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CN-113799968-B - Aviation composite structure and physical state monitoring method and system thereof

CN113799968BCN 113799968 BCN113799968 BCN 113799968BCN-113799968-B

Abstract

An aerospace composite structure includes a joint between structural members, a multi-core optical fiber having at least two fiber cores, the multi-core optical fiber being integrated in the joint along a longitudinal direction of the joint and including two fiber ends, each fiber end coinciding with an end of the joint, and a connector on at least each fiber end, the connector being configured to connect each fiber to an interrogation unit for measuring at least one parameter of the joint, each core of the multi-core optical fiber being configured to transmit a predefined pulse of light in accordance with at least one parameter to be measured for monitoring a physical state of the joint between structural members. A method and system for monitoring the physical state of a bonded portion in an aerospace composite structure are also disclosed. The invention also discloses an aircraft comprising the aviation composite structure.

Inventors

  • Jacinto Enrique Rodriguez Serrano
  • Carlos Miguel hilardo

Assignees

  • 空中客车西班牙运营有限责任公司

Dates

Publication Date
20260505
Application Date
20210610
Priority Date
20200615

Claims (13)

  1. 1. An aeronautical composite structure (1) comprising a joining portion (2) between structural members (3, 4, 5), the aeronautical composite structure (1) further comprising: -a plurality of multicore optical fibers (6), a plurality of said multicore optical fibers (6) being integrated in the joint (2) along a longitudinal direction (X-X') of the joint (2), each multicore optical fiber (6) comprising at least two cores (9) and two fiber ends (6.1, 6.2) embedded within the optical fiber (6), namely a first fiber end (6.1) and a second fiber end (6.2), each fiber end (6.1, 6.2) coinciding with an end (2.1) of the joint (2), and A connector (17) at least on each fiber end (6.1, 6.2) configured to connect each optical fiber (6) with its respective fiber end (6.1, 6.2) to an interrogation unit (18) for measuring at least one parameter of the bonded portion (2) in the aeronautical composite structure (1), Wherein, the Each connector (17) being adapted to connect each core (9) to the interrogation unit (18), and Each core (9) of each multicore optical fiber (6) is configured to transmit a predefined light pulse from the interrogation unit (18) from the first fiber end (6.1) to the second fiber end (6.2) along an optical fiber extension through the bonding portion (2) according to the at least one parameter to be measured for monitoring the physical state of the bonding portion (2) between structural parts (3, 4, 5).
  2. 2. An aerospace composite structure (1) according to claim 1, wherein the bonding portion (2) comprises an adhesive line into which at least the multicore optical fibre (6) is embedded.
  3. 3. The aeronautical composite structure (1) according to any preceding claim, wherein at least one core (9) of the multicore optical fiber (6) is a multimode core configured to provide raman scattering when the connector (17) is connected to the interrogation unit (18).
  4. 4. The airborne composite structure (1) according to any of the preceding claims, wherein at least one core (9) of the multi-core optical fiber (6) is a single core configured to provide rayleigh scattering when the connector (17) is connected to the interrogation unit (18).
  5. 5. The aeronautical composite structure (1) according to any preceding claim, wherein at least one core (9) of the multicore optical fiber (6) is a single-core comprising a bragg grating.
  6. 6. The aeronautical composite structure (1) according to any preceding claim, wherein a multicore optical fiber (6) comprises a distribution of single cores, wherein multiplexed bragg grating sensors are written in different cores (9) of the multicore optical fiber (6).
  7. 7. The aircraft composite structure (1) according to any one of the preceding claims, wherein the aircraft composite structure is a leading edge (16) of a vertical tail (13), the leading edge comprising the following structural components: -an inner panel base laminate (4) with a plurality of omega stringers (3), and An outer panel (5), At least the outer panel (5) is joined to the inner panel base laminate (4) by means of an adhesive line between one side of the outer panel (5) and the head of each omega-shaped stringer (3) such that at least a multicore optical fiber (6) is embedded in the adhesive line.
  8. 8. The aerospace composite structure (1) according to claim 7, wherein each Ω -shaped stringer (3) is joined to the inner panel base laminate (4) by means of an adhesive line arranged between one side of the inner panel base laminate (4) and each foot of the Ω -shaped stringer (3) such that at least multicore optical fibers (6) are embedded in each of the adhesive lines.
  9. 9. A system for monitoring a physical state of a bonded portion (2) in an aerospace composite structure (1), the system comprising: -an aeronautical composite structure (1) according to any of the previous claims 1 to 8, and -An interrogation unit (18) connected to the connector (17) of the aeronautical composite structure (1) and configured to measure at least one parameter in the bonding portion (2) of the aeronautical composite structure (1) in order to monitor the physical state of the bonding portion (2); Wherein the interrogation unit (18) comprises: a light source configured to emit light pulses through a first fiber end (6.1) of the multi-core optical fiber (6); a receiver configured to detect or sense the emitted light pulses passing through the second fiber end (6.2) of the multi-core optical fiber (6), and A processor configured to process the sensed light pulses.
  10. 10. A method for monitoring the physical state of a joining portion (2) in an aeronautical composite structure (1) according to any of the previous claims 1 to 8, comprising the steps of: a) An interrogation unit (18) is provided, the interrogation unit (18) comprising a light source, a receiver and a processor, B) -connecting the interrogation unit (18) to a connector (17) located on each multicore optical fiber end (6.1, 6.2) of the aeronautical composite structure (1), and C) -interrogating at least one multi-core optical fiber (6) between connectors (17) by transmitting predefined pulses of light through at least two cores (9) of said multi-core optical fiber (6) according to parameters to be measured for monitoring the physical state of said bonding portion (2) between structural components (3, 4, 5); Wherein said step c) comprises: i. -emitting predefined light pulses by said light source from a first fiber end (6.1) of said multi-core optical fiber (6) through at least one core (9) of said multi-core optical fiber (6), Measuring the received light pulses by the receiver at a second fiber end (6.2) of the multi-core optical fiber (6), and Processing the measured light pulses by the processor in order to monitor the physical state of the binding portion (2) in the aeronautical composite structure (1).
  11. 11. The method according to claim 10, further comprising monitoring the temperature in the joining portion (2) of the aeronautical composite structure (1) by interrogating the at least one multicore optical fiber (6) in step c) to measure the temperature in this joining portion (2) while this aeronautical composite structure (1) is in the course of a curing cycle.
  12. 12. The method according to any one of claims 10 to 11, further comprising monitoring damage in the bonded portion (2) of the aeronautical composite structure (1) by measuring strain or deformation in this bonded portion (2) by interrogating the at least one multicore optical fiber (6) in step c).
  13. 13. An aircraft (12) comprising an aeronautical composite structure (1) according to any one of claims 1 to 8.

Description

Aviation composite structure and physical state monitoring method and system thereof Technical Field The present invention relates to an aeronautical composite structure intended for monitoring the physical state of a joint between structural components. The invention also relates to such a method and system for monitoring the physical state of a bonded part in an aeronautical composite structure. More particularly, the present invention relates to a structure and method for monitoring the physical state of a bonded portion of an aerospace composite structure from the manufacture of the aerospace composite structure to its use in flight. Background Aerospace composite structures are often integrated with stiffeners (such as stringers) to improve the stiffness or buckling resistance of these composite structures. The stringers may be joined by adhesive lines between the structural components of the composite structure referred to. That is, this composite structural technology field typically utilizes adhesive joints suitable for composite structures during the manufacturing and assembly stages. The adhesive joint is typically subjected to a curing process after the composite structure has been entered into the autoclave, where temperature is one of the key parameters to be considered in the manufacture of such structures. The temperature in the autoclave is nowadays mainly controlled by conventional thermocouples, which need to be in direct contact with the surface of the composite material being cured and only provide on-time measurements throughout the process. This solution for monitoring the temperature of some composite structures is cumbersome for the complex parts of the structure and leaves room for improvement. The quality control is continuously upgraded in the aviation field, mainly to improve production and safety, while reducing final wastage and reprocessing costs. Currently, once an aircraft is put into service and requires periodic inspection or experience of an accident, an operator must disassemble the aircraft to find possible damage to the adhesive lines of joints in the composite structure. Existing inspections are considered to be complex manual work procedures and certainly time consuming. Today, the adhesive lines in composite structures undergo quality control by non-destructive inspection (NDI), such as ultrasonic pulse echo. Once the aircraft returns to the ground, NDI control requires specific tools, certified inspectors, access to the inspection area, which also means that economies and scheduling are affected. Furthermore, it is known that the actual conventional procedure for monitoring the temperature at the time of manufacture depends on the skill of the operator and requires a large time investment. There are known systems for monitoring the integrity of an adhesive within a cured bond line of a bonded structural component, similar to the system described in patent application US 8812251 B2. Heretofore, the adhesive lines have been monitored using a network of electrical sensors disposed inside the adhesive lines. In addition, these systems for monitoring the status of the adhesive include a power source for providing power to the electrical sensor network in order to check the integrity of the adhesive lines as needed. The system is made by interpreting the changes measured directly in the cured bond line. However, such known apparatus and methods provide monitoring only after the adhesive lines are manufactured, but not during the manufacturing process. In another technical field, it is further known that optical fibers made of a plurality of cores positioned along the diameter of the cladding layer, which can be made to respond to temperature changes or strain changes. These variations are typically measured by selecting the fiber optic material (especially core, cladding and coating) and the spacing and shape of the cores within the fiber. Today, the temperature variation measured by optical fiber optimization can be provided over a wide range. Furthermore, there are known composite structures integrated with optical fibers capable of locating lesions along the composite structure, and methods for manufacturing the composite structures integrated with lesion locating capability. It is known to provide optically aligned fiber optic connection devices with minimal insertion loss that can be placed anywhere on the composite surface to collect in-service parameters from the connected optical fibers and to continue to collect information. Accordingly, the present invention addresses the need in the art for an improved method of monitoring the physical state of a bonded portion in an aerospace composite structure, wherein such improved method provides manufacturing advantages over systems and methods known in the art. In addition, the present invention further provides a system for monitoring temperature, strain or deformation during in-service inspection and in-service operat