JPH09273906A - Damage and breakage position detector - Google Patents

Abstract

PROBLEM TO BE SOLVED: To immediately detect the damage position of an airframe during ground maintenannce and flying by providing an optical fiber cqble along the surface layer surface of the airframe or in the layer, disposing it in a matrix state, radially reducing the thickness of a core, and forming a stress sensitive unit to raise the sensitivity of a strain. SOLUTION: An optical fiber cable constitutes the element of a damage or breakage position detector 2. A stress sensitive unit for constituting a sensor partly applies lengthwise tension and heat to the cable to form a concave curved surface-like reduceddiameter part, and provided at a certain interval along the lengthwise direction of the cable. When a radial external force is applied to the unit, a stress is generated and the transmittin amount of the light which transmits the core is changed according to the bending angle. The unit detects the change of the transmitting light amount with high sensitivity. Such a cable is embedded, for example, on the surface layer surface or in the surface layer of a horizontal stabilizer 13, and formed in a matrix to cover substantially the entire surface of the stabilizer 13.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a damage / damage position detecting device, and more particularly to a damage / damage position detecting device for detecting a damage position such as strain, crack or loss of an aircraft body and a damage state. It is a thing.

[0002]

2. Description of the Related Art Aircraft that fly at high speed in the atmosphere are constantly navigating with strong wind pressure applied to their bodies. Therefore,
In particular, there is a high probability that distortion, cracks, etc. will occur in proportion to the flight time on the surfaces of the wing and tail that overhang sideways. Conventionally, the grasping of the damage state due to the deformation and cracks of the body of the aircraft has been carried out mainly by visual inspection by maintenance personnel. Further, it may be performed by a mechanical inspection method using an ultrasonic flaw detector or a magnetic particle flaw detector.

[0003]

SUMMARY OF THE INVENTION As described above, the finding of the failure state of the airframe that occurs in the conventional aircraft of this type has been mainly done by sight. Therefore, the skill level of the maintenance staff was required. In addition to the above, there is a mechanical inspection method as described above, but in any case, it was carried out on the ground such as an airfield or a maintenance hangar. Therefore, the conventional damage / damage position detection method has a problem that the flight of the aircraft must be stopped for a long time.

The present invention has been made in order to solve the drawbacks of the conventional damage / damage detection system, and can immediately detect the damaged portion of the aircraft not only during maintenance on the ground but also during flight. It is intended to realize a damage / damage position detection device in an aircraft.

[0005]

SUMMARY OF THE INVENTION The present invention comprises a damaged / damaged position detecting device in which an optical fiber cable is provided on the surface layer or in the surface layer of a machine body. Further, the optical fiber cable is arranged in a matrix to constitute a damaged / damaged position detecting device. Further, a damage / breakage position detecting device is configured in which a stress sensitive portion is formed in order to increase the sensitivity to strain by reducing the wall thickness of the core of the optical fiber cable.

When a part of the fuselage of an aircraft is damaged by an external force due to an external force, bending stress is generated in the X-component and Y-component optical fiber cables arranged in the area of the matrix of the detection device. Therefore, the transmitted light amount of the measurement light transmitted through the stress sensitive portions of both X and Y components forming the matrix is attenuated according to the bending angle, and this attenuation amount is detected by the light receiver. Therefore, it is determined that the region corresponding to the attenuated plural components of X and Y has been damaged by the depression. At this time, the damage state in the damage area is also estimated based on the amount of attenuation of a plurality of X and Y components.

Further, when a part of the machine body is damaged due to peeling or missing, the transmitted light of the stress sensitive parts of the X and Y components leading to the damaged part is scattered or reflected. The part surrounded by the X and Y components that received the scattered light and the reflected light is judged to have been destroyed and dropped.

[0008]

BEST MODE FOR CARRYING OUT THE INVENTION

Embodiment 1 FIG. 1 is a configuration explanatory view of an aircraft to which Embodiment 1 of the present invention is applied, and FIG. 2 is an enlarged explanatory view of a part of FIG. 1 and 2, reference numeral 1 is an aircraft. Reference numeral 10 is a fuselage of the aircraft 1, 11 is a main wing, 12 is an engine, and 13 and 14 are horizontal and vertical tails. In the figure, a twin-engine aircraft having left and right engines 12 is illustrated as the aircraft 1. Reference numeral 15 is an auxiliary wing, 16 is a flap, 17 is a spoiler, 18 is an elevator provided on a horizontal stabilizer, and 19 is a rudder.

In FIG. 1, reference numeral 2 is a damaged / damaged position detecting device (hereinafter referred to as a detecting device) according to the present invention, and 3 in FIG. 3 and 4 are perspective views showing the specific structure of the optical fiber cable 3 and the sensor portion in an enlarged manner. 3 and 4, 3a is a coating,
31 is a core of the optical fiber cable 3, 32 is a clad,
33 is a boundary layer. The clad 32 is the central core 31
Surrounds the.

In FIGS. 3 and 4, reference numeral 34 is a stress sensitive portion (stress sensor) which constitutes a sensor. The stress-sensitive portion 34 is formed by a concave portion having a small diameter by partially applying longitudinal tension and heat to the optical fiber cable 3, and as shown in the drawing, has a gap along the length direction of the optical fiber cable 3. Are separated from each other. And the optical fiber cable 3
When an external force in the radial direction is applied to the stress sensitive portion 34, a stress σ corresponding to this external force is generated, and the amount of light L transmitted through the core 31 in the arrow direction changes according to the bending angle.
The relationship between the transmitted light amount v of the light L and the stress σ at this time is shown in FIG. As shown in FIG. 5, the transmitted light amount v has a characteristic that it decreases almost in proportion to the increase of the stress σ.

Generally, when an external force is applied to this type of optical fiber, stress is generated inside the optical fiber to cause "mode conversion".
The phenomenon occurs. When the "mode conversion" phenomenon occurs, the total reflection angle of the boundary layer changes. As a result, the light transmitted through the inside of the optical fiber is diffusely reflected by the boundary layer or transmitted through the outer cladding and is dissipated to the outside, resulting in a loss of light energy called radiation loss. Further, when pressure is applied from the axial direction of the optical fiber, the axial center is bent by about several μm, and the boundary layer becomes uneven. Then, the amount of transmitted light changes based on the similar "mode conversion" phenomenon, and optical loss called "micro-bending loss" occurs.

The small-diameter stress sensitive portion 34 provided in the optical fiber cable 3 detects with high sensitivity the change in the transmitted light amount v due to the loss of light energy as described above. On the other hand, when the optical fiber cable 3 that transmits the light L into the core 31 is cut in the middle (see point p in FIG. 6), the light L in progress is scattered or reflected by the cut surface. As a result, the scattered light and the reflected light on the cut surface travel in the optical fiber cable 3 in the opposite directions and are returned to the incident side.

Such an optical fiber cable 3 is, for example, as shown in FIG. 2, a horizontal stabilizer 13 made of a composite material or the like.
Embedded in or on the surface layer of. The illustrated horizontal optical fiber cable 3 has an X component and a vertical direction has a Y component. Both components intersect in rows and columns in the X direction and the Y direction to form a matrix M so as to cover almost the entire surface of the horizontal stabilizer 13. Reference numeral 30 in FIG. 2 denotes a Y component optical fiber bundle, and the Y component input / output optical fiber cables 3 are put together.

6 and 7 are block diagrams showing the configuration of the detection apparatus according to the first embodiment of the present invention in which the optical fiber cable 3 composed of the X component and the Y component is arranged. In FIG.
X1, X2 ... Xm are optical fiber cables 3 of X component for forming the matrix M, and Y1, Y2 ... Yn.
Is the Y component (however, m and n are positive integers), C is both components X1
, X2 ... Xm and intersections of Y1, Y2 ... Yn [xm, yn
] (Hereinafter, also refer to FIG. 8). And the above X,
A large number of stress sensitive parts 34 in both Y components are respectively formed at respective intersections C.
It is arranged in the vicinity of [xm, yn] and each intersection C [xm,
yn], the change of the transmitted light amount v is detected.

6 and 7, 4 is a control processor for the X component, 5 is a control processor for the Y component, 6 is an integrated processor, and 7 is a flight controller. 41 is X
The component side light source, 42 is an optical amplifier, and 43 is a branching device.
44 is an optical coupler, 45 is a light receiver, and 46 is a light receiving element such as a photo diode. Each of the optical coupler 44, the light receiver 45, and the light receiving element 46 is composed of m pieces corresponding to the number of the X component side optical fiber cables 3 forming the matrix M. Inside the control processor 4, a converter and the like for mutually converting an electric signal and an optical signal on the X component side are provided.

FIG. 7 shows a block diagram of the Y component side. Since the Y component side has a similar structure to the X component side, the reference numerals corresponding to the light sources 51 and the like are given and detailed description thereof is omitted.
However, the Y component of the optical fiber cable 3 is n and constitutes the vertical axis of the diagram of the matrix M. The output terminal of the control processor 5 is connected to the output terminal of the control processor 4 and the flight controller etc. 7 via the integrated processor 6 of FIG. S1 is a deformation detecting unit such as unevenness, and S2 is a damage detecting unit such as peeling or missing of the machine body.

The operation of the present invention having the above configuration will be described below with reference to FIGS. A light source 4 such as a laser diode based on a command from the control processor 4 on the X component side.
The measuring light L emitted from the optical fiber 1 is amplified by the optical amplifier 42 and is branched into m pieces by the branching device 43. Light L divided into m
After passing through the polarization-independent optical coupler 44, the light is projected inside the optical fiber cables 3 of the X components X1, X2 ... Xm in the matrix M, respectively. Similarly, Y
The measurement light L emitted from the light source 51 by the control processor 5 on the component side is also amplified by the optical amplifier 52 and the branching device 53 and then branched into n pieces to be transmitted from the optical coupler 54 through the insides of the respective components Y1, Y2 ... Yn. To do.

Here, a case will be described in which, in a part of the matrix M as shown in FIG. 8A, an external force is applied and the region a surrounded by the dotted line is damaged by the depression. When an area a surrounded by a dotted line is damaged by an external force due to an external force, bending stress σ is generated in the optical fiber cables 3 of the X component and the Y component arranged in the area. Therefore, as described above, both components X
2, X3, X4 and Y3, Y4, Y5 transmitted light amount v of the measurement light L transmitted through the stress sensitive portion 34 of the optical fiber cable 3
Will be attenuated according to the bending angle.

On the X component side, both X1 and X5 are normal and X
2 and X3 and X4 are attenuated respectively. In the Y component, Y1, Y2, Y6 and Y7 are normal, and Y3 and Y
4 and Y5 decay. Then, the attenuated transmitted light amounts v of X2 to 4 and Y3 to 5 are received by the corresponding light receiving elements 46 and 56 of the deformation detecting section S1. Therefore, X2-4
It is judged that the inside of the sensitive matrix including the area a surrounded by Y3 and Y3-5 has been damaged by the depression. In this case, when the depression at the center of the region a is deepest, the attenuation amounts of the components X3 and Y4 are maximum. Then, from the magnitude of the relative amount of light received by the other four light-receiving elements 46 and 56 of the light-receiving elements 46 and 56 connected to the components X3 and Y4, the area a
The degree of damage and the condition of damage in the area are also estimated.

Next, in FIG. 8 (b), if the inside of the solid closed curve (region b) is damaged by peeling or missing, X1, X6
Then, the light L normally passes through the stress sensitive portion 34, and the component X2,
Broken fiber optic cable 3 for X3, X4, X5
Scatter and reflect on the cut surface of. The scattered light and the reflected light on the cut surface travel in the opposite direction through the half mirror of the optical coupler 44 and are received by the four light receivers 45 of the breakage detecting section S2.
Similarly, the same number of optical receivers 55 corresponding to the components Y3 to 6 are provided.
Scattered light and reflected light are incident on. Therefore, the component X2
It is judged that the region b surrounded by .about.5 and Y3.about.6 was destroyed and dropped.

The photodetectors 45 and 55 and the photodetectors 46 and 56 that have detected the transmitted light amount v in the abnormal state in this way output their detection signals to the control processors 4 and 5, respectively. The X-axis and Y-axis data processed by the control processors 4 and 5 are further transmitted to the flight controller 7 through the integrated processor 6.

In the first embodiment described above, the stress-sensitive portion composed of the concave portion having a small diameter is exemplified, but the shape is gradually reduced or expanded on both sides of the uniform small diameter portion having a constant width. You may comprise the stress response part etc. which were pinched | interposed by the taper. Further, the case where the matrix of the optical fiber sensor is provided on the upper surface of the wing has been illustrated and described, but a more effective detection device can be configured by arranging the matrix on the lower surface of the wing or on the surface layer of the entire body such as the body can do.

[0023]

According to the present invention, a damage / damage position detecting device is provided which is provided with an optical fiber cable along the surface of the body or along the surface thereof. In addition, a damage / damage position detection device was constructed by arranging optical fiber cables in a matrix. Furthermore, a damage / breakage position detection device was constructed in which the stress-sensitive portion was formed to reduce the thickness of the core of the optical fiber cable and increase the sensitivity to strain.

As a result, the change in the amount of transmitted light due to scattering or reflection of the transmitted light of the X-axis and Y-axis optical fiber sensors forming the matrix is measured, and the following aircraft failure is judged. . (1) When the fuselage is damaged by depression, the damage condition such as the damaged position, the range of damage, and the depth of the recess is clarified. (2) When a part of the fuselage is destroyed due to peeling, dropping, etc., the state of destruction such as the position and range of damage of the fuselage becomes clear.

Therefore, the inspection of metal fatigue and damage on the ground in the hangar or the like can be made accurate, accidents can be prevented in advance, the time required for the inspection can be remarkably shortened, and the operational efficiency of the aircraft can be improved. it can. Further, even during flight, the failure status of the airframe is clearly grasped, and the output signal of the integrated processor is transmitted to a flight controller or the like which controls the entire aircraft. For this reason, normal flaps, spoilers, auxiliary wings, etc., which are not damaged, are corrected, and emergency control such as adjustment of the rudder and elevator of the rear part of the fuselage and output adjustment of the left and right engines is performed. It is possible to continue a safe flight close to normal even if you receive a.

Therefore, according to the present invention, it is possible to provide a damaged / damaged position detecting device for an aircraft which can immediately detect a damaged portion of the machine body not only during maintenance on the ground but also during flight.

[Brief description of drawings]

FIG. 1 is a configuration explanatory diagram of an aircraft to which a first embodiment of the present invention is applied.

FIG. 2 is an enlarged explanatory diagram of a part of FIG.

FIG. 3 is a structural explanatory view of the optical fiber cable according to the first embodiment of the present invention.

FIG. 4 is an enlarged perspective view of a circle portion of FIG.

FIG. 5 is a characteristic diagram of stress and transmitted light amount.

FIG. 6 is a block diagram showing the configuration of the first embodiment of the present invention.

FIG. 7 is a block diagram showing a configuration of a Y component.

FIG. 8 is an operation explanatory diagram of the first embodiment of the present invention.

[Explanation of symbols]

 1 Aircraft 2 Damage / Break Position Detection Device 3 Optical Fiber Cable 4, 5 Control Processor 6 Integrated Processor 7 Flight Controller etc. 10 Body 11 Main Wing 31 Core 32 Cladding 33 Boundary Layer 34 Stress Sensing Part (Stress Sensor) 41, 51 Light Source 42,52 Optical amplifier 43,53 Splitter 44,54 Optical coupler 45,55 Light receiver 46,56 Light receiving element C Matrix intersection point L Light M Matrix S1 Deformation detection section S2 Damage detection section V Transmitted light quantity X1, X2… Xm X Component Y1, Y2 ... Yn Y component σ stress

 ─────────────────────────────────────────────────── ─── Continuation of front page (72) Inventor Eiji Shimizu 1-1-11-103 Higashimachi, Akishima-shi, Tokyo (72) Inventor ▲ Kunihiko Mano 1-7-12 Toranomon, Minato-ku, Tokyo Oki Electric Industry Co., Ltd. Incorporated (72) Inventor Fumio Nakamura 1-7-12 Toranomon, Minato-ku, Tokyo Oki Electric Industry Co., Ltd. (72) Inventor Tetsuro Kuno 10 Oemachi, Minato-ku, Nagoya Mitsubishi Heavy Industries Ltd. (72) Inventor, Hisako Yasui, 10 Oemachi, Minato-ku, Nagoya City Mitsubishi Heavy Industries Ltd., Nagoya Aerospace Systems Factory

Claims (3)

[Claims] 1. A damage / damage position detection device for an aircraft, characterized in that an optical fiber cable is provided on the surface layer of the airframe or in the surface layer. 2. A damage / damage position detecting device in an aircraft, wherein the optical fiber cables are arranged in a matrix. 3. The damage / damage position detecting device for an aircraft according to claim 1, wherein the stress-sensitive portion is formed by reducing the wall thickness of the core of the optical fiber cable.