuator 10 and transmits it to the core controller 5, which displays it on the status display screen 3 in real time.

[0076] Step S2 specifically includes:

[0077] The manual remote control lift actuator 10 raises the aircraft until the tires are off the ground. At this time, the fixed support 11 bears the weight of the aircraft, the pressure detector 7 reads the current data value and transmits it to the core controller 5. The core controller 5 records various parameters of the current protection and support adjustment module, including the pressure value of the pressure detector 7, the position parameters of the actuator 8, the lift actuator 10, and the actuator 12 support, and the aircraft gravity recorded by the fixed support 11. At this time, the various parameter values ​​are referred to as the original gravity values ​​of the aircraft in this system. At this time, the aircraft is under balanced forces, which is equivalent to the state when the aircraft is on the ground, and is called the original gravity state.

[0078] Step S3 specifically includes:

[0079] When the aircraft is in its initial gravity state, if the pitch and roll values ​​in the aircraft's internal system are not at zero, the elevator actuator 10 is manually controlled to lift the aircraft. When the pitch and roll values ​​in the aircraft's internal system are both at zero, the control is stopped. At this time, the roll sensor 13 and pitch sensor 14 in the attitude detection module 6 detect and record the current aircraft roll and pitch values ​​and transmit them to the core controller 5. The parameter values ​​at this time are called the aircraft's initial attitude values ​​in this system, or the initial attitude state.

[0080] The pitch and roll values ​​in the aircraft's internal systems may be 0, but the sensor readings may not be 0. However, if the change in the original attitude value reaches a set threshold, the system will activate.

[0081] Step S4 specifically includes:

[0082] When the aircraft is in its original gravity and attitude states, the original gravity and attitude values ​​are transmitted to the core controller 5. The operator manually selects the automatic mode using the mode switching knob 2, and the system switches to automatic control. The elevator actuator 10 and the actuator support 12 then begin to work. That is, the elevator actuator 10 and the actuator support 12 work according to the different values ​​of Gb, the original attitude values, and the ratio of Gs to Gy.

[0083] Where: Gb = Gs - Gy;

[0084] In the formula, Gb is the change in structural force, Gs is the real-time detection data, and Gy is the original gravity value in KG. When the Gb value is greater than 200 and the original attitude value remains unchanged, the actuator support 12 will start to work to make the Gb value ≤ 200. After the aircraft is lifted up, staff members walk and work on it. Under normal circumstances, the weight of the three people is about 200KG.

[0085] When the change in the original attitude value reaches 0.02, regardless of the Gb value, the lifting actuator cylinder 10 and the actuator support 12 will start working until the attitude value is consistent with the original attitude value and then stop working.

[0086] Once the aircraft has climbed to a position where its tilt and pitch values ​​are both 0, they generally will not change. However, when the initial attitude values ​​begin to change until they reach ≥0.02, it indicates that the aircraft's attitude has changed and will affect the structure. At this point, the system needs to work to correct the attitude.

[0087] The system automatically compares the Gs value with the Gy value. When the ratio of Gs to Gy exceeds a certain value (Gs:Gy = 1.2), the system alarms. At the same time as the alarm, the actuator support 12 starts working again until the Gs value reaches the Gy value and then stops.

[0088] When the actuator feedback detection module 9 detects a system malfunction in the lifting actuator cylinder 10, the system automatically stops and runs and issues an alarm. At this time, the manual operation mode conversion knob 2 selects the manual mode.

[0089] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A system for monitoring changes in stress in an aircraft structure, characterized by, The application relates to a ground operation module, a plurality of protection and support adjustment modules and a plurality of attitude detection modules (6). The ground operation module, the plurality of protection and support adjustment modules and the attitude detection module (6) are connected through wireless signals; the number of the protection and support adjustment modules and the attitude detection module (6) is set according to the type of the airplane. The ground operation module comprises a manual operation panel (1), a mode conversion knob (2), a state display screen (3), an audible and light alarm (4) and a core controller (5). The mode conversion knob (2) is electrically connected with the core controller (5) and is used for selecting the operation mode of the system in manual mode, automatic mode and emergency mode. The manual operation panel (1) is electrically connected with the core controller (5) and is used for operating the protection and support adjustment module through the manual operation panel (1) only when the mode conversion knob (2) selects the manual mode. When the mode conversion knob (2) is manually selected in the automatic mode, the protection and support adjustment module is controlled by the system. In an emergency, the mode conversion knob (2) is manually selected in the emergency mode, the protection and support adjustment module is stopped by the system according to the preset program, the system is stopped and the current state is kept. The state display screen (3) is electrically connected with the core controller (5) and is used for displaying the working parameters of the protection and support adjustment module and the attitude detection module (6) in real time. The audible and light alarm (4) is electrically connected with the core controller (5) and sends out sound and light to the outside when the system parameters exceed a certain threshold value or the system fails. The protection and support adjustment module comprises a pressure detector (7), an execution actuating mechanism (8), an execution actuating cylinder feedback detector (9) and a lifting actuating cylinder (10). The pressure detector (7) is fixedly installed in the execution actuating mechanism (8), the upper end of the execution actuating cylinder feedback detector (9) is fixedly connected with the execution actuating mechanism (8) and the upper end of the lifting actuating cylinder (10) is connected with the execution actuating mechanism (8). The execution actuating mechanism (8) comprises a fixed support (11) and an actuating mechanism support (12), the pressure detector (7) detects and reads the pressure data of the fixed support (11) and feeds the data to the core controller (5) for the working control of the execution actuating mechanism (8). The attitude detection module (6) is installed on the upper part of the airplane body through a suction cup and comprises an inclination sensor (13) and a pitching sensor (14); the two sensors detect the inclination state and the pitching state parameters of the airplane and transmit the parameters to the core controller (5) for the control of the system.

2. The system for monitoring changes in stress in an aircraft structure according to claim 1, wherein, The application further discloses a method for operating the airplane.

3. A method of using a system for monitoring changes in stress in an aircraft structure according to any one of claims 1-2, characterized in that, The method comprises the following steps: S1: manual mode operation; S2: adjusting the state of the airplane to a gravity original state; S3: adjusting the state of the airplane to an attitude original state; S4: working state of the automatic mode and system failure.

4. The method of using a system for monitoring changes in stress in an aircraft structure of claim 3, wherein, Step S1 specifically comprises the following steps: The mode conversion knob (2) is selected in the manual mode, the lifting actuating cylinder (10) is manually remotely controlled through the manual operation panel (1) to ascend, the execution actuating mechanism (8) and the airplane body surface are adhered, the displacement of the lifting actuating cylinder (10) is collected by the execution actuating mechanism feedback detector (9) and is transmitted to the core controller (5) and displayed on the state display screen (3) in real time.

5. The method of using a system for monitoring changes in stress in an aircraft structure of claim 4, wherein, Step S2 specifically includes: Manual remote control lifting cylinder (10) will be lifted to the tire off the ground, at this time the fixed support (11) to bear the weight of the aircraft, pressure detector (7) read the current data value and transmission to the core controller (5), the core controller (5) record the current protection and support adjustment module parameters, including the pressure value of the pressure detector (7), the position parameters of the execution mechanism (8), lifting cylinder (10), actuator support (12) support, fixed support (11) record the weight of the aircraft; At this time the parameter value in the system is called the original value of the aircraft gravity, at this time the aircraft is balanced, which is equivalent to the state of the aircraft on the ground, called the original state of gravity.

6. The method of using a system for monitoring changes in stress in an aircraft structure of claim 5, wherein, Step S3 specifically includes: When the aircraft is in the original state of gravity, if the pitch value, tilt value and horizontal reading in the internal system of the aircraft are not zero, manual remote control lifting cylinder (10) will be lifted, when the pitch value, tilt value and horizontal reading in the internal system of the aircraft are zero, stop control, at this time the tilt sensor (13) and pitch sensor (14) in the attitude detection module (6) detect and record the current tilt and pitch value of the aircraft and transmit it to the core controller (5), at this time the parameter value in the system is called the original value of the aircraft attitude, called the original attitude state.

7. The method of using a system for monitoring changes in stress in an aircraft structure of claim 6, wherein, Step S4 specifically includes: When the aircraft is in the original state of gravity and attitude, the original value of the aircraft gravity and the original value of the aircraft attitude are transmitted to the core controller (5), the mode conversion knob (2) is selected to the automatic mode by manual operation, the system is automatically controlled, and the lifting cylinder (10) and the actuator support (12) start to work; That is, the lifting cylinder (10) and the actuator support (12) work according to the value of Gb, the original value of the attitude, the ratio of Gs and Gy; Where: Gb=Gs-Gy; In the formula, Gb is the structural stress change, Gs is the real-time detection data, Gy is the original value of gravity, unit is KG, when the value of Gb is greater than 200 and the original value of attitude is unchanged, the actuator support (12) will start to work, so that the value of Gb is less than or equal to 200; When the change of the original value of attitude reaches 0.02, no matter what the value of Gb is, the lifting cylinder (10) and the actuator support (12) will start to work until the attitude value is consistent with the original attitude value; The system automatically compares the value of Gs with the value of Gy, when the ratio of Gs and Gy exceeds a certain value, the system alarms, at the same time the actuator support (12) works again until the value of Gs reaches the value of Gy; When the execution actuator feedback detector (9) detects that the lifting cylinder (10) system fails, the system automatically stops and runs and issues an alarm, at this time the mode conversion knob (2) is selected to the manual mode by manual operation.

Images