JP2889952B2 - Damage / breakage position detection device - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a damaged / damaged position detecting device, and more particularly, to a damaged / damaged position detecting device for detecting a damaged position such as a distortion, a crack or a chip of an aircraft body, and a damaged state. Things.

[0002]

2. Description of the Related Art An aircraft that flies at high speed in the atmosphere always travels under strong wind pressure on its body. Therefore,
In particular, there is a high probability that distortion, cracks, and the like will occur on the surfaces of the main wing and tail wing that protrude laterally in proportion to the flight time. Heretofore, in order to grasp the state of damage due to such deformation or cracking of the body of an aircraft, it has been performed artificially mainly by visual inspection of maintenance personnel. Further, the inspection may be performed by a mechanical inspection method using an ultrasonic inspection device, a magnetic particle inspection device, or the like.

[0003]

SUMMARY OF THE INVENTION As described above, the discovery of a fault condition of an airframe occurring in a conventional aircraft of this type has been performed mainly visually and artificially. Therefore, the skill of the maintenance staff was required. In addition to the above, there are mechanical inspection methods as described above, but in all cases, the inspection methods are performed on the ground such as an airfield or a maintenance hangar. For this reason, the conventional method for detecting a damaged / damaged position has a problem that the flight of the aircraft must be suspended for a long time.

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

[0005]

SUMMARY OF THE INVENTION The present invention relates to an apparatus for detecting a damaged / damaged position in an aircraft in which an optical fiber cable is provided on a surface layer or in a surface layer of an airframe and the optical fiber cables are arranged in a matrix. And a damaged / damaged position detecting device for an aircraft provided with a stress-sensitive portion formed of a concave curved small-diameter portion and having a mode conversion function. Further, in the above-mentioned means, damage and damage to the aircraft arranged near the intersection of the optical fiber cables in which the stress sensitive portions are arranged in a matrix.
This constitutes a damage position detecting device.

[0006] When a dent is damaged in a part of the fuselage of an aircraft by an external force, a 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. For this reason, the transmitted light amount of the measuring light passing through the stress sensitive portions of both the X and Y components forming the matrix is attenuated according to the bending angle, and the amount of attenuation is detected by the light receiver. Therefore, the region corresponding to the attenuated plural components of X and Y is determined to have received the dent damage. At this time, the damage state in the damage area is also estimated based on the amount of attenuation of the multiple components of X and Y.

Further, when a part of the fuselage is damaged due to peeling or missing, the transmitted light of the X- and Y-component stress-sensitive parts leading to the damaged part is scattered or reflected. The portion surrounded by the X and Y components that receive the scattered light and the reflected light is determined to be broken and dropped.

[0008]

BEST MODE FOR CARRYING OUT THE INVENTION

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

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

In FIGS. 3 and 4, reference numeral 34 denotes a stress sensitive portion (stress sensor) constituting the sensor. The stress responsive portion 34 is formed of a concave-shaped small-diameter portion by partially applying tension and heat in the longitudinal direction to the optical fiber cable 3, and has a certain interval along the length direction of the optical fiber cable 3 as shown in the figure. Are provided at intervals. 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 the external force is generated, and the amount of light L transmitted through the core 31 in the direction of the arrow changes according to the bending angle.
FIG. 5 shows the relationship between the transmitted light amount v of the light L and the stress σ at this time. As shown in FIG. 5, the transmitted light amount v has such a characteristic that it decreases almost in proportion to the increase in the stress σ.

Generally, when an external force is applied to this type of optical fiber, stress is generated inside the optical fiber and "mode conversion" occurs.
The phenomenon occurs. When the "mode conversion" phenomenon occurs, the total reflection angle of the boundary layer changes. As a result, light passing through the inside of the optical fiber is irregularly reflected at the boundary layer or transmitted through the outer cladding and is scattered to the outside, resulting in a loss of light energy called radiation loss. Also, when pressure is applied from the axial direction of the optical fiber, a bend of about several μm is formed in the axial center, and unevenness occurs in the boundary layer. Then, the amount of transmitted light changes based on a similar “mode conversion” phenomenon, and an optical loss called “microbending loss” occurs.

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

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

FIGS. 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.
.., Xm are X-component optical fiber cables 3 for forming the matrix M. Y1, Y2,.
Is the Y component (where m and n are positive integers), and C is both components X1
, X2... Xm and Y1, Y2.
(See also FIG. 8). And the X,
A number of stress-sensitive parts 34 in both Y components
[Xm, yn], each intersection C [xm,
[yn] is detected.

In FIGS. 6 and 7, reference numeral 4 denotes a control processor for the X component, 5 denotes a control processor for the Y component, 6 denotes an integrated processor, and 7 denotes a flight controller and the like. 41 is X
A light source on the component side, 42 is an optical amplifier, and 43 is a branching device.
44 is an optical coupler, 45 is a light receiver, 46 is a light receiving element such as a photodiode. 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 optical fiber cables 3 on the X component side constituting the matrix M. Inside the control processor 4, a converter or the like for mutually converting an electric signal and an optical signal on the X component side is provided.

FIG. 7 shows a block diagram of the Y component side. Since the Y component side has a configuration similar to that of the X component side, reference numerals corresponding to the light source 51 and the like are assigned and detailed description is omitted.
However, the Y component of the optical fiber cable 3 is n and forms the vertical axis of the matrix M. The output terminal of the control processor 5 is connected to an output terminal of the control processor 4 and a flight controller 7 via the integrated processor 6 of FIG. S1 is a deformation detecting unit for irregularities or the like, and S2 is a damage detecting unit for peeling or missing the 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 a control processor 4 on the X component side.
The measurement light L output from 1 is amplified by the optical amplifier 42 and split into m pieces by the splitter 43. Light L divided into m pieces
Are transmitted through the polarization-independent optical coupler 44 and then projected into the respective optical fiber cables 3 of the X components X1, X2,. 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 an optical amplifier 52 and a branching device 53 and then branched into n pieces, and transmitted from an optical coupler 54 through the components Y1, Y2,. I do.

Here, a case where an external force is applied to a part of the matrix M as shown in FIG. 8A and a region a surrounded by a dotted line is damaged by dents will be described. When the region a surrounded by the dotted line is damaged by an external force, a bending stress σ is generated in the X- and Y-component optical fiber cables 3 arranged in the region. Therefore, as described above, both components X
2, X3, X4 and Y3, Y4, Y5, the transmitted light amount v of the measuring 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, X1 and X5 are both normal and X
2 and X3 and X4 respectively attenuate. In the Y component, Y1, Y2, Y6 and Y7 are normal and Y3 and Y7 are normal.
4 and Y5 decay. Then, the attenuated transmitted light amount v of X2-4 and Y3-5 is received by the corresponding light receiving elements 46 and 56 of the deformation detecting section S1. Therefore, X2-4
And the sensitive matrix including the area a surrounded by Y3 to Y5 is determined to have received dent damage. In this case, when the depression at the center of the region a is the deepest, the attenuation of the components X3 and Y4 becomes maximum. Then, based on the relative light receiving amounts of the other four light receiving elements 46 and 56 connected to the components X3 and Y4, the area a
The extent and status of damage within the building is also estimated.

Next, in FIG. 8 (b), if the inside of the solid closed curve (region b) is broken due to peeling or missing, X1, X6
In this case, the light L normally passes through the stress sensitive part 34, and the components X2,
X3, X4, X5 damaged optical fiber cable 3
Scattered and reflected on the cut surface. The scattered light and the reflected light of the cut surface travel in the reverse direction through the half mirror of the optical coupler 44 and the like, and are received by the four light receivers 45 of the damage detection unit S2.
Similarly, the same number of receivers 55 corresponding to the components Y3 to Y6 are used.
The scattered light and the reflected light are incident on the. Therefore, the component X2
It is determined that the area b surrounded by .about.5 and Y3 .about.6 has been destroyed and dropped.

The light receivers 45 and 55 and the light receiving elements 46 and 56 which have detected the abnormally transmitted light amount v output respective detection signals to the control processors 4 and 5. The X-axis and Y-axis data processed by the control processors 4 and 5 are further transmitted to a flight controller 7 via an integrated processor 6.

In the above-described first embodiment, the stress sensitive portion made of a concavely curved small diameter portion has been described as an example. However, a tapered shape and a gradually expanded shape are formed on both sides of a uniform small diameter portion having a constant width. You may comprise the stress response part etc. which were pinched by the taper. Although the case where the matrix of the optical fiber sensor is provided on the upper surface of the wing is illustrated and described, a more effective detection device can be configured by arranging the matrix on the surface of the entire body such as the lower surface of the wing or the fuselage. can do.

[0023]

According to the present invention, there is provided a device for detecting a damaged / damaged position in an aircraft in which an optical fiber cable is provided on a surface layer or in a surface layer of an airframe, and the optical fiber cable is arranged in a matrix. A damaged / damaged position detecting device for an aircraft provided with a stress-sensitive portion formed of a small-diameter portion having a mode conversion function. Further, in the above configuration, a damage / breakage position detecting device for an aircraft is arranged near the intersection of the optical fiber cables in which the stress sensing portions are arranged in a matrix.

As a result, a 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 determined. . (1) When the airframe is damaged by depression, the damage position, the range of the damage, the depth of the dent, and the like are clarified. (2) When a part of the fuselage is broken due to peeling, falling off, or the like, a broken state such as a position or a range of damage of the fuselage becomes clear.

Therefore, it is possible to accurately inspect metal fatigue and damage on the ground in the hangar or the like, thereby preventing an accident from occurring, and significantly reducing the time required for the inspection, thereby improving the operational efficiency of the aircraft. it can. Further, even during the flight, the failure state of the aircraft is clearly grasped, and the output signal of the integrated processor is transmitted to a flight controller or the like that controls the entire aircraft. For this reason, emergency control such as correction of rudder and elevator of the normal wings and other flaps, spoilers, auxiliary wings or the fuselage rear that is not damaged, and adjustment of the output of the right and left engines are performed. It is possible to continue a safe flight close to normal even if it is received.

Thus, 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 airframe during flight as well as during maintenance on the ground.

[Brief description of the drawings]

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

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

FIG. 3 is an explanatory diagram of a configuration of an optical fiber cable according to the first embodiment of the present invention.

FIG. 4 is an enlarged perspective view of a circle of FIG. 3;

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

FIG. 6 is a block diagram showing a configuration of Embodiment 1 of the present invention.

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

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

[Explanation of symbols]

 DESCRIPTION OF SYMBOLS 1 Aircraft 2 Damage / breakage position detecting device 3 Optical fiber cable 4,5 Control processor 6 Integrated processor 7 Flight controller etc. 10 Fuselage 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 Branching device 44,54 Optical coupler 45,55 Light receiving device 46,56 Intersection of light receiving element C matrix L Light M matrix S1 Deformation detecting unit S2 Damage detecting unit V Transmitted light amount X1, X2 ... Xm X Component Y1, Y2 ... Yn Y component σ stress

──────────────────────────────────────────────────続 き Continuing on the front page (72) Inventor ▲ Kunihiko Nono 1-7-12 Toranomon, Minato-ku, Tokyo Oki Electric Industry Co., Ltd. (72) Fumio Nakamura 1-7-7 Toranomon, Minato-ku, Tokyo No. 12 Oki Electric Industry Co., Ltd. Nagoya Aerospace Systems Works, Mitsubishi Heavy Industries, Ltd. (56) References JP-A-63-208748 (JP, A) JP-A-62-231142 (JP, A) JP-A-63-285448 (JP, A) JP-A-7-174527 (JP, A) JP-A-8-54343 (JP, A) JP-A-60-165504 (JP, A) JP-A-60-174960 (JP, A) JP-A-63-109306 (J , A) JP Akira 63-210604 (JP, A) PCT National Akira 58-500421 (JP, A) (58 ) investigated the field (Int.Cl. ^6, DB name) G01B 11/00 - 11/30 102 G01M 13/00-13/04 G01M 19/00-19/02 G01N 21/84-21/91

Claims (2)

(57) [Claims] 1. A damage / breakage position detecting apparatus for an aircraft, wherein an optical fiber cable is provided on a surface layer of an airframe or in a surface layer thereof, and the optical fiber cable is arranged in a matrix. A damaged / damaged position detecting device for an aircraft, comprising a stress sensitive part formed of a diameter part and having a mode conversion function. 2. The damage / breakage detecting device for an aircraft according to claim 1, wherein said stress sensitive portions are arranged near intersections of optical fiber cables arranged in a matrix.