JPH08104296A - Main wing of supersonic aircraft - Google Patents
Main wing of supersonic aircraftInfo
- Publication number
- JPH08104296A JPH08104296A JP6266329A JP26632994A JPH08104296A JP H08104296 A JPH08104296 A JP H08104296A JP 6266329 A JP6266329 A JP 6266329A JP 26632994 A JP26632994 A JP 26632994A JP H08104296 A JPH08104296 A JP H08104296A
- Authority
- JP
- Japan
- Prior art keywords
- wing
- winglet
- drag
- chord length
- supersonic aircraft
- 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.)
- Granted
Links
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C23/00—Influencing air flow over aircraft surfaces, not otherwise provided for
- B64C23/06—Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices
- B64C23/065—Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices at the wing tips
- B64C23/069—Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices at the wing tips using one or more wing tip airfoil devices, e.g. winglets, splines, wing tip fences or raked wingtips
- B64C23/072—Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices at the wing tips using one or more wing tip airfoil devices, e.g. winglets, splines, wing tip fences or raked wingtips the wing tip airfoil devices being moveable in their entirety
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C23/00—Influencing air flow over aircraft surfaces, not otherwise provided for
- B64C23/06—Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices
- B64C23/065—Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices at the wing tips
- B64C23/069—Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices at the wing tips using one or more wing tip airfoil devices, e.g. winglets, splines, wing tip fences or raked wingtips
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/10—Drag reduction
Landscapes
- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Toys (AREA)
Abstract
Description
【０００１】[0001]
【産業上の利用分野】本発明は、超音速航空機の主翼に
係り、特に翼端部にウイングレットを備えた主翼に関す
るものである。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a wing of a supersonic aircraft, and more particularly to a wing having a winglet at a wing tip.
【０００２】[0002]
【従来の技術】一般に、航空機の抗力は、機体の揚力発
生に伴う抗力（揚力依存抗力）と、揚力に無関係な抗力
（零揚力抗力）に分けられる。亜音速航空機の場合は、
零揚力抗力は空気の粘性に基づく摩擦抗力、揚力依存抗
力は翼後方に吐き出される渦に基づく渦抗力から構成さ
れる。2. Description of the Related Art In general, the drag of an aircraft is divided into a drag associated with the generation of lift of the airframe (lift-dependent drag) and a drag irrelevant to lift (zero lift drag). For subsonic aircraft,
The zero lift drag consists of a friction drag based on the viscosity of air, and the lift dependent drag consists of a vortex drag based on a vortex discharged toward the rear of the blade.
【０００３】一方、超音速航空機の場合は、零揚力抗力
は摩擦抗力の他に、機体の体積に基づく造波抗力が加わ
り、また揚力依存抗力にも渦抗力の他に揚力発生に伴う
造波抗力が加わる。On the other hand, in the case of a supersonic aircraft, in addition to the frictional drag force, the zero drag force is added with a wave drag force based on the volume of the airframe, and the lift force dependent drag force is generated in addition to the vortex drag force due to the lift force generation. Drag is added.
【０００４】一般に、亜音速の翼理論によると、揚力依
存抗力換言すれば渦抗力を低減するためには主翼を横方
向（流れに直角な方向）に長い翼とすることが有効であ
る。従って、空力的には限りなく横に細長い主翼が最も
効率の良い翼となる。しかし、実際にはその構造的な問
題により、翼幅には制限が与えられる。In general, according to the subsonic wing theory, it is effective to make the main wing long in the lateral direction (direction perpendicular to the flow) in order to reduce lift-dependent drag, in other words, eddy drag. Therefore, the aerodynamically long and slender main wing becomes the most efficient wing. However, in practice, the structural issues limit the span.
【０００５】そこで、通常の亜音速旅客機では、さらに
渦抗力を低減する方法の１つとして、翼端に小翼を取り
付け、それを上方側或いは下方側に折り曲げて、翼幅の
延長を軽減し、これにより翼根部の曲げモーメントに関
する構造的な制限を満たしつつ、空力的には翼幅延長と
同等な効果を生み出すウイングレットが開発され、既に
実機にも採用されている（先行技術文献として、航空機
技術水準の向上に関する研究調査（１９７９年３月）Ｎ
ｏ．４小型機のウイングレット方式適用に関する研究
（Ｐ．１〜９）社団法人 日本航空宇宙工業会、流体力
学ハンドブック（Ｐ．２４８〜２４９）日本流体力学会
編 丸善株式会社、応用空気力学「次世代航空機をめざ
して」（Ｐ．３６〜３９）相原康彦・森下脱生著 東京
大学出版会等がある）。[0005] Therefore, in a conventional subsonic airliner, as one of the methods for further reducing the vortex drag, a small wing is attached to the wing tip and bent upward or downward to reduce the extension of the wing width. A winglet has been developed that satisfies the structural limitations on the bending moment of the wing root, and produces an aerodynamic effect equivalent to the wing span extension. Research Survey on Improvement of Technology Level (March 1979) N
o. Research on application of winglet method to 4 small aircraft (P.1-9) Japan Aerospace Industry Association, Fluid Dynamics Handbook (P.248-249) Japan Fluid Dynamics Society Maruzen Co., Ltd. (P. 36-39), written by Yasuhiko Aihara and Kasei Morishita.
【０００６】尚、ここで注意すべき点は、ウイングレッ
トは空力，構造の両面を考慮して、その有効性が発揮さ
れるものであるという点である。つまり、空力的観点だ
けでは翼端延長に勝るものはないからである。It should be noted that the winglet is effective in consideration of both aerodynamics and structure. In other words, there is nothing better than the wing tip extension from the aerodynamic point of view alone.
【０００７】然し乍ら、超音速航空機の場合には、新た
に造波抗力が加わるため、このウイングレット付き主翼
形状に大幅な変更が余儀なくされる。なぜなら、一般
に、この造波抗力は超音速理論によると、主翼を縦方向
（流れ方向）に長くするほど低減されることが判ってい
るからである。即ち、超音速航空機の揚力依存抗力を低
減するためには、その主翼は横に長く、且つ縦にも長
い、相矛盾する要求を満たす工夫が必要となる。However, in the case of a supersonic aircraft, a new wave drag is added, so that the shape of the winglet-equipped wing has to be greatly changed. This is because, generally, according to the supersonic theory, it is known that the wave drag decreases as the length of the main wing increases in the longitudinal direction (flow direction). In other words, in order to reduce the lift-dependent drag of a supersonic aircraft, it is necessary to devise a device that satisfies contradictory requirements because its main wing is long horizontally and vertically.
【０００８】このようなことから、一般の設計技術で
は、超音速航空機の主翼には、先ず亜音速航空機に比べ
て縦に細長い翼を採用し、横に長くという要求を犠牲に
するのが普通である。但し、その平面形の工夫によって
は、渦抗力をある程度低減することも可能で、これまで
のＮＡＳＡを始めとする研究から、やじりに似たアロー
翼と言われる平面形が有望視されている。しかし、この
翼は外翼部だけを取り出すと、その部分は横に長い翼と
等価になるため、構造的には外翼部付け根の曲げモーメ
ントがきつくなる傾向がある。そのため、さらに翼端延
長を行うことはできない。[0008] For this reason, in general design technology, the main wing of a supersonic aircraft generally employs a vertically long and narrow wing as compared to a subsonic aircraft, at the cost of sacrificing the requirement of being long horizontally. It is. However, it is possible to reduce the vortex drag force to some extent by devising the plane shape, and from the studies up to now, such as NASA, a plane shape called an arrow wing similar to a whirlpool is considered promising. However, if only the outer wing portion is taken out of this wing, that portion becomes equivalent to a laterally long wing, so that the bending moment of the root of the outer wing portion tends to be structurally strong. Therefore, further wing tip extension cannot be performed.
【０００９】[0009]
【発明が解決しようとする課題】そこで本発明は、超音
速航空機に特有の縦に細長い翼の抗力低減をさらに進め
るために、翼端延長せずに、翼端延長同等の渦抗力低減
効果を生み出せるウイングレットを備えた超音速航空機
の主翼を提供しようとするものである。SUMMARY OF THE INVENTION Therefore, in order to further reduce the drag of a vertically elongated wing peculiar to a supersonic aircraft, the present invention provides a vortex drag reduction effect equivalent to the wing tip extension without extending the wing tip. The aim is to provide a supersonic aircraft wing with a winglet that can be produced.
【００１０】[0010]
【課題を解決するための手段】上記課題を解決するため
の本発明の超音速航空機の主翼は、マッハ１〜５の超音
速航空機の主翼の翼端部に後縁を一致させて非平面的に
ウイングレットが取り付けられていて、このウイングレ
ットの平面形は低アスペクト比の台形タイプとなし、前
縁後退角ΛLE. winglet をΛLE. winglet ≧ΛM ≡ COS
-1（１／Ｍ）とし、キャント角φを９０°＞φ≧７０の
範囲とし、取付角θを０〜１°とし、ウイングレットの
翼根部翼弦長Ｃr.w を主翼翼端部の翼弦長Ｃｔより小さ
くその半分より大きいＣｔ≧Ｃr.w ≧Ｃｔ／２の範囲と
し、ウイングレットの翼根部翼弦長Ｃr.w と翼端翼弦長
Ｃt.w との比λを 0.5≧λ≧0.2 の範囲とし、ウイング
レットに捩り分布を設けない形状としたことを特徴とす
るものである。According to the present invention, there is provided a supersonic aircraft having a non-planar wing having a trailing edge coinciding with a wing tip of a supersonic aircraft of Mach 1-5. A winglet is attached to the winglet. The flat shape of this winglet is a low aspect ratio trapezoid type, and the leading edge receding angle Λ LE . Winglet Λ LE . Winglet ≧ Λ M ≡ COS
-1 (1 / M), the cant angle φ is in the range of 90 °> φ ≧ 70, the mounting angle θ is 0 to 1 °, and the wing root chord length Cr.w of the wing is the wing tip of the wing. Ct ≧ Cr.w ≧ Ct / 2, which is smaller than the chord length Ct and larger than half thereof, and the ratio λ between the wing root chord length Cr.w and the wing tip chord length Ct.w of the winglet is 0.5 ≧ λ ≧ It is characterized in that it has a shape in which the winglet has no torsional distribution in the range of 0.2.
【００１１】かかる構造の超音速航空機の主翼は、前記
ウイングレットのキャント角φを可変可能とし、超音速
飛行時９０°＞φ≧７０°とし、亜音速飛行時φ＝０°
としたものもある。主翼自体の形状は、デルタ翼、アロ
ー翼でも良いが、クランクト・アロー翼が好ましい。The main wing of the supersonic aircraft having such a structure can change the cant angle φ of the winglets, 90 °> φ ≧ 70 ° during supersonic flight, and φ = 0 ° during subsonic flight.
There is also one. The shape of the main wing itself may be a delta wing or an arrow wing, but a cranked arrow wing is preferable.
【００１２】本発明の超音速航空機の主翼に於いて、主
翼の翼端部に後縁を一致させて非平面的にウイングレッ
トを取り付けている理由は、翼端渦を拡散し、フラップ
無しの主翼の誘導抵抗を減少するためである。In the main wing of the supersonic aircraft of the present invention, the reason why the winglet is mounted non-planarly with the trailing edge coincident with the wing tip of the wing is that the wing tip vortex is diffused and the winglet without flap is used. This is to reduce the induced resistance of the.
【００１３】また、このウイングレットの平面形を低ア
スペクト比の台形タイプとした理由は、まず造波抗力を
低減すること、次にもしクランクト・アロー翼タイプと
した場合、構造的に厳しくなるからである。The reason why the plane shape of the winglet is a trapezoidal type having a low aspect ratio is that firstly, the wave drag is reduced, and if a cranked arrow blade type is used, the structure becomes severer. is there.
【００１４】また、このウイングレットの前縁後退角Λ
LE. winglet をΛLE. winglet ≧ΛM ≡ COS-1（１／
Ｍ）とした理由は、前縁をマッハ円錐の内側に入れるこ
とで、局所的には前縁に垂直方向の流入速度が亜音速に
なり、その結果前縁での衝撃波の発生が抑制され、造波
抗力の増加を抑えることができるためである。Further, the leading edge receding angle こ の of this winglet
LE . Winglet Λ LE . Winglet ≧ Λ M ≡ COS -1 (1 /
The reason for M) is that by inserting the leading edge inside the Mach cone, the inflow velocity in the direction perpendicular to the leading edge locally becomes subsonic, and as a result, the generation of shock waves at the leading edge is suppressed, This is because an increase in wave-making drag can be suppressed.
【００１５】また、このウイングレットのキャット角φ
を９０°＞φ≧７０の範囲とした理由は、超音速航空機
でのウイングレット部分の迎角を小さく抑えることによ
って揚力に伴う造波抗力の増加を抑えるためである。The winglet has a cat angle φ.
Is set to the range of 90 °> φ ≧ 70 in order to suppress the increase in wave drag due to lift by suppressing the angle of attack of the winglet portion in a supersonic aircraft.
【００１６】また、ウイングレットの取付角θを０〜１
°とした理由は、前面投影面積を少なくし、超音速流へ
の擾乱をできるだけ小さくすることで造波抗力の増加を
緩和するためである。Also, the mounting angle θ of the winglet is 0 to 1
The reason for making ° is to reduce the frontal projected area and minimize disturbance to the supersonic flow to mitigate the increase in wave drag.
【００１７】また、ウイングレットの翼根部翼弦長Ｃr.
w を主翼翼端部の翼弦長Ｃｔより小さくその半分より大
きいＣｔ≧Ｃr.w ≧Ｃｔ／２の範囲とした理由は、まず
上限についてはウイングレットの取り付け部分の確保の
観点からであり、次に下限についてはウイングレット効
果を最大限に生かす観点からあまり小さいウイングレッ
トでは不適当となるためである。The wing root chord length of the winglet, Cr.
The reason why w is set in the range of Ct ≧ Cr.w ≧ Ct / 2 which is smaller than the chord length Ct of the main wing tip and larger than half thereof is first from the viewpoint of securing the mounting portion of the winglet. This is because the lower limit is unsuitable for a winglet that is too small from the viewpoint of maximizing the winglet effect.
【００１８】また、ウイングレットの翼根部翼弦長Ｃr.
w と翼端翼弦長Ｃt.w との比λを0.5 ≧λ≧0.2 の範囲
とした理由は、ウイングレットの前縁後退角に関する条
件を満たし易くし、かつ構造的にも厳しくならないよう
にするためである。The wing root chord length of the winglet, Cr.
The reason that the ratio λ between the w and the tip chord length Ct.w is in the range of 0.5 ≧ λ ≧ 0.2 is to make it easier to satisfy the condition regarding the sweepback angle of the leading edge of the winglet and to prevent the structure from becoming severe. That's why.
【００１９】また、ウイングレットに捩り分布を設けな
い形状とした理由は、ウイングレットの構造上の複雑さ
を緩和するためである。The reason why the winglets are not provided with a twist distribution is to reduce the structural complexity of the winglets.
【００２０】然してまた、本発明の超音速航空機の主翼
に於いて、ウイングレットのキャント角φを可変可能と
し、超音速飛行時９０°＞φ≧７０°とし、亜音速飛行
時φ＝０°とした理由は、超音速航空機といえども離着
陸時と陸地上空に於いては、亜音速飛行を行うので、全
体の経済性を考慮した場合、亜音速飛行時の空力特性を
向上させるためである。即ち、亜音速飛行時では、揚力
依存抗力は渦抗力しか存在しないので、渦抗力を低減す
る最も有効な方法は翼幅を伸ばすことである。そのた
め、亜音速飛行時にはウイングレットを倒して、翼端延
長状態にすることで、渦抗力低減効果を取り出すように
している。However, in the main wing of the supersonic aircraft of the present invention, the winglet cant angle φ can be varied so that 90 °> φ ≧ 70 ° at the time of supersonic flight and φ = 0 ° at the time of subsonic flight. The reason for this is that even a supersonic aircraft performs subsonic flight during take-off and landing and over land and ground, so that considering the overall economics, the aerodynamic characteristics during subsonic flight are improved. That is, during subsonic flight, the only effective lift-dependent drag is vortex drag, and the most effective way to reduce vortex drag is to increase the wing span. For this reason, during subsonic flight, the winglets are defeated to extend the wing tip, thereby extracting the eddy drag reduction effect.
【００２１】[0021]
【作用】上記のように構成された本発明の超音速航空機
の主翼は、翼端延長せずに、翼端延長同等に渦抗力を低
減でき、超音速航空機特有の縦に細長い翼の抗力低減を
進めることができる。即ち、超音速航空機の場合、造波
抗力の存在のため翼幅を際限なく大きくすれば良いとは
限らず、最適な翼幅が予想される。このことは、純粋に
空力的観点だけからも翼端にウイングレットを付加した
ことにより、単に翼幅方向に延長した場合よりも有効で
あり、しかもウイングレットの形状を前述のように限定
したことにより、翼端渦が効率良く低減されると共に拡
散が大きくなり、渦抗力が低減される。従って、誘導抗
力が減少するので揚抗比を高く改善することができる。
また、構造的観点からも超音速航空機の最適な主翼とい
われるアロー翼にとっては、外翼部の構造的に厳しい翼
端延長を回避できる。そして、ウイングレットのキャン
ト角φを可変可能とし、超音速飛行時９０°＞φ≧７０
°とし、亜音速飛行時φ＝０°とした本発明の他の超音
速航空機の主翼は、構造重量の増加をさほど招くことな
く、離着陸時と陸地上空に於ける亜音速飛行の空力特性
が向上する。The main wing of the supersonic aircraft of the present invention configured as described above can reduce the vortex drag equivalent to the wing tip extension without extending the wing tip, thereby reducing the drag of the vertically elongated wing peculiar to the supersonic aircraft. Can proceed. That is, in the case of a supersonic aircraft, it is not always necessary to increase the wing width indefinitely due to the existence of the wave drag, and the optimum wing width is expected. This is more effective than simply extending in the wing span direction by adding a winglet to the wing tip from a purely aerodynamic point of view, and by limiting the shape of the winglet as described above, The tip vortex is efficiently reduced, the diffusion is increased, and the vortex drag is reduced. Therefore, the induced drag decreases, so that the lift-drag ratio can be improved.
Also, from the structural point of view, the structurally severe wing tip extension of the outer wing can be avoided for the arrow wing, which is said to be the optimal main wing of the supersonic aircraft. The cant angle φ of the winglet is variable, and at supersonic flight 90 °> φ ≧ 70
And the wing of another supersonic aircraft of the present invention in which subsonic flight φ = 0 ° has aerodynamic characteristics of subsonic flight during takeoff and landing and over land without causing a significant increase in structural weight. improves.
【００２２】[0022]
【実施例】本発明の超音速航空機の主翼の実施例を図に
よって説明する。図１はマッハ１〜５の超音速旅客機
（ＳＳＴ）の主翼で、超音速巡航での揚抗比を極限まで
高めた形態としてのクランクト・アロー翼１を示す。こ
のクランクト・アロー翼１の翼端部に図２のａ，ｂ，ｃ
に示すように後縁を一致させて非平面的にウイングレッ
ト２が取り付けられている。このウイングレット２の平
面形は低アスペクト比（ＡＲ）１〜２、本例の場合ＡＲ
＝１．５の台形タイプであり、前縁後退角がΛLE. wing
let ≧９０−（マッハ角）≡ COS-1（１／Ｍ）、本例の
場合５５度とし、キャント角φはウイングレット２の非
平面性を示す角度として、通常の主翼の上反角及び下反
角に対応（φ＝０°はウイングレット２を倒した状態、
φ＝９０°は上方に直角に曲げた状態）するもので、こ
こでは９０°＞φ≧７０°の範囲、本例では８０°に定
めている。ウイングレット２の取付角θは０〜１°、本
例の場合１°とし、ウイングレット２の翼根部翼弦長Ｃ
r.w を主翼１の翼端部の翼弦長Ｃｔより小さく、その半
分より大きいＣｔ≧Ｃr.w ≧Ｃｔ／２の範囲、本例の場
合Ｃr.w ＝Ｃｔ／２とし、ウイングレット２の翼根部翼
弦長Ｃr.w と翼端部翼弦長Ｃt.w との比λ＝Ｃt.w ／Ｃ
r.w を 0.5≧λ≧0.2 の範囲、本例ではλ＝0.4 とし、
そしてウイングレット２の捩り分布を設けない捩り無し
の形状とした。DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of the main wing of a supersonic aircraft according to the present invention will be described with reference to the drawings. FIG. 1 shows a cranked arrow wing 1 of a main wing of a supersonic airliner (SST) of Mach 1 to 5 in which the lift-drag ratio in supersonic cruising is maximized. 2a, 2b and 2c are attached to the tip of the cranked arrow blade 1.
The winglets 2 are mounted non-planarly with their trailing edges aligned as shown in FIG. The winglet 2 has a low aspect ratio (AR) of 1 to 2, in the case of this example,
= 1.5 trapezoidal type, leading edge receding angle Λ LE .wing
let ≧ 90− (Mach angle) ≡COS −1 (1 / M), 55 degrees in this example, and the cant angle φ is an angle indicating the non-planarity of the winglet 2 and is the dihedral angle and the lower angle of a normal wing Corresponds to the opposite angle (φ = 0 ° with winglet 2 down,
φ = 90 ° is bent upward at a right angle), and is set in the range of 90 °> φ ≧ 70 °, in this example, 80 °. The installation angle θ of the winglet 2 is 0 to 1 °, and in this example, 1 °, and the blade root chord length C of the winglet 2 is set to 1 °.
The rw is smaller than the chord length Ct of the blade tip of the main wing 1 and is larger than half thereof in the range of Ct ≧ Cr.w ≧ Ct / 2, and in this example, Cr.w = Ct / 2, and the blade root of the winglet 2 is set. Ratio of chord length Cr.w to wing tip chord length Ct.w λ = Ct.w / C
rw is in the range of 0.5 ≧ λ ≧ 0.2, and in this example, λ = 0.4,
Then, the winglet 2 was formed into a non-twisted shape with no twist distribution.
【００２３】このように構成した実施例の超音速旅客機
のクランクト・アロー翼１は、翼端延長せずに、翼端に
ウイングレットを取り付けているので、翼端延長した場
合と同等に渦抗力を低減でき、超音速航空機特有の縦に
細長い翼の抗力低減を進めることができる。しかもウイ
ングレットの形状を実施例のように限定したことによ
り、翼端渦が効率良く低減されると共に拡散が大きくな
り、渦抗力が低減される。従って、誘導抗力が減少する
ので、揚抗比を高く改善することができる。また構造上
クランクト・アロー翼１にとって外翼部の厳しい翼端延
長を回避できる。Since the cranked arrow wing 1 of the supersonic passenger aircraft of the embodiment thus configured has the winglet attached to the wing tip without extending the wing tip, the vortex drag force is the same as when the wing tip is extended. It is possible to reduce the drag force of the vertically elongated slender wing peculiar to the supersonic aircraft. Moreover, by limiting the shape of the winglet as in the embodiment, the vortex on the blade tip is efficiently reduced, the diffusion is increased, and the vortex drag is reduced. Therefore, the induced drag decreases, and the lift-drag ratio can be improved. In addition, it is possible to avoid a severe wing tip extension of the outer wing portion for the cranked arrow wing 1 due to its structure.
【００２４】本発明の超音速航空機の主翼の他の実施例
は、前記実施例のクランクト・アロー翼１に於けるウイ
ングレット２のキャント角φを可変可能とし、超音速飛
行時図３のａに示すように９０°＞φ≧７０°とし、亜
音速飛行時図３のｂに示すようにφ＝０°としたもので
ある。Another embodiment of the main wing of the supersonic aircraft according to the present invention is shown in FIG. 3A in which the cant angle φ of the winglet 2 in the cranked arrow wing 1 of the above embodiment can be varied. 90 °> φ ≧ 70 ° as shown, and φ = 0 ° during subsonic flight as shown in FIG. 3B.
【００２５】このように構成した実施例の超音速旅客機
のクランクト・アロー翼１は、構造上翼端部でのウイン
グレット２の可変機構の付加にさほどの重量増加を招く
ことがなく、キャント角φを０°とすることにより離着
陸時と陸地上空に於ける亜音速飛行の空力特性が向上す
る。超音速飛行時にはキャント角φを９０°＞φ≧７０
°の範囲で適宜変化させることにより、適切に造波抗力
を低減し、誘導抗力も減少できる。The cranked arrow wing 1 of the supersonic airliner of the embodiment constructed in this way does not cause a significant increase in weight due to the addition of the variable mechanism of the winglet 2 at the wing end, and the cant angle φ Is set to 0 °, the aerodynamic characteristics of subsonic flight during take-off and landing and over land and ground are improved. During supersonic flight, the cant angle φ is 90 °> φ ≧ 70
By appropriately changing the angle within the range of °, the wave drag and the induced drag can be reduced appropriately.
【００２６】尚、上記実施例は超音速航空機の主翼が、
クランクト・アロー翼１であるが、これに限るものでは
なく、デルタ翼やアロー翼の場合でも略同様に造波抗力
が低減され、誘導抗力も減少する。In the above embodiment, the main wing of the supersonic aircraft is
The cranked arrow blade 1 is not limited to this, and the wave drag and the induced drag are reduced substantially similarly in the case of the delta blade and the arrow blade.
【００２７】[0027]
【発明の効果】以上の説明で判るように本発明の超音速
航空機の主翼は、超音速航空機特有の縦に細長い翼の抗
力低減をさらに進めるために翼端延長せずに特定の形状
にウイングレットを備えたので、翼端を延長したと同等
の渦抗力低減効果が得られ、従って、誘導抗力が減少す
るので、揚抗比を高く改善することができる。また、上
記ウイングレットを可変化した主翼にあっては、可変機
構の付加にさほどの重量増加を招くことなく、キャント
角φを０°とすることにより、翼端延長によって離着陸
時と陸地上空に於ける亜音速飛行の空力特性が向上す
る。As can be seen from the above description, the main wing of the supersonic aircraft of the present invention has a winglet having a specific shape without extending the wing tip in order to further reduce the drag of the vertically elongated slender wing peculiar to the supersonic aircraft. As a result, the vortex drag reduction effect equivalent to that when the blade tip is extended is obtained, and therefore the induced drag is reduced, so that the lift-drag ratio can be improved to a high level. In the case of the main wing with variable winglets, the cant angle φ is set to 0 ° without adding a variable mechanism to the weight, so that the wing tip can be extended for take-off and landing and for ground and ground. Aerodynamic characteristics of subsonic flight in
【図１】クランクト・アロー翼の超音速旅客機を示す平
面図である。FIG. 1 is a plan view showing a cranked arrow wing supersonic airliner.
【図２】図１のクランクト・アロー翼にウイングレット
を設けた本発明の超音速航空機の主翼の一実施例に於け
る要部を示すもので、ａは正面図、ｂは側面図、ｃは平
面図である。2 shows essential parts of an embodiment of a main wing of a supersonic aircraft of the present invention in which a winglet is provided on the cranked arrow wing of FIG. 1, a is a front view, b is a side view, and c is It is a top view.
【図３】可変可能なウイングレットを設けた本発明の超
音速航空機の主翼の他の実施例を示すもので、ａは超音
速飛行時の状態、ｂは亜音速飛行時の状態である。FIG. 3 shows another embodiment of the main wing of the supersonic aircraft of the present invention provided with a variable winglet, where a is a state during supersonic flight and b is a state during subsonic flight.
１ クランクト・アロー翼（主翼） ２ ウイングレット 1 Cranked Arrow Wing (Main Wing) 2 Winglet
Claims (2)
いて、該主翼の翼端部に後縁を一致させて非平面的にウ
イングレットが取り付けられ、該ウイングレットの平面
形は低アスペクト比の台形タイプとなし、前縁後退角Λ
LE. wingletをΛLE. winglet ≧ΛM ≡ COS-1（１／
Ｍ）とし、キャント角φを９０°＞φ≧７０の範囲と
し、取付角θを０〜１°とし、ウイングレットの翼根部
翼弦長Ｃr.w を主翼翼端部の翼弦長Ｃｔより小さくその
半分より大きいＣｔ≧Ｃr.w ≧Ｃｔ／２の範囲とし、ウ
イングレットの翼根部翼弦長Ｃr.w と翼端翼弦長Ｃt.w
との比λを 0.5≧λ≧0.2 の範囲とし、ウイングレット
に捩り分布を設けない形状としたことを特徴とする超音
速航空機の主翼。In a wing of a supersonic aircraft of Mach 1 to 5, a winglet is attached in a non-planar manner with a trailing edge coincident with a wing tip of the wing, and the winglet has a low aspect ratio. With or without trapezoidal type, leading edge receding angle.
LE . Winglet Λ LE . Winglet ≧ Λ M ≡ COS -1 (1 /
M), the cant angle φ is in the range of 90 °> φ ≧ 70, the mounting angle θ is 0 to 1 °, and the wing root chord length Cr.w of the winglet is smaller than the chord length Ct of the main wing tip. Ct ≧ Cr.w ≧ Ct / 2, which is larger than half of the range, and the wing root chord length Cr.w and the tip chord length Ct.w of the winglet
A wing of a supersonic aircraft characterized by having a ratio λ of 0.5 ≧ λ ≧ 0.2 and a winglet having no twist distribution.
いて、ウイングレットのキャント角φを可変可能とし、
超音速飛行時９０°＞φ≧７０°とし、亜音速飛行時φ
＝０°としたことを特徴とする超音速航空機の主翼。2. The wing of a supersonic aircraft according to claim 1, wherein the winglet cant angle φ is variable.
90 °> φ ≧ 70 ° for supersonic flight and φ for subsonic flight
The main wing of a supersonic aircraft, wherein = 0 °.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP6266329A JP2622670B2 (en) | 1994-10-05 | 1994-10-05 | Supersonic aircraft wing |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP6266329A JP2622670B2 (en) | 1994-10-05 | 1994-10-05 | Supersonic aircraft wing |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPH08104296A true JPH08104296A (en) | 1996-04-23 |
| JP2622670B2 JP2622670B2 (en) | 1997-06-18 |
Family
ID=17429422
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP6266329A Expired - Fee Related JP2622670B2 (en) | 1994-10-05 | 1994-10-05 | Supersonic aircraft wing |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JP2622670B2 (en) |
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2008540240A (en) * | 2005-05-19 | 2008-11-20 | エアバス・ドイチュラント・ゲーエムベーハー | Variable wing concept to reduce lateral loads, reduce lateral and longitudinal combined loads, and enhance the performance of moving means |
| JP2009501678A (en) * | 2005-07-21 | 2009-01-22 | ザ・ボーイング・カンパニー | Integrated wingtip extension for jet transport aircraft and other types of aircraft |
| JP2010280252A (en) * | 2009-06-02 | 2010-12-16 | Keiji Shigemiya | Vertical tail of aircraft |
| JP2011506160A (en) * | 2007-12-10 | 2011-03-03 | エアバス・オペレーションズ・ゲーエムベーハー | Small wing extension for reducing aircraft tip vortices |
| US8757555B2 (en) | 2005-05-19 | 2014-06-24 | Airbus Operations Gmbh | Concept of a variable winglet for lateral load reduction for combined lateral and vertical load reduction, and for improving the performance of means of locomotion |
| US9162755B2 (en) | 2009-12-01 | 2015-10-20 | Tamarack Aerospace Group, Inc. | Multiple controllable airflow modification devices |
| JP2016097863A (en) * | 2014-11-25 | 2016-05-30 | 直美 菊池 | Flight method for aircraft |
| WO2017118832A1 (en) * | 2016-01-05 | 2017-07-13 | Airbus Operations Limited | An aircraft wing with a moveable wing tip device for load alleviation |
| US9764825B2 (en) | 2009-12-01 | 2017-09-19 | Tamarack Aerospace Group, Inc. | Active winglet |
-
1994
- 1994-10-05 JP JP6266329A patent/JP2622670B2/en not_active Expired - Fee Related
Non-Patent Citations (1)
| Title |
|---|
| THEORY OF WING SECTIONS\\\9800563=1959 * |
Cited By (15)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2008540240A (en) * | 2005-05-19 | 2008-11-20 | エアバス・ドイチュラント・ゲーエムベーハー | Variable wing concept to reduce lateral loads, reduce lateral and longitudinal combined loads, and enhance the performance of moving means |
| US8757555B2 (en) | 2005-05-19 | 2014-06-24 | Airbus Operations Gmbh | Concept of a variable winglet for lateral load reduction for combined lateral and vertical load reduction, and for improving the performance of means of locomotion |
| JP2009501678A (en) * | 2005-07-21 | 2009-01-22 | ザ・ボーイング・カンパニー | Integrated wingtip extension for jet transport aircraft and other types of aircraft |
| JP2011506160A (en) * | 2007-12-10 | 2011-03-03 | エアバス・オペレーションズ・ゲーエムベーハー | Small wing extension for reducing aircraft tip vortices |
| JP2010280252A (en) * | 2009-06-02 | 2010-12-16 | Keiji Shigemiya | Vertical tail of aircraft |
| US9969487B2 (en) | 2009-12-01 | 2018-05-15 | Tamarack Aerospace Group, Inc. | Multiple controllable airflow modification devices |
| US9764825B2 (en) | 2009-12-01 | 2017-09-19 | Tamarack Aerospace Group, Inc. | Active winglet |
| US9162755B2 (en) | 2009-12-01 | 2015-10-20 | Tamarack Aerospace Group, Inc. | Multiple controllable airflow modification devices |
| US10486797B2 (en) | 2009-12-01 | 2019-11-26 | Tamarack Aerospace Group, Inc. | Active winglet |
| US11111006B2 (en) | 2009-12-01 | 2021-09-07 | Tamarack Aerospace Group, Inc. | Multiple controlloable airflow modification devices |
| US11912398B2 (en) | 2009-12-01 | 2024-02-27 | Tamarack Aerospace Group, Inc. | Multiple controllable airflow modification devices |
| JP2016097863A (en) * | 2014-11-25 | 2016-05-30 | 直美 菊池 | Flight method for aircraft |
| WO2017118832A1 (en) * | 2016-01-05 | 2017-07-13 | Airbus Operations Limited | An aircraft wing with a moveable wing tip device for load alleviation |
| EP3862261A1 (en) * | 2016-01-05 | 2021-08-11 | Airbus Operations Limited | An aircraft wing with a moveable wing tip device for load alleviation |
| US11203410B2 (en) | 2016-01-05 | 2021-12-21 | Airbus Operations Limited | Aircraft wing with a moveable wing tip device for load alleviation |
Also Published As
| Publication number | Publication date |
|---|---|
| JP2622670B2 (en) | 1997-06-18 |
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