阅读说明：本技术 航空复合结构及其物理状态监测方法和系统 (Aviation composite structure and physical state monitoring method and system thereof ) 是由 哈辛托·恩里克·罗德里格斯塞拉诺 卡洛斯·米格尔-希拉尔多 于 2021-06-10 设计创作，主要内容包括：本发明公开了一种航空复合结构,包括位于结构部件之间的结合部分；具有至少两个纤维芯的多芯光导纤维,多芯光导纤维沿着结合部分的纵向方向被集成在结合部中、并且包括两个纤维端部,每个纤维端部与结合部分的端部重合；以及至少位于每个纤维端部上的连接器,连接器被配置为将每个纤维连接到询问单元以便测量结合部分的至少一个参数,多芯光导纤维的每个芯被配置为根据为了监测结构部件之间的结合部分的物理状态而要测量的至少一个参数来传输预定义的光脉冲。本发明还公开了一种用于监测航空复合结构中的结合部分的物理状态的方法和系统。本发明还公开了一种包括此航空复合结构的飞行器。(The invention discloses an aeronautical composite structure, comprising a combination part positioned between structural components; a multicore optical fiber having at least two fiber cores, the multicore optical fiber being integrated in the joint portion along a longitudinal direction of the joint portion and including two fiber ends, each fiber end coinciding with an end of the joint portion; and a connector located at least on an end of each fibre, the connector being configured to connect each fibre to an interrogation unit for measuring at least one parameter of the bonded portion, each core of the multi-core optical fibre being configured to transmit a predefined pulse of light in dependence on the at least one parameter to be measured for monitoring the physical state of the bonded portion between the structural components. A method and system for monitoring the physical state of a bonded section in an aerospace composite structure is also disclosed. The invention also discloses an aircraft comprising the aviation composite structure.)

1. -an aeronautical composite structure (1) comprising a joining portion (2) between structural components (3, 4, 5), this aeronautical composite structure (1) further comprising:

-a multicore optical fiber (6) having at least two fiber cores (9), the multicore optical fiber (6) being integrated in the joint (2) along a longitudinal direction (X-X') of this joint (2) and comprising two fiber ends (6.1, 6.2), each fiber end coinciding with an end (2.1) of the joint (2), and

-a connector at least on each fiber end (6.2) configured to connect each fiber (6) to an interrogation unit (18) for measuring at least one parameter of a bonding portion (2) in the aeronautical composite structure (1),

wherein each core (9) of the multi-core optical fiber (6) is configured to transmit a predefined light pulse in dependence of the at least one parameter to be measured for monitoring the physical state of the joint section (2) between the structural components (3, 4, 5).

2. -an aerospace composite structure (1) according to claim 1, wherein the bonding portion (2) comprises a line of adhesive, at least the multicore optical fibres (6) being embedded in the line of adhesive.

3. -the aerospace composite structure (1) according to any of the preceding claims, further comprising a plurality of multicore optical fibers (6) integrated in the bonding portion (2), wherein one of the connectors (17) is located at a first multicore optical fiber end (6.1) and the other connector is located at a second multicore optical fiber end (6.2).

4. -the aerospace 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 multi-mode core configured to provide raman scattering when the connector (17) is connected to the interrogation unit (18).

5. -an aerospace composite structure (1) according to any one of the preceding claims, wherein at least one core (9) of the multi-core optical fibre (6) is a single mode core configured to provide rayleigh scattering when the connector (17) is connected to the interrogation unit (18).

6. -the aerospace 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 comprising bragg gratings.

7. The aerospace composite structure (1) according to any of the preceding claims, wherein the multicore optical fiber (6) comprises a single core profile, wherein multiplexed bragg grating sensors are written in different cores (9) of the multicore optical fiber (6).

8. -an aeronautical composite structure (1) according to any of the preceding claims, wherein the aeronautical composite structure is a leading edge (16) of a vertical tail (13), comprising the following structural components:

-an inner panel base laminate (4) with a plurality of omega-shaped 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 a line of adhesive between one side of the outer panel (5) and the head of each omega-shaped stringer (3), so that at least a multicore optical fibre (6) is embedded in the line of adhesive.

9. -the aerospace composite structure (1) of claim 8, wherein each omega-shaped stringer (3) is joined to the inner panel base laminate (4) by a line of adhesive arranged between one side of the inner panel base laminate (4) and each foot of the omega-shaped stringer (3), such that at least a multicore optical fibre (6) is embedded in each of the lines of adhesive.

10. -a system for monitoring the physical state of a bonding part (2) in an aeronautical composite structure (1), said system comprising:

-an aeronautical composite structure (1) according to any one of the preceding claims, and

-an interrogation unit (18) connected to a connector (17) of the aeronautical composite structure (1) and configured to measure at least one parameter in a bonded portion (2) of the aeronautical composite structure (1) in order to monitor a physical state of the bonded portion (2).

11. -a method for monitoring the physical state of a bonding part (2) in an aeronautical composite structure (1) according to any one of claims 1 to 9, comprising the steps of:

a) an interrogation unit (18) is provided,

b) connecting the interrogation unit (18) to a connector located at each multicore optical fibre end (6.1, 6.2) of the aeronautical composite structure (1), and

c) interrogating the multi-core optical fibre (6) between connectors (17) by transmitting a predefined light pulse through at least two cores (9) of the multi-core optical fibre (6) in dependence on a parameter to be measured for monitoring a physical state of the bonding portion (2) between structural components (3, 4, 5).

12. -the method according to claim 11, wherein said step c) comprises:

i. emitting predefined light pulses by a light source through at least one core (9) of the multi-core optical fiber (6),

measuring the received light pulse, and

processing the measured light pulses in order to monitor the physical state of the bonding part (2) in the aeronautical composite structure (1).

13. -the method according to any one of claims 11 and 12, further comprising monitoring the temperature in the bonded portion (2) of the aeronautical composite structure (1) by measuring the temperature in this bonded portion (2) by interrogating the multicore optical fibre (6) in step c) while this aeronautical composite structure (1) is in the course of a curing cycle.

14. -the method according to any one of claims 11 to 13, further comprising monitoring damage in the bonded portion (2) of the aeronautical composite structure (1) by interrogating the multicore optical fibre (6) in step c) to measure strain or deformation in this bonded portion (2).

15. -an aircraft (12) comprising an aeronautical composite structure (1) according to any of claims 1 to 9.

Technical Field

The present invention relates to an aeronautical composite structure intended for monitoring the physical state of the joints 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. These stringers, and the structural components of the composite structure mentioned, may be joined by adhesive lines. That is, this composite structure technology field typically utilizes adhesive joints suitable for composite structures during the manufacturing and assembly stages.

The adhesive joint is typically cured after the composite structure is placed in an autoclave, where temperature is one of the key parameters to be considered in the manufacture of such a structure. The temperature in the autoclave is today mainly controlled by conventional thermocouples, which need to be in direct contact with the surface of the composite material being cured and which only provide a punctual measurement during the whole 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 aerospace industry is constantly upgrading quality control, primarily to improve production and safety while reducing ultimate waste and rework costs. Currently, once an aircraft is put into service and periodic inspections are required or unexpected events are experienced, operators must disassemble the aircraft to discover the possible damage experienced by the adhesive lines of the joints in the composite structure. The existing examinations are considered to be a complicated manual work process and certainly time consuming.

Today, the adhesive lines in composite structures undergo quality control by non-destructive inspection (NDI), such as ultrasonic pulse-echo. NDI control requires specific tools, certified inspectors, access to the inspection area once the aircraft returns to the ground, which also means that economy and scheduling are impacted. 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 investment of time.

There are known systems for monitoring the integrity of the adhesive within the cured bondline of a bonded structural assembly, similar to the system described in patent application US 8812251B 2. Heretofore, the adhesive line was monitored using a network of electrical sensors disposed inside the adhesive line. In addition, these systems for monitoring the status of the adhesive include a power source for providing power to a network of electrical sensors in order to check the integrity of the adhesive line as needed. The system is made by interpreting the changes measured directly within the cured bond line. However, such known apparatus and methods provide monitoring only after the adhesive line is manufactured, but not during the manufacturing process.

In another technical field, further known are optical fibers made of a plurality of cores positioned along the diameter of the cladding, which can be made to respond to temperature or strain changes. These variations are typically measured by the selection of the optical fiber materials (especially the core, cladding and coating) and the spacing and shape of the core within the fiber. Today, the temperature variations measured by optical fiber optimization can be provided over a wide range.

Furthermore, there are known composite structures integrated with optical fibres capable of localizing damaged along the composite structure, and methods for manufacturing such composite structures integrated with the ability to localize damaged. It is well known to provide optical fiber connection devices that provide optical alignment with minimal insertion loss, and which may be placed anywhere on the composite surface to collect in-service parameters and continue to collect information from the connected optical fibers.

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 an improved method provides manufacturing advantages over systems and methods known in the art. Additionally, the present invention further provides a system for monitoring temperature, strain or deformation during in-service inspection and in-service operation of a composite structure.

Disclosure of Invention

The present invention provides a solution to the above-mentioned problems by an aeronautical composite structure according to the first inventive aspect, a system according to the second inventive aspect, a method for monitoring the physical state of a joint in an aeronautical composite structure according to the third inventive aspect, and an aircraft according to the fourth inventive aspect. Preferred embodiments of the invention are defined in the dependent claims.

In a first aspect of the invention, the invention provides an aerospace composite structure including a joint between structural members, the aerospace composite structure further comprising:

-a multicore optical fiber having at least two fiber cores, the multicore optical fiber being integrated in the joint along a longitudinal direction of this joint portion and comprising two fiber ends, each fiber end coinciding with an end of the joint portion, and

-a connector on at least an end of each fiber, the connector being configured to connect each fiber to an interrogation unit for measuring at least one parameter of a bonded portion in the aerospace composite structure,

wherein each core of the multi-core optical fiber is configured to transmit a predefined light pulse in accordance with the at least one parameter to be measured for monitoring the physical state of the bonded portion between structural components.

Aircraft are made up of a plurality of aerospace composite structures formed by the joining of various structural components. For example, a torsion box for a tailplane of an aircraft is formed, among other things, by structural components such as panels, stiffeners (i.e., stringers, ribs … …), frames, and skins that are joined together such that, for each joint, the deployed composite structure includes a bonded portion.

These joints between structural components are vital joints in composite structures, and their physical state is of particular concern to monitor. In particular, from the time of manufacture of the aerospace composite structure (i.e., assembly of the structural components) until it is integrated in the aircraft, including also during the operational life of the aerospace composite structure.

The present invention provides improved configurations for joints between these structural components of integrated aerospace composite structures to improve monitoring of the physical condition of these joints. In particular, the invention proposes to provide at least an optical fibre integrated in a joining portion of the composite structure, which joining portion corresponds to a joining portion between said structural components.

An optical fiber is a multicore optical fiber comprising at least two fiber cores embedded within the fiber. The optical fiber is covered by a conventional cladding and includes two fiber ends. The multicore optical fibres are integrated in the joint portion of the composite structure along the longitudinal direction of the same joint portion. I.e. this longitudinal direction corresponds to the direction in which the joint portion follows in the joint between the structural parts of the composite structure. The multicore 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.

The composite structure further comprises at least a connector which can connect each optical fibre with its respective fibre end to an interrogation unit intended to monitor the physical state of the composite structure, in particular of the bonded part. More particularly, the connector is adapted to connect each fibre optic core to an interrogation unit. The connector is fan-out to access each fiber core of a multi-core optical fiber. The connector distributes light from the optical fiber to each core of the multi-core optical fiber. The fact that the connectors are connected to both ends of the optical fibre enables greater flexibility in application and interrogation at both ends of the optical fibre. In addition to providing a connector, the two fibre ends further allow interrogation of the fibre core to measure a physical parameter (such as Brillouin scattering) using a technique that requires connection at the two fibre ends.

A physical state must be understood in the context of the present invention as every physically distinguishable case or form obtained by measuring certain characteristics that the composite structure may adopt in its temporal evolution. That is, in the bonded portion of the composite structure that undergoes changes, the physical state is anything possible due to these changes. Examples of such physical states of the composite structure are temperature, strain, deformation, damage, load, vibration and fire detection.

In particular, providing a connector connecting the fiber end to an interrogation unit allows measuring at least one parameter of the bonded portion in order to monitor the physical state of the composite structure. In this way, the fiber core integrated in the multi-core optical fiber transmits the predefined light pulses from the interrogation unit along the optical fiber extension through the coupling portion. I.e. the optical fibre transmits a predefined light pulse through the fibre core depending on the parameter to be measured. Thus, by monitoring the measured parameter, the physical state in the bonded portion of the composite structure can be determined.

In another particular embodiment, the connector is a direct connector configured to connect to a general-purpose interrogator. This card edge connector includes a precision handling system of fiber cores that allows for simultaneous alignment of all fiber cores, where insertion loss values and return loss values are compatible with the dynamic range of each technology applied to the fiber cores. The general purpose interrogator is configured to interrogate all fiber cores using different techniques (e.g., FBG, Raman (Raman), Rayleigh (Rayleigh) or brillouin).

The present invention advantageously allows monitoring of the physical state of bonded portions in a composite structure during the manufacturing process of the composite structure, during in-service inspection, and during the operational life of the composite structure.

Furthermore, the presence of multiple cores in the optical fiber enables the optical monitoring technique of interest to be selected and applied at each lifetime step of the composite structure. That is, the fiber core may be intended to measure different parameters of interest or the same parameter with respect to the bonded portion. This parameter may be temperature or strain or deformation or damage or load or vibration or fire detection.

Advantageously, the proposed solution aims at supporting and improving the quality of control of the bonding portions during the curing process (manufacturing step). Since the optical fibers are embedded inside the bonded part of the composite structure, the measured temperature is more reliable than the external sensors in prior art solutions. In addition, the provision of optical fibres enables measurements along the length of the optical fibres rather than at on-time locations as in prior art solutions, and therefore control quality and temperature mapping can be done more extensively.

For structural testing or inspection and in-service operation, the provision of optical fibers further enables monitoring of structural performance by strain measurement and even the presence of damage, such as disbonding, on the bonded portions. That is, the present invention allows for the detection of damage in the bonded portion without the need to adjust the aircraft and access the affected area on the composite structure.

Thus, the multi-core fiber optic acts as a permanent sensor mounted inside the joint between the structural members, this fiber optic being interrogated according to maintenance and operator requirements.

Accordingly, the invention provides mainly the following advantages with respect to conventional solutions:

improved quality control of the joint between the structural parts by means of a multicore optical fibre controlling the longer surface of the composite structure compared to prior art solutions.

-monitoring the temperature and strain of the bonded portion during manufacturing. This also facilitates on-line inspection during manufacturing time and aims to detect possible defects, even so that they can be corrected before the manufacturing process is completed.

Monitoring the in-service mechanical behaviour by providing an optical fibre integrated in the joint, thus allowing to control this joint and detect possible damages without touching the aircraft.

In a particular embodiment, the bonding portion comprises a binder thread into which at least the multicore optical fibers are embedded. This line of adhesive corresponds to the joining means arranged between the two structural parts for their joint. In this embodiment, the optical fibers are embedded within the adhesive line. The use of integrated optical fibres in the adhesive line advantageously eliminates the need for drilling the strength members for the joint, thus simplifying the manufacturing process and is desirable from a structural point of view.

The type of adhesive used for the adhesive lines depends on the operating conditions, the application requirements, and the materials to be joined. The operating conditions and application requirements may be temperature range, dynamic or static load conditions, necessary chemical resistance, durability, application time and curing time. The materials to be joined to accommodate the composite structure are typically metals, polymers, or ceramic materials.

In a more particular embodiment, the adhesive line is an epoxy adhesive. Epoxy resins are preferred adhesives in view of the materials, required resistance, temperature range and working time of known aerospace composite structures.

In another particular embodiment, the adhesive line is a silicone, cyanoacrylate, polyurethane or phenolic adhesive.

In a particular embodiment, the bonding portion comprises a plurality of adhesive lines.

In a particular embodiment, the aerospace composite structure comprises a plurality of multicore optical fibers integrated into the bonding portion, wherein one of the connectors is located at a first multicore optical fiber end and another connector is located at a second multicore optical fiber end.

Advantageously, the plurality of multicore fibers provides redundancy, which provides the possibility of having spare fibers in case of a fiber failure. In addition, a greater number of parameters can be monitored simultaneously by a plurality of multicore fibers. Furthermore, the accuracy of the measurement is also improved, since multiple cores/fibers can monitor the same parameters.

In a particular embodiment, the multicore optical fiber comprises a coating. The known optical fibre is covered by a cladding layer. The coating is thus an additional coating to the cladding layer and provides mechanical protection to the multicore optical fiber.

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

In a particular embodiment, at least one core of the multi-core optical fiber is a multi-core configured to provide raman scattering when the connector is connected to the interrogation unit. By multi-mode core is meant that the fiber is configured to propagate several optical modes at once.

In a particular embodiment, at least one core of the multi-core optical fiber is a single core 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 optical mode at a time.

In a particular embodiment, at least one core of the multi-core optical fiber is a single core comprising a Bragg grating. More particularly, the multi-core optical fiber may be a single mode core including a multiplexed bragg grating sensor.

In a particular embodiment, the multi-core optical fiber comprises a distribution of single core cores, wherein multiplexed bragg grating sensors are written in different cores of the multi-core optical fiber. Advantageously, providing multiplexed bragg grating sensors in the fiber core allows matching multiple spatial resolution requirements over hundreds of meters of optical fiber, while requiring the installation of a unique fiber.

The fiber core diameters, wavelengths, light sources, and bandwidths of the single mode fibers and the multimode fibers are different from each other. The core diameter of the single mode fiber is smaller than the core diameter of the multimode fiber. The wavelength of the multimode fiber is less than the wavelength of the single mode fiber. In addition, 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 at a time. Furthermore, single mode fibers are suitable for long distance applications, while multimode fibers are designed for short distances.

Considering that a multi-core optical fiber integrates multiple cores (single mode and multi-mode) for a single optical fiber, it allows:

reducing the number of fibers to be installed and simplifying it, since a single optical fiber may comprise a plurality of cores and each core comprises a plurality of sensors or distributed sensors. From the viewpoint of aircraft applications, this aspect is very important, for example, in temperature measurement applications with critical spatial resolution (in the mm or cm order) and requiring hundreds of meters over the entire plane, and in damage detection applications requiring damage detection resolution of the order of a few mm; and-increasing the spatial resolution in those applications where there is a limit to the distance between the fiber ends.

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

In a particular embodiment, the aeronautical composite structure is a leading edge of a vertical tail, the leading edge comprising the following structural components:

-an inner panel base laminate with a plurality of omega-shaped stringers, and

-an outer panel which is,

at least the outer panel is joined to the inner panel base laminate by a line of adhesive between one side of the outer panel and the head of each omega stringer, such that at least a multicore optical fibre is embedded in the line of adhesive.

In a more particular embodiment, each omega stringer is joined to the inner panel base laminate by a line of adhesive arranged between one side of the inner panel base laminate and each foot of the omega stringer, such that at least a multicore optical fibre is embedded in each of the lines of adhesive.

In a second inventive aspect, the present invention provides a system for monitoring the physical state of a bonded section in an aerospace composite structure, the system comprising:

-an aeronautical composite structure according to the first inventive aspect, and

-an interrogation unit connected to a connector of the aeronautical composite structure and configured to measure a parameter in a bonded portion of the aeronautical composite structure in order to monitor a physical state of the bonded portion.

That is, the main function of this interrogation unit is to interrogate the fibres integrated in the joint between the structural components to measure the parameter of interest in order to determine the physical state in which the joint is located. An interrogation unit connected to the connector of the composite structure allows monitoring of the physical state of the bonded portions in the composite structure during manufacture of the composite structure and during the operational life thereof.

In a more particular embodiment, the interrogation unit comprises:

a light source configured to emit light pulses through a first fiber end of a multi-core optical fiber,

a receiver configured to detect or sense the emitted light pulses through the second fiber end, an

A processor configured to process the sensed light pulses.

The interrogation unit is responsible for interrogating the optical fibre by emitting light pulses through the end of the optical fibre by the light source and detecting these pulses through the end of the optical fibre by the receiver. Further, by this interrogation, the processor of the interrogation unit performs an analysis comparing the light pulse output (light pulses detected by the receiver) with the light pulse input (light pulses emitted by the light source).

Depending on the parameters to be measured (temperature, strain, deformation, damage, load, vibration and fire detection), the light pulses are configured with certain characteristics so that, based on the above-mentioned comparative analysis, the physical state of the binding moiety can be determined based on said parameters.

In a third inventive aspect, the present invention provides a method for monitoring the physical state of a bonded section in an aerospace composite structure according to the first inventive aspect, the method comprising the steps of:

a) an interrogation unit is provided which is adapted to interrogate the device,

b) connecting the interrogation unit to the connectors located on each multicore optical fibre end of the aeronautical composite structure, an

c) Interrogating the multi-core optical fibre between connectors by transmitting a predefined light pulse through at least two cores of the multi-core optical fibre in dependence on a parameter to be measured for monitoring a physical state of the bonding portion between structural components.

By means of the method of the invention, it is possible to monitor the physical state of the joints between the components of an aeronautical composite structure both during the manufacturing and during the operational life of the aircraft in which said composite structure is installed.

In a particular embodiment, said step c) comprises:

i. emitting a predefined light pulse by a light source through at least one core of the multi-core optical fiber,

measuring the received light pulse, and

processing the measured light pulses so as to monitor a physical state of the bonded portion in the aerospace composite structure.

In a particular embodiment, the method further comprises monitoring the temperature in the bonded portion of the aerospace composite structure by interrogating the multicore optical fiber in step c) to measure the temperature in the bonded portion while the aerospace composite structure is during a cure cycle.

In a particular embodiment, the method further comprises monitoring damage in the bonded portion of the aerospace composite structure by interrogating the multicore optical fibre in step c) to measure strain or deformation in this bonded portion.

In a fourth inventive aspect, the invention provides an aircraft comprising an aerospace composite structure according to the first inventive aspect.

In more particular embodiments, the aircraft includes a plurality of aerospace composite structures, such as horizontal stabilizers, vertical stabilizers, and wings.

The provision of aircraft manufactured with aeronautical composite structures equipped with multicore optical fibres integrated on the joined portions of these structures advantageously allows monitoring the condition of these joined portions by inspection of the aircraft.

Drawings

These and other features and advantages of the present invention will be more clearly seen from the following detailed description of preferred embodiments, given purely by way of illustrative and non-limitative example, with reference to the accompanying drawings.

FIG. 1 this figure shows a perspective view of an aerospace composite structure, in accordance with embodiments of the invention.

Figure 2 this figure shows an exploded view of the aerospace composite structure shown in figure 1.

FIG. 3 this figure illustrates a perspective view of a bonded portion of an aerospace composite structure, according to an embodiment of the invention.

Fig. 4a to 4c these figures show cross-sectional views of a multi-core optical fiber according to an embodiment of the present invention.

FIG. 5 this figure shows a schematic view of a monitoring system according to an embodiment of the invention.

Figure 6 this figure shows a side view of an aircraft including an aerospace composite structure according to an embodiment of the invention.

Detailed Description

As will be appreciated by one skilled in the art, aspects of the present invention may be embodied in an aerospace composite structure, a system or method for monitoring the physical state of a bonded portion of such an aerospace composite structure.

The invention provides an aeronautical composite structure (1) provided with at least a multicore optical fibre (6) integrated in a joining portion (2) between structural components (3, 4, 5) of said aeronautical composite structure (1). This configuration of the multicore optical fibres (6) allows monitoring the physical state of the bonded portions (2) during the manufacture of the aeronautical composite structure and also during inspection of the aeronautical composite structure in service.

Fig. 1 shows a perspective view of an aeronautical composite structure (1) corresponding to the leading edge of a lifting surface (for example a horizontal tail). This leading edge (1) is formed by a structural component being an inner panel base laminate (4) and a plurality of stringers (3). In particular, the inner panel base (4) is mounted on various ribs (7.1, 7.2, 7.3, 7.4) to shape the aeronautical composite structure or leading edge (1) into a semi-elliptical shape. The base of the leading edge (1) is covered by a film (10) made of composite material.

More particularly, the leading edge (1) comprises two end ribs (7.1; 7.3) located at both ends of the aeronautical 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) (as shown in figure 1). Furthermore, a plurality of stiffening ribs (7.4) placed inside the inner panel base laminate (4) give the laminate (4) rigidity and shape retention in the leading edge (1).

Both parts of the inner panel base laminate (4) comprise omega-shaped stringers (3) which are positioned along the leading edge (1) and lie parallel to each other along the surface of the laminate (4). In particular, these omega-shaped stringers (3) are joined to the inner panel base laminate (4) (shown in fig. 2) by lines of adhesive in the joint (2) between the inner panel laminate (4) and the outer panel (5). This adhesive line (not shown in fig. 1 and 2) is located between one side of the inner panel base laminate (4) and each foot of the omega-shaped stringer (3) so as to embed at least one multicore optical fibre (6) in each bonding portion (2).

Fig. 2 shows an exploded view of the leading edge (1) of fig. 1. More particularly, this figure 2 shows an outer panel (5) to be mounted on the inner panel base laminate (4) to cover the entire inner panel base laminate (4). In particular, as shown in fig. 3, the adhesive lines to be placed on each of the plurality of omega-shaped stringers (3) join the mentioned outer panel (5) to the inner panel base laminate (4) providing resistance and fixation to the structural components of the leading edge (1).

In a preferred embodiment, adhesive strands such as adhesives made from epoxy, silicone, cyanoacrylate, polyurethane, phenolic, and the like are used, depending on the type of materials to be assembled.

Figure 3 shows a perspective view of a portion of an aerospace composite structure (1) (such as the structure shown in figure 2) with an inner panel base laminate (4), omega stringers (3) and outer panel (5). At one foot of the omega-shaped stringer (3), the omega-shaped stringer (3) is attached to the inner panel base laminate (4) over the portion identified as the bonding portion (2). In addition, the outer panel (5) is attached to the head of the omega-shaped stringer (3), in particular to its outer surface, on the portion also identified as coupling portion (2). These binding moieties (2) are arranged along a longitudinal direction (X-X'). Contact between the structural components (3, 4, 5) in these bonding portions (2) is ensured via lines of adhesive of a predetermined bonding width. A plurality of multicore optical fibers (6) are integrated into each long coupling portion (2). Each multicore optical fibre (6) comprises two fibre ends (6.1; 6.2) coinciding with the ends of the bonding portion (2) at each end of the multicore optical fibre (6).

Each fibre end (6.1; 6.2) is provided with a connector (not shown) for connecting the fibre (6) to an interrogation unit (not shown) for measuring a parameter of the bonded part (2) in the aeronautical composite structure (1), such as temperature, deformation or strain. Each multi-core optical fibre (6) comprises at least two fibre cores (9) transmitting predefined optical pulses according to previously set parameters to be measured for monitoring the physical state of the bonding part (2).

In a specific example, the interrogation unit (18) emits a light pulse from the first fiber end (6.1) through the multi-core optical fiber (6) by means of a light source (not shown in the figure). This emitted light is then detected at the second fibre end (6.2) by a receiver of the interrogation unit (18). Upon sensing the emitted light pulse, the sensed light pulse is processed by a processor included in the interrogation unit (18).

In a particular example, the multicore optical fiber (6) comprises at least two cores (9) integrated inside the cladding (11), preferably spaced apart at a pitch of 35 to 70 microns.

Fig. 4a to 4c show cross-sectional views of a multi-core optical fiber (6), wherein the arrangement of the cores (9) inside the multi-core optical fiber (6) is star-shaped (fig. 4b) or hexagonal (fig. 4a and 4 c). More precisely, one multicore optical fiber (6) comprises seven cores (9) (fig. 4a), thirteen cores (9) (fig. 4b) and nineteen cores (9) (fig. 4 c).

Fig. 4a to 4c further show a multicore optical fiber (6) which is covered by a coating (8) in addition to the cladding (11), which coating provides mechanical protection for the multicore optical fiber (6). In a particular example, the coating (8) is made of polyamide for temperature and deformation measurements, metal for fire detection and reinforced polymer for further determined measurements.

One core (9) may be single-mode to perform bragg scattering, brillouin scattering or rayleigh scattering, or multimode to perform raman scattering. In a preferred example, each multicore optical fibre (6) comprises at least each type of core. On the one hand, this preferred construction reduces the number of substance cores (9) and simplifies them when producing the multicore optical fiber (6). On the other hand, the preferred structure also improves the spatial resolution to reduce the distance between the sensors of the multicore optical fibers (6) on the aeronautical composite structure (1).

In a further preferred embodiment, at least one multimode core (9) of the multicore optical fiber (6) is integrated so as to provide raman scattering when connected to the interrogation unit (18). Furthermore, at least one single core (9) of the multicore optical fiber (6) is provided for providing rayleigh scattering when connected to the interrogation unit (18). Finally, at least one single core (9) of the multi-core optical fiber (6) is integrated for performing bragg grating sensing measurements.

Fig. 5 shows a system for monitoring the physical state of a bonded part (2) in an aeronautical composite structure (1). Multicore optical fibres (6) are embedded in the joint (2) between the structural components (4) corresponding to the panel laminate base of the composite structure (1). The system further comprises an interrogation unit (18) connected to the multicore optical fibre (6) by a connector (17). In particular, each connector (17) is attached to each fiber end (6.1, 6.2), allowing a connection between the interrogation unit (18) and both fiber ends (6.1, 6.2) of the multicore optical fiber (6). The interrogation unit (18) transmits predefined light pulses through at least two cores (9) of a multi-core optical fibre (6) in dependence on a parameter to be measured for monitoring a physical state of a bonding portion (2) between two panel laminate bases (4).

Fig. 6 shows a side view of an aircraft (12) comprising a vertical tail (13), a horizontal tail (14) and wings (15), all of which are aeronautical composite structures (1) according to embodiments of the invention. Each of the empennage (13, 14) and the wing (15) presents a leading edge (16) comprising a multicore optical fibre (6) integrated on the joining portion (2) for providing the capability of measuring the physical state of the structure. The leading edge (16) is the first component of each aeronautical composite structure (1) that comes into contact with the oncoming airflow while the aircraft (12) is in service.

Method for monitoring the physical state of a bonding part (2) in an aeronautical composite structure (1)

The invention further provides a method for monitoring the physical state of a bonded section (2) in an aeronautical composite structure (1), such as the vertical tail shown in figures 1 and 2.

The monitoring method comprises the following steps:

a) an interrogation unit (18) is provided,

b) connecting an interrogation unit (18) to a connector located at each multicore optical fibre end (6.1, 6.2) of the aeronautical composite structure (1), and

c) interrogating the multi-core optical fibre (6) between the connectors (17) by transmitting a predefined light pulse through at least two cores (9) of the multi-core optical fibre (6) in dependence on a parameter to be measured for monitoring the physical state of the joint (2) between the structural components (3, 4, 5).

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

Once the connector (17) of the multi-core optical fibre (6) is connected to the interrogation unit (18), step c) starts an interrogation of the multi-core optical fibre (6). For this step c), a predefined light pulse is emitted through the optical fiber according to the parameter to be measured on the binding moiety (2). Such parameters may be temperature, strain, deformation, damage, load, vibration and fire detection. The light pulses are thus configured with characteristics based on the parameter to be measured.

Once the light pulses have been set, the light source emits these light pulses in step i) through an optical fibre integrated on the binding portion (2) to be sensed by the receiver in step ii). Both the light source and the receiver are included in an interrogation unit (18).

In step iii), the light pulses that have been sensed are processed by a processor also included in the interrogation unit (18). The processor compares the light pulses corresponding to the light pulses detected by the receiver with the input light pulses corresponding to the light pulses emitted by the light source. Based on this light pulse comparison, the processor is able to determine the physical state of the adhesive lines (2) in the composite structure (1). Thus, through such a comparative analysis, the method allows monitoring the physical state of the bonding portions (2) in the aeronautical composite structure (1). At least one multi-core (9) of the multi-core optical fiber (6) is integrated to perform raman scattering in a specific example, following parameters intended to be measured in the bonding section (2). Raman scattering is an inelastic process caused by molecular vibrations. Incident light is scattered into two components, namely Stokes (Stokes) at higher wavelengths and anti-Stokes at lower wavelengths. The ratio between the anti-stokes light intensity and the stokes light intensity is a direct measure of temperature. When the multicore optical fiber (6) is connected to an interrogation unit (18), the raman scattered components are compared in different time stamps in the direction of light inside the multicore optical fiber's multimode core (9).

In another particular example, at least one multimode core (9) of the multicore optical fiber (6) is integrated to perform rayleigh scattering. This is thatElastic scatteringWherein the frequency of the scattered light remains unchanged with respect to the input light. The analysis and correlation of the changes in backscatter at different stages is tracked on the core of a multi-core optical fibre (6). As a result, temperature and/or strain may be monitored.

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

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

For example, during the manufacture of an aerospace composite structure (1), it is of interest to monitor the temperature of the bonded portions (2) while the composite structure is in a curing cycle. The system of the invention allows interrogating the multicore optical fibres (6) integrated on each of the bonding portions (2) in order to determine the temperature of the multicore optical fibres.

In another example, during the operational lifetime of an aircraft (12) having several composite structures (1), it is of interest to determine possible damage in the bonded portion (2). For this purpose, the joining sections (2) are interrogated by means of a multicore optical fibre (6) in order to measure the strain or deformation in these joining sections (2).