JP3529911B2 - Aircraft leading edge structure and method for manufacturing leading edge - Google Patents
Aircraft leading edge structure and method for manufacturing leading edgeInfo
- Publication number
- JP3529911B2 JP3529911B2 JP22826495A JP22826495A JP3529911B2 JP 3529911 B2 JP3529911 B2 JP 3529911B2 JP 22826495 A JP22826495 A JP 22826495A JP 22826495 A JP22826495 A JP 22826495A JP 3529911 B2 JP3529911 B2 JP 3529911B2
- Authority
- JP
- Japan
- Prior art keywords
- warm air
- outer skin
- skin
- aircraft
- wall
- 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.)
- Expired - Lifetime
Links
Classifications
-
- Y02T50/433—
Landscapes
- Moulding By Coating Moulds (AREA)
Abstract
Description
【０００１】[0001]
【発明の属する技術分野】本発明は、航空機の主翼、尾
翼、エンジンカウル・インレット部等の前縁部の防氷構
造を強化繊維樹脂で成形する場合の製造法及び前縁構造
に関する。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a manufacturing method and a leading edge structure for molding an anti-icing structure of a leading edge portion such as a main wing, a tail wing and an engine cowl inlet of an aircraft with a reinforced fiber resin.
【０００２】[0002]
【従来の技術】従来、航空機の主翼、尾翼、エンジンカ
ウル・インレット部、昇降舵、方向舵のホーンバランス
先端部等の前縁には、エンジンコンプレッサ等から高温
のブリードエアを導いて前縁部の内部空間室に流通させ
るような防氷構造が採用される。このような前縁構造と
して、例えば米国特許５，０１１，０９８号とか、米国
Reg.Number：Ｈ６４８とか、米国特許４，７３８，４１
６号とか、特開昭６１−９４８９８号のような技術が知
られている。そして、米国特許５，０１１，０９８号の
場合は、前縁部の構造として、外皮の裏側に内皮と隔壁
を設けて暖気室を区画形成するとともに、この暖気室を
暖気噴出部と暖気通路部に分け、主として外皮と隔壁で
囲まれる空間部を暖気噴出部とし、外皮と内皮で囲まれ
る空間部を暖気通路部とするようにしている。そして、
内皮の断面形状を凹凸形状にすることで、暖気通路部の
内部空間を複数の整流壁で仕切り、複数の暖気通路を形
成するようにしている。2. Description of the Related Art Conventionally, high temperature bleed air is introduced from an engine compressor or the like to the leading edge of the main wing, tail, engine cowl / inlet, elevator and rudder horn balance tip of the aircraft. An anti-icing structure is adopted so that it can be distributed to the internal space room. Examples of such a leading edge structure include US Pat. No. 5,011,098 and US Pat.
Reg.Number: H648 or US Pat. No. 4,738,41
Techniques such as No. 6 and JP-A No. 61-94898 are known. In the case of US Pat. No. 5,011,098, as the structure of the front edge portion, the inner wall and the partition wall are provided on the back side of the outer skin to partition and form the warm air chamber, and the warm air chamber is formed by the warm air ejection part and the warm air passage part. The space surrounded by the outer skin and the partition wall is mainly used as the warm air jetting part, and the space surrounded by the outer skin and the inner skin is used as the warm air passage part. And
By making the cross-sectional shape of the inner skin uneven, the inner space of the warm air passage portion is partitioned by a plurality of flow regulating walls to form a plurality of warm air passages.
【０００３】また、米国Reg.Number：Ｈ６４８の場合
は、例えばエンジンカウル・インレット部の前縁構造に
関し、外皮の内部に、隔壁を結合せしめたフランジ部品
を結合し、この隔壁と外皮によって暖気噴出室を区画形
成するとともに、前記フランジ部品の一部に暖気通路と
なる波形部を形成するようにしている。また、米国特許
４，７３８，４１６号、及び特開昭６１−９４８９８号
の場合も、外皮の前縁内部を隔壁で区画して暖気噴出室
を形成し、この暖気噴出室から噴出した暖気を外皮の裏
面に沿って流動させるような技術を開示している。In the case of US Reg.Number: H648, for example, regarding the front edge structure of the engine cowl inlet part, a flange part in which a partition wall is connected is connected to the inside of the outer skin, and warm air is ejected by the partition wall and the outer skin. The chamber is partitioned and a corrugated portion serving as a warm air passage is formed in a part of the flange component. In the case of U.S. Pat. No. 4,738,416 and Japanese Patent Laid-Open No. 61-94898 as well, the inside of the front edge of the outer skin is partitioned by a partition wall to form a warm air ejection chamber, and the warm air ejected from this warm air ejection chamber is Disclosed is a technique of flowing along the back surface of the outer skin.
【０００４】[0004]
【発明が解決しようとする課題】ところが、上記のよう
な前縁部の防氷構造において、外皮の裏側に内皮と隔壁
を固着して暖気室を形成する際、例えば外皮と内皮と隔
壁がアルミ合金等の金属素材であるような場合には、内
皮等を結合するためのリベット等の結合部品が外皮の外
側表面に露出し、前縁部の外表面に沿って流れる空気抵
抗が増えるため、空気抵抗を一層減らして空力特性の向
上を図ることが望まれるところである。またリベット等
の結合部品は重量増加を招くため、部品削減による重量
の軽減が望まれるところである。また、前縁防氷部に高
温の暖気を流通させた際に外皮が熱変形すると前縁部の
気流が乱れるため熱的強度の高い構造が望まれるところ
である。更に、結合部品用の貫通孔等を無くし、耐クラ
ック破壊、耐腐食性等の特性を向上させることが望まれ
るところである。However, in the above-mentioned front edge anti-icing structure, when the inner skin and the partition are fixed to the back side of the outer skin to form the warm air chamber, for example, the outer skin, the inner skin and the partition are made of aluminum. In the case of a metal material such as an alloy, the connecting parts such as rivets for connecting the inner skin are exposed on the outer surface of the outer skin, and the air resistance flowing along the outer surface of the front edge increases, It is desired to further reduce the air resistance to improve the aerodynamic characteristics. Further, since a connecting component such as a rivet causes an increase in weight, it is desired to reduce the weight by reducing the number of components. Further, when the outer skin is thermally deformed when hot air is circulated through the leading edge anti-icing portion, the air flow at the leading edge is disturbed, so that a structure having high thermal strength is desired. Furthermore, it is desired to eliminate the through holes and the like for joint parts and improve the properties such as crack fracture resistance and corrosion resistance.
【０００５】そこで、前縁部を強化繊維樹脂で成形すれ
ば、結合部品が外皮表面に露出するような問題はなくな
り、空気抵抗も少なくなるばかりでなく、熱膨張係数も
小さいため（例えばアルミ合金の熱膨張係数は２３であ
るに対して、ＣＲＦＰは３〜５）熱的強度も高くなって
好都合であるが、例えば強化繊維樹脂の外皮と内皮と隔
壁等の各パーツを別々に成形し、これを接着等で組み立
てるような方法では、夫々別個の成形型が必要であり、
費用と手間がかかることに加えて、特に隔壁と内皮の接
合部に隙間等が生じる虞れがあった。そしてこの接合部
に隙間等が生じて気密状に保たれないような場合は、そ
の隙間部分から暖気が漏れて防氷効果を損なうという問
題があった。そこで、強化繊維樹脂を用いた航空機の前
縁構造において、特に防氷効果を左右する隔壁と内皮の
接合部を気密状にし、しかも、全体の構造を可能な限り
強固にするとともに、成形容易でコストもかからない前
縁構造が望まれていた。Therefore, if the front edge portion is formed of a reinforced fiber resin, the problem that the joint component is exposed on the surface of the outer cover is eliminated, not only the air resistance is reduced but also the thermal expansion coefficient is small (for example, aluminum alloy). Has a thermal expansion coefficient of 23, while CRFP has a high thermal strength of 3 to 5), which is convenient. For example, the outer skin of the reinforced fiber resin, the inner skin, and the partition wall are molded separately, In the method of assembling this by bonding etc., separate molds are required,
In addition to the cost and labor, there is a possibility that a gap or the like may be generated especially at the joint between the partition wall and the inner skin. In the case where a gap or the like is formed at the joint and the airtightness cannot be maintained, there is a problem that warm air leaks from the gap and impairs the anti-icing effect. Therefore, in the aircraft leading edge structure using reinforced fiber resin, in particular, the joint between the partition wall and the endothelium, which influences the anti-icing effect, is made airtight, and the overall structure is made as strong as possible and easy to mold. A leading-edge structure that does not cost much has been desired.
【０００６】[0006]
【課題を解決するための手段】上記課題を解決するため
本発明は、請求項１において、外皮と内皮と隔壁で囲ま
れる前縁部の内部空間部を暖気室として形成し、この暖
気室に複数の整流壁で仕切られる複数の暖気通路を形成
するようにした航空機の前縁構造において、前記外皮と
整流壁を、夫々の強化繊維樹脂を一体に接合硬化させて
成形し、この上から予め強化繊維樹脂から一部品として
成形した内皮と隔壁を所定部位に接着固定するように
し、前記強化繊維樹脂がビスマレイミド変性樹脂或いは
ポリイミド系樹脂からなるようにした。そして、外皮と
整流壁の２部品を１つの型から一体に接合硬化させ、ま
た内皮と隔壁を１つの型から一部品として成形すること
で手間が省かれ、しかも構造全体の強度を高めることが
出来る。しかも内皮と隔壁は一部品であるため両者間は
完全に気密状態に保たれ、暖気が漏れるような不具合が
ない。また、請求項２では、外皮と内皮と隔壁部で囲ま
れる前縁部の内部空間部を暖気室として形成し、この暖
気室に複数の整流壁で仕切られる複数の暖気通路を形成
するようにした航空機の前縁構造において、前記外皮と
整流壁は、夫々の強化繊維樹脂を一体に接合硬化させた
ものであり、前記内皮と隔壁部は、予め強化繊維樹脂か
ら一部品として成形したものを所定部位に接着固定した
ものであり、前記内皮の後方には後方隔壁が設けられ、
この後方隔壁と前記隔壁部によって暖気排出室が区画形
成され、この暖気排出室に入った暖気が翼端の排出口か
ら大気に放出されるようにした。 In order to solve the above-mentioned problems, the present invention provides a warm air chamber in which the inner space of the front edge surrounded by the outer skin, the inner skin and the partition wall is formed as a warm air chamber. In the leading edge structure of an aircraft configured to form a plurality of warm air passages partitioned by a plurality of straightening walls, the outer skin and the straightening walls are formed by integrally bonding and hardening the respective reinforcing fiber resins, and then from above The inner wall and partition wall molded as one part from the reinforced fiber resin are adhered and fixed to a predetermined portion, and the reinforced fiber resin is a bismaleimide-modified resin or
It is made of a polyimide resin . Then, two parts of the outer skin and the flow regulating wall are integrally bonded and cured from one mold, and the inner skin and the partition wall are molded as one part from one mold, which saves labor and enhances the strength of the entire structure. I can. Moreover, since the inner skin and the partition wall are a single component, the space between the inner skin and the partition wall is completely kept airtight, and there is no problem that hot air leaks. Moreover, in Claim 2, it is surrounded by an outer skin, an inner skin, and a partition wall.
The interior space of the front edge is formed as a warm air chamber,
Multiple warm air passages are formed in the air chamber that are partitioned by multiple flow control walls
In the leading edge structure of the aircraft,
The straightening wall was made by joining and hardening the respective reinforced fiber resins integrally.
The inner wall and the partition wall are made of reinforced fiber resin in advance.
The one that was molded as one part was fixed to the specified area by adhesion.
A rear partition is provided behind the endothelium,
This rear partition wall and the partition wall partition the warm air discharge chamber
The warm air that is generated and enters the warm air discharge chamber
To be released into the atmosphere.
【０００７】また請求項３では、外皮と内皮と隔壁で囲
まれる前縁部の内部空間部を暖気室として形成し、この
暖気室に複数の整流壁で仕切られる複数の暖気通路を形
成するようにした航空機の前縁部の製造方法において、
前縁部の形状を模した工具に外皮となる未硬化状態の強
化繊維樹脂製の積層体を位置決めし、この積層体の内側
面上に、前記暖気通路を形成するための形状保持具と、
この形状保持具の端部に配置された断面Ｌ字型の未硬化
状態の強化繊維樹脂の整流壁を複数並べて位置決めする
とともに、この上を真空バッグで覆って内部を真空引き
し加熱・加圧するようにした。そして整流壁の一端側を
前記積層体に一体に接合硬化させ、形状保持具を引き出
した後、予め強化繊維樹脂から一部品として成形した内
皮と隔壁を整流壁の他端側に当接させて所定部に接着固
定するようにした。Further, according to a third aspect of the present invention , an inner space portion of a front edge surrounded by the outer skin, the inner skin and the partition wall is formed as a warm air chamber, and a plurality of warm air passages partitioned by a plurality of flow regulating walls are formed in the warm air chamber. In the manufacturing method of the leading edge of the aircraft,
A reinforced fiber resin laminate in an uncured state that serves as an outer skin is positioned on a tool simulating the shape of the front edge, and a shape retainer for forming the warm air passage on the inner surface of the laminate,
A plurality of rectifying walls of an uncured reinforcing fiber resin having an L-shaped cross section arranged at the end of this shape-retaining device are aligned and positioned, and a vacuum bag is covered thereover to evacuate and heat and pressurize the inside. I did it. Then, one end side of the straightening wall is integrally bonded and cured to the laminated body, and after the shape-retaining tool is drawn out, the inner wall and the partition wall formed in advance as one component from the reinforced fiber resin are brought into contact with the other end side of the straightening wall. The adhesive was fixed to a predetermined portion.
【０００８】また、請求項４では、請求項２記載の航空
機の前縁部の製造方法において、形状保持具をシリコー
ンゴムとした。このシリコーンゴムとは、シロキサン結
合の繰り返し（Ｓi-o）nを主鎖とし、側鎖にアルキル、
アリール基等を持つ重合体であり、耐熱性に優れ、硬化
した積層体との離脱性も良好な弾性体である。According to a fourth aspect of the present invention, in the method of manufacturing a front edge portion of an aircraft according to the second aspect, the shape retainer is made of silicone rubber. This silicone rubber has a repeating siloxane bond (Si-o) n as the main chain and an alkyl side chain,
It is a polymer having an aryl group and the like, and is an elastic body having excellent heat resistance and good releasability from the cured laminate.
【０００９】[0009]
【発明の実施の形態】本発明の実施の形態の一例につい
て添付した図面に基づき説明する。ここで図１は本発明
の航空機の前縁構造を示す斜視図、図２は同断面図、図
３乃至図６は前縁部の製造方法を示す工程図である。航
空機の主翼等の前縁部には氷等が付着するのを防止する
ための防氷構造が採用され、この防氷構造は、例えばエ
ンジンコンプレッサー等から引出された温度の高いブリ
ードエアを、外皮で囲まれる前縁部の内部空間部に導
き、外皮の裏側に沿って流動させることで外皮の表面温
度を高め前縁部表面に氷等が付着するのを防止するよう
にしている。BEST MODE FOR CARRYING OUT THE INVENTION An example of an embodiment of the present invention will be described with reference to the accompanying drawings. Here, FIG. 1 is a perspective view showing a leading edge structure of an aircraft of the present invention, FIG. 2 is a sectional view of the same, and FIGS. 3 to 6 are process drawings showing a method of manufacturing a leading edge portion. An anti-icing structure is used to prevent ice from adhering to the front edge of the main wing of an aircraft.This anti-icing structure uses, for example, high-temperature bleed air drawn from an engine compressor, etc. It is guided to the inner space of the front edge surrounded by and is made to flow along the back side of the outer skin to raise the surface temperature of the outer skin and prevent ice and the like from adhering to the surface of the front edge.
【００１０】そして、本発明の航空機の前縁構造は、機
体の前縁部をビスマレイミド変性樹脂／炭素繊維複合材
からなる強化繊維樹脂にて成形しており、図１、図２に
示すように、上記強化繊維樹脂からなる外皮１の内側の
前縁部寄りに同強化繊維樹脂からなる内部壁２が設けら
れ、この外皮１と内部壁２によって暖気室４が区画形成
されている。そして内部壁２は従来構造の内皮に対応す
る内皮部２ａ、２ａと隔壁に対応する隔壁部２ｂを備え
ている。また、内部壁２の後方には後方隔壁５が設けら
れ、この後方隔壁５と内部壁２によって暖気排出室６を
区画形成するとともに、この後方隔壁５の後方の室を温
度緩衝室７として区画し、この後方の温度緩衝室７内に
外気を導入して翼の後方部に内装したインテグラルタン
ク等の過熱を防止するようにしている。In the leading edge structure of the aircraft of the present invention, the leading edge portion of the body is molded with a reinforcing fiber resin composed of a bismaleimide modified resin / carbon fiber composite material, as shown in FIGS. 1 and 2. Further, an inner wall 2 made of the same reinforced fiber resin is provided near the front edge portion inside the outer wrap 1 made of the reinforced fiber resin, and the warm air chamber 4 is defined by the outer skin 1 and the inner wall 2. The inner wall 2 is provided with inner wall portions 2a, 2a corresponding to the inner wall and a partition wall portion 2b corresponding to the partition wall of the conventional structure. Further, a rear partition wall 5 is provided behind the inner wall 2, and a warm air discharge chamber 6 is defined by the rear partition wall 5 and the inner wall 2, and a chamber behind the rear partition wall 5 is partitioned as a temperature buffer chamber 7. However, the outside air is introduced into the temperature buffer chamber 7 at the rear side to prevent overheating of the integral tank or the like installed inside the rear portion of the blade.
【００１１】前記暖気室４は、内部壁２の隔壁部２ｂを
境にして前方の暖気噴出部４ａと後方の暖気通路部４ｂ
に分けられ、前記暖気噴出部４ａ内には、ブリードエア
を流通させるピッコロチューブ８が収容され、このピッ
コロチューブ８には多数のエア噴出口８ａ（図１）が設
けられている。また、暖気通路部４ｂの外皮１には、長
手方向に沿って所定間隔置きに断面Ｌ字型の複数の整流
フィンＦ、…が一体に接合硬化されている。そして、こ
の整流フィンＦ、…によって暖気通路部４ｂ内の空間を
複数の暖気通路Ｔ、…に仕切っている。The warm air chamber 4 has a front warm air blowout portion 4a and a rear warm air passage portion 4b with a partition wall portion 2b of the inner wall 2 as a boundary.
A piccolo tube 8 for circulating bleed air is housed in the warm air jetting portion 4a, and the piccolo tube 8 is provided with a large number of air jets 8a (FIG. 1). Further, a plurality of straightening fins F having an L-shaped cross section are integrally bonded and hardened to the outer skin 1 of the warm air passage portion 4b at predetermined intervals along the longitudinal direction. The rectifying fins F, ... Divide the space in the warm air passage portion 4b into a plurality of warm air passages T ,.
【００１２】そして、ピッコロチューブ８のエア噴出口
８ａから噴出した高温のブリードエアは、暖気噴出部４
ａから暖気通路部４ｂの暖気通路Ｔ、…を流動し、暖気
排出室６に入った後、翼端の排出口から大気に放出され
る。そしてこのように前縁部の外皮１の内面に沿って高
温のブリードエアを流動させることで、翼前面に氷等が
付着するのを防止する。The high-temperature bleed air ejected from the air ejection port 8a of the piccolo tube 8 is heated by the warm air ejection portion 4
After flowing from a into the warm air passage T of the warm air passage portion 4b, and entering the warm air discharge chamber 6, it is discharged to the atmosphere from the discharge port of the blade tip. Then, by causing the high temperature bleed air to flow along the inner surface of the outer skin 1 at the front edge portion, ice and the like are prevented from adhering to the front surface of the blade.
【００１３】以上のような前縁構造の製造法について図
３乃至図６に基づき説明する。まず、図３（Ａ）に示す
ように、前縁部の形状を模したアウタースキンツール型
１０に対し、例えば炭素繊維にビスマレイシド変性樹脂
を含浸させた複合中間材料を積層した積層体としてのシ
ート状のアウタースキンプリプレグ１１を位置決めす
る。このアウタースキンプリプレグ１１は、予め所定の
サイズにセットされて適当な粘着性と柔らかさ、なじみ
やすさが与えられており、アウタースキンツール型１０
に押し付けることで型面に粘着し、型面形状に倣って成
形される。A method of manufacturing the above leading edge structure will be described with reference to FIGS. First, as shown in FIG. 3 (A), a sheet as a laminated body in which a composite intermediate material obtained by impregnating carbon fiber with a bismaleide modified resin is laminated on an outer skin tool mold 10 simulating the shape of the front edge portion. Position the outer skin prepreg 11 in the shape of a circle. The outer skin prepreg 11 is set to a predetermined size in advance to give suitable adhesiveness, softness, and adaptability.
It is adhered to the mold surface by being pressed against and is molded following the shape of the mold surface.
【００１４】次いで図３（Ｂ）に示すように、このアウ
タースキンプリプレグ１１の上下傾斜面に、形状保持具
としての矩形状のシリコーンブロック１２、…と断面Ｌ
字型の整流フィンプリプレグ１３、…を順次並べて位置
決めする。この際、断面Ｌ字型の整流フィンプリプレグ
１３は、図４に示すように、シリコーンブロック１２の
片側端部のコーナ部を囲うように位置決めセットされて
おり、この整流フィンプリプレグ１３が位置決めセット
されたシリコーンブロック１２を長手方向に沿って密着
状に並べるとともに、整流フィンプリプレグ１３の一面
側がアウタースキンプリプレグ１１の内面に当接するよ
うにしている。Next, as shown in FIG. 3 (B), a rectangular silicone block 12, serving as a shape retainer, and a cross section L are formed on the vertically inclined surface of the outer skin prepreg 11.
The character-shaped straightening fin prepregs 13, ... Are sequentially arranged and positioned. At this time, the straightening fin prepreg 13 having an L-shaped cross section is positioned and set so as to surround the corner portion of one end of the silicone block 12, as shown in FIG. 4, and this straightening fin prepreg 13 is positioned and set. The silicone blocks 12 are arranged in close contact with each other along the longitudinal direction, and one side of the flow regulating fin prepreg 13 is brought into contact with the inner surface of the outer skin prepreg 11.
【００１５】次に、図３（Ｃ）に示すように、この上を
真空バッグ１４で覆い、この真空バッグ１４内を減圧す
る。すると、各プリプレグ１１、１３内の空気が脱気さ
れ、アウタースキンプリプレグ１１はアウタースキンツ
ール型１０に対して正確になじむとともに、アウタース
キンプリプレグ１１と整流フィンプリプレグ１３の当接
面が密着する。そして、アウタースキンツール型１０ご
とオートクレーブに入れて、所定の加熱・加圧パターン
で加熱・加圧を行えば、アウタースキンプリプレグ１１
は硬化して外皮１となり、整流フィンプリプレグ１３、
…は硬化して整流フィンＦになるとともに、外皮１と整
流フィンＦ、…の接合部は自己接着力によって一体に強
固に接合する。Next, as shown in FIG. 3 (C), a vacuum bag 14 is covered thereover, and the inside of the vacuum bag 14 is depressurized. Then, the air inside the prepregs 11 and 13 is degassed, the outer skin prepreg 11 conforms to the outer skin tool mold 10 accurately, and the contact surfaces of the outer skin prepreg 11 and the rectifying fin prepreg 13 are brought into close contact with each other. Then, the outer skin prepreg 11 is put in an autoclave together with the outer skin tool mold 10 and heated and pressed according to a predetermined heating and pressing pattern.
Hardens to form the outer skin 1, and the straightening fin prepreg 13,
Is cured to become the flow-rectifying fins F, and the joint portion between the outer cover 1 and the flow-rectifying fins F is integrally and firmly bonded by self-adhesive force.
【００１６】その後、シリコーンブロック１２、…を引
張って抜き出すと、シリコーンブロック１２、…が引抜
かれた箇所が暖気通路Ｔ、…部分として形成される。ま
た、シリコーンブロック１２は、引張ることで幅方向の
寸法が縮まるため、抜き勾配がなくても容易に引き出す
ことが出来る。Then, when the silicone blocks 12, ... Are pulled out and pulled out, the portions where the silicone blocks 12, ... Are pulled out are formed as warm air passages T ,. Further, since the width of the silicone block 12 is reduced by pulling, the silicone block 12 can be easily pulled out without a draft.
【００１７】一方、内部壁２は、以上の手順とほぼ同様
な要領で別成形にて成形される。すなわち、不図示のイ
ンナースキンツール型に対して、炭素繊維にビスマレイ
シド変性樹脂を含浸させてなるインナースキンプリプレ
グをセットし、これを真空バッグで覆って減圧した後、
オートクレーブに入れて加熱・加圧し硬化させる。そし
て、このインナースキンプリプレグを硬化させれば内皮
部２ａ、２ａと隔壁部２ｂを備えた内部壁２が成形され
る。On the other hand, the inner wall 2 is formed separately by a procedure similar to the above procedure. That is, for an inner skin tool type (not shown), an inner skin prepreg made by impregnating a carbon fiber with a bismaleside modified resin is set, and after covering this with a vacuum bag to reduce the pressure,
Put in an autoclave and heat / pressurize to cure. Then, by curing the inner skin prepreg, the inner wall 2 including the inner skin portions 2a, 2a and the partition wall portion 2b is formed.
【００１８】そして、このように別工程にて成形された
内部壁２は、図５に示すように外皮１に固着される。す
なわち、例えば内部壁２の両端部を左右一対のリブ１
５、１５に接着等で固定し、このリブ１５、１５を外皮
１に接着等で固定して内皮部２ａと整流フィンＦ、…を
密接させる。この際、例えば図６に示すように、整流フ
ィンＦ、…の先端部ｆ、…に内皮部２ａを当接させる前
に、予め整流フィンＦ、…の先端部ｆ、…にシール剤を
塗布しておき、整流フィンＦ、…と内皮部２ａの当接部
をシールする。The inner wall 2 thus formed in a separate step is fixed to the outer cover 1 as shown in FIG. That is, for example, a pair of left and right ribs 1 is provided on both ends of the inner wall 2.
The ribs 15 and 15 are fixed to the outer skin 1 by adhesion or the like so that the inner skin portion 2a and the rectifying fins F are brought into close contact with each other. At this time, for example, as shown in FIG. 6, before the inner skin portion 2a is brought into contact with the tip end portions f of the flow regulating fins F, ..., A sealing agent is applied to the tip end portions f of the flow regulating fins F ,. Then, the contact portions of the rectifying fins F, ... And the inner skin portion 2a are sealed.
【００１９】以上のような手順で成形された航空機の前
縁構造は、内部壁２の内皮部２ａと隔壁部２ｂの隙間が
ないため、従来のように暖気室４の暖気が暖気排出室６
に漏れ出すような不具合がなく、しかも、内皮部２ａと
隔壁部２ｂが一部品から形成されているため構造が強固
である。In the leading edge structure of an aircraft formed by the above procedure, since there is no gap between the inner wall portion 2a of the inner wall 2 and the partition wall portion 2b, the warm air in the warm air chamber 4 is different from the warm air exhaust chamber 6 as in the conventional case.
There is no problem of leakage to the inside, and since the inner skin portion 2a and the partition wall portion 2b are formed from one part, the structure is strong.
【００２０】因みに、本実施形態で複合材樹脂としてビ
スマレイミド変性樹脂を採用した理由は、吹き付けられ
るブリードエアの温度が１８０℃前後であるのに対し
て、この高温下での物性を満足する複合材樹脂として、
ビスマレイミド変性樹脂（硬化温度１８０〜１９０℃）
とポリイミド系樹脂（硬化温度３７０〜４００℃）が考
えられ、このうちポリイミド系樹脂は繊維と樹脂が別々
にしか入手出来ないのに対して、ビスマレイミド変性樹
脂はプリプレグ状態で入手可能で、且つ硬化温度がポリ
イミド系樹脂より低いため設備コストがより安価に済む
からである。勿論、実施の形態に限定されるものではな
い。また、適用箇所も主翼に限定されるものではなく、
その他、エンジンカウル・インレット部の前縁部、エン
ジン・サポートアーム部の先端部、水平・垂直尾翼前縁
部、昇降舵、方向舵ホーンバランス先端部等にも適用可
能である。Incidentally, the reason why the bismaleimide-modified resin is adopted as the composite material resin in this embodiment is that the temperature of the bleed air to be sprayed is around 180 ° C., whereas the composite material satisfying the physical properties at this high temperature is used. As material resin,
Bismaleimide modified resin (curing temperature 180-190 ℃)
And a polyimide resin (curing temperature of 370 to 400 ° C.) are conceivable. Among them, the polyimide resin is available only as a fiber and a resin, whereas the bismaleimide modified resin is available in a prepreg state, and This is because the curing temperature is lower than that of the polyimide resin, so that the facility cost can be reduced. Of course, it is not limited to the embodiment. Also, the application site is not limited to the wing,
In addition, it is also applicable to the leading edge of the engine cowl inlet, the tip of the engine support arm, the leading edge of the horizontal / vertical tail, the elevator, and the rudder horn balance tip.
【００２１】[0021]
【発明の効果】以上のように本発明は、請求項１におい
て、外皮と内皮と隔壁で囲まれる前縁部の内部空間部を
暖気室として形成するとともに、この暖気室に複数の整
流壁を設けるようにした航空機の前縁構造において、外
皮と整流壁を一体に接合硬化せしめるとともに、内皮と
隔壁を一部品として成形するようにしたため、前縁構造
全体の曲げ剛性が高くなり、しかも暖気室を気密状に保
持することが出来る。しかも作業に手間がかからない。
また、前縁部をビスマレイミド変性樹脂或いはポリイミ
ド系樹脂からなる強化繊維樹脂で成形しているので、結
合部材が外皮表面に露出するような問題はなくなり、空
気抵抗も少なくなるばかりでなく、熱膨張係数も小さい
ため熱的強度も高くなって好都合である。更に、請求項
２のように、内皮の後方に後方隔壁を設け、後方隔壁と
隔壁部によって暖気排出室を区画形成し、暖気排出室に
入った暖気を翼端の排出口から大気に放出させるように
すれば、前縁部の外皮の内面に沿って高温のブリードエ
アを流動させ、翼前面に氷等が付着するのを防止出来
る。そして、請求項３のような製造方法を採用すれば手
際良く作業を行うことが出来、また請求項４のように、
形状保持具をシリコーンゴムとすれば、引張って抜き出
す際に幅方向の寸法が縮み、抜き勾配がなくても硬化し
た積層体から抜き出すことが出来、作業性が良い。As described above, according to the present invention, in claim 1, the inner space portion of the front edge portion surrounded by the outer skin, the inner skin and the partition wall is formed as a warm air chamber, and a plurality of straightening walls are provided in the warm air chamber. In the aircraft's leading edge structure, the outer skin and flow straightening wall were integrally bonded and hardened, and the inner skin and the bulkhead were molded as one part, so the bending rigidity of the entire leading edge structure was increased and the warm air chamber Can be kept airtight. Moreover, it does not take time to work.
In addition, the front edge is bismaleimide-modified resin or polyimid.
Since it is molded with reinforced fiber resin made of
There is no problem that the composite material is exposed on the outer skin surface,
Not only the air resistance decreases, but also the coefficient of thermal expansion is small
Therefore, the thermal strength is high, which is convenient. Further, as in claim 2, a rear partition is provided behind the inner skin, and the warm air discharge chamber is defined by the rear partition and the partition portion, and the warm air that has entered the warm air discharge chamber is discharged to the atmosphere from the outlet of the blade tip. By doing so, high-temperature bleed air can be caused to flow along the inner surface of the outer skin of the front edge portion, and ice or the like can be prevented from adhering to the front surface of the blade. Then, if the manufacturing method as claimed in claim 3 is adopted, the work can be done smoothly , and as in claim 4 ,
If the shape-retaining tool is made of silicone rubber, the dimension in the width direction shrinks when pulled out by pulling, and it can be pulled out from the cured laminate without any draft, so that workability is good.
【図１】本発明の航空機の前縁構造を示す斜視図FIG. 1 is a perspective view showing a leading edge structure of an aircraft of the present invention.
【図２】同断面図FIG. 2 is a sectional view of the same.
【図３】前縁部の製造方法を示す工程図FIG. 3 is a process drawing showing the method for manufacturing the front edge portion.
【図４】シリコーンブロックと整流フィンプリプレグの
配置の説明図FIG. 4 is an explanatory diagram of an arrangement of a silicone block and a flow-regulating fin prepreg.
【図５】内部壁を接着する工程の説明図FIG. 5 is an explanatory view of a process of adhering the inner wall.
【図６】内皮部と外皮の接合状態を示す説明図FIG. 6 is an explanatory view showing a joined state of an inner skin portion and an outer skin.
１…外皮、２…内部壁、２ａ…内皮部、２ｂ…隔壁部、
４…暖気室、１０…アウタースキンツール型、１１…ア
ウタースキンプリプレグ、１２…シリコーンブロック、
１３…整流フィンプリプレグ、１４…真空バッグ、Ｆ…
整流フィン、Ｔ…暖気通路。1 ... outer skin, 2 ... inner wall, 2a ... inner skin part, 2b ... partition part,
4 ... Warm air chamber, 10 ... Outer skin tool type, 11 ... Outer skin prepreg, 12 ... Silicone block,
13 ... rectifying fin prepreg, 14 ... vacuum bag, F ...
Straightening fins, T ... Warm air passage.
フロントページの続き (72)発明者 中山 晴夫 埼玉県和光市中央１丁目４番１号 株式 会社本田技術研究所内 (72)発明者 白石 浩司 埼玉県和光市中央１丁目４番１号 株式 会社本田技術研究所内 (72)発明者 松本 敬三 埼玉県和光市中央１丁目４番１号 株式 会社本田技術研究所内 (56)参考文献 特開 昭62−157898（ＪＰ，Ａ) 特開 昭61−160395（ＪＰ，Ａ) 特開 平３−272829（ＪＰ，Ａ) (58)調査した分野(Int.Cl.7，ＤＢ名) B64D 15/04 B64C 3/28 Front page continued (72) Inventor Haruo Nakayama 1-4-1 Chuo, Wako-shi, Saitama Prefectural Technology Research Institute (72) Inventor Koji Shiraishi 1-4-1 Chuo, Wako-shi, Saitama Honda Motor Co., Ltd. In the laboratory (72) Inventor Keizo Matsumoto 1-4-1 Chuo, Wako-shi, Saitama Incorporated in Honda R & D Co., Ltd. (56) References JP-A-62-157898 (JP, A) JP-A-61-160395 (JP) , A) JP-A-3-272829 (JP, A) (58) Fields investigated (Int.Cl. 7 , DB name) B64D 15/04 B64C 3/28
Claims (4)
部空間部を暖気室として形成し、この暖気室に複数の整
流壁で仕切られる複数の暖気通路を形成するようにした
航空機の前縁構造において、前記外皮と整流壁は、夫々
の強化繊維樹脂を一体に接合硬化させたものであり、前
記内皮と隔壁は、予め強化繊維樹脂から一部品として成
形したものを所定部位に接着固定したものであり、前記
強化繊維樹脂がビスマレイミド変性樹脂或いはポリイミ
ド系樹脂からなることを特徴とする航空機の前縁構造。1. An aircraft in which an inner space portion of a front edge surrounded by an outer skin, an inner skin and a partition wall is formed as a warm air chamber, and a plurality of warm air passages partitioned by a plurality of straightening walls are formed in the warm air chamber. In the leading edge structure, the outer skin and the straightening wall are obtained by integrally bonding and hardening the respective reinforcing fiber resins, and the inner skin and the partition wall are molded in advance from the reinforcing fiber resin as a single component and bonded to a predetermined portion. der is, the thing that fixed
Reinforcing fiber resin is bismaleimide modified resin or polyimid
Leading edge structure of an aircraft, characterized in Rukoto such a de-based resin.
内部空間部を暖気室として形成し、この暖気室に複数の
整流壁で仕切られる複数の暖気通路を形成するようにし
た航空機の前縁構造において、前記外皮と整流壁は、夫
々の強化繊維樹脂を一体に接合硬化させたものであり、
前記内皮と隔壁部は、予め強化繊維樹脂から一部品とし
て成形したものを所定部位に接着固定したものであり、
前記内皮の後方には後方隔壁が設けられ、この後方隔壁
と前記隔壁部によって暖気排出室が区画形成され、この
暖気排出室に入った暖気が翼端の排出口から大気に放出
されることを特徴とする航空機の前縁構造。2. A form an internal space of the front edge portion surrounded by outer skin and endothelium and the partition section as the warm chamber and to form a plurality of warm air passages partitioned by a plurality of rectification walls on the warm chamber aircraft In the leading edge structure of, the outer skin and the rectifying wall are obtained by integrally bonding and hardening the respective reinforcing fiber resins,
The endothelium and the partition wall portion is state, and are those those molded bonded and fixed to a predetermined portion as one piece from a previously reinforced fiber resin,
A rear septum is provided behind the inner skin, and the rear septum is
And the partition wall section forms a warm air discharge chamber.
Warm air that has entered the warm air discharge chamber is released to the atmosphere from the outlet of the wing tip.
Is the leading edge structure of an aircraft, characterized in Rukoto.
部空間部を暖気室として形成し、この暖気室に複数の整
流壁で仕切られる複数の暖気通路を形成するようにした
航空機の前縁部の製造方法において、前縁部の形状を模
した工具に外皮となる未硬化状態の強化繊維樹脂製の積
層体を位置決めする工程と、この積層体の内側面上に、
前記暖気通路を形成するための形状保持具と、この形状
保持具の端部に配置された断面Ｌ字型の未硬化状態の強
化繊維樹脂の整流壁を複数並べて位置決めする工程と、
この上を真空バッグで覆って内部を真空引きし加熱・加
圧することで前記整流壁の一端側を前記積層体に一体に
接合硬化させる工程と、前記形状保持具を引き出す工程
と、予め強化繊維樹脂から一部品として成形した内皮と
隔壁を前記整流壁の他端側に当接させて所定部に接着固
定する工程からなることを特徴とする航空機の前縁部の
製造方法。3. An aircraft in which an inner space portion of a front edge surrounded by an outer skin, an inner skin and a partition wall is formed as a warm air chamber, and a plurality of warm air passages partitioned by a plurality of straightening walls are formed in the warm air chamber. In the manufacturing method of the front edge portion, a step of positioning a laminate made of reinforced fiber resin in an uncured state to be the outer skin on a tool simulating the shape of the front edge portion, and on the inner surface of the laminate,
A shape-retaining tool for forming the warm air passage, and a step of arranging a plurality of rectifying walls of an uncured reinforced fiber resin having an L-shaped cross-section and arranged at the end of the shape-retaining tool side by side.
Covering this with a vacuum bag, vacuuming the inside to heat and pressurize to integrally bond and harden one end side of the flow straightening wall to the laminate, and pulling out the shape retainer; A method for manufacturing a front edge portion of an aircraft, which comprises a step of bringing an inner skin and a partition formed of resin as one part into contact with the other end side of the rectifying wall and adhering and fixing the same to a predetermined portion.
法において、前記形状保持具はシリコーンゴムであるこ
とを特徴とする航空機の前縁部の製造方法。4. The method for manufacturing a front edge portion of an aircraft according to claim 3 , wherein the shape retainer is made of silicone rubber.
Priority Applications (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP22826495A JP3529911B2 (en) | 1995-09-05 | 1995-09-05 | Aircraft leading edge structure and method for manufacturing leading edge |
| US08/711,678 US5807454A (en) | 1995-09-05 | 1996-09-04 | Method of maufacturing a leading edge structure for aircraft |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP22826495A JP3529911B2 (en) | 1995-09-05 | 1995-09-05 | Aircraft leading edge structure and method for manufacturing leading edge |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPH0971299A JPH0971299A (en) | 1997-03-18 |
| JP3529911B2 true JP3529911B2 (en) | 2004-05-24 |
Family
ID=16873756
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP22826495A Expired - Lifetime JP3529911B2 (en) | 1995-09-05 | 1995-09-05 | Aircraft leading edge structure and method for manufacturing leading edge |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JP3529911B2 (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2011111405A1 (en) * | 2010-03-08 | 2011-09-15 | 三菱重工業株式会社 | Deicing device for leading edge of wing of aircraft, and aircraft main wing |
| WO2012029782A1 (en) * | 2010-08-30 | 2012-03-08 | 三菱重工業株式会社 | Aircraft ice protection system and aircraft provided with same |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110550182A (en) * | 2019-09-19 | 2019-12-10 | 中国航空工业集团公司西安飞机设计研究所 | Leading edge structure of airplane |
-
1995
- 1995-09-05 JP JP22826495A patent/JP3529911B2/en not_active Expired - Lifetime
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2011111405A1 (en) * | 2010-03-08 | 2011-09-15 | 三菱重工業株式会社 | Deicing device for leading edge of wing of aircraft, and aircraft main wing |
| JP2011183922A (en) * | 2010-03-08 | 2011-09-22 | Mitsubishi Heavy Ind Ltd | Anti-icing and deicing device at wing leading edge part in aircraft and main wing of aircraft |
| WO2012029782A1 (en) * | 2010-08-30 | 2012-03-08 | 三菱重工業株式会社 | Aircraft ice protection system and aircraft provided with same |
| JP2012046151A (en) * | 2010-08-30 | 2012-03-08 | Mitsubishi Heavy Ind Ltd | Aircraft ice protection system and aircraft with the same |
| RU2529927C1 (en) * | 2010-08-30 | 2014-10-10 | Мицубиси Хеви Индастрис, Лтд. | Aircraft de-icing system and aircraft with such system |
Also Published As
| Publication number | Publication date |
|---|---|
| JPH0971299A (en) | 1997-03-18 |
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