JP2013163480A - Anti-icing device and aircraft main wing - Google Patents

Anti-icing device and aircraft main wing Download PDF

Info

Publication number
JP2013163480A
JP2013163480A JP2012028270A JP2012028270A JP2013163480A JP 2013163480 A JP2013163480 A JP 2013163480A JP 2012028270 A JP2012028270 A JP 2012028270A JP 2012028270 A JP2012028270 A JP 2012028270A JP 2013163480 A JP2013163480 A JP 2013163480A
Authority
JP
Japan
Prior art keywords
front edge
extraction air
main wing
aircraft
leading edge
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
JP2012028270A
Other languages
Japanese (ja)
Inventor
Kurodo Ichikawa
玄人 市川
Yoichi Kamifuji
陽一 上藤
Nobuyasu Sakata
展康 坂田
Keiichi Sato
恵一 佐藤
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Mitsubishi Heavy Industries Ltd
Original Assignee
Mitsubishi Heavy Industries Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Mitsubishi Heavy Industries Ltd filed Critical Mitsubishi Heavy Industries Ltd
Priority to JP2012028270A priority Critical patent/JP2013163480A/en
Publication of JP2013163480A publication Critical patent/JP2013163480A/en
Pending legal-status Critical Current

Links

Images

Abstract

PROBLEM TO BE SOLVED: To provide an anti-icing device for a front edge part in an aircraft, configured to reliably and efficiently prevent icing at the front edge of the aircraft with a simple structure.SOLUTION: An anti-icing device 40 includes an extraction air nozzle 41 where extraction air to be supplied to a main-wing front edge part 22 flows, and a rear wall 32 that is provided on a flow passage of the extraction air and with which the extraction air collides. The extraction air nozzle 41 is provided on the side closer to the main-wing front edge 22 than the rear wall 32. The extraction air nozzle 41 has a blow hole 41a formed opposite the rear wall 32 to blow the extraction air to the rear opposite a tip part 22a of the main-wing front edge part 22. The extraction air blown from the blow hole 41a collides with the rear wall 32 and flows from the rear of the main-wing front edge part 22 toward the tip part 22a along an inner wall surface 31, thereby preventing local overheating of the inner wall surface 31 of the tip part 22a, and heating the main-wing front edge part 22 entirely and more uniformly.

Description

本発明は、航空機の前縁の着氷を防止する防氷装置および航空機主翼に関する。   The present invention relates to an anti-icing device and an aircraft main wing that prevent icing of the leading edge of an aircraft.

航空機が雲の中を飛行する際に、翼の前縁の外側表面に、空気中に存在する過冷却の水滴の衝突による着氷が生じやすい。この着氷現象は、航空機に様々な悪影響を及ぼす。
そのため、航空機の翼の前縁には、加熱手段により着氷を防止するための防氷装置が装備されている。一般的に、タービン・エンジンを装備する航空機では、エンジンの圧縮機から高温かつ高圧のエアを抽出し、この抽出されたエア（抽出エア）を加熱手段として用いている。このような防氷装置では、抽出エアを翼の内部まで導き、抽出エアが流れる抽出エア供給管に設けられた多数のノズルまたはスリットから、翼の前縁の先端部に向けて抽出エアを吹き付けることにより、翼の前縁を加熱している。この場合、抽出エアが吹きつけられた先端部付近では高い熱伝達率が得られるため、少量の抽出エア量で加熱することができる。その一方で、先端部付近から離れた領域では、先端部付近ほどの高い熱伝達率を得ることができない。したがって、翼の前縁を全体にわたって均一に加熱することができず、翼の前縁への着氷を適正に防止できないことが問題となっていた。 When an aircraft flies in the clouds, icing is likely to occur on the outer surface of the leading edge of the wing due to the collision of supercooled water droplets present in the air. This icing phenomenon has various adverse effects on the aircraft.
For this reason, an anti-icing device for preventing icing by the heating means is provided at the leading edge of the wing of the aircraft. In general, an aircraft equipped with a turbine engine extracts high-temperature and high-pressure air from a compressor of the engine and uses the extracted air (extracted air) as a heating means. In such an anti-icing device, the extraction air is guided to the inside of the blade, and the extraction air is blown toward the tip of the leading edge of the blade from a number of nozzles or slits provided in the extraction air supply pipe through which the extraction air flows. As a result, the leading edge of the wing is heated. In this case, since a high heat transfer coefficient is obtained in the vicinity of the tip portion where the extraction air is blown, it can be heated with a small amount of extraction air. On the other hand, in a region far from the vicinity of the tip portion, a heat transfer coefficient as high as that near the tip portion cannot be obtained. Therefore, the leading edge of the blade cannot be heated uniformly over the entire surface, and it has been a problem that icing on the leading edge of the blade cannot be prevented properly.

この問題を解決するために、特許文献１では、図４に示すように、航空機の主翼前縁部１１の内側における抽出エア１５の流路に沿ってガイド板１２を取り付け、主翼前縁部１１の内面を二重構造とし、主翼前縁部１１の後方に向けて抽出エア１５を流す暖気通路１３を設けた防氷装置１４が提案されている。この構造により、抽出エア１５が暖気通路１３を通って下流側まで確実に流れるため、主翼前縁部１１の内面における熱伝達率が平均化され、主翼前縁部１１を全体にわたって適正に加熱することができる。
しかしながら、特許文献１に記載の防氷装置１４は、構造が複雑で重量が重いため、航空機の機体重量が増大する。
また、この防氷装置１４では、主翼前縁部１１のみならず、ガイド板１２も抽出エア１５により加熱されるため、熱ロスが大きい。そのため、主翼前縁部１１全体を適切に加熱するのに必要な抽出エア１５の量が増大し、結果として航空機の燃費の悪化につながる。 In order to solve this problem, in Patent Document 1, as shown in FIG. 4, a guide plate 12 is attached along the flow path of the extracted air 15 inside the main wing leading edge 11 of the aircraft, and the main wing leading edge 11. An anti-icing device 14 has been proposed that has a double structure on the inner surface thereof and provided with a warm air passage 13 through which the extraction air 15 flows toward the rear of the leading edge 11 of the main wing. With this structure, the extraction air 15 surely flows to the downstream side through the warm air passage 13, so that the heat transfer coefficient on the inner surface of the main wing leading edge 11 is averaged, and the main wing leading edge 11 is appropriately heated throughout. be able to.
However, since the anti-icing device 14 described in Patent Document 1 has a complicated structure and a heavy weight, the weight of the aircraft increases.
In this anti-icing device 14, not only the main wing leading edge 11 but also the guide plate 12 is heated by the extraction air 15, so that heat loss is large. As a result, the amount of extracted air 15 required to appropriately heat the entire main wing leading edge 11 increases, resulting in deterioration in fuel consumption of the aircraft.

さらに、特許文献１の防氷装置１４では、抽出エア１５が吹きつけられる主翼前縁部１１の先端部１１ａの熱伝達率が局所的に高くなり、先端部１１ａの内面が焼損する恐れがある。
これに対し、特許文献２では、主翼前縁部の外側に粗い表面を有するバンドを取り付けて、主翼前縁部の外側を流れる層流を乱流に遷移させることにより主翼前縁部を冷却させ、局所的に高くなった熱伝達率を低下させる防氷装置を提案している。この防氷装置によれば、抽出エアが吹き付けられる主翼前縁部の焼損を防ぐことができる。しかしながら、翼の外側表面に部材を設けると、飛行中の空気抵抗が増し、航空機の燃費の悪化につながる。 Furthermore, in the anti-icing device 14 of Patent Document 1, the heat transfer coefficient of the tip portion 11a of the main wing leading edge portion 11 to which the extraction air 15 is blown is locally increased, and the inner surface of the tip portion 11a may be burned out. .
On the other hand, in Patent Document 2, a band having a rough surface is attached outside the leading edge of the main wing, and the laminar flow flowing outside the leading edge of the main wing is changed to turbulent flow to cool the leading edge of the main wing. They have proposed an anti-icing device that lowers the heat transfer rate that is locally high. According to this anti-icing device, it is possible to prevent burning of the leading edge of the main wing to which the extraction air is blown. However, when a member is provided on the outer surface of the wing, air resistance during flight increases, leading to deterioration in aircraft fuel consumption.

特開２０１１−１８３９２２号公報JP 2011-183922 A 特表２００９−５２３６７号公報Special table 2009-52367

本発明は、このような技術的課題に基づいてなされたもので、簡易な構造で、航空機の前縁における着氷を確実かつ効率的に防ぐことができる、航空機における前縁部の防氷装置を提供することを目的とする。   The present invention has been made based on such a technical problem, and has a simple structure and can reliably and efficiently prevent icing at the leading edge of the aircraft. The purpose is to provide.

かかる目的のもと、本発明の航空機における前縁部の防氷装置は、湾曲形状をなす前縁部の内周面に抽出エアを供給することで前縁部の外周面への着氷を防止する装置であって、前縁部とは反対方向である後方に向けて抽出エアを吹き付ける抽出エア供給部と、抽出エア供給部から吹き付けられた抽出エアが衝突する衝突部と、を有し、衝突部に衝突した抽出エアが前縁部の内周面に向けて供給されることを特徴としている。
このように、抽出エアを前縁部の反対方向である後方に向けて吹き付ける抽出エア供給部を設け、一旦衝突部に衝突させてから前縁部の内周面に供給するという簡易な構成により、前縁部の内周面に抽出エアが直接的に衝突することによりその衝突部分が局所的に過熱されることを防ぐことができる。また、これにより、前縁部と、そこから離れた後方との温度差が小さくなるため、前縁部の内周面を全体にわたって均一に加熱することができ、前縁部の外周面への着氷を確実かつ効率的に防止することができる。 For this purpose, the anti-icing device for the front edge portion of the aircraft of the present invention prevents the icing on the outer peripheral surface of the front edge portion by supplying extracted air to the inner peripheral surface of the front edge portion having a curved shape. An apparatus for preventing, having an extraction air supply part that blows extraction air toward the rear that is opposite to the front edge part, and a collision part that collides with extraction air blown from the extraction air supply part The extracted air that has collided with the colliding part is supplied toward the inner peripheral surface of the front edge part.
In this way, the extraction air supply part that blows the extraction air toward the rear, which is the opposite direction of the front edge part, is provided, and after having collided with the collision part once, it is supplied to the inner peripheral surface of the front edge part. In addition, when the extracted air directly collides with the inner peripheral surface of the front edge portion, it is possible to prevent the collision portion from being overheated locally. Also, this reduces the temperature difference between the front edge and the rear away from it, so that the inner peripheral surface of the front edge can be uniformly heated over the entire surface, Icing can be reliably and efficiently prevented.

本発明では、航空機の構造部材が衝突部を兼ねる構成とすることができる。これにより、衝突部用の部材を別途設ける必要がなく、防氷装置をさらに簡易な構造とし、軽量化を実現することができる。   In the present invention, the structure member of the aircraft can also serve as the collision portion. Thereby, it is not necessary to separately provide a member for the collision part, and the anti-icing device can be further simplified in structure and light weight can be realized.

また、衝突部に遮熱コーティングを施すことができる。これにより、衝突部の部材表層の熱伝導率を下げ、衝突部における熱ロスを防ぐことができ、前縁部の内周面をより効率的に加熱することができる。また、衝突部の耐熱性を向上することができ、衝突部およびその周囲の構造部材の過熱による焼損をより確実に防ぐことができる。   Moreover, a thermal barrier coating can be applied to the collision part. Thereby, the heat conductivity of the member surface layer of a collision part can be reduced, the heat loss in a collision part can be prevented, and the internal peripheral surface of a front edge part can be heated more efficiently. In addition, the heat resistance of the collision part can be improved, and burning caused by overheating of the collision part and the surrounding structural member can be more reliably prevented.

また、本発明では、前縁部を翼の前縁部とし、衝突部に衝突した抽出エアが翼を構成するパネルに沿って前縁部の内周面に供給されるようにすることができる。航空機の翼の前縁部に本発明の防氷装置を設けることで、翼の前縁部の外側に生じる着氷を確実かつ効果的に防止することができる。   Further, in the present invention, the leading edge can be used as the leading edge of the blade, and the extracted air that has collided with the collision portion can be supplied to the inner peripheral surface of the leading edge along the panel constituting the blade. . By providing the anti-icing device of the present invention at the front edge of the wing of an aircraft, icing that occurs outside the front edge of the wing can be reliably and effectively prevented.

本発明は、また、上述の防氷装置を備えた航空機の主翼とすることもできる。   The present invention can also be a main wing of an aircraft provided with the above-described anti-icing device.

本発明によれば、簡易な構造で、航空機の前縁における着氷を確実かつ効率的に防ぐことができる、航空機における前縁部の防氷装置を提供することができる。   ADVANTAGE OF THE INVENTION According to this invention, the anti-icing apparatus of the front edge part in an aircraft which can prevent the icing in the front edge of an aircraft reliably and efficiently with a simple structure can be provided.

本実施形態における防氷装置が適用された主翼の要部を表す図であり、（ａ）はその断面図、（ｂ）は主翼の前縁部の内側を示す斜視図である。It is a figure showing the principal part of the main wing to which the anti-icing apparatus in this embodiment was applied, (a) is the sectional drawing, (b) is a perspective view which shows the inner side of the front edge part of a main wing. 本実施形態の防氷装置の後壁に遮熱コーティングを施した場合の主翼の要部の断面図である。It is sectional drawing of the principal part of the main wing at the time of providing thermal barrier coating to the rear wall of the anti-icing apparatus of this embodiment. 本実施形態の防氷装置の内壁面に表面処理を施した場合の主翼の要部を表す図であり、（ａ）はその断面図、（ｂ）は主翼の前縁部の内側を示す斜視図である。It is a figure showing the principal part of the main wing at the time of surface-treating to the inner wall surface of the anti-icing device of this embodiment, (a) is the sectional view, (b) is the perspective view which shows the inner side of the front edge part of the main wing FIG. 従来の航空機の主翼の前縁部の防氷装置が適用された主翼の要部を表す断面図である。It is sectional drawing showing the principal part of the main wing to which the anti-icing device of the front edge part of the conventional main wing of an aircraft was applied.

以下、添付図面に示す実施形態に基づいてこの発明を詳細に説明する。なお、これらの実施形態により本発明が限定されるものではない。   Hereinafter, the present invention will be described in detail based on embodiments shown in the accompanying drawings. Note that the present invention is not limited to these embodiments.

＜主翼２０＞
図１（ａ）に示すように、本実施形態による主翼２０は、その外殻が例えばアルミニウム合金からなる翼パネル（スキン）２１によって形成されている。翼パネル２１は、まげ加工により、主翼２０の前縁において湾曲している。これにより、主翼２０の前縁の近傍で翼長方向に沿った主翼前縁部２２は湾曲形状となっている。
なお、翼パネル２１は、炭素繊維と樹脂との複合材料であるＣＦＲＰ（Ｃａｒｂｏｎ Ｆｉｂｅｒ Ｒｅｉｎｆｏｒｃｅｄ Ｐｌａｓｔｉｃｓ）やガラス繊維と樹脂との複合材料であるＧＦＲＰ（Ｇｌａｓｓ Ｆｉｂｅｒ Ｒｅｉｎｆｏｒｃｅｄ Ｐｌａｓｔｉｃｓ）から形成したものを用いることもできる。
主翼２０内側の暖気室３０は、翼パネル２１と、主翼前縁部２２の後方に主翼２０の翼長方向に沿って設けられる後壁（衝突部）３２とに囲まれることにより形成されている。なお、暖気室３０は、翼長方向に区分されていてもよい。 <Main wing 20>
As shown in FIG. 1A, the main wing 20 according to the present embodiment is formed by a wing panel (skin) 21 whose outer shell is made of, for example, an aluminum alloy. The wing panel 21 is curved at the front edge of the main wing 20 by a curling process. Thereby, the main wing leading edge 22 along the wing length direction in the vicinity of the leading edge of the main wing 20 has a curved shape.
The wing panel 21 may be made of CFRP (Carbon Fiber Reinforced Plastics), which is a composite material of carbon fiber and resin, or GFRP (Glass Fiber Reinforced Plastics), which is a composite material of glass fiber and resin. it can.
The warm air chamber 30 inside the main wing 20 is formed by being surrounded by a wing panel 21 and a rear wall (impact portion) 32 provided along the blade length direction of the main wing 20 behind the main wing leading edge portion 22. . The warm air chamber 30 may be divided in the blade length direction.

＜防氷装置４０＞
防氷装置４０は、主翼前縁部２２を加熱する抽出エアを供給する抽出エアノズル（抽出エア供給部）４１と、抽出エアノズル４１に連結した図示しない抽出エア供給管と、抽出エアが吹き付けられる後壁３２と、から構成される。防氷装置４０は、暖気室３０に設けられている。 <Anti-icing device 40>
The anti-icing device 40 includes an extraction air nozzle (extraction air supply unit) 41 that supplies extraction air for heating the leading edge 22 of the main wing, an extraction air supply pipe (not shown) connected to the extraction air nozzle 41, and after the extraction air is sprayed And a wall 32. The anti-icing device 40 is provided in the warm air chamber 30.

図１（ａ）に示すように、抽出エアノズル４１は、主翼前縁部２２の先端部２２ａから離れた位置であって、後壁３２よりも先端部２２ａ側に、主翼２０の翼長方向に沿って配設されている。抽出エアノズル４１は、暖気室３０内に適宜の手段により固定されている。抽出エアノズル４１には、図示しないエンジンの圧縮機から抽出された高温（２００℃近傍）のエアを供給する図示しない抽出エア供給管が連結されている。
抽出エアノズル４１には、後壁３２に向けて開口する複数の吹き出し孔４１ａが形成されている。つまり、吹き出し孔４１ａは、後壁３２に対向して開口し、先端部２２ａとは反対方向である後方に向けて抽出エアが吹き出されるように形成されている。吹き出し孔４１ａは、抽出エアノズル４１の長手方向に所定の間隔を隔てて直線上に並んでいる。吹き出し孔４１ａからは、抽出エア供給管から抽出エアノズル４１に流入した抽出エアが、高温かつ高速のジェット噴流として後壁３２に向けて吹き付けられる。
後壁３２は、抽出エアが直接的に衝突する衝突点３２ｃを境に、その上下が前方に向けて湾曲しており、吹き付けられた抽出エアを先端部２２ａに向けた流れに形成しやすくなっている。ただし、これは好ましい形状であり、湾曲していない平坦な形状とすることもできる。
なお、本実施形態では、主翼２０を構成する構造部材（例えばスパー）が後壁３２を兼ねているが、後壁３２を構造部材とは別体として作製し、翼パネル２１の内側に適宜の手段で取り付けることもできる。 As shown in FIG. 1 (a), the extraction air nozzle 41 is located away from the tip 22a of the main wing leading edge 22 and closer to the tip 22a than the rear wall 32 in the blade length direction of the main wing 20. It is arranged along. The extraction air nozzle 41 is fixed in the warm air chamber 30 by an appropriate means. The extraction air nozzle 41 is connected to an extraction air supply pipe (not shown) that supplies high-temperature (around 200 ° C.) air extracted from a compressor of an engine (not shown).
The extraction air nozzle 41 is formed with a plurality of blowing holes 41 a that open toward the rear wall 32. That is, the blowout hole 41a is formed so as to open facing the rear wall 32, and the extraction air is blown out toward the rear, which is the direction opposite to the front end portion 22a. The blowout holes 41a are arranged in a straight line at a predetermined interval in the longitudinal direction of the extraction air nozzle 41. From the blowing hole 41a, the extraction air that has flowed into the extraction air nozzle 41 from the extraction air supply pipe is blown toward the rear wall 32 as a high-temperature and high-speed jet jet.
The rear wall 32 is curved forward and downward at a collision point 32c where the extracted air directly collides, and the blown extracted air is easily formed into a flow toward the tip 22a. ing. However, this is a preferred shape and can be a flat shape that is not curved.
In this embodiment, the structural member (for example, a spar) constituting the main wing 20 also serves as the rear wall 32. However, the rear wall 32 is manufactured separately from the structural member, and is appropriately placed inside the wing panel 21. It can also be attached by means.

＜抽出エアの流れＡＦ＞
図１を参照しながら、抽出エアの流れＡＦを説明する。
図示しないエンジンから抽出された抽出エアＡＦは、抽出エア供給管を介して抽出エアノズル４１に供給され、複数の吹き出し孔４１ａから高温かつ高速のジェット噴流として後壁３２に吹き出つけられる。抽出エアは後壁３２に衝突し、衝突点３２ｃを中心に分岐して後壁３２に沿って上下方向に流れ、主翼前縁部２２の後方から主翼前縁部２２の先端部２２ａに向けて内壁面３１に沿って流れる。これにより、主翼前縁部２２の後方から先端部２２ａにかけて全体が加熱される。なお、主翼前縁部２２の先端部２２ａで合流した抽出エアは、主翼前縁部２２の後方に流れ、暖気室３０から図示しない排出ポートを通して航空機の機外に排出される。 <Flow of extraction air AF>
The extraction air flow AF will be described with reference to FIG.
Extracted air AF extracted from an engine (not shown) is supplied to an extracted air nozzle 41 via an extracted air supply pipe, and blown out to the rear wall 32 as a high-temperature and high-speed jet jet from a plurality of blowing holes 41a. The extracted air collides with the rear wall 32, branches about the collision point 32 c, flows in the vertical direction along the rear wall 32, and extends from the rear of the main wing leading edge 22 toward the tip 22 a of the main wing leading edge 22. It flows along the inner wall surface 31. Thereby, the whole is heated from the rear of the main wing leading edge 22 to the tip 22a. The extracted air merged at the tip 22a of the main wing leading edge 22 flows behind the main wing leading edge 22, and is discharged from the warm air chamber 30 to the outside of the aircraft through a discharge port (not shown).

＜本実施形態の効果＞
このように、本実施形態では、先端部２２ａとは反対の方向に吹き付けられ後壁３２に衝突して上下方向に分岐した抽出エアが、主翼前縁部２２の後方から先端部２２ａに向けて流れる循環流として主翼前縁部２２内側の内壁面３１全体に供給されることにより、主翼前縁部２２の外周面が加熱される。したがって、主翼前縁部の先端部の内周面に直接抽出エアを吹き付ける従来の防氷装置を適用した場合のように、先端部の内周面が局所的に過熱されることがない。また、主翼前縁部２２の先端部２２ａと、そこから離れた主翼前縁部２２の後方との温度差が小さくなり、主翼前縁部２２全体を均一かつ効率的に加熱することができる。 <Effect of this embodiment>
As described above, in this embodiment, the extracted air that is blown in the direction opposite to the tip 22a and collides with the rear wall 32 and branches in the vertical direction is directed from the rear of the main wing leading edge 22 toward the tip 22a. By supplying the entire inner wall surface 31 inside the main wing leading edge 22 as a circulating flow, the outer peripheral surface of the main wing leading edge 22 is heated. Therefore, the inner peripheral surface of the tip is not locally overheated as in the case of applying a conventional anti-icing device that blows extracted air directly on the inner peripheral surface of the tip of the leading edge of the main wing. Further, the temperature difference between the front end 22a of the main wing leading edge 22 and the rear of the main wing leading edge 22 away from the leading end 22a is reduced, and the entire main wing leading edge 22 can be heated uniformly and efficiently.

さらに、主翼２０の構造部材である後壁３２に抽出エアを吹き付ける抽出エアノズル４１を設けるだけで済むため、簡易かつ軽量な防氷装置とすることができる。また、主翼前縁部２２の外側に部材を設ける必要もないので、空気抵抗が増えることによる燃費の悪化を招かない。   Furthermore, since it is only necessary to provide the extraction air nozzle 41 for blowing the extraction air to the rear wall 32 that is a structural member of the main wing 20, it is possible to provide a simple and lightweight anti-icing device. Moreover, since it is not necessary to provide a member outside the main wing leading edge 22, the fuel consumption is not deteriorated due to the increase in air resistance.

このような構成により、本実施形態の防氷装置では、主翼前縁部２２をより効率的に加熱することができる。その結果、主翼前縁部２２の外周面への着氷を確実かつ効率的に防止することができる。   With such a configuration, in the anti-icing device of this embodiment, the main wing leading edge 22 can be heated more efficiently. As a result, icing on the outer peripheral surface of the main wing leading edge 22 can be reliably and efficiently prevented.

なお、本実施形態の防氷装置では、図２に示すように、後壁３２の表面に遮熱コーティング３２ａを施すことができる。遮熱コーティングは通常の方法により形成することができ、例えばセラミックス塗料を後壁３２の表面に塗布する方法等を適用することができる。
そうすることにより、後壁３２の熱伝導率が下がり後壁３２における熱ロスを防ぐことができるため、主翼前縁部２２をより効率的に加熱することができる。また、後壁３２の耐熱性が向上するため、後壁３２自体はもちろん、その周囲の構造部材の過熱による焼損をより確実に防ぐことができる。 In the anti-icing device of the present embodiment, a thermal barrier coating 32a can be applied to the surface of the rear wall 32 as shown in FIG. The thermal barrier coating can be formed by an ordinary method. For example, a method of applying a ceramic paint to the surface of the rear wall 32 can be applied.
By doing so, the thermal conductivity of the rear wall 32 is lowered and heat loss in the rear wall 32 can be prevented, so that the main wing leading edge 22 can be heated more efficiently. In addition, since the heat resistance of the rear wall 32 is improved, not only the rear wall 32 itself but also the surrounding structural members can be more reliably prevented from being burned out.

さらに、本実施形態の防氷装置では、図３に示すように、内壁面３１全体に熱伝達率を向上する表面処理部３１ａを設けることもできる。この表面処理部３１ａは、例えば、ショットブラストによって、梨地に加工することができる。梨地以外でも、表面処理部３１ａの表面粗さが、表面処理が施されない場合の内壁面３１の表面粗さよりも大きくなればよい。なお、ショットブラスト以外の表面処理方法としては、エッチングを適用することもできる。また、梨地加工等により表面処理部３１ａの表面粗さを大きくする処理の他の方法を用いても、表面処理部３１ａに凹凸等が形成されれば同様の効果を得ることができる。
そうすることにより、内壁面３１全体の熱伝達率が向上し、後壁３２を介して供給される抽出エアにより主翼前縁部２２をより効果的に加熱することができる。 Furthermore, in the anti-icing device of this embodiment, as shown in FIG. 3, a surface treatment portion 31 a that improves the heat transfer coefficient can be provided on the entire inner wall surface 31. The surface treatment unit 31a can be processed into a satin finish by, for example, shot blasting. The surface roughness of the surface treatment part 31a should just become larger than the surface roughness of the inner wall surface 31 when not surface-treating also except a satin. Etching can also be applied as a surface treatment method other than shot blasting. Moreover, even if another method for increasing the surface roughness of the surface treatment portion 31a by using a satin finish or the like is used, the same effect can be obtained if the surface treatment portion 31a is formed with irregularities.
By doing so, the heat transfer coefficient of the entire inner wall surface 31 is improved, and the main wing leading edge 22 can be more effectively heated by the extracted air supplied through the rear wall 32.

なお、上記では、本発明の航空機における前縁部の防氷装置を航空機の主翼に適用して説明したが、主翼に限らず、尾翼など、他の翼に適用してもよい。また、翼の他にも、エンジンの空気取り入れ口の前縁部など、航空機の前縁であって着氷しやすい場所に適用することもできる。
また、上記では、複数の吹き出し孔４１ａを備える抽出エアノズル４１を例に説明したが、抽出エアを所望する箇所に供給できるものであればよく、例えば、吹き出し孔を一つだけ有するノズルを複数設けてもよい。さらに、上記では、複数の吹き出し孔４１ａを抽出エアノズル４１の長手方向の直線上に並べる例に説明したが、吹き出し孔４１ａから吹き出された抽出エアが後壁３２に直接吹き付けられるように配置されていればよく、例えば、吹き出し孔を抽出エアノズル４１の長手方向に千鳥状に設けたり、複列に配置することもできる。
これ以外にも、本発明の主旨を逸脱しない限り、上記実施の形態で挙げた構成を取捨選択したり、他の構成に適宜変更することが可能である。 In the above description, the anti-icing device for the leading edge in the aircraft according to the present invention has been applied to the main wing of the aircraft, but the present invention is not limited to the main wing, and may be applied to other wings such as the tail wing. In addition to the wing, the present invention can also be applied to a place where the front edge of the aircraft, such as the front edge portion of the air intake port of the engine, is icing.
In the above description, the extraction air nozzle 41 having a plurality of blowing holes 41a has been described as an example. However, any extraction air nozzle that can supply extraction air to a desired location may be used. For example, a plurality of nozzles having only one blowing hole are provided. May be. Further, in the above description, the example in which the plurality of blowing holes 41a are arranged on a straight line in the longitudinal direction of the extraction air nozzle 41 has been described. For example, the blowing holes may be provided in a staggered manner in the longitudinal direction of the extraction air nozzle 41, or may be arranged in a double row.
In addition to this, as long as it does not depart from the gist of the present invention, the configuration described in the above embodiment can be selected or changed to another configuration as appropriate.

２０…主翼、２１…翼パネル、２２…主翼前縁部、２２ａ…先端部、３０…暖気室、３１…内壁面、３１ａ…表面処理部、３２…後壁（衝突部）、３２ａ…遮熱コーティング、３２ｃ…衝突点、４０…防氷装置、４１…抽出エアノズル（抽出エア供給部）、４１ａ…吹き出し孔   20 ... main wing, 21 ... wing panel, 22 ... main wing leading edge, 22a ... tip, 30 ... warm air chamber, 31 ... inner wall surface, 31a ... surface treatment part, 32 ... rear wall (impact part), 32a ... heat shield Coating, 32c ... collision point, 40 ... anti-icing device, 41 ... extraction air nozzle (extraction air supply part), 41a ... blowout hole

Claims (5)

湾曲形状をなす前縁部の内周面に抽出エアを供給することで前記前縁部の外周面への着氷を防止する装置であって、
前記前縁部とは反対方向である後方に向けて抽出エアを吹き付ける抽出エア供給部と、
前記抽出エア供給部から吹き付けられた抽出エアが衝突する衝突部と、を備え、
前記衝突部に衝突した前記抽出エアは、前記前縁部の前記内周面に向けて供給されることを特徴とする
航空機における前縁部の防氷装置。 An apparatus for preventing icing on the outer peripheral surface of the front edge by supplying extraction air to the inner peripheral surface of the front edge having a curved shape,
An extraction air supply unit that blows extraction air toward the rear that is opposite to the front edge, and
A collision part with which the extraction air blown from the extraction air supply part collides,
The anti-icing device for a front edge portion in an aircraft, wherein the extracted air that has collided with the collision portion is supplied toward the inner peripheral surface of the front edge portion. 前記航空機の構造部材が前記衝突部を兼ねる、
請求項１に記載の航空機における前縁部の防氷装置。 The structural member of the aircraft also serves as the collision part,
The anti-icing device for the front edge of the aircraft according to claim 1. 前記衝突部は、遮熱コーティングが施されている、
請求項１に記載の航空機における前縁部の防氷装置。 The collision part is provided with a thermal barrier coating,
The anti-icing device for the front edge of the aircraft according to claim 1. 前記前縁部は、航空機の翼の前縁部であり、
前記衝突部に衝突した前記抽出エアは、前記翼を構成するパネルに沿って前記前縁部の前記内周面に供給される、
請求項１に記載の航空機における前縁部の防氷装置。 The leading edge is the leading edge of an aircraft wing;
The extracted air that has collided with the colliding part is supplied to the inner peripheral surface of the front edge part along a panel constituting the wing.
The anti-icing device for the front edge of the aircraft according to claim 1. 請求項１〜４のいずれか一項に記載された防氷装置を備えた、航空機の主翼。   The main wing of an aircraft provided with the anti-icing device as described in any one of Claims 1-4.
JP2012028270A 2012-02-13 2012-02-13 Anti-icing device and aircraft main wing Pending JP2013163480A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
JP2012028270A JP2013163480A (en) 2012-02-13 2012-02-13 Anti-icing device and aircraft main wing

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
JP2012028270A JP2013163480A (en) 2012-02-13 2012-02-13 Anti-icing device and aircraft main wing

Publications (1)

Publication Number Publication Date
JP2013163480A true JP2013163480A (en) 2013-08-22

Family

ID=49175127

Family Applications (1)

Application Number Title Priority Date Filing Date
JP2012028270A Pending JP2013163480A (en) 2012-02-13 2012-02-13 Anti-icing device and aircraft main wing

Country Status (1)

Country Link
JP (1) JP2013163480A (en)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN106319422A (en) * 2016-09-28 2017-01-11 晋西工业集团有限责任公司 Method for spraying thermal barrier coating onto empennage
CN106435433A (en) * 2016-09-28 2017-02-22 晋西工业集团有限责任公司 Thermal barrier coating spraying method applied to empennage
CN106498331A (en) * 2016-09-28 2017-03-15 晋西工业集团有限责任公司 A kind of spraying method of empennage thermal barrier coating
US9809298B2 (en) 2015-02-13 2017-11-07 Mitsubishi Aircraft Corporation Drainage system
CN113844659A (en) * 2021-09-30 2021-12-28 中航通飞华南飞机工业有限公司 Double-skin anti-icing cavity structure of airplane and heat exchange method

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9809298B2 (en) 2015-02-13 2017-11-07 Mitsubishi Aircraft Corporation Drainage system
CN106319422A (en) * 2016-09-28 2017-01-11 晋西工业集团有限责任公司 Method for spraying thermal barrier coating onto empennage
CN106435433A (en) * 2016-09-28 2017-02-22 晋西工业集团有限责任公司 Thermal barrier coating spraying method applied to empennage
CN106498331A (en) * 2016-09-28 2017-03-15 晋西工业集团有限责任公司 A kind of spraying method of empennage thermal barrier coating
CN113844659A (en) * 2021-09-30 2021-12-28 中航通飞华南飞机工业有限公司 Double-skin anti-icing cavity structure of airplane and heat exchange method

Similar Documents

Publication Publication Date Title
US10393020B2 (en) Injector nozzle configuration for swirl anti-icing system
US8061657B2 (en) Method and apparatus for aircraft anti-icing
JP2013163480A (en) Anti-icing device and aircraft main wing
US6354538B1 (en) Passive control of hot air injection for swirling rotational type anti-icing system
EP3216697B1 (en) Shielded anti-icing system and methods
JP6839920B2 (en) De-icing equipment and aircraft
EP2851299B1 (en) Concentric nozzles for enhanced mixing of fluids
US6131855A (en) Device for removing hot air for a jet engine air inlet cowl with a de-icing circuit
JP6666159B2 (en) Anti-icing equipment and aircraft
US20170210475A1 (en) Nozzle and vane system for nacelle anti-icing
CA2928267C (en) Anti-icing system for an aircraft
EP2894096B1 (en) Aircraft anti-icing systems having deflector vanes
EP2250090B1 (en) Icing protection for aircraft air inlet scoops
JP2017031977A (en) Method for cooling turbo-engine component and turbo-engine component
US20010003897A1 (en) Device for discharging hot air for a jet engine air inlet cowl, with a deicing circuit
US10167086B2 (en) Anti-icing system for an aircraft
JP2017136893A (en) Anti-icing device and aircraft
BRPI0617040A2 (en) aircraft double-flow turbocharger
JP2009507179A (en) Infrared suppression system
JP2013163479A (en) Anti-icing structure and aircraft main wing
US10723464B2 (en) Injector nozzle configuration for swirl anti-icing system
CN106678873B (en) A kind of supporting plate tail portion double oil circuits integration after-burner
CN110318883A (en) A kind of aero-engine calotte single hole impingement heat transfer structure in helical curve channel
US11459087B2 (en) Leading-edge device for an aircraft
CN106762146B (en) A kind of hot air anti-icing structure of engine guide vane