A device and method for monitoring the state of a repaired composite helicopter skin

Abstract

This application provides a method for monitoring the condition of helicopter composite skin after repair. The method includes the following steps: distributing sensors in the patch area and the intact area near the patch; the sensors collect and store vibration wave information in real time, and the recorded vibration wave information is used for data comparison; the vibration wave information is transmitted to a storage processor through a signal shielded cable; comparing the vibration wave information obtained at different points to determine whether the interface has begun to fail, and the comparison is performed between vibration wave information collected by multiple pairs of sensors, divided into a reference group and an interface group; by comparing the vibration wave characteristic difference data of the interface group and the reference group, the condition monitoring of helicopter composite skin after repair is realized; at the same time, this application also provides a device for monitoring the condition of helicopter composite skin after repair.

Description

Technical Field

[0001] This application belongs to the field of composite material failure monitoring technology, and in particular relates to a device and method for monitoring the condition of helicopter composite material skin after repair. Background Technology

[0002] Helicopters, as multi-purpose aircraft, are widely used in various fields. However, their unique flight characteristics and mission scenarios make them more susceptible to damage than other aircraft. In emergencies, necessary damaged parts are repaired before the helicopter can immediately return to the mission. Currently, some advanced helicopters use composite materials accounting for over 50% of their structure; therefore, emergency repairs for helicopters primarily target composite materials damaged by external impacts.

[0003] For composite material structures in helicopters, emergency repair methods after damage generally rely on riveting and adhesive bonding of composite patches. Due to the time constraints of missions, rapid repair of damaged structures is required, preventing the high-reliability maintenance achieved through routine checks. Temporary emergency repair structures inevitably carry a probability of failure during flight. Although emergency repair techniques have seen improvements in reliability through extensive research, for composite structures with aerodynamic functions, such as helicopter horizontal stabilizers, failure at the repaired site not only affects mission performance but also poses a significant threat to flight safety.

[0004] Therefore, after emergency repairs, key areas should be dynamically monitored to prevent sudden failures without warning. Normally, this monitoring process is only detected by the pilot through visual observation or by the impact of the repaired area's failure on the helicopter control system, which is highly unstable and delayed. Therefore, there is a need for a low-cost, highly reliable method and device for early warning of failures in composite material repair structures, capable of providing early warnings when failure is imminent or when the quality of emergency repairs changes, thereby improving helicopter flight safety. Summary of the Invention

[0005] Purpose of the invention: To conduct low-cost and highly reliable failure state monitoring of helicopter composite material skin structures after emergency repairs, thereby enabling early warning and response to failures in emergency repaired parts.

[0006] In a first aspect, this application provides a method for monitoring the condition of a helicopter composite skin after repair, the method comprising the following steps:

[0007] Sensors are distributed in the patch area and the intact area near the patch;

[0008] The sensor collects and stores vibration wave information in real time, and the recorded vibration wave information is used for data comparison.

[0009] The vibration wave information is transmitted to the storage processor via a signal shielded cable;

[0010] The vibration wave information obtained from different points is compared to determine whether the interface has begun to fail. The comparison is carried out between vibration wave information collected by multiple pairs of sensors, and is divided into a reference group and an interface group.

[0011] By comparing the vibration wave characteristic difference data of the interface group and the reference group, the condition monitoring of the helicopter composite skin after repair can be realized.

[0012] Preferably, the vibration wave characteristic difference data includes the amplitude, phase, coherence, and envelope characteristics of the vibration wave.

[0013] Preferably, the sensor is a high-precision piezoelectric ceramic sensor, and its sensing core has a shear-type structure.

[0014] Preferably, the distributed sensors in the patch area and the intact area near the patch include:

[0015] In the patch area, vibration characteristic information of only one target point is collected, which is located at the geometric center of the repair area;

[0016] The distribution rule for the target collection points in the intact area near the patch is as follows: they are distributed along the boundary outline between the edge of the patch and the intact area of ​​the patch. The number of points varies depending on the size of the emergency repair, but the total number is odd and the distribution is equidistant.

[0017] Preferably, the reference group is generated from the vibration wave information collected by the sensor in the intact area near the patch; the interface group is generated between the vibration wave information collected by the sensor in the patch area and the vibration wave information collected by the sensor in the intact area near the patch.

[0018] Preferably, the reference group represents the vibration wave characteristics of the undamaged part; the interface group represents the vibration wave characteristics between different sampling points across the interface.

[0019] Secondly, this application also provides a condition monitoring device for helicopter composite skin after repair, the device comprising:

[0020] Sensors are distributed in the patch area and the intact area near the patch;

[0021] The sensor is used to collect and store vibration wave information in real time, and the recorded vibration wave information is used for data comparison.

[0022] The storage processor acquires and stores vibration wave information, and compares the vibration wave information obtained from different points to determine whether the interface has begun to fail. The comparison is performed between vibration wave information collected by multiple pairs of sensors, divided into a reference group and an interface group. By comparing the vibration wave characteristic difference data between the interface group and the reference group, the condition monitoring of the helicopter composite skin after repair is realized.

[0023] Preferably, the device further includes:

[0024] The signal shielded cable transmits vibration wave information to the storage processor.

[0025] The beneficial technical effects of this application are as follows:

[0026] This application provides a device and method for monitoring the condition of helicopter composite material skin after emergency repair. It proposes a method and device for monitoring the failure of emergency repaired areas by monitoring the vibration waves generated by the helicopter itself during flight. Vibration wave monitoring sensors are arranged at specific locations, and the characteristics of the vibration waves obtained at different points are compared to determine whether interface failure has begun. This comparison is performed between vibration waves collected by multiple pairs of sensors. The comparison is divided into a reference group and an interface group. The reference group consists of sensor vibration wave signals generated in the intact area near the patch, while the interface group consists of sensor vibration waves generated in the patch area and sensor vibration waves generated in the intact area near the patch. By comparing the differences in vibration wave characteristics between the interface group and the reference group, a low-cost and highly reliable method and device for early warning of composite material repair structure failure is achieved. Attached Figure Description

[0027] Figure 1 This is a flowchart of the emergency repair post-repair status monitoring method provided in a specific embodiment of the present invention;

[0028] Figure 2 This is a diagram illustrating the deployment of vibration wave acquisition points according to a specific embodiment of the present invention;

[0029] Figure 3 This is a schematic diagram of a vibration wave data comparison method provided in a specific embodiment of the present invention. Detailed Implementation

[0030] Please see Figures 1-3 A device and method for monitoring the condition of helicopter composite skin after emergency repair are provided, comprising the following steps:

[0031] Step one: The basic principle of this device originates from the vibrations generated by the helicopter's own engine and transmission system during flight, as well as the vibrations caused by the flapping and swaying of the rotor blades and the action of airflow. These vibrations are transmitted to various structural parts of the fuselage, thus causing corresponding vibrations.

[0032] Step two: As we can understand from step one, most of the vibrations in the helicopter's fuselage structure are not spontaneous vibrations, but rather vibrations transmitted from the engine and rotor components. Therefore, these vibration waves are transmitted throughout most parts of the fuselage structure.

[0033] Step three: The conclusion in step two that vibration waves are transmitted across most of the fuselage structure establishes the following fundamental fact: Damaged composite material structures repaired through riveting, bonding, or other patching methods, compared to undamaged areas, inevitably affect vibration wave transmission due to the presence of repair interface connections and geometric changes. If these repair interfaces begin to show signs of failure during helicopter flight, this failure will be reflected in the vibration wave transmission characteristics. By detecting these vibration wave transmission characteristics, the initial changes in interface failure can be detected in advance, thus providing early warning of failure at emergency repair sites.

[0034] Step four: Firstly, a high-precision piezoelectric ceramic sensor is preferred, with a shear-type structure for its sensing core.

[0035] Step 5: The high-precision piezoelectric ceramic sensor is only used to collect the vibration characteristics of the target points of the structure. For a locally repaired structure, the single-point monitoring location is in the patch distribution area and the intact area near the patch.

[0036] Step six: The distribution of sensors in the patch area and the intact area near the patch is different. In the patch area, vibration characteristic information of only one target point is collected. This target point is located at the geometric center of the repair area.

[0037] Step 7, further, the distribution rule of the sampling target points in the intact area near the patch is as follows: they are distributed along the boundary outline between the edge of the patch and the intact area of ​​the patch. The number of points varies according to the size of the emergency repair, but the total number is odd and the distribution is equidistant.

[0038] Step 8: The data collected by each vibration sensor is the vibration wave at the collection point, which is then stored and recorded. The recorded vibration wave information is used for data comparison.

[0039] Step 9: The vibration wave from each sensor is transmitted to the storage processor via a signal shielded cable. There are no special requirements for the position of the sensor and the length of the cable, and the layout can be adjusted arbitrarily according to the actual situation.

[0040] Step 10: The storage processor extracts and processes the feature data of the vibration wave from each sensor. The main feature data includes the amplitude, phase, coherence, and envelope characteristics of the real-time vibration wave.

[0041] Step 11: The vibration wave characteristics obtained at different points need to be compared to determine whether the interface has begun to fail. The comparison is performed between vibration waves collected by multiple pairs of sensors. It is divided into a reference group and an interface group. The reference group is generated from the sensor vibration wave signals in the intact area near the patch, while the interface group is generated between the sensor vibration waves in the patch area and the sensor vibration waves in the intact area near the patch.

[0042] Step Twelve: Further, the data from the benchmark group uses the difference in vibration wave characteristics between two adjacent acquisition points, where the connecting line between them does not pass through the repair interface, as a feature data point. This is to obtain the vibration wave propagation characteristics of the undamaged area. (See attached...) Figure 3 As shown, A1 to A5 are five sets of baseline data, representing the differences in vibration wave characteristics between different adjacent acquisition points.

[0043] Step Thirteen: The data for the interface group comes from the characteristic differences between the vibration waves collected at the single vibration wave acquisition point within the patch area and at each acquisition point in the intact area near the patch. (See attached image.) Figure 3 As shown, R1 to R5 are five sets of interface data, representing the differences in vibration wave characteristics between different sampling points at the interface.

[0044] Step Fourteen: By comparing the vibration wave characteristic difference data of the interface group and the reference group, the main differences in the characteristic data to be compared are the differences in the amplitude, phase, coherence and envelope characteristics of the vibration waves. Due to the different structural polygons and emergency repair methods, the failure threshold needs to be determined according to the situation, but the principle of the threshold being the difference in vibration wave characteristics between the two groups remains unchanged.

[0045] Step 15: Through the above steps, low-cost and highly reliable real-time monitoring of the emergency repair area can be completed, thereby preventing the sudden occurrence of unexpected failures in the emergency repair area during helicopter flight.

[0046] This invention belongs to the field of composite material failure monitoring. It discloses a device and method for monitoring the condition of helicopter composite material skin after emergency repair. It proposes a method and device for monitoring the failure of emergency repair parts by monitoring the vibration waves generated by the helicopter itself during flight.

[0047] This application utilizes vibration wave monitoring sensors deployed at specific locations to compare the vibration wave characteristics obtained at different points, thereby determining whether interface failure has begun. This comparison is performed between vibration waves collected by multiple pairs of sensors. The comparison is divided into a reference group and an interface group. The reference group consists of vibration wave signals generated from sensors in the intact area near the patch, while the interface group consists of vibration waves generated from sensors within the patch area and from sensors in the intact area near the patch. By comparing the differences in vibration wave characteristics between the interface group and the reference group, a low-cost, high-reliability method and device for early warning of composite material repair structure failure is achieved.

Claims

1. A method of monitoring the post-repair condition of a composite skin of a helicopter, characterized in that, The method includes the following steps: Sensors are distributed in the patch area and the intact area near the patch; The sensor collects and stores vibration wave information in real time, and the recorded vibration wave information is used for data comparison. The vibration wave information is transmitted to the storage processor via a signal shielded cable; The vibration wave information obtained from different points is compared to determine whether the interface has begun to fail. The comparison is carried out between vibration wave information collected by multiple pairs of sensors, and is divided into a reference group and an interface group. By comparing the vibration wave characteristic difference data of the interface group and the reference group, the condition monitoring of the helicopter composite skin after repair can be realized. The sensors distributed in the patch area and the intact area near the patch include: In the patch area, vibration characteristic information of only one target point is collected, which is located at the geometric center of the repair area; The distribution rule for the target points in the intact area near the patch is as follows: they are distributed along the boundary outline between the edge of the patch and the intact area of ​​the patch. The number of points varies depending on the size of the emergency repair, but the total number is odd and the distribution is equidistant. The data for the reference group is generated from vibration wave information collected by sensors in the intact area near the patch; the data for the interface group is generated from vibration wave information collected by sensors in the patch area and vibration wave information collected by sensors in the intact area near the patch.

2. The method of claim 1, wherein, The vibration wave characteristic difference data includes the amplitude, phase, coherence, and envelope characteristics of the vibration wave.

3. The method of claim 1, wherein, The preferred sensor is a high-precision piezoelectric ceramic sensor, and its sensing core has a shear-type structure.

4. The method according to claim 1, characterized in that, The data from the reference group represent the vibration wave characteristics of undamaged areas; the data from the interface group represent the vibration wave characteristics between different sampling points across the interface.

5. A device for monitoring the condition of a helicopter composite material skin after repair, characterized in that, The apparatus is used to implement the method as described in any one of claims 1-4, the apparatus comprising: Sensors are distributed in the patch area and the intact area near the patch; The sensor is used to collect and store vibration wave information in real time, and the recorded vibration wave information is used for data comparison. The storage processor acquires and stores vibration wave information, and compares the vibration wave information obtained from different points to determine whether the interface has begun to fail. The comparison is performed between vibration wave information collected by multiple pairs of sensors, divided into a reference group and an interface group. By comparing the vibration wave characteristic difference data between the interface group and the reference group, the condition monitoring of the helicopter composite skin after repair is realized.

6. The apparatus according to claim 5, characterized in that, The device further includes: The signal shielded cable transmits vibration wave information to the storage processor.

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