Aerospace composite structures and their physical condition monitoring methods and systems

Abstract

This invention discloses an aerospace composite structure, comprising a joint portion located between structural components; a multi-core optical fiber having at least two fiber cores, the multi-core optical fiber being integrated within the joint portion along its longitudinal direction and including two fiber ends, each fiber end coinciding with an end of the joint portion; and a connector located at least at each fiber end, the connector being configured to connect each fiber to an interrogation unit for measuring at least one parameter of the joint portion, each core of the multi-core optical fiber being configured to transmit a predefined optical pulse according to at least one parameter to be measured for monitoring the physical state of the joint portion between the structural components. This invention also discloses a method and system for monitoring the physical state of the joint portion in an aerospace composite structure. Furthermore, this invention discloses an aircraft including this aerospace composite structure.

Description

Technical Field

[0001] This invention relates to an aerospace composite structure intended for monitoring the physical state of joints between structural components. The invention also relates to a method and system for monitoring the physical state of joints in an aerospace composite structure. More specifically, the invention relates to a structure and method for monitoring the physical state of joints in an aerospace composite structure from its manufacture to its use in flight. Background Technology

[0002] Aerospace composite structures typically integrate stiffeners (such as stringers) to improve their stiffness or buckling resistance. These stringers, and the structural components of the composite structure in question, can be connected via adhesive lines. That is, this field of composite structure technology typically utilizes adhesive joints suitable for composite structures during the manufacturing and assembly phases.

[0003] The adhesive joints are typically cured after the composite structure enters the autoclave, where temperature is one of the key parameters to consider in manufacturing such structures. Currently, the temperature in the autoclave is primarily controlled by conventional thermocouples, which need to be in direct contact with the surface of the curing composite material and provide only timely measurements throughout the process. This solution for monitoring the temperature of some composite structures is cumbersome for complex parts of the structure and leaves room for improvement.

[0004] The aviation industry is constantly upgrading its quality control, primarily to improve production and safety while reducing waste and reprocessing costs. Currently, once an aircraft enters service and requires periodic inspections or experiences an unexpected event, operators must disassemble the aircraft to detect potential damage to the adhesive lines at joints in composite structures. Existing inspections are considered complex manual processes and are undoubtedly time-consuming.

[0005] Currently, the adhesive lines in composite structures undergo quality control through non-destructive testing (NDI) (e.g., ultrasonic pulse-echo testing). Once the aircraft returns to Earth, NDI control requires specific tools, certified inspectors, and proximity to the inspection area, which also implies economic and scheduling impacts. Furthermore, the actual routine procedures for monitoring temperature during manufacturing are known to depend on operator skill and require a significant time investment.

[0006] Known systems exist for monitoring the integrity of the adhesive within the cured bond line of a bonded structural assembly, similar to those described in patent application US 8812251 B2. To date, the adhesive line has been monitored using a network of electrical sensors arranged within it. Furthermore, these systems for monitoring the adhesive state include a power source to supply power to the electrical sensor network so that the integrity of the adhesive line can be checked as needed. These systems are made by interpreting changes measured directly within the cured bond line. However, such known devices and methods only provide monitoring after the adhesive line has been manufactured, not during the manufacturing process.

[0007] In another area of ​​expertise, it is further known to be optical fibers made of multiple cores positioned along the diameter of a cladding layer, which can be manufactured to respond to changes in temperature or strain. These changes are typically measured by selecting the optical fiber materials (particularly the cores, cladding, and coating) and the spacing and shape of the cores within the fiber. Currently, temperature variations that can be measured through optical fiber optimization can be provided over a wide range.

[0008] Furthermore, there are known composite structures integrating optical fibers capable of locating damage along the composite structure, and methods for manufacturing such composite structures with integrated damage-locating capabilities. It is well known that optical fiber connectors provide optical alignment with minimal insertion loss; these connectors can be placed anywhere on the composite surface to collect in-service parameters from the connected optical fibers and continue collecting information.

[0009] Accordingly, the present invention addresses the need in the prior art for improved methods for monitoring the physical state of joints in aerospace composite structures, wherein such improved methods offer manufacturing advantages over known systems and methods in the prior art. Furthermore, the present invention provides a system for monitoring temperature, strain, or deformation during in-service inspection and operation of composite structures. Summary of the Invention

[0010] The present invention provides solutions to the aforementioned problems through an aerospace composite structure according to a first aspect, a system according to a second aspect, a method for monitoring the physical state of joints in an aerospace composite structure according to a third aspect, and an aircraft according to a fourth aspect. Preferred embodiments of the invention are defined in the dependent claims.

[0011] In a first aspect, the present invention provides an aerospace composite structure including a joint portion located between structural components, the aerospace composite structure further comprising:

[0012] - A multi-core optical fiber having at least two fiber cores, said multi-core optical fiber being integrated in the joint along the longitudinal direction of this joint portion and including two fiber ends, each fiber end coinciding with the end of said joint portion, and

[0013] - Connectors located at least at the ends of each fiber, the connectors being configured to connect each fiber to an interrogation unit for measuring at least one parameter of the bonded portion in the aerospace composite structure.

[0014] Each core of the multi-core optical fiber is configured to transmit a predefined optical pulse according to at least one parameter to be measured for monitoring the physical state of the joint between structural components.

[0015] An aircraft is composed of multiple aerospace composite structures formed by the connection of various structural components. For example, the anti-torsion box for the horizontal tail of an aircraft is formed in particular by structural components such as panels, stiffeners (i.e., stringers, ribs, etc.), frames, and skins that are connected together, such that for each joint, the configured composite structure includes a connecting part.

[0016] These joints between structural components are crucial connections in composite structures, and their physical state deserves particular attention and monitoring. This is especially true from the time the aerospace composite structure is manufactured (i.e., the assembly of structural components) until it is integrated into the aircraft, and also throughout the operational life of the aerospace composite structure.

[0017] This invention provides an improved configuration for the joints between these structural components of an integrated aerospace composite structure, thereby improving the monitoring of the physical condition of these joints. In particular, the invention proposes providing at least an optical fiber integrated into the joint portion of the composite structure, which corresponds to the connection portion between the structural components.

[0018] An optical fiber is a multi-core optical fiber comprising at least two fiber cores embedded within a fiber. The optical fiber is covered by a conventional cladding layer and includes two fiber ends. The multi-core optical fibers are integrated into the joint portion of the composite structure along the longitudinal direction of the same joint portion. That is, this longitudinal direction corresponds to the direction followed by the joint portion in the connection between structural components of the composite structure. The multi-core optical fibers are arranged along the entire length of the joint portion such that the fiber ends of the optical fibers coincide with the ends of the joint portion.

[0019] The composite structure further includes at least connectors that connect each optical fiber at its respective fiber end to an interrogation unit designed to monitor the physical state of the composite structure, particularly the joint portion. More specifically, the connectors are adapted to connect each optical fiber core to the interrogation unit. This connector is a fan-shaped spread to reach each core of the multi-core optical fiber. The connectors distribute light from the optical fibers to each core of the multi-core optical fiber. The fact that the connectors connect to both ends of the optical fiber allows for greater flexibility in application and enables interrogation at both ends of the optical fiber. In addition to providing connectors, the two fiber ends further allow interrogation of the fiber cores to utilize techniques requiring connection at both fiber ends for measuring physical parameters (e.g., Brillouin scattering).

[0020] In the context of this invention, the physical state must be understood as every physically distinguishable condition or form that can be obtained by measuring certain characteristics that the composite structure may adopt during its temporal evolution. That is, in the combined portions of the composite structure undergoing changes, the physical state is any possible condition due to these changes. Examples of such physical states of the composite structure are temperature, strain, deformation, damage, load, vibration, and fire detection.

[0021] Specifically, the connector providing the fiber end to the interrogation unit allows for the measurement of at least one parameter of the joint portion to monitor the physical state of the composite structure. In this way, the fiber core integrated in the multi-core optical fiber transmits a predefined optical pulse from the interrogation unit along the optical fiber extension passing through the joint portion. That is, the optical fiber transmits the predefined optical pulse through the fiber core according to the parameter to be measured. Therefore, by monitoring the measured parameter, the physical state of the joint portion of the composite structure can be determined.

[0022] In another specific embodiment, the connector is a direct connector configured to connect to a universal interrogator. This direct connector includes a precision processing system for the fiber cores that allows for simultaneous alignment of all fiber cores, wherein the insertion loss and return loss values ​​are compatible with the dynamic range of each technology applied to the fiber cores. The universal interrogator is configured to interrogate all fiber cores using different technologies (e.g., FBG, Raman, Rayleigh, or Brillouin).

[0023] The present invention advantageously allows for monitoring of the physical state of the joints in the composite structure during the manufacturing process, during in-service inspection, and throughout the operational life of the composite structure.

[0024] Furthermore, the presence of multiple cores within the optical fiber allows for the selection and application of optical monitoring techniques of interest at each stage of the composite structure's lifespan. That is, the fiber cores can be designed to measure different or the same parameters of interest regarding the composite portion. These parameters could be temperature, strain, deformation, damage, load, vibration, or fire detection.

[0025] Advantageously, the proposed solution aims to support and improve the control quality of the bonded portion during the curing process (manufacturing step). Because the optical fiber is embedded within the bonded portion of the composite structure, the measured temperature is more reliable than that of external sensors in prior art solutions. Furthermore, the provision of the optical fiber allows for measurement along the length of the fiber, rather than at a precise location as in prior art solutions, thus enabling broader control quality and temperature mapping.

[0026] For structural testing or inspection, as well as in-service operation, the availability of optical fibers further enables the monitoring of structural performance via strain measurements, and even the detection of damage, such as disintegration, at joints. In other words, this invention allows for the detection of damage at joints without adjusting the aircraft or touching the affected areas of the composite structure.

[0027] Therefore, multi-core optical fibers act as permanent sensors installed inside the joints between structural components, and these optical fibers are inquired about according to the requirements of maintenance and operation.

[0028] Accordingly, compared to conventional solutions, the present invention mainly offers the following advantages:

[0029] - Compared to existing technology solutions, controlling the longer surface of the composite structure through multi-core optical fibers improves the quality control of the joints between structural components.

[0030] - Monitor the temperature and strain of the joints during manufacturing. This also facilitates online inspections during manufacturing and aims to detect potential defects, even allowing them to be corrected before the manufacturing process is complete.

[0031] - By providing optical fibers integrated into the joint to monitor the behavior of in-service machinery, this allows for control of the joint and detection of potential damage without touching the aircraft.

[0032] In a particular embodiment, the bonding portion includes an adhesive line in which at least the multi-core optical fiber is embedded. This adhesive line corresponds to a connecting device arranged between two structural components for their joint. In this embodiment, the optical fiber is embedded within the adhesive line. Using integrated optical fiber in the adhesive line advantageously eliminates the need for drilling holes in the reinforcement for the joint, thus simplifying the manufacturing process and being preferable from a structural perspective.

[0033] The type of adhesive used in adhesive lines depends on the working conditions, application requirements, and the materials to be bonded. Working conditions and application requirements can include temperature range, dynamic or static load conditions, necessary chemical resistance, durability, application time, and curing time. The materials to be bonded to accommodate the composite structure are typically metals, polymers, or ceramics.

[0034] In a more specific embodiment, the adhesive line is an epoxy resin adhesive. Considering the materials, required resistance, temperature range, and operating time of known aerospace composite structures, epoxy resin is the preferred adhesive.

[0035] In another specific embodiment, the adhesive line is an adhesive made of silicone, cyanoacrylate, polyurethane, or phenolic resin.

[0036] In a particular embodiment, the bonding portion includes multiple adhesive lines.

[0037] In a particular embodiment, the aerospace composite structure includes a plurality of multi-core optical fibers integrated in the bonding portion, wherein one connector of the connectors is located at the end of a first multi-core optical fiber and the other connector is located at the end of a second multi-core optical fiber.

[0038] Advantageously, the multiple core fibers provide redundancy, offering the possibility of having backup fibers in case of fiber failure. Furthermore, the multiple core fibers allow for the simultaneous monitoring of a greater number of parameters. Consequently, the accuracy of the measurements is also improved because multiple cores / fibers can monitor the same parameters.

[0039] In a particular embodiment, the multicore optical fiber includes a coating. Known optical fibers are covered by a cladding layer. Therefore, the aforementioned coating is an additional coating to the cladding layer and provides mechanical protection for the multicore optical fiber.

[0040] The material of the coating depends on the operating and temperature conditions of the optical fiber. For damage and temperature measurements, the coating material is polyamide. For fire detection, the coating material is metal. And for specific operating conditions, the coating material may be a reinforcing polymer. In certain embodiments, the coating has a thickness starting from 200 μm.

[0041] In a particular embodiment, at least one core of the multi-core optical fiber is a multimode core, which is configured to provide Raman scattering when the connector is connected to the interrogation unit. A multimode core means that the fiber is configured to propagate several light modes simultaneously.

[0042] In a particular embodiment, at least one core of the multi-core optical fiber is a single-mode core, which is configured to provide Rayleigh scattering when the connector is connected to the interrogation unit. A single-mode core means that the fiber can only propagate one light mode at a time.

[0043] In a particular embodiment, at least one core of the multi-core optical fiber is a single-mode core comprising a Bragg grating. More specifically, the multi-core optical fiber may be a single-mode core comprising a multiplexed Bragg grating sensor.

[0044] In a particular embodiment, the multi-core optical fiber comprises a distribution of single-mode cores, wherein multiplexed Bragg grating sensors are incorporated into the different cores of the multi-core optical fiber. Advantageously, providing multiplexed Bragg grating sensors within the fiber cores allows for matching multiple spatial resolution requirements over hundreds of meters of optical fiber while necessitating the installation of a unique fiber.

[0045] Single-mode and multimode fibers differ in their core diameter, wavelength, source, and bandwidth. The core diameter of a single-mode fiber is smaller than that of a multimode fiber. The wavelength of a multimode fiber is shorter than that of a single-mode fiber. Furthermore, the bandwidth of a multimode fiber is limited by its source mode, while the bandwidth of a single-mode fiber is theoretically unlimited because it allows one optical mode to pass through at a time. Additionally, single-mode fibers are suitable for long-distance applications, while multimode fibers are designed for short-distance applications.

[0046] Given that multi-core optical fibers integrate multiple cores (single-mode and multi-mode) for a single optical fiber, the following is allowed:

[0047] - Reduce the number of fibers to be installed and simplify the process, as a single optical fiber can include multiple cores, and each core includes multiple sensors or distributed sensors. This is important from the perspective of aircraft applications, such as temperature measurement applications requiring hundreds of meters across the entire plane with critical spatial resolution (mm or cm level) and damage detection applications requiring damage detection resolution of approximately a few millimeters; and - improve spatial resolution in applications where there are limitations on the distance between fiber ends.

[0048] Advantageously, the distribution of single-mode cores within the multi-core optical fiber provides consistency and uniformity of the measured parameters along the entire monitored aerospace composite structure.

[0049] In a particular embodiment, the aerospace composite structure is the leading edge of a vertical tail fin, the leading edge comprising the following structural components:

[0050] - An inner panel base laminate with multiple Ω-shaped struts, and

[0051] -Outer panel,

[0052] At least the outer panel is connected to the inner panel base laminate by an adhesive line located between one side of the outer panel and the head of each Ω-shaped stringer, such that at least a multi-core optical fiber is embedded in the adhesive line.

[0053] In a more particular embodiment, each Ω-shaped strut is connected to the inner panel base laminate by an adhesive line arranged between one side of the inner panel base laminate and each foot of the Ω-shaped strut, such that at least a multi-core optical fiber is embedded in each of the adhesive lines.

[0054] In a second aspect, the present invention provides a system for monitoring the physical state of joint portions in an aerospace composite structure, the system comprising:

[0055] -Aerospace composite structure according to the first aspect of the invention, and

[0056] - An interrogation unit, which is connected to the connector of the aerospace composite structure and configured to measure parameters in the joint portion of the aerospace composite structure in order to monitor the physical state of the joint portion.

[0057] In other words, the primary function of this interrogation unit is to interrogate the fibers integrated into the joints between structural components to measure parameters of interest in order to determine the physical state of the joints. The interrogation unit, connected to the connector of the composite structure, allows for monitoring of the physical state of the joints within the composite structure during manufacturing and throughout its operational life.

[0058] In a more specific embodiment, the interrogation unit includes:

[0059] - A light source configured to emit light pulses through the first fiber end of a multi-core optical fiber.

[0060] - A receiver configured to detect or sense emitted light pulses passing through the end of a second fiber, and

[0061] - A processor configured to process the sensed light pulses.

[0062] The interrogation unit is responsible for interrogating the optical fiber by transmitting light pulses emitted by a light source through the end of the optical fiber and detecting these pulses through the end of the optical fiber by a receiver. Further, through this interrogation, the processor of the interrogation unit performs analysis, comparing the output light pulses (the light pulses detected by the receiver) with the input light pulses (the light pulses emitted by the light source).

[0063] Depending on the parameters to be measured (temperature, strain, deformation, damage, load, vibration, and fire detection), the light pulses are configured with certain characteristics that allow the physical state of the joint to be determined based on the parameters, according to the comparative analysis described above.

[0064] In a third aspect of the invention, the present invention provides a method for monitoring the physical state of a bonded portion in an aerospace composite structure according to a first aspect of the invention, the method comprising the following steps:

[0065] a) Provide an inquiry unit.

[0066] b) Connect the interrogation unit to the connector located at the end of each multi-core optical fiber in the aerospace composite structure, and

[0067] c) By transmitting a predefined optical pulse through at least two cores of the multi-core optical fiber according to parameters to be measured for monitoring the physical state of the joint between structural components, the multi-core optical fiber is interrogated between the connectors.

[0068] The method of the present invention allows for monitoring the physical condition of the joints between components of the aerospace composite structure during both the manufacturing and operational lifespan of an aircraft equipped with the composite structure.

[0069] In a particular embodiment, step c) includes:

[0070] i. A predefined light pulse emitted by a light source passes through at least one core of the multi-core optical fiber.

[0071] ii. Measure the received light pulse, and

[0072] iii. Process the measured light pulses to monitor the physical state of the bonding portion in the aerospace composite structure.

[0073] In a particular embodiment, the method further includes monitoring the temperature in the bonded portion of the aerospace composite structure by interrogating the multi-core optical fiber in step c) while the aerospace composite structure is in the curing cycle.

[0074] In a particular embodiment, the method further includes monitoring damage in the joint portion of the aerospace composite structure by interrogating the multi-core optical fiber in step c) to measure strain or deformation in this joint portion.

[0075] In a fourth aspect of the invention, the present invention provides an aircraft comprising an aeronautical composite structure according to the first aspect of the invention.

[0076] In a more specific embodiment, the aircraft includes multiple aeronautical composite structures, such as a horizontal tail, a vertical tail, and wings.

[0077] Providing aircraft manufactured with aerospace composite structures (configured with multi-core optical fibers integrated on the joints of these structures) advantageously allows for monitoring the condition of these joints by inspecting the aircraft. Attached Figure Description

[0078] These and other features and advantages of the invention will become clearer from the following detailed description of preferred embodiments provided with reference to the accompanying drawings, which are provided as illustrative rather than limiting examples only.

[0079] Figure 1 This figure shows a perspective view of an aerospace composite structure according to an embodiment of the present invention.

[0080] Figure 2 This diagram shows Figure 1 An exploded view of the aerospace composite structure shown.

[0081] Figure 3 This figure shows a perspective view of the joint portion of an aerospace composite structure according to an embodiment of the present invention.

[0082] Figures 4a to 4c These figures show cross-sectional views of multi-core optical fibers according to embodiments of the present invention.

[0083] Figure 5 This figure shows a schematic diagram of a monitoring system according to an embodiment of the present invention.

[0084] Figure 6 This figure shows a side view of an aircraft including an aeronautical composite structure according to an embodiment of the present invention. Detailed Implementation

[0085] As those skilled in the art will recognize, aspects of the present invention can be embodied in aerospace composite structures, systems or methods for monitoring the physical state of the joint portions of such aerospace composite structures.

[0086] The present invention provides an aerospace composite structure (1) having at least a multi-core optical fiber (6) integrated in a joint portion (2) between structural components (3, 4, 5) of the aerospace composite structure (1). This configuration of the multi-core optical fiber (6) allows for monitoring of the physical condition of the joint portion (2) during the manufacture of the aerospace composite structure and during inspections while the aerospace composite structure is in service.

[0087] Figure 1A perspective view of an aerospace composite structure (1) is shown, corresponding to the leading edge of a lifting surface (e.g., a horizontal tail). This leading edge (1) is formed by structural components serving as an inner panel base laminate (4) and multiple stringers (3). Specifically, the inner panel base (4) is mounted on multiple ribs (7.1, 7.2, 7.3, 7.4) to shape the aerospace composite structure or leading edge (1) into a semi-elliptical shape. The base of the leading edge (1) is covered by a membrane (10) made of composite material.

[0088] More specifically, the leading edge (1) includes two end ribs (7.1; 7.3) located at both ends of the aerospace composite structure (1), and an interface rib (7.2) located between the two end ribs (7.1, 7.3). This interface rib (7.2) provides a rigid connection between the two parts of the standard-sized inner panel base laminate (4) (e.g., Figure 1 (As shown). In addition, multiple reinforcing ribs (7.4) placed inside the inner panel base laminate (4) provide rigidity and shape retention of the laminate (4) at the leading edge (1).

[0089] Both parts of the inner panel base laminate (4) include Ω-shaped stringers (3), which are positioned along the leading edge (1) and placed parallel to each other along the surface of the laminate (4). In particular, these Ω-shaped stringers (3) are connected to the inner panel base laminate (4) by adhesive lines in the joint portion (2) between the inner panel laminate (4) and the outer panel (5). Figure 2 (As shown in the image). This adhesive line (in...) Figure 1 and Figure 2 (Not shown) is located between one side of the inner panel base laminate (4) and each foot of the Ω-shaped stringer (3) so as to embed at least one multi-core optical fiber (6) in each joint (2).

[0090] Figure 2 It shows Figure 1 An exploded view of the leading edge (1). More specifically, this Figure 2 An outer panel (5) is shown to be mounted on the inner panel base laminate (4) to cover the entire inner panel base laminate (4). Specifically, as Figure 3 As shown, the adhesive line to be placed on each of the plurality of Ω-shaped struts (3) connects the aforementioned outer panel (5) to the inner panel base laminate (4), thereby providing resistance and fixation for the structural components of the leading edge (1).

[0091] In a preferred embodiment, depending on the type of materials to be assembled, an adhesive line is used, such as an adhesive made of epoxy resin, silicone resin, cyanoacrylate, polyurethane, phenolic resin, etc.

[0092] Figure 3An aerospace composite structure (1) is shown (e.g.) Figure 2 A perspective view of a portion of the structure shown, comprising an inner panel base laminate (4), an Ω-shaped stringer (3), and an outer panel (5). At one foot of the Ω-shaped stringer (3), the Ω-shaped stringer (3) is attached to the inner panel base laminate (4) at a portion designated as a joint (2). Additionally, the outer panel (5) is attached to the head of the Ω-shaped stringer (3) at a portion also designated as a joint (2), particularly to its outer surface. These joints (2) are arranged along a longitudinal direction (X-X'). Contact between structural components (3, 4, 5) in these joints (2) is ensured by an adhesive line of predetermined joint width. Multiple multi-core optical fibers (6) are integrated on each long joint (2). Each multi-core optical fiber (6) includes two fiber ends (6.1; 6.2), which coincide with the end of the joint (2) at each end of the multi-core optical fiber (6).

[0093] Each fiber end (6.1; 6.2) has a connector (not shown) for connecting the fiber (6) to an interrogation unit (not shown) to measure parameters such as temperature, deformation, or strain of the joint (2) in the aerospace composite structure (1). Each multi-core optical fiber (6) includes at least two fiber cores (9) that transmit predefined optical pulses according to previously set parameters to be measured in order to monitor the physical state of the joint (2).

[0094] In a specific example, the interrogation unit (18) emits a light pulse from the first fiber end (6.1) through a light source (not shown) through the multi-core optical fiber (6). The emitted light is then detected at the second fiber end (6.2) by a receiver in the interrogation unit (18). Once the emitted light pulse is sensed, it is processed by a processor included in the interrogation unit (18).

[0095] In a particular example, the multi-core optical fiber (6) includes at least two cores (9) integrated inside the cladding layer (11), preferably spaced apart at a pitch of 35 to 70 micrometers.

[0096] Figures 4a to 4c A cross-sectional view of a multi-core optical fiber (6) is shown, wherein the cores (9) are arranged in a star shape inside the multi-core optical fiber (6). Figure 4b ) or hexagon ( Figure 4a and Figure 4c More precisely, a multi-core optical fiber (6) comprises seven cores (9). Figure 4a ), thirteen cores (9)( Figure 4b ) and nineteen cores (9)( Figure 4c).

[0097] Figures 4a to 4c A multi-core optical fiber (6) is further shown covered in addition to the overlay layer (11) by a coating (8) that provides mechanical protection for the multi-core optical fiber (6). In a particular example, the coating (8) is made of polyamide for temperature and deformation measurements, metal for fire detection, and reinforcing polymer for other determined measurements.

[0098] A core (9) can be single-mode to perform Bragg scattering, Brillouin scattering, or Rayleigh scattering, or multimode to perform Raman scattering. In a preferred example, each multi-core optical fiber (6) includes at least one of each type of core. On the one hand, this preferred structure reduces the number of material cores (9) and simplifies their fabrication when manufacturing the multi-core optical fiber (6). On the other hand, the preferred structure also improves spatial resolution to reduce the distance between sensors of the multi-core optical fiber (6) on the aerospace composite structure (1).

[0099] In another preferred embodiment, at least one multimode core (9) of the multi-core optical fiber (6) is integrated to provide Raman scattering when connected to the interrogation unit (18). Furthermore, at least one single-mode core (9) of the multi-core optical fiber (6) is provided to provide Rayleigh scattering when connected to the interrogation unit (18). Finally, at least one single-mode core (9) of the multi-core optical fiber (6) is integrated to perform Bragg grating sensing measurements.

[0100] Figure 5 A system for monitoring the physical state of a joint portion (2) in an aerospace composite structure (1) is shown. Multi-core optical fibers (6) are embedded in the joint portion (2) between structural components (4), which correspond to the panel laminate base of the composite structure (1). This system further includes interrogation units (18) connected to the multi-core optical fibers (6) via connectors (17). Specifically, each connector (17) is attached to each fiber end (6.1, 6.2), thereby allowing connection between the interrogation unit (18) and two fiber ends (6.1, 6.2) of the multi-core optical fiber (6). The interrogation unit (18) transmits predefined optical pulses through at least two cores (9) of the multi-core optical fiber (6) according to parameters to be measured for monitoring the physical state of the joint portion (2) between the two panel laminate bases (4).

[0101] Figure 6A side view of an aircraft (12) is shown, comprising a vertical tail (13), a horizontal tail (14), and a wing (15), all of which are aerospace composite structures (1) according to an embodiment of the invention. Each tail (13, 14) and wing (15) has a leading edge (16) comprising a multi-core optical fiber (6) integrated on a junction (2) for providing the ability to measure the physical state of the structure. The leading edge (16) is the first component of each aerospace composite structure (1) to come into contact with oncoming airflow during the service of the aircraft (12).

[0102] A method for monitoring the physical state of the joint (2) in an aerospace composite structure (1).

[0103] The present invention further provides a method for monitoring the joint portion (2) in an aerospace composite structure (1) (e.g. Figure 1 and Figure 2 The method for determining the physical state of the vertical tail fin (as shown).

[0104] This monitoring method includes the following steps:

[0105] a) Provide an inquiry unit (18),

[0106] b) Connect the interrogation unit (18) to the connector located at the end (6.1, 6.2) of each multi-core optical fiber in the aerospace composite structure (1), and

[0107] c) By transmitting a predefined optical pulse through at least two cores (9) of the multi-core optical fiber (6) according to parameters to be measured for monitoring the physical state of the joint portion (2) between structural components (3, 4, 5), the multi-core optical fiber (6) between the connectors (17) is interrogated.

[0108] During the manufacture of the composite structure or once it has been manufactured and / or installed on the aircraft (12), starting with the composite structure (1) object to be monitored, the interrogation unit (18) provided in step a) is then connected to the connector (17) on each fiber end (6.1, 6.2) in step b). These fiber ends (6.1, 6.2) correspond to the ends of the multi-core optical fibers (6) embedded in each joint (2) of the composite structure (1). Monitoring is performed independently for each joint (2) of the composite structure (1).

[0109] Once the connector (17) of the multi-core optical fiber (6) is connected to the interrogation unit (18), step c) begins the interrogation of the multi-core optical fiber (6). For this step c), a predefined light pulse is emitted through the optical fiber according to the parameter to be measured at the connection part (2). This parameter can be temperature, strain, deformation, damage, load, vibration, and fire detection. Therefore, the light pulse is configured with characteristics based on the parameter to be measured.

[0110] Once the light pulses have been set, the light source emits these light pulses in step i) through the optical fiber integrated on the junction (2) so that they can be sensed by the receiver in step ii). Both the light source and the receiver are included in the interrogation unit (18).

[0111] In step iii), the sensed light pulses are processed by a processor also included in the interrogation unit (18). The processor compares the light pulse corresponding to the light pulse detected by the receiver with the input light pulse corresponding to the light pulse emitted by the light source. Based on this light pulse comparison, the processor is able to determine the physical state of the adhesive line (2) in the composite structure (1). Thus, through this comparative analysis, this method allows for monitoring the physical state of the bonding portion (2) in the aerospace composite structure (1). In accordance with the parameters intended to be measured in the bonding portion (2), at least one multimode core (9) of the multicore optical fiber (6) is integrated to perform Raman scattering in a particular example. Raman scattering is an inelastic process caused by molecular vibrations. Incident light is scattered into two components, namely, Stokes at a higher wavelength and anti-Stokes at a lower wavelength. The ratio between the anti-Stokes light intensity and the Stokes light intensity is a direct measure of temperature. When the multi-core optical fiber (6) is connected to the interrogation unit (18), the Raman scattering components are compared at different timestamps in the direction of light within the multimode core (9) of the multi-core optical fiber.

[0112] In another specific example, at least one multimode core (9) of the multi-core optical fiber (6) is integrated to perform Rayleigh scattering. This is Elastic scattering The frequency of the scattered light remains constant relative to the input light. Analysis and correlation of backscattering variations at different stages were performed on the core of a multi-core optical fiber (6). As a result, temperature and / or strain can be monitored.

[0113] In another example, at least one, but preferably more than one, single-mode core (9) of the multi-core optical fiber (6) includes a multiplexed Bragg grating sensor. When strain is applied to the multi-core optical fiber (6) or a temperature change is detected, the change in reflected wavelength is detected by the interrogation unit (18).

[0114] In addition, the monitoring and comparison of engineering parameters measured by each core of the multi-core fiber enables the improvement of the accuracy of individual engineering parameters (such as temperature or strain), and thus enhances the compensation and isolation of coupling effects in the measurement of unique parameters.

[0115] For example, in the manufacturing process of an aerospace composite structure (1), it is important to monitor the temperature of the bonding portion (2) while the composite structure is in a curing cycle. The system of the present invention allows querying the multi-core optical fiber (6) integrated on each bonding portion (2) in order to determine the temperature of the multi-core optical fiber.

[0116] In another example, during the operational life of an aircraft (12) having several composite structures (1), the focus is on determining potential damage in the joints (2). To this end, the joints (2) are interrogated via multi-core optical fibers (6) to measure strain or deformation in these joints (2).

Claims

1. An aerospace composite structure (1) including a connecting portion (2) located between structural components (3, 4, 5), the aerospace composite structure (1) further comprising: - A plurality of multi-core optical fibers (6), wherein the plurality of multi-core optical fibers (6) are integrated in the joint portion (2) along the longitudinal direction (X-X'), each multi-core optical fiber (6) including at least two cores (9) and two fiber ends (6.1, 6.2) embedded therein, namely a first fiber end (6.1) and a second fiber end (6.2), each fiber end (6.1, 6.2) coinciding with the end (2.1) of the joint portion (2), and - Connectors (17) located at least at each fiber end (6.1, 6.2), the connectors being configured to connect each optical fiber (6) to the interrogation unit (18) at its respective fiber end (6.1, 6.2) in order to measure at least one parameter of the joint portion (2) in the aerospace composite structure (1), in, Each connector (17) is adapted to connect each core (9) to the interrogation unit (18); and Each core (9) of each multi-core optical fiber (6) is configured to transmit a predefined light pulse from the interrogation unit (18) along the optical fiber extension passing through the joint (2) from the first fiber end (6.1) to the second fiber end (6.2) according to the at least one parameter to be measured in order to monitor the physical state of the joint portion (2) between the structural components (3, 4, 5).

2. The aerospace composite structure (1) according to claim 1, wherein, The bonding portion (2) includes an adhesive line, into which at least the multi-core optical fiber (6) is embedded.

3. The aerospace composite structure (1) according to any one of the preceding claims, wherein, At least one core (9) of the multi-core optical fiber (6) is a multimode core, which is configured to provide Raman scattering when the connector (17) is connected to the interrogation unit (18).

4. The aerospace composite structure (1) according to any one of the preceding claims, wherein, At least one core (9) of the multi-core optical fiber (6) is a single-mode core, which is configured to provide Rayleigh scattering when the connector (17) is connected to the interrogation unit (18).

5. The aerospace composite structure (1) according to any one of the preceding claims, wherein, At least one core (9) of the multi-core optical fiber (6) is a single-mode core including a Bragg grating.

6. The aerospace composite structure (1) according to any one of the preceding claims, wherein, The multi-core optical fiber (6) includes a distribution of single-mode cores, wherein multiplexed Bragg grating sensors are written into different cores (9) of the multi-core optical fiber (6).

7. The aerospace composite structure (1) according to any one of the preceding claims, wherein, The aerospace composite structure is the leading edge (16) of the vertical tail (13), which includes the following structural components: - An inner panel base laminate (4) having multiple Ω-shaped struts (3), and -Outer panel (5) At least the outer panel (5) is connected to the inner panel base laminate (4) by an adhesive line located between one side of the outer panel (5) and the head of each Ω-shaped stringer (3), such that at least a multi-core optical fiber (6) is embedded in the adhesive line.

8. The aerospace composite structure (1) according to claim 7, wherein, Each Ω-shaped strut (3) is connected to the inner panel base laminate (4) by an adhesive line arranged between one side of the inner panel base laminate (4) and each foot of the Ω-shaped strut (3), such that at least a multi-core optical fiber (6) is embedded in each of the adhesive lines.

9. A system for monitoring the physical state of a joint (2) in an aerospace composite structure (1), the system comprising: -The aerospace composite structure (1) according to any one of claims 1 to 8, and - Interrogation unit (18), which is connected to the connector (17) of the aerospace composite structure (1) and is configured to measure at least one parameter in the joint portion (2) of the aerospace composite structure (1) in order to monitor the physical state of the joint portion (2); The interrogation unit (18) includes: A light source configured to emit light pulses through the first fiber end (6.1) of the multi-core optical fiber (6). A receiver configured to detect or sense 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. A method for monitoring the physical state of a joint portion (2) in an aerospace composite structure (1) according to any one of claims 1 to 8, the method comprising the steps of: a) Provide an interrogation unit (18), the interrogation unit (18) including a light source, a receiver and a processor, b) Connect the interrogation unit (18) to the connector (17) located at the ends (6.1, 6.2) of each multi-core optical fiber in the aerospace composite structure (1), and c) By transmitting a predefined optical pulse through at least two cores (9) of the multi-core optical fiber (6) according to parameters to be measured for monitoring the physical state of the joint portion (2) between structural components (3, 4, 5), at least one multi-core optical fiber (6) between connectors (17). Step c) includes: i. A predefined light pulse is emitted by the light source and passes through at least one core (9) of the multi-core optical fiber (6) from the first fiber end (6.1) of the multi-core optical fiber (6). ii. The received optical pulse is measured at the second fiber end (6.2) of the multi-core optical fiber (6) by the receiver, and iii. The measured light pulses are processed by the processor to monitor the physical state of the bonding portion (2) in the aerospace composite structure (1).

11. The method according to claim 10, further comprising, while the aerospace composite structure (1) is in the curing cycle, measuring the temperature in the bonding portion (2) by interrogating the at least one multi-core optical fiber (6) in step c), and monitoring the temperature in the bonding portion (2) of the aerospace composite structure (1).

12. The method according to any one of claims 10 to 11, further comprising monitoring damage in the joint portion (2) of the aerospace composite structure (1) by measuring strain or deformation in the joint portion (2) by interrogating the at least one multi-core optical fiber (6) in step c).

13. An aircraft (12) comprising an aeronautical composite structure (1) according to any one of claims 1 to 8.

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