JPH05280495A - Fan moving blade - Google Patents
Fan moving bladeInfo
- Publication number
- JPH05280495A JPH05280495A JP7786692A JP7786692A JPH05280495A JP H05280495 A JPH05280495 A JP H05280495A JP 7786692 A JP7786692 A JP 7786692A JP 7786692 A JP7786692 A JP 7786692A JP H05280495 A JPH05280495 A JP H05280495A
- Authority
- JP
- Japan
- Prior art keywords
- blade
- long
- short
- blades
- pressure loss
- 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.)
- Withdrawn
Links
Abstract
Description
【０００１】[0001]
【産業上の利用分野】本発明は、ターボファンエンジン
の空気取入口に配設されたファン動翼に関し、詳しく
は、遷音速領域（亜音速から超音速までの領域）の入口
相対速度においても高効率に空気圧縮が可能なものに関
する。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a fan rotor blade arranged at an air intake of a turbofan engine, and more specifically, at an inlet relative velocity in a transonic region (region from subsonic speed to supersonic speed). The present invention relates to a device capable of highly efficient air compression.
【０００２】[0002]
【従来の技術】図１は、ターボファンエンジンの概略を
示すものであり、空気取入口１から取り入れられた流入
空気は、まずファン動翼２で圧縮される。そして、ファ
ン動翼２から送り出された圧縮空気の一部は、ファン空
気排出ダクト３を通過して直接エンジン外へバイパス推
力として噴出される。一方、ファン動翼２から送りださ
れた圧縮空気は、圧縮機４でさらに高圧圧縮され、その
後、燃焼室５に送られて供給燃料と混合されて燃焼工程
が行われる。そして、燃焼室５から排気された排気ガス
は、圧縮機タービン６及びファンタービン７を回転させ
るとともに、排気ダクト８よりコア推力として噴出され
る構造になっている。2. Description of the Related Art FIG. 1 shows an outline of a turbofan engine. Inflow air taken in from an air intake 1 is first compressed by a fan rotor blade 2. Then, a part of the compressed air sent out from the fan rotor blade 2 passes through the fan air exhaust duct 3 and is directly ejected outside the engine as a bypass thrust. On the other hand, the compressed air sent from the fan rotor blades 2 is further compressed by the compressor 4 to a higher pressure, and then sent to the combustion chamber 5 where it is mixed with the supplied fuel to perform a combustion process. Exhaust gas exhausted from the combustion chamber 5 rotates the compressor turbine 6 and the fan turbine 7, and is ejected from the exhaust duct 8 as a core thrust.
【０００３】そして、前述したファン動翼２は、ハブ
（根元）９ａ側から翼端９ｂに向かうにつれてねじれを
付けた複数枚の翼９…が、ファン回転軸１０に放射状に
植設されて構成されている。The above-described fan rotor blade 2 is formed by radially arranging a plurality of blades 9 ... Twisted from the hub (root) 9a side toward the blade tip 9b. Has been done.
【０００４】ここで、各翼９…は、図４に示すように、
翼中央部９ｃから翼端９ｂまでの形状が、同一の翼弦長
Ｃで、かつ同一の翼断面形状とされた翼が使用され、そ
れら翼９…が所定のピッチＳをあけて等間隔に配列され
て、所定のソリディティ値（ｃ／ｓ）を有する翼列１２
を形成している。なお、各翼９…は、ファン動翼２の矢
印Ａ方向の回転方向に対して、翼９の翼前縁１３が回転
方向の前方側、翼後縁１４が回転方向の後方側となるよ
うにねじりが付けられている。Here, each of the wings 9 ...
The blades having the same blade chord length C and the same blade cross-sectional shape from the blade central portion 9c to the blade tip 9b are used, and the blades 9 ... Are spaced at a predetermined pitch S at equal intervals. Cascade 12 arranged and having a predetermined solidity value (c / s)
Is formed. Each blade 9 is arranged so that the blade leading edge 13 of the blade 9 is on the front side in the rotational direction and the blade trailing edge 14 is on the rear side in the rotational direction with respect to the rotational direction of the fan rotor blade 2 in the direction of arrow A. Is twisted.
【０００５】そして、ファン動翼２が所定の周速度Ｕ1
で回転し、上流側１６から所定の入口絶対速度Ｖ1で空
気が取り入れられることにより、ファン動翼２は入口相
対速度Ｗ1の空気を圧縮するようになっている。Then, the fan rotor blade 2 has a predetermined peripheral velocity U 1
The fan rotor blade 2 compresses the air at the inlet relative velocity W 1 by rotating at the inlet side and taking in air from the upstream side 16 at a predetermined inlet absolute velocity V 1 .
【０００６】[0006]
【発明が解決しようとする課題】ところで、上記従来の
ファン動翼２においては、上流側１６から超音速の入口
相対速度Ｗ1で空気が流れ込むと、圧力損失の発生によ
り圧縮効率が低下してしまう。その要因としては、第１
には衝撃波自身の存在によるもの、第２には衝撃波の存
在によって翼面に境界層が急激に発達することが要因で
ある。In the conventional fan blade 2 described above, when air flows in from the upstream side 16 at the supersonic inlet relative velocity W 1 , the compression efficiency decreases due to the pressure loss. I will end up. The first factor is
Is due to the existence of the shock wave itself, and secondly is due to the rapid development of the boundary layer on the blade surface due to the existence of the shock wave.
【０００７】そして、入口相対速度Ｗ1がマッハ数１.６
程度までの領域では、上述した第２の要因により圧力損
失が圧倒的に高くなることが知られている。すなわち、
図４に示すように、マッハ数１.６の入口相対速度Ｗ1に
より、全ての翼９…の翼前縁１３に衝撃波１８が存在
し、この衝撃波１８より下流側１７の翼上面に急激に境
界層２０が発達する。そして、この発達した境界層２０
の剥離によりエントロピーの上昇、すなわち圧力損失が
発生してファン動翼２の圧縮効率が著しく低下してしま
う。The inlet relative velocity W 1 is Mach number 1.6.
It is known that the pressure loss becomes overwhelmingly high due to the above-mentioned second factor in a region up to a certain degree. That is,
As shown in FIG. 4, due to the inlet relative velocity W 1 of Mach number 1.6, shock waves 18 are present on the blade leading edges 13 of all the blades 9 ... The boundary layer 20 develops. And this developed boundary layer 20
Due to the separation, the entropy rises, that is, pressure loss occurs, and the compression efficiency of the fan rotor blade 2 is significantly reduced.
【０００８】本発明は、上記事情に鑑みてなされたもの
で、相対入口流速が超音速領域であっても、衝撃波の存
在による圧力損失を極力抑さえて高効率圧縮が可能なフ
ァン動翼を提供することを目的としている。The present invention has been made in view of the above circumstances, and provides a fan rotor blade capable of highly efficient compression by suppressing pressure loss due to the presence of shock waves as much as possible even when the relative inlet flow velocity is in the supersonic region. It is intended to be provided.
【０００９】[0009]
【課題を解決するための手段】本発明のファン動翼は、
ターボファンエンジンの空気取入口に配設され、放射状
に配設された複数の翼の回転により流入空気を圧縮して
下流側へ送り出すファン動翼であって、各翼を、翼中央
部から翼端までの形状が、所定の翼弦長を有する長翼
と、長翼より短い翼弦長を有する短翼とで構成し、かつ
これら長翼及び短翼を互い違いに放射状に配列するとと
もに、長翼の翼前縁を短翼の翼前縁より上流側に位置さ
せてなることを特徴とするものである。A fan rotor blade according to the present invention comprises:
A fan rotor blade that is installed at the air intake of a turbofan engine and that compresses the inflowing air by the rotation of multiple blades that are radially arranged and sends it out to the downstream side. The shape to the end is composed of long blades having a predetermined chord length and short blades having a chord length shorter than the long blades, and these long blades and short blades are arranged in a staggered radial pattern and The blade is characterized in that the blade leading edge is located upstream of the blade leading edge of the short blade.
【００１０】[0010]
【作用】本発明のファン動翼によれば、上流側から超音
速の入口相対速度で空気が流れ込むと、長翼の翼前縁に
のみ衝撃波が存在し、この衝撃波よって下流側の長翼の
翼上面に厚い境界層が発達する。一方、衝撃波より下流
側の長翼間の領域は、亜音速（マッハ数０.６〜０.８）
流入となるので、短翼の翼上面での境界層は発達が抑さ
えられ、長翼面上のものと比べると非常に薄い。これに
より、短翼面上に発達する境界層による圧力損失は、長
翼面上に発達する境界層による圧力損失と比較して無視
できる程小さい。したがって、ファン動翼全体の圧力損
失が減少し、効率良く空気を圧縮して下流側に送り出す
ことができる。According to the fan rotor blade of the present invention, when air flows in from the upstream side at a supersonic inlet relative velocity, a shock wave exists only at the leading edge of the blade of the long blade. A thick boundary layer develops on the upper surface of the wing. On the other hand, the region between the long wings on the downstream side of the shock wave has a subsonic velocity (Mach number of 0.6 to 0.8).
Since it is an inflow, the development of the boundary layer on the upper surface of the short blade is suppressed, and it is very thin compared to that on the long blade surface. As a result, the pressure loss due to the boundary layer developing on the short blade surface is negligibly small as compared with the pressure loss due to the boundary layer developing on the long blade surface. Therefore, the pressure loss of the entire fan rotor blade is reduced, and the air can be efficiently compressed and sent to the downstream side.
【００１１】[0011]
【実施例】本発明のファン動翼の一実施例について、図
１ないし図３を参照して説明する。なお、図４に示した
構成と同一構成部分には、同一符号を付してその説明を
省略する。DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of the fan rotor blade of the present invention will be described with reference to FIGS. The same components as those shown in FIG. 4 will be assigned the same reference numerals and explanations thereof will be omitted.
【００１２】本実施例のファン動翼３０は、各翼３１…
が、中央部３１ｃから翼端３１ｂまでの形状を、所定の
翼弦長Ｃ1を有する長翼３１Ａと、この長翼３１Ａより
短い翼弦長Ｃ2を有する短翼３１Ｂとで構成されてい
る。そして、これら長翼３１Ａ、短翼３１Ｂは、互い違
いに放射状に配列されているとともに、長翼３１Ａの翼
前縁３２が、短翼３１Ｂの翼前縁３３より上流側１６に
位置している。The fan rotor blade 30 of the present embodiment includes each blade 31 ...
Is composed of a long blade 31A having a predetermined chord length C 1 and a short blade 31B having a chord length C 2 shorter than the long blade 31A in the shape from the central portion 31c to the blade tip 31b. .. The long blades 31A and the short blades 31B are alternately arranged in a radial pattern, and the blade leading edge 32 of the long blade 31A is located on the upstream side 16 of the blade leading edge 33 of the short blade 31B.
【００１３】また、長翼３１Ａ、短翼３１Ｂ間が同一の
ピッチＳ1に設定され、全体としてソリディティ値の平
均値は変えずに翼列３５が形成されている。そして、フ
ァン動翼３０が所定の周速度Ｕ1で回転し、上流側１６
から所定の入口絶対速度Ｖ1で空気が取り入れられるこ
とにより、ファン動翼３０は入口相対速度Ｗ1の空気を
圧縮するようになっている。Further, the long blades 31A and the short blades 31B are set to the same pitch S 1 , and the blade row 35 is formed without changing the average value of the solidity value as a whole. Then, the fan rotor blade 30 rotates at a predetermined peripheral velocity U 1 , and the upstream side 16
From the above, air is taken in at a predetermined inlet absolute velocity V 1 so that the fan rotor blade 30 compresses the air at the inlet relative velocity W 1 .
【００１４】上記構成からなるファン動翼３０に、上流
側１６から超音速の入口相対速度Ｗ2で空気が流れ込む
と、図２に示すように、長翼３１Ａの翼前縁３２にのみ
衝撃波１８が存在し、この衝撃波１８よって下流側１７
の長翼３１Ａの翼上面に境界層２０が発達する。そし
て、衝撃波１８より下流側の長翼３１Ａ、３１Ａ間の領
域は、亜音速（マッハ数０.６〜０.８）流入となるの
で、短翼３１Ｂの翼上面への境界層発達は抑さえられ
る。これにより、厚い境界層２０は、長翼３１Ａの翼上
面にのみ発達して圧力損失が発生するだけなので、ファ
ン動翼３０全体の圧力損失は減少し、効率良く空気を圧
縮して下流１７側のファン空気排出ダクト３、圧縮機４
に送り出すことができる。When air flows into the fan rotor blade 30 having the above structure from the upstream side 16 at the supersonic inlet relative velocity W 2 , as shown in FIG. 2, the shock wave 18 is generated only at the blade leading edge 32 of the long blade 31A. Exists, and this shock wave 18 causes the downstream side 17
The boundary layer 20 develops on the upper surface of the long blade 31A. The region between the long wings 31A, 31A on the downstream side of the shock wave 18 has a subsonic velocity (Mach number of 0.6 to 0.8), so that the development of the boundary layer on the upper surface of the short blade 31B is suppressed. Be done. As a result, the thick boundary layer 20 only develops on the blade upper surface of the long blade 31A to generate a pressure loss, so that the pressure loss of the entire fan moving blade 30 is reduced, and the air is efficiently compressed to the downstream 17 side. Fan air exhaust duct 3, compressor 4
Can be sent to.
【００１５】従って、本実施例のファン動翼３０は、各
翼３１…を、中央部３１ｃから翼端３１ｂまでの形状
が、所定の翼弦長Ｃ1を有する長翼３１Ａと、この長翼
３１Ａより短い翼弦長Ｃ2を有する短翼３１Ｂとで構成
し、かつこれら長翼３１Ａ、短翼３１Ｂを互い違いに放
射状に配列するとともに、長翼３１Ａの翼前縁３２を、
短翼３１Ｂの翼前縁３３より上流側１６に位置させた構
造としており、上流側１６から超音速の入口相対速度Ｗ
2で空気が流れ込んだ場合、長翼３１Ａの翼前縁３２に
のみ衝撃波１８が存在して翼上面に境界層２０が発達
し、短翼３１Ｂへ流れ込む空気は亜音速流入となり翼上
面への境界層発達が抑さえられるので、図４に示したフ
ァン動翼２と比較して、ファン動翼全体の圧力損失を減
少させたファン動翼を提供することができる。Therefore, in the fan rotor blade 30 of the present embodiment, each blade 31 ... Has a long blade 31A whose shape from the central portion 31c to the blade tip 31b has a predetermined chord length C1, and this long blade 31A. The short blades 31B having a shorter chord length C 2 and the long blades 31A and the short blades 31B are radially arranged alternately, and the blade leading edge 32 of the long blade 31A is
The structure is located on the upstream side 16 from the blade leading edge 33 of the short blade 31B, and the inlet relative velocity W of supersonic speed from the upstream side 16 is set.
When air flows in at 2 , the shock wave 18 exists only on the blade leading edge 32 of the long blade 31A and the boundary layer 20 develops on the blade upper surface, and the air flowing into the short blade 31B becomes a subsonic inflow and becomes a boundary to the blade upper surface. Since the layer development is suppressed, it is possible to provide a fan rotor blade in which the pressure loss of the entire fan rotor blade is reduced as compared with the fan rotor blade 2 shown in FIG.
【００１６】[0016]
【発明の効果】以上説明したように、本発明のファン動
翼は、各翼を、翼中央部から翼端までの形状が、所定の
翼弦長を有する長翼と、この長翼より短い翼弦長を有す
る短翼とで構成し、かつこれら長翼及び短翼を互い違い
に放射状に配列するとともに、長翼の翼前縁を、短翼の
翼前縁より上流側に位置させた構造としており、上流側
から超音速の入口相対速度で空気が流れ込んだ場合であ
っても、長翼の翼前縁にのみ衝撃波が存在して長翼の翼
上面に境界層が発達し、短翼へ流れ込む空気は亜音速流
入となり翼上面への境界層発達が抑さえられるため、フ
ァン動翼全体の圧力損失が減少し、高効率に空気圧縮が
可能なファン動翼を得ることができる。As described above, in the fan moving blade of the present invention, each blade has a long blade whose shape from the blade central portion to the blade tip has a predetermined chord length, and is shorter than this long blade. Structure consisting of short blades having a chord length, these long blades and short blades being arranged radially in a staggered manner, with the leading edge of the long blade positioned upstream from the leading edge of the short blade. Therefore, even when air flows in from the upstream side at a supersonic inlet relative velocity, shock waves exist only at the leading edge of the long blade and a boundary layer develops on the upper surface of the long blade, resulting in a short blade. Since the air flowing into the fan becomes subsonic inflow and the development of the boundary layer on the upper surface of the blade is suppressed, the pressure loss of the entire fan blade is reduced, and a fan blade capable of highly efficient air compression can be obtained.
【図１】ターボファンエンジンを示す概略図である。FIG. 1 is a schematic diagram showing a turbofan engine.
【図２】本発明のファン動翼の翼列を示す断面図であ
る。FIG. 2 is a cross-sectional view showing a row of fan moving blades of the present invention.
【図３】本発明のファン動翼が配設されたターボファン
エンジンの空気取入口を示す概略図である。FIG. 3 is a schematic view showing an air intake of a turbofan engine provided with a fan rotor blade of the present invention.
【図４】従来のファン動翼の翼列を示す断面図である。FIG. 4 is a cross-sectional view showing a blade row of a conventional fan rotor blade.
１ 空気取入口 ３ ファン空気排気ダクト ４ 圧縮機 １０ ファン回転軸 ３０ ファン動翼 ３１ 翼 ３１ｂ 翼端 ３１ｃ 翼中央部 ３１Ａ 長翼 ３１Ｂ 短翼 ３２ 長翼の翼前縁 ３３ 長翼の翼前縁 ３５ 翼列 Ｃ1 長翼の翼弦長 Ｃ2 短翼の翼弦長 Ｓ1 ピッチ Ｕ1 ファン動翼の回転周速度 Ｖ1 空気の入口絶対速度 Ｗ1 空気の入口相対速度1 Air Intake 3 Fan Air Exhaust Duct 4 Compressor 10 Fan Rotating Shaft 30 Fan Blade 31 Blade 31b Blade Tip 31c Blade Central Part 31A Long Blade 31B Short Blade 32 Long Blade Leading Edge 33 Long Blade Leading Edge 35 Cascade C 1 Long blade chord length C 2 Short blade chord length S 1 Pitch U 1 Rotating peripheral velocity of fan blade V 1 Air inlet absolute velocity W 1 Air inlet relative velocity
Claims (1)
設され、放射状に配設された複数の翼の回転により流入
空気を圧縮して下流側へ送り出すファン動翼であって、 各翼を、翼中央部から翼端までの形状が、所定の翼弦長
を有する長翼と、当該長翼より短い翼弦長を有する短翼
とで構成し、かつこれら長翼及び短翼を互い違いに放射
状に配列するとともに、長翼の翼前縁を短翼の翼前縁よ
り上流側に位置させてなることを特徴とするファン動
翼。1. A fan rotor blade that is arranged at an air intake of a turbofan engine and that compresses inflowing air by the rotation of a plurality of radially arranged blades and sends the compressed air to a downstream side. The shape from the blade central part to the blade tip is composed of long blades having a predetermined chord length and short blades having a chord length shorter than the long blades, and these long blades and short blades are radially alternated. The fan blade is characterized in that the leading edge of the long blade is located upstream of the leading edge of the short blade.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP7786692A JPH05280495A (en) | 1992-03-31 | 1992-03-31 | Fan moving blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP7786692A JPH05280495A (en) | 1992-03-31 | 1992-03-31 | Fan moving blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPH05280495A true JPH05280495A (en) | 1993-10-26 |
Family
ID=13645980
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP7786692A Withdrawn JPH05280495A (en) | 1992-03-31 | 1992-03-31 | Fan moving blade |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH05280495A (en) |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2006291955A (en) * | 2005-04-07 | 2006-10-26 | General Electric Co <Ge> | Low solidity turbofan |
| GB2443082A (en) * | 2006-10-19 | 2008-04-23 | Rolls Royce Plc | Suction surface profile for a gas turbine engine transonic fan blade |
| JP2009168024A (en) * | 2008-01-10 | 2009-07-30 | Snecma | Two aerofoil type blade including spacer strip |
| CN109989943A (en) * | 2019-04-18 | 2019-07-09 | 江苏大学 | A kind of multistage pump return guide vane splitterr vanes and design method |
-
1992
- 1992-03-31 JP JP7786692A patent/JPH05280495A/en not_active Withdrawn
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2006291955A (en) * | 2005-04-07 | 2006-10-26 | General Electric Co <Ge> | Low solidity turbofan |
| GB2443082A (en) * | 2006-10-19 | 2008-04-23 | Rolls Royce Plc | Suction surface profile for a gas turbine engine transonic fan blade |
| GB2443082B (en) * | 2006-10-19 | 2010-07-21 | Rolls Royce Plc | A fan blade arrangement |
| US7997872B2 (en) | 2006-10-19 | 2011-08-16 | Rolls-Royce Plc | Fan blade |
| JP2009168024A (en) * | 2008-01-10 | 2009-07-30 | Snecma | Two aerofoil type blade including spacer strip |
| CN109989943A (en) * | 2019-04-18 | 2019-07-09 | 江苏大学 | A kind of multistage pump return guide vane splitterr vanes and design method |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| JP3578769B2 (en) | Flow orientation assembly for the compression region of rotating machinery | |
| US7144221B2 (en) | Method and apparatus for assembling gas turbine engines | |
| US3505819A (en) | Gas turbine power plant | |
| US6292763B1 (en) | Methods for designing variable cycle gas turbine engines | |
| US6195983B1 (en) | Leaned and swept fan outlet guide vanes | |
| EP1098092B1 (en) | Stator blade | |
| JP5177959B2 (en) | Low solidity turbofan | |
| JP3062199B2 (en) | Gas turbine engine | |
| JP3735116B2 (en) | Gas turbine airfoil clocking | |
| US7594388B2 (en) | Counterrotating turbofan engine | |
| US11300136B2 (en) | Aircraft fan with low part-span solidity | |
| US5480284A (en) | Self bleeding rotor blade | |
| US20100232970A1 (en) | Fan rotating blade for turbofan engine | |
| EP1731716A2 (en) | Forward tilted turbine nozzle | |
| US7874794B2 (en) | Blade row for a rotary machine and method of fabricating same | |
| US3692425A (en) | Compressor for handling gases at velocities exceeding a sonic value | |
| US20080118362A1 (en) | Transonic compressor rotors with non-monotonic meanline angle distributions | |
| JP2007536459A (en) | Extraction of shock wave induced boundary layer of transonic gas turbine | |
| JP2001271792A (en) | Flow path for compressor with flute | |
| US4199296A (en) | Gas turbine engines | |
| US10385871B2 (en) | Method and system for compressor vane leading edge auxiliary vanes | |
| US4874289A (en) | Variable stator vane assembly for a rotary turbine engine | |
| US10519976B2 (en) | Fluid diodes with ridges to control boundary layer in axial compressor stator vane | |
| US20210372288A1 (en) | Compressor stator with leading edge fillet | |
| JPH05280495A (en) | Fan moving blade |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| A300 | Withdrawal of application because of no request for examination |
Free format text: JAPANESE INTERMEDIATE CODE: A300 Effective date: 19990608 |