JPH0774036B2 - Aircraft forward wing - Google Patents
Aircraft forward wingInfo
- Publication number
- JPH0774036B2 JPH0774036B2 JP62156480A JP15648087A JPH0774036B2 JP H0774036 B2 JPH0774036 B2 JP H0774036B2 JP 62156480 A JP62156480 A JP 62156480A JP 15648087 A JP15648087 A JP 15648087A JP H0774036 B2 JPH0774036 B2 JP H0774036B2
- Authority
- JP
- Japan
- Prior art keywords
- shock wave
- wing
- blade
- trailing edge
- forward wing
- 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
Landscapes
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Description
【発明の詳細な説明】 〔産業上の利用分野〕 本発明は航空機（無人機を含む）の前進翼に関する。The present invention relates to a forward wing of an aircraft (including an unmanned aerial vehicle).
従来の航空機の前進翼101の翼根形状は第２図に示すよ
うに一定の変化率で即ち直線状にその翼弦長Ｓが変化す
るものが一般的である。As shown in FIG. 2, the wing root shape of a conventional aircraft forward wing 101 generally has a constant rate of change, that is, its chord length S changes linearly.
上記従来の前進翼には次の不具合があった。 The conventional forward wing described above has the following problems.
すなわち、遷音速、超音速飛行時には前進翼では、外翼
部における流れの圧縮効果と内翼における流れの境界条
件の効果とが重なりあい、内翼部において通常、前進角
の浅い（あるいは無い）衝撃波103を生ずる。この衝撃
波の強さは衝撃波の前進角が浅いために比較的強いもの
となるのが一般的であり、これは、 （１） 衝撃波による運動量損失からくる造波抵抗の増
大 （２） 強い衝撃波による圧力上昇によって引き起こさ
れる境界層の衝撃波剥離105（Shock−Induced Separati
on）と、それに伴なう抵抗増大、揚力減少、バフエツト 等を引き起こし、航空機（無人機を含む）の性能に悪影
響をおよぼす。That is, in transonic and supersonic flight, in the forward wing, the effect of flow compression on the outer wing overlaps with the effect of boundary conditions of the flow on the inner wing, and the inner wing usually has a small forward angle (or none). A shock wave 103 is generated. The strength of this shock wave is generally relatively strong due to the shallow advancing angle of the shock wave. (1) Increase in wave-making resistance resulting from momentum loss due to shock wave (2) Due to strong shock wave Boundary layer shock wave separation caused by pressure rise 105 (Shock-Induced Separati
on) and the accompanying increase in resistance, decrease in lift, buffering, etc., adversely affecting the performance of the aircraft (including unmanned aerial vehicles).
本発明は上記問題点解決の手段として、機体中心に近づ
くに従って翼弦長が１以上の逓倍率で長くなる翼根後縁
を有することを特徴とする航空機前進翼を提供しようと
するものである。As a means for solving the above problems, the present invention is to provide an aircraft forward wing having a blade root trailing edge whose chord length increases at a multiplication factor of 1 or more as it approaches the center of the airframe. .
本発明は上記のように構成するので次の作用を有する。 Since the present invention is configured as described above, it has the following effects.
すなわち、翼根後縁部の翼弦長が機体中心に近づくに従
って１以上の逓倍率で長くなるので翼根部に生じる衝撃
波の前進角が加速度的に大きくなり、それに伴ない衝撃
波の強さはより弱いものとなる。これにより、 （１） 衝撃波による運動量損失の減少と、それに伴う
造波抵抗の減少 （２） 衝撃波による圧力上昇が小さくなることによる
境界層の衝撃波剥離の防止とそれに伴なう抵抗減少、揚
力増大、バフエツト防止 等の作用が果たされ、航空機（無人機を含む）の性能が
向上する。That is, as the chord length of the trailing edge of the blade root increases with a multiplication factor of 1 or more as it approaches the center of the airframe, the advancing angle of the shock wave generated at the blade root increases at an accelerating rate, and the strength of the shock wave accompanying it increases Will be weak. As a result, (1) the momentum loss due to the shock wave is reduced and the wave-making resistance accompanying it is reduced. (2) The shock wave separation in the boundary layer is prevented due to the decrease in the pressure rise due to the shock wave, and the accompanying decrease in resistance and increase in lift force. , Functions such as buffering prevention are performed, and the performance of aircraft (including unmanned aerial vehicles) is improved.
本発明の一実施例について第１図により説明する。 An embodiment of the present invention will be described with reference to FIG.
図は一枚の前進翼（前進角を有する翼）を上から見たも
ので、前進翼１の前縁２は前進角θを有しかつ、翼根後
縁４が機体中心に近づくに従って逓倍的（加速度的）に
長くなるように構成されている。The figure shows a single advancing wing (a wing with an advancing angle) seen from above. The leading edge 2 of the advancing wing 1 has an advancing angle θ, and the blade root trailing edge 4 is multiplied as it approaches the airframe center. It is configured to be longer (acceleratively).
従って、たとば翼弦長S2は同S1より、それが機体中心に
近づいた割合より長い関係にあり、衝撃波３はたとえば
従来の、カーブした、かつ機体中心線近傍では殆ど前進
角を持たない衝撃波103を適正に補正して、図のように
充分な前進角αを持つことになるので、流れの向きに垂
直な同一断面に音のエネルギーが蓄積されるという好ま
しくない状態が回避され、衝撃波３は充分に弱いものに
なる。従って従来のような衝撃波103による造波抵抗は
減少し、衝撃波剥離も抑制される。Therefore, the chord chord length S 2 is longer than that of S 1 and is closer to the center of the airframe, and the shock wave 3 has, for example, a conventional, curved, and almost advancing angle near the airframe centerline. Since the shock wave 103 which is not present is properly corrected to have a sufficient advancing angle α as shown in the figure, an unfavorable state in which sound energy is accumulated in the same cross section perpendicular to the flow direction is avoided. The shock wave 3 becomes weak enough. Therefore, the conventional wave-making resistance due to the shock wave 103 is reduced, and shock wave separation is also suppressed.
第３図は従来例の前進翼101と本発明の実施例の前進翼
１との翼根近傍断面における圧力分布等の比較図であ
る。前進翼１の翼根後縁は機体中心に近づくに従って逓
倍的に翼弦長が長くなるのでその分、従来例に比し、後
縁が延長される。図中ではその部分を『後縁延長』と呼
んでいる。FIG. 3 is a comparison diagram of a pressure distribution and the like in a blade root vicinity cross section of the conventional advancing blade 101 and the advancing blade 1 of the embodiment of the present invention. The blade root trailing edge of the advancing blade 1 has a blade chord length that gradually increases as it approaches the center of the airframe, and therefore the trailing edge is extended as compared with the conventional example. In the figure, that part is called "extension of trailing edge".
図より明らかなように衝撃波剥離の生じない前進翼１で
は翼上面圧力分布面積が大きくなり、大きな揚力を創出
していること等が分る。As is clear from the figure, it can be seen that in the advancing blade 1 in which shock wave separation does not occur, the pressure distribution area on the blade upper surface becomes large and a large lift is created.
なお、本発明の適用される前進翼は主翼に限定されるも
のではなく、水平尾翼，垂直尾翼，カナード等何れが対
象となってもよい。The forward wing to which the present invention is applied is not limited to the main wing, and any of a horizontal tail, a vertical tail, a canard, etc. may be used.
又、翼根後縁が逓倍的に長くなる範囲には逓倍率の定ま
ったもの、そうでないもの何れも含まれ、たとえば直線
状後縁から或る変曲点（大きな曲率の）をもって翼根後
縁が機体中心に向ってスタートし、スタート後は機体中
心に至る迄、逓倍率が１以上の曲線の場合も含むものと
する。その場合は後縁の巨視的平面形状は略近似折線と
なる。In addition, the range where the trailing edge of the blade root gradually becomes longer includes both those with a fixed multiplying factor and those with no such multiplying factor. For example, from a straight trailing edge to a certain inflection point (having a large curvature) The case where the edge starts toward the center of the aircraft and after the start reaches the center of the aircraft is also a curve with a multiplication factor of 1 or more. In that case, the macroscopic planar shape of the trailing edge is approximately a polygonal line.
この発明は次のような効果を有する。すなわち、翼根部
に生じる衝撃波の前進角が確保され、それに伴ない衝撃
波の強さが確実に弱いものとなることにより、 （１） 衝撃波による運動量損失の減少と、それに伴な
う造波抵抗の減少が果される。This invention has the following effects. In other words, the advancing angle of the shock wave generated at the blade root is secured, and the strength of the shock wave is surely weakened accordingly. (1) The decrease in momentum loss due to the shock wave and the accompanying wave-making resistance Reduction is achieved.
（２） 衝撃波による圧力上昇が小さくなることによる
境界層の衝撃波剥離の防止とそれに伴なう抵抗減少，揚
力増大，バフエツト防止等が果される。(2) Preventing shock wave separation in the boundary layer due to the decrease in pressure rise due to shock waves, and the accompanying reduction in resistance, increase in lift, and prevention of buffering.
第１図は本発明の一実施例の平面図、第２図は従来例の
平面図、第３図は従来例と実施例の前進翼の翼断面上の
流れパターンと翼断面上圧力分布の一例を比較して示し
た図（翼のある翼幅方向位置での断面に対応させて示
す）である。 １……前進翼、２……前縁、３……衝撃波、４……翼根
後縁。FIG. 1 is a plan view of an embodiment of the present invention, FIG. 2 is a plan view of a conventional example, and FIG. 3 is a flow pattern on the blade cross section of a forward blade of the conventional example and the pressure distribution on the blade cross section. It is the figure which compared and showed one example (it shows corresponding to the cross section in the wing width direction position with a wing). 1 ... Forward wing, 2 ... Leading edge, 3 ... Shock wave, 4 ... Blade root trailing edge.
Claims (1)
の逓倍率で長くなる翼根後縁を有することを特徴とする
航空機前進翼。1. An aircraft forward wing having a blade root trailing edge whose chord length increases with a multiplication factor of 1 or more as it approaches the center of the airframe.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP62156480A JPH0774036B2 (en) | 1987-06-25 | 1987-06-25 | Aircraft forward wing |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP62156480A JPH0774036B2 (en) | 1987-06-25 | 1987-06-25 | Aircraft forward wing |
Publications (3)
| Publication Number | Publication Date |
|---|---|
| JPS641696A JPS641696A (en) | 1989-01-06 |
| JPH011696A JPH011696A (en) | 1989-01-06 |
| JPH0774036B2 true JPH0774036B2 (en) | 1995-08-09 |
Family
ID=15628677
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP62156480A Expired - Lifetime JPH0774036B2 (en) | 1987-06-25 | 1987-06-25 | Aircraft forward wing |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH0774036B2 (en) |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS5938198A (en) * | 1982-08-25 | 1984-03-01 | 富士重工業株式会社 | Variable propelling plane |
-
1987
- 1987-06-25 JP JP62156480A patent/JPH0774036B2/en not_active Expired - Lifetime
Also Published As
| Publication number | Publication date |
|---|---|
| JPS641696A (en) | 1989-01-06 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| EXPY | Cancellation because of completion of term |