Ice detection device

Abstract

This invention provides an icing detection device that is less susceptible to electromagnetic interference. The icing detection device comprises: an aircraft structure; an exposed component connected to the aircraft structure and exposed on the exterior of the aircraft structure; a fiber optic sensor connected to the exposed component and covered by at least one of the aircraft structure and the exposed component; and a measuring device for measuring light obtained from the fiber optic sensor.

Description

Technical Field

[0001] This invention relates to an ice accumulation detection device. Background Technology

[0002] Patent Document 1 discloses an icing sensor for detecting ice buildup on the surface of an aircraft. The icing sensor in Patent Document 1 detects ice buildup by detecting the resonant frequency of a finger generated by electrical excitation.

[0003] Existing technical documents

[0004] Patent documents

[0005] Patent Document 1: Japanese Patent Application Publication No. 2005-145453 Summary of the Invention

[0006] Technical issues

[0007] Electrical ice accumulation sensors, as described in Patent Document 1, are susceptible to electromagnetic interference.

[0008] Therefore, the purpose of this invention is to provide an ice accumulation detection device that is less susceptible to electromagnetic interference.

[0009] Technical solution

[0010] To address the aforementioned issues, an ice accumulation detection device according to one embodiment of the present invention includes:

[0011] Aircraft structure;

[0012] Exposed components are connected to the aircraft structure and are exposed on the exterior of the aircraft structure;

[0013] A fiber optic sensor, connected to the exposed component and covered by at least one of the aircraft structure and the exposed component; and

[0014] A measuring device that measures the light obtained from the fiber optic sensor.

[0015] Invention Effects

[0016] According to the present invention, an ice accumulation detection device that is less susceptible to electromagnetic interference can be provided. Attached Figure Description

[0017] Figure 1 This is a simplified perspective view of an aircraft according to one embodiment of the present invention.

[0018] Figure 2 This is a simplified structural diagram of the ice accumulation detection device according to this embodiment.

[0019] Figure 3This diagram shows the state in which the exposed part of this embodiment is covered with ice.

[0020] Figure 4 This is a diagram showing the first modified example of the hammer.

[0021] Figure 5 This is a diagram showing a second variation of the hammer.

[0022] Figure 6 This is a diagram showing the third variation of the hammer.

[0023] Figure 7 This is a diagram showing the fourth variation of the hammer.

[0024] Figure 8 This is a simplified structural diagram of an ice accumulation detection device according to another embodiment.

[0025] Figure 9 This is a diagram showing the state of an exposed part with ice attached in another embodiment.

[0026] Symbol Explanation

[0027] AS aircraft structure

[0028] SU outer surface

[0029] 1. Aircraft

[0030] 100 Ice Detection Device

[0031] 110 Exposed components

[0032] 111 pillars

[0033] 113 hammers

[0034] 120 Fiber Optic Sensor

[0035] 130 Measuring device

[0036] 200 Ice Detection Device

[0037] 211 Flexible components

[0038] 220 Fiber Optic Sensor Detailed Implementation

[0039] Hereinafter, embodiments of the present invention will be described in detail with reference to the accompanying drawings. The specific dimensions, materials, values, etc., shown in these embodiments are merely illustrative for ease of understanding of the invention and, unless otherwise specified, are not intended to limit the scope of the invention. It should be noted that in this specification and the accompanying drawings, elements having substantially the same function or structure are labeled with the same symbols, and repeated descriptions are omitted. Furthermore, elements not directly related to the present invention are omitted from the illustrations.

[0040] [1. Overall Structure of the Aircraft]

[0041] First, refer to Figure 1 The overall structure of an aircraft 1 according to one embodiment of the present invention will be described. Figure 1 This is a simplified perspective view of an aircraft 1 according to one embodiment of the present invention.

[0042] like Figure 1 As shown, the aircraft 1 has a fuselage 3, main wings 5, horizontal tail 7, and vertical tail 9. Hereinafter, the main wings 5, horizontal tail 7, and vertical tail 9 will be referred to as wings.

[0043] The fuselage 3 is the central structural component of the aircraft 1, and its length in the longitudinal direction (roll axis direction) is longer than its length in the lateral direction (pitch axis direction) and vertical direction (yaw axis direction). The interior of the fuselage 3 forms a passenger space and houses various devices such as engines, fuel tanks, flight control systems, and measuring instruments.

[0044] A pair of main wings 5, 5 are arranged on the left and right sides of the central part of the fuselage 3. The pair of main wings 5, 5 are configured to extend from the central part of the fuselage 3 to the left and right. The main wings 5 ​​generate upward lift for the aircraft 1.

[0045] A pair of horizontal stabilizers 7, 7 are provided on the left and right sides of the rear of the fuselage 3. The pair of horizontal stabilizers 7, 7 are configured to extend from the rear of the fuselage 3 to the left and right. The horizontal stabilizers 7 have the function of maintaining the stability of the aircraft 1 about the pitch axis.

[0046] A vertical tail 9 is provided on the upper rear side of the fuselage 3. The vertical tail 9 is configured to extend upward from the rear of the fuselage 3. The vertical tail 9 has the function of maintaining the stability of the aircraft 1 about the yaw axis.

[0047] When such an aircraft 1 is in flight or in cold regions, icing may sometimes occur. For example, ice may sometimes adhere to the surface of the nose section of the fuselage 3 (hereinafter referred to as the aircraft structure AS). If ice adheres to the surface of the aircraft 1, there is a problem of increased weight and reduced lift of the aircraft 1.

[0048] Therefore, the aircraft 1 in this embodiment is equipped with an icing detection device 100 for detecting the adhesion of ice on the surface of the aircraft structure AS. It should be noted that in this embodiment, the aircraft structure AS is described as the nose section of the fuselage 3, but it is not limited to this. For example, the aircraft structure AS could also be a main wing 5, a horizontal tail 7, a vertical tail 9, etc.

[0049] [2. Structure of the ice accumulation detection device]

[0050] Figure 2 This is a simplified structural diagram of the ice accumulation detection device 100 according to this embodiment. Figure 2 As shown, the icing detection device 100 includes an aircraft structure AS, exposed components 110, fiber optic sensor 120, and measuring device 130.

[0051] The exposed component 110 is connected to the outer surface SU of the aircraft structure AS. Therefore, the exposed component 110 is exposed on the outside of the aircraft structure AS. The exposed component 110 includes a support column 111 and a hammer 113.

[0052] A support column 111 is erected on the outer surface SU of the aircraft structure AS, covering at least a portion of the fiber optic sensor 120. One end of the support column 111 is connected to the outer surface SU, and the other end is connected to a hammer 113. The hammer 113 is a weight adjustment component for adjusting the weight of the exposed component 110.

[0053] Fiber optic sensor 120 is embedded in support column 111, located at the center of support column 111. Fiber optic sensor 120 extends along the length of support column 111. Fiber optic sensor 120 is covered by exposed component 110 and is integrally formed with exposed component 110. Thus, fiber optic sensor 120 is connected to exposed component 110. Fiber optic sensor 120 penetrates the interior of aircraft structure AS and is connected to measuring device 130.

[0054] The measuring device 130 is disposed inside the aircraft structure AS. The measuring device 130 illuminates light into the fiber optic sensor 120 and detects the light received from the fiber optic sensor 120. In this embodiment, the measuring device 130 measures the backscattered light of the light traveling within the fiber optic sensor 120.

[0055] If strain or deformation occurs in the fiber optic sensor 120, the frequency of the backscattered light generated at the strained portion changes. The measuring device 130 measures the change in the frequency of the backscattered light traveling within the fiber optic sensor 120. However, it is not limited to this; the measuring device 130 can also measure the change in the frequency of the reflected light traveling within the fiber optic sensor 120.

[0056] Furthermore, if strain or deformation occurs in the fiber optic sensor 120, the loss of light traveling within the fiber optic sensor 120 increases in the strained portion, and the backscattered light generated in the strained portion also decreases. Therefore, the measuring device 130 can also measure the change in the amount of backscattered light traveling within the fiber optic sensor 120.

[0057] The natural frequency f of the exposed component 110 integrated with the fiber optic sensor 120 is given by f = 1 / (2π) × (k / m). 1 / 2 In this context, k is the spring constant of the exposed component 110, and m is the weight of the exposed component 110. That is, the natural frequency f of the exposed component 110 integrated with the fiber optic sensor 120 can be tuned by adjusting the spring constant k and the weight m of the exposed component 110.

[0058] Figure 3 This diagram illustrates the state of the exposed part 110 in this embodiment, where ice has adhered to it. Figure 3 As shown, if ice adheres to the exposed part 110, the weight and stiffness of the exposed part 110 change. If the weight and stiffness of the exposed part 110 change, the vibration frequency of the exposed part 110 changes.

[0059] If the vibration frequency of the exposed component 110 changes, the strain and / or deformation of the fiber optic sensor 120 changes, and the frequency change of the backscattered light is measured by the measuring device 130. The measuring device 130 can detect the presence of ice on the exposed component 110 by measuring the change in the frequency of the backscattered light.

[0060] Thus, the ice detection device 100 according to this embodiment, by having an optical fiber sensor 120 covered by the exposed part 110, can detect that ice is attached to the exposed part 110 and then to the aircraft 1.

[0061] The fiber optic sensor 120 is less susceptible to electromagnetic interference. Therefore, the ice detection device 100 according to this embodiment can also be applied to locations affected by electromagnetic interference.

[0062] Furthermore, electrical ice-accumulation sensors like those in Patent Document 1 use magnetic materials and / or crystal oscillators. Therefore, in electrical ice-accumulation sensors like those in Patent Document 1, it is difficult to freely tune the resonant frequency according to the application environment and the application structure. On the other hand, according to the exposed member 110 of this embodiment, for example, by changing the weight of the hammer 113, the natural frequency (resonant frequency) of the exposed member 110 integrated with the fiber optic sensor 120 can be freely tuned.

[0063] In this embodiment, the fiber optic sensor 120 is covered by the exposed component 110. Therefore, compared to the case where the fiber optic sensor 120 is exposed externally to the aircraft structure AS, the reduction in the durability of the fiber optic sensor 120 can be suppressed.

[0064] Figure 4 This is a diagram showing a first modified example of hammer 113. (See diagram below.) Figure 4As shown, the hammer 113A of the first modified example has a helical shape. The hammer 113A of the first modified example has a plurality of portions 150 separated, for example, in the direction R of the central axis of the support 111. A gap Sa is formed between the plurality of portions 150. Thus, the hammer 113A of the first modified example has a gap Sa formed between the plurality of portions.

[0065] Figure 5 This is a diagram showing a second modified example of hammer 113. (See diagram below.) Figure 5 As shown, the hammer 113B of the second modified example has a conical helical shape. The hammer 113B of the second modified example has a plurality of portions 160, for example, separated in the direction R of the central axis of the support 111. A gap Sb is formed between the plurality of portions 160. Thus, the hammer 113B of the second modified example has a gap Sb formed between the plurality of portions.

[0066] Figure 6 This is a diagram showing a third variation of hammer 113. (See diagram below.) Figure 6 As shown, the hammer 113C of the third modification has a spider web shape. The hammer 113C of the third modification has a plurality of radially extending portions 170 and a plurality of arcuate portions 171 connected in an arc shape between the plurality of extending portions 170. A gap Sc is formed between the plurality of extending portions 170 and the plurality of arcuate portions 171. Thus, the hammer 113C of the third modification has gaps Sc formed between multiple portions.

[0067] Figure 7 This is a diagram showing the fourth variation of hammer 113. (See diagram for example.) Figure 7 As shown, the hammer 113D of the fourth modification has a rhomboid shape. The hammer 113D of the fourth modification has, for example, an extension 180 extending along the central axis direction R of the support column 111, an orthogonal portion 181 extending along a direction orthogonal to the extension 180, and a plurality of inclined portions 183 inclined in both the central axis direction R and the orthogonal direction V. The plurality of inclined portions 183 are separated from each other and connected to the extension 180 and the orthogonal portion 181. A gap Sd is formed between the extension 180, the orthogonal portion 181, and the plurality of inclined portions 183. Thus, the hammer 113D of the fourth modification has a gap Sd formed between multiple portions.

[0068] Hammers 113A, 113B, 113C, and 113D have gaps Sa, Sb, Sc, and Sd formed between multiple parts, making it easier for ice to adhere compared to hammer 113 in the above embodiment. Because of the ice adhesion, the weight of the exposed part 110 can easily increase, resulting in the ice accumulation being easily detected by the measuring device 130.

[0069] [Another implementation method]

[0070] Figure 8This is a simplified structural diagram of an ice accumulation detection device 200 according to another embodiment. Components substantially the same as those in the ice accumulation detection device 100 of the above embodiment are labeled with the same reference numerals and their descriptions are omitted. The ice accumulation detection device 200 includes an exposed component 210 and a fiber optic sensor 220. The exposed component 210 includes a flexible component 211. The fiber optic sensor 220 is connected to the exposed component 210 (flexible component 211).

[0071] The flexible component 211 is, for example, an optical fiber cable. One end of the flexible component 211 is connected to the optical fiber sensor 220, and the other end is connected to the hammer 113. The flexible component 211 is wound around the support column 111 with at least a portion detached from it. Therefore, the flexible component 211 is configured in a flexed state on the outside of the aircraft structure AS. In the flexed state, almost no load is transmitted from the flexible component 211 to the optical fiber sensor 220.

[0072] The fiber optic sensor 220 is embedded in the aircraft structure AS and located at the center of the aircraft structure AS. The fiber optic sensor 220 is covered by the aircraft structure AS and is integrally formed with the aircraft structure AS. The fiber optic sensor 220 is connected to the measuring device 130.

[0073] In the above embodiment, the fiber optic sensor 120 is covered by the exposed component 110. On the other hand, the fiber optic sensor 220 is covered by the aircraft structure AS. In this respect, the two differ. It should be noted that the fiber optic sensors 120 and 220 may also be covered by both the exposed component 110 and the aircraft structure AS. Thus, the fiber optic sensors 120 and 220 are covered by at least one of the aircraft structure AS and the exposed component 110.

[0074] Figure 9 This diagram illustrates the state in which ice adheres to the exposed part 210 in another embodiment. (See diagram for example.) Figure 9 As shown, if ice adheres to the exposed part 210, the flexible part 211 becomes integrated with the support 111 through the ice.

[0075] Therefore, during the flight of aircraft 1, if the support column 111 and hammer 113 vibrate, the flexible component 211 will also vibrate, and the load will be transmitted to the fiber optic sensor 220. If the load is transmitted to the fiber optic sensor 220, the fiber optic sensor 220 will experience strain or deformation, and the frequency of the backscattered light measured by the measuring device 130 will change. The measuring device 130 can detect the presence of ice on the exposed component 210 by measuring the change in the frequency of the backscattered light.

[0076] As described above, the fiber optic sensor 220 is covered by the aircraft structure AS. Therefore, compared to the case where the fiber optic sensor 220 is exposed on the outside of the aircraft structure AS, the reduction in the durability of the fiber optic sensor 220 can be suppressed.

[0077] Furthermore, because the flexible component 211 is provided exposed on the exterior of the aircraft structure AS, vibrations and loads caused by icing on the exposed component 210 can be effectively transmitted to the fiber optic sensor 220. Additionally, the effects described in the above embodiment can be obtained.

[0078] The embodiments of the present invention have been described above with reference to the accompanying drawings, but the present invention is by no means limited to these embodiments. Those skilled in the art will obviously be able to conceive of various modifications or alterations within the scope of the claims, and these are also within the technical scope of the present invention.

[0079] In the above embodiment, an example of providing a hammer 113 in the exposed parts 110 and 210 was described. However, the hammer 113 is not a necessary structure, and it is also possible not to provide a hammer 113 in the exposed parts 110 and 210.

Claims

1. An ice accumulation detection device, characterized in that, have: Aircraft structure; Exposed components are connected to the aircraft structure and are exposed on the exterior of the aircraft structure; An optical fiber sensor is connected to the exposed component and is covered by at least one of the aircraft structure and the exposed component; as well as A measuring device that measures the light received from the fiber optic sensor. The exposed component includes: A support column is erected on the outer surface of the aircraft structure and covers the fiber optic sensor; as well as A flexible component, which is an optical fiber cable, is connected to the optical fiber sensor and is wound around the support column in a state where at least a portion is detached from the support column.

2. The ice accumulation detection device according to claim 1, characterized in that, The exposed component has a hammer for adjusting the weight of the exposed component.

3. The ice accumulation detection device according to claim 2, characterized in that, The hammer has gaps formed between multiple parts.

4. The ice accumulation detection device according to claim 3, characterized in that, The hammer can be cylindrical spiral, conical spiral, spider web, or rhomboid in shape.

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