JP2001012390A - Compressor blade of gas turbine - Google Patents
Compressor blade of gas turbineInfo
- Publication number
- JP2001012390A JP2001012390A JP11177781A JP17778199A JP2001012390A JP 2001012390 A JP2001012390 A JP 2001012390A JP 11177781 A JP11177781 A JP 11177781A JP 17778199 A JP17778199 A JP 17778199A JP 2001012390 A JP2001012390 A JP 2001012390A
- Authority
- JP
- Japan
- Prior art keywords
- compressor
- compressor blade
- blade
- hole
- airfoil
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Landscapes
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
Description
【０００１】[0001]
【発明の属する技術分野】本発明はガスタービンの圧縮
機翼、および該圧縮機翼を搭載したガスタービンに関す
る。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a compressor blade of a gas turbine and a gas turbine equipped with the compressor blade.
【０００２】[0002]
【従来の技術】一般にガスタービンには、空気を圧縮し
て燃焼器へ送るための圧縮機が設けられている。圧縮機
内部には、ガスタービンの中心軸周りに回転する圧縮機
ロータが設けられ、このロータに固定されたコンプレッ
サディスクに圧縮機翼が埋め込まれている。2. Description of the Related Art Generally, a gas turbine is provided with a compressor for compressing air and sending it to a combustor. Inside the compressor, a compressor rotor that rotates around the central axis of the gas turbine is provided, and compressor blades are embedded in a compressor disk fixed to the rotor.
【０００３】このようなガスタービンの圧縮機翼に関す
る従来の技術としては、特開平7−180502号公報には、
エアフォイル先端を面取りし、その面取り部を含むエア
フォイル先端にピーニングを行って、疲労寿命の向上を
図ることが提案されている。また、特開平9−209708号
公報にはシュラウドを有する蒸気タービン静翼に孔を開
け、孔の内部に挿入物を挿入し、孔と挿入物の摩擦によ
り、翼の減衰を大きくする方法が提案されている。As a prior art relating to such a compressor blade of a gas turbine, Japanese Patent Application Laid-Open No. 7-180502 discloses
It has been proposed that the tip of the airfoil be chamfered and the tip of the airfoil including the chamfered portion be peened to improve the fatigue life. Japanese Patent Application Laid-Open No. 9-209708 proposes a method of making a hole in a steam turbine vane having a shroud, inserting an insert into the hole, and increasing the damping of the blade by friction between the hole and the insert. Have been.
【０００４】[0004]
【発明が解決しようとする課題】ガスタービンの運転中
においては圧縮機ロータは高い回転速度で回転するた
め、この回転速度に応じて圧縮機翼には、遠心力による
引張応力、空気を圧縮するためのガス反力、圧力変動に
よる振動荷重が加わっている。この中の振動荷重は、回
転数の整数倍の周波数成分、圧縮機翼の前後の数段の静
翼通過周波数成分を有し、圧縮機翼が振動荷重と一致も
しくは極めて近い周波数を有する場合、圧縮機翼に高い
振動応力が発生し、翼が疲労破壊する恐れがある。During operation of the gas turbine, the compressor rotor rotates at a high rotational speed, so that the compressor blades compress the tensile stress and air due to centrifugal force according to the rotational speed. Vibration force due to gas reaction force and pressure fluctuation. The vibration load therein has a frequency component that is an integral multiple of the number of revolutions, several stages of stationary blade passing frequency components before and after the compressor blade, and when the compressor blade has a frequency that matches or is very close to the vibration load, High vibration stress is generated in the compressor blade, which may cause fatigue failure of the blade.
【０００５】圧縮機翼は、アメリカ機会学会論文96-GT-
145のFig。6に示されているように高次の局所的な振動
モードを有している。前述の振動荷重の周波数が高次の
固有値と一致もしくは近接した場合には、圧縮機翼は図
１０に示すように振動モードの節ａで高い応力が発生す
る可能性がある。[0005] Compressor blades are based on the American Opportunity Society paper 96-GT-
Fig. 145. As shown in FIG. 6, it has a higher-order local vibration mode. When the frequency of the vibration load matches or approaches a higher-order eigenvalue, the compressor blade may generate high stress at the node a in the vibration mode as shown in FIG.
【０００６】圧縮効率を高めるために、圧縮機翼のエア
フォイル先端は圧縮機ケーシングと微小な間隔を保ちな
がら回転する。このため、圧縮機入り口から吸い込んだ
異物をエアフォイル先端と圧縮機ケーシングとの微小な
間隔に挟み込み易く、万一、異物を挟み込んでしまう
と、エアフォイル先端ｂが損傷することがある。損傷を
受けた状態で、高次の振動モードによる共振が起こった
ときには、図１０に示した振動モードの節ａで高応力が
発生し、エアフォイルｂの損傷部からき裂が発生し進展
する可能性がある。[0006] In order to increase the compression efficiency, the tip of the airfoil of the compressor blade rotates while maintaining a small distance from the compressor casing. For this reason, foreign matter sucked from the inlet of the compressor is easily trapped in the minute gap between the tip of the airfoil and the compressor casing, and if the foreign matter is trapped, the tip b of the airfoil may be damaged. When resonance occurs due to a higher-order vibration mode in a damaged state, high stress is generated at a node a of the vibration mode shown in FIG. 10, and a crack may be generated from a damaged portion of the airfoil b and propagated. There is.
【０００７】圧縮機翼の破損は、破損した破片が後流の
ロータ、ケーシングを損傷させる可能性があり、圧縮機
翼の高次の振動モードによる破損は絶対に避けなければ
ならない。[0007] Damage to the compressor blades may cause the broken pieces to damage the rotor and casing downstream of the compressor blades, and damage to the compressor blades caused by higher vibration modes must be avoided.
【０００８】前述の従来技術のうち特開平7−180502号
公報では、ピーニングにより翼の疲労強度特性を高めて
いるものの、共振による応力の低減に関する配慮がなさ
れていない。また、特開平9−209708号公報では、減衰
材を封入することで共振応力を低減させることはできる
ものの、シュラウドを有する静翼のみ適用でき、一般に
シュラウドを有さないガスタービン翼への適用ができな
い。[0008] Of the above-mentioned prior art, Japanese Patent Application Laid-Open No. Hei 7-180502 discloses that although the fatigue strength characteristics of a blade are enhanced by peening, no consideration is given to reduction of stress due to resonance. Further, in Japanese Patent Application Laid-Open No. 9-209708, although the resonance stress can be reduced by enclosing the damping material, it can be applied only to a vane having a shroud, and is generally applied to a gas turbine blade having no shroud. Can not.
【０００９】本発明の目的は、ガスタービン圧縮機翼に
おいて、振動応力を低減することが可能なガスタービン
圧縮機翼、および該圧縮機翼を搭載したガスタービン翼
を提供することにある。An object of the present invention is to provide a gas turbine compressor blade capable of reducing vibration stress in a gas turbine compressor blade, and a gas turbine blade equipped with the compressor blade.
【００１０】[0010]
【課題を解決するための手段】上記の課題を解決する手
段として、請求項１、２、３、４は圧縮機翼に孔を設け
た。これにより、翼の剛性を下げ、翼の固有振動数を励
振流体の周波数より離すことで、翼の振動応力を低減さ
せることができる。また、請求項２、３では、穿孔した
孔内に減衰材、該孔内をすべる挿入物を設け、翼に減衰
特性を付与することで、翼の振動応力を低減することが
できる。As means for solving the above problems, claims 1, 2, 3, and 4 provide holes in the compressor blades. Thus, the vibration stress of the blade can be reduced by lowering the rigidity of the blade and separating the natural frequency of the blade from the frequency of the excitation fluid. According to the second and third aspects, the vibration stress of the blade can be reduced by providing a damping material and an insert that slides in the hole to provide damping characteristics to the blade.
【００１１】請求項５はダブテイルとコンプレッサディ
スクの側面の接触部に隙間を設け、圧縮機翼の減衰を高
くすることにより達成される。A fifth aspect of the present invention is achieved by providing a gap at the contact portion between the dovetail and the side surface of the compressor disk to increase the damping of the compressor blade.
【００１２】上記課題を解決する手段として以下のこと
が期待できる。The following can be expected as means for solving the above problems.
【００１３】請求項１は圧縮機翼エアフォイル部の板厚
の中心の翼長方向にエアフォイル先端より孔を開け、圧
縮機翼の剛性を下げることで、運転時の励振力の周波数
と圧縮機翼の固有値とを近接させないようにすることが
でき、圧縮機翼の高次モードでの応答を下げることが可
能になる。さらに、孔を開けた箇所にボルトをきり、内
部にボルトを挿入し、エアフォイル先端を平坦にしてや
ることで、効率を下げることがなく剛性を下げてやるこ
とができる。According to the first aspect of the present invention, a hole is formed from the tip of the airfoil in the blade length direction at the center of the thickness of the compressor airfoil portion to reduce the rigidity of the compressor blade, thereby reducing the frequency of the excitation force during operation and the compression. The eigenvalues of the blades can be kept away from each other, and the response of the compressor blades in higher modes can be reduced. Furthermore, the rigidity can be reduced without lowering the efficiency by cutting the bolt at the location where the hole is formed, inserting the bolt inside, and flattening the tip of the airfoil.
【００１４】請求項２、３は穿孔した孔内に減衰材や該
孔内をすべる挿入物を設けることにより、圧縮機翼に減
衰を与えることができ、運転時の励振力の周波数での圧
縮機翼の応答を下げることが可能になる。According to the second and third aspects of the present invention, by providing a damping material or an insert that slides in the hole, a damper can be given to the compressor blade, and the compression at the frequency of the exciting force during operation is provided. The response of the wing can be reduced.
【００１５】請求項４は圧縮機翼エアフォイル部の高次
の固有値の節の部分にエアフォイル板厚方向に孔を開
け、その開けた孔にエアフォイルの材料よりも剛性の低
い材料を挿入し、圧縮機翼の剛性を下げることで、運転
時の励振力の周波数と圧縮機翼の固有値とを近接させな
いようにすることができ、圧縮機翼の高次モードでの応
答を下げることが可能になる。According to a fourth aspect of the present invention, a hole is made in a thickness direction of the airfoil plate at a node of a higher eigenvalue of the compressor blade airfoil portion, and a material having a lower rigidity than the material of the airfoil is inserted into the hole. By reducing the rigidity of the compressor blade, the frequency of the excitation force during operation and the eigenvalue of the compressor blade can be kept away from each other, and the response of the compressor blade in higher-order modes can be reduced. Will be possible.
【００１６】請求項５はダブテイルとコンプレッサディ
スクの側面の接触面に隙間を設けることで圧縮機翼の減
衰を高めることができ、運転時の励振力の周波数での圧
縮機翼の応答を下げることが可能となる。According to a fifth aspect of the present invention, the damping of the compressor blade can be increased by providing a gap at the contact surface between the dovetail and the side surface of the compressor disk, and the response of the compressor blade at the frequency of the exciting force during operation is reduced. Becomes possible.
【００１７】[0017]
【発明の実施の形態】図１に本発明の一実施例を示す。
図に示す実施例では圧縮機翼エアフォイル２の曲げ応力
が最も小さい板厚中央部にエアフォイル２先端部分より
孔４を設け、その孔４の途中までボルト５を挿入し、エ
アフォイル２先端を平坦に仕上げる構造になっている。
これによりエアフォイル２に剛性の低い箇所を設けるこ
とができ、圧縮機翼１の固有値を運転時の励振力の周波
数より低い方向に定めることが可能となる。また、この
実施例では据え付け後に圧縮機翼１の固有値を測定し、
運転時の励振力の周波数と圧縮機翼１の固有値とが近接
していてかつ運転時の励振力の周波数よりも圧縮機翼１
の固有値が高い場合には、エアフォイル２端より孔４を
開ける事ができるため、据え付けた後でも補修が可能で
ある。本実施例では、穿孔した孔をボルト５で封止して
あるために圧縮機の性能を下げることもない。また、穿
孔した孔が性能上問題とならない圧縮機においては孔を
封止しなくてもよい。FIG. 1 shows an embodiment of the present invention.
In the embodiment shown in the figure, a hole 4 is provided from the tip of the airfoil 2 at the center of the plate thickness where the bending stress of the compressor blade airfoil 2 is the smallest, and a bolt 5 is inserted halfway through the hole 4 so that the tip of the airfoil 2 Has a structure that finishes flat.
As a result, a portion having low rigidity can be provided in the airfoil 2, and the eigenvalue of the compressor blade 1 can be determined in a direction lower than the frequency of the excitation force during operation. In this embodiment, the eigenvalue of the compressor blade 1 is measured after installation,
The frequency of the excitation force during operation and the eigenvalue of the compressor blade 1 are close to each other, and are higher than the frequency of the excitation force during operation.
Is high, the hole 4 can be opened from the end of the airfoil 2 so that repair can be performed even after installation. In this embodiment, the performance of the compressor is not degraded because the perforated holes are sealed with the bolts 5. Further, in a compressor in which the perforated holes do not pose a problem in performance, the holes need not be sealed.
【００１８】本発明の応力低減のメカニズムを以下に示
す。図１２に示すガスタービンの運転時の励振力の周波
数と圧縮機翼の固有値の関係で、運転時の励振力の周波
数と圧縮機翼の固有値との近接を避けるために、運転時
の励振力の周波数よりも高い側の周波数ｂか低い側の周
波数ａのどちらかに圧縮機翼の固有値を定める必要があ
る。図１３には圧縮機翼を模擬して加振によるき裂進展
試験を行った時のき裂長さと周波数の関係を示す。図に
示すようにき裂が進展とともに固有値が下がっているの
がわかる。図１２に示すように圧縮機翼の固有値を運転
時の励振力の周波数より高いｂの方に定めると据え付け
時にはＡの大きさだった応答が、経年劣化あるいは前述
したようにき裂が発生した場合には、圧縮機翼の固有値
が下がり、図中の点線で示す方向に圧縮機翼の固有値が
変化し、運転時の励振力の周波数と圧縮機翼の固有値と
が近接あるいは一致し、応答がＡ２のように大きくなる
方向にあり、圧縮機翼エアフォイルの破損を招き易くな
る。逆に、圧縮機翼の固有値を運転時の励振力の周波数
より低いａの方に定めると据え付け時にはＡだった応答
が、経年劣化により固有値が下がっても図中の点線で示
す方向に固有値が変化するため、応答がＡ１のように小
さくなる方向になる。また仮にき裂が発生し、き裂が進
展したとしても、前述したように圧縮機翼の固有値がき
裂の進展とともに下がるため圧縮機翼の応答が小さくな
り、き裂が停留する方向になる。The mechanism of stress reduction according to the present invention will be described below. In order to avoid the proximity between the frequency of the excitation force during operation and the eigenvalue of the compressor blade, the relationship between the frequency of the excitation force during operation of the gas turbine and the eigenvalue of the compressor blade shown in FIG. , It is necessary to determine the eigenvalue of the compressor blade to either the frequency b on the higher side or the frequency a on the lower side. FIG. 13 shows the relationship between the crack length and the frequency when a crack propagation test by vibration was performed by simulating a compressor blade. As shown in the figure, it can be seen that the eigenvalue decreases as the crack progresses. As shown in FIG. 12, when the eigenvalue of the compressor blade is set to the value of b higher than the frequency of the excitation force during operation, the response having the magnitude of A at the time of installation deteriorates over time or cracks occur as described above. In this case, the eigenvalue of the compressor blade decreases, the eigenvalue of the compressor blade changes in the direction indicated by the dotted line in the figure, and the frequency of the excitation force during operation and the eigenvalue of the compressor blade become close to or coincide with each other, and the response In the direction of A2, which tends to cause damage to the compressor airfoil. Conversely, if the eigenvalue of the compressor blade is set to a which is lower than the frequency of the excitation force during operation, the response that was A at the time of installation, but the eigenvalue in the direction indicated by the dotted line in the figure will decrease even if the eigenvalue decreases due to aging. Because of the change, the response tends to be smaller like A1. Even if a crack occurs and the crack propagates, as described above, the eigenvalue of the compressor blade decreases with the progress of the crack, so that the response of the compressor blade decreases and the crack stops.
【００１９】圧縮機翼は前述したように運転時の励振力
の周波数よりも圧縮機翼の固有値を低く定め応答を下げ
てやることが有効である。しかしこれは、圧縮機翼を据
え付ける前では流体力の励振周波数を正確に特定するこ
とが難しいこともあるため、圧縮機翼を据え付けた後に
固有値を測定し、もし運転時の励振力の周波数と圧縮機
翼の固有値が近接していてかつ運転時の励振力の周波数
よりも圧縮機翼の固有値の方が高い場合には、運転時の
励振力の周波数よりも圧縮機翼の固有値の方が低くなる
ように工夫をする必要がある。また、定期点検時には圧
縮機翼の固有値を測定し、運転時の励振力の周波数と圧
縮機翼の固有値とが近接していてかつ運転時の励振力の
周波数よりも圧縮機翼の固有値が高い場合には、圧縮機
翼の固有値を下げ、運転時の励振力の周波数より圧縮機
翼の固有値を低くする必要もある。As described above, it is effective to set the eigenvalue of the compressor blade lower than the frequency of the exciting force during operation, as described above, to lower the response. However, since it is sometimes difficult to accurately determine the excitation frequency of the fluid force before installing the compressor blades, the eigenvalue is measured after the compressor blades are installed, and if the frequency of the excitation force during operation is If the eigenvalue of the compressor blade is close and the eigenvalue of the compressor blade is higher than the frequency of the excitation force during operation, the eigenvalue of the compressor blade is higher than the frequency of the excitation force during operation. It is necessary to devise it so that it is lower. Also, during the periodic inspection, the eigenvalue of the compressor blade is measured, and the eigenvalue of the compressor blade is higher than the frequency of the excitation force during operation because the frequency of the excitation force during operation and the eigenvalue of the compressor blade are close to each other. In such a case, it is necessary to lower the eigenvalue of the compressor blade so that the eigenvalue of the compressor blade is lower than the frequency of the excitation force during operation.
【００２０】以下に本特許の実施例を説明する。本特許
では圧縮機翼の高次での応答を下げるために以下の点に
着目した。一つは運転時の励振力の周波数と圧縮機翼の
固有値を近接あるいは一致させずかつ圧縮機翼の固有値
を運転時の励振力の周波数より低くする方法。もう一つ
は圧縮機翼の運転時の励振力の周波数に対する応答を下
げる方法である。以下に上記方法を用いた実施例を示
す。An embodiment of the present invention will be described below. In this patent, the following points were focused on in order to reduce the high-order response of the compressor blades. One is a method in which the frequency of the excitation force during operation and the eigenvalue of the compressor blade do not approach or match each other, and the eigenvalue of the compressor blade is lower than the frequency of the excitation force during operation. The other is to reduce the response of the compressor blades to the frequency of the excitation force during operation. An example using the above method will be described below.
【００２１】図２は請求項２の実施例である。図１で示
した孔の部分に圧縮機翼の材料よりも剛性が低く、減衰
が大きくなる材質のもの６を孔４の途中まで挿入し、エ
アフォイル２上部よりボルト５を挿入し、エアフォイル
２先端を平坦に仕上げる構造になっている。これにより
エアフォイル２に剛性の低い箇所を設けることができ、
圧縮機翼１の固有値を運転時の励振力の周波数より低い
方向に定めることが可能となる。さらに、運転時に翼が
励振されても、減衰材６により、圧縮機翼１の応答を下
げることが可能になる。減衰が大きくなる材質として
は、エアフォイル２より剛性の低い鉛や、アルミなどの
金属のほかに、αゲルのような粘性体を入れることも有
効である。FIG. 2 shows a second embodiment of the present invention. In the hole shown in FIG. 1, a material 6 having a lower rigidity and a larger damping than the material of the compressor blade is inserted halfway through the hole 4, and a bolt 5 is inserted from above the airfoil 2. 2 It has a structure to finish the tip flat. This allows the airfoil 2 to be provided with a low rigidity portion,
It is possible to determine the eigenvalue of the compressor blade 1 in a direction lower than the frequency of the exciting force during operation. Furthermore, even if the blades are excited during operation, the damping material 6 can reduce the response of the compressor blades 1. As a material having a large attenuation, it is also effective to add a viscous material such as α-gel in addition to lead having a lower rigidity than the airfoil 2 and a metal such as aluminum.
【００２２】図３は請求項３の実施例である。図１で示
した孔の部分に棒６を孔４の途中まで挿入し、エアフォ
イル２上部よりボルト５を挿入し、エアフォイル２先端
を平坦に仕上げる構造になっている。これによりエアフ
ォイル２に剛性の低い箇所を設けることができ、圧縮機
翼１の固有値を運転時の励振力の周波数より低い方向に
定めることが可能となる。また、孔４の内面と棒６の表
面に運転に伴う摩擦が生じ、その摩擦により圧縮機翼１
の減衰が大きくなり、運転時の励振力の周波数に対する
圧縮機翼１の応答を下げることが可能となる。 さら
に、孔４に挿入する棒６の表面を図に示すように粗くし
て、孔４と棒６との摩擦を大きくし、圧縮機翼１の減衰
を大きくすることで運転時の励振力の周波数に対する圧
縮機翼１の応答を下げることが可能である。棒６の材質
としては金属系の材質の物にすることも有効であるが、
磁石のような磁性体を挿入することで、振動エネルギー
をジュール熱に変換し、圧縮機翼１の減衰を大きくさせ
ることも可能である。FIG. 3 shows a third embodiment of the present invention. The rod 6 is inserted into the hole 4 shown in FIG. 1 halfway through the hole 4, the bolt 5 is inserted from above the airfoil 2, and the end of the airfoil 2 is finished flat. As a result, a portion having low rigidity can be provided in the airfoil 2, and the eigenvalue of the compressor blade 1 can be determined in a direction lower than the frequency of the excitation force during operation. In addition, friction during operation occurs between the inner surface of the hole 4 and the surface of the rod 6, and the friction causes the compressor blade 1 to rotate.
And the response of the compressor blade 1 to the frequency of the exciting force during operation can be reduced. Further, the surface of the rod 6 to be inserted into the hole 4 is roughened as shown in the figure to increase the friction between the hole 4 and the rod 6 and increase the damping of the compressor blade 1 to reduce the excitation force during operation. It is possible to reduce the response of the compressor blade 1 to the frequency. It is also effective to use a metal material as the material of the rod 6,
By inserting a magnetic material such as a magnet, the vibration energy can be converted into Joule heat, and the damping of the compressor blade 1 can be increased.
【００２３】図４には図２、３のボルト部の拡大図を示
す。図に示すようにボルト部をダブルボルト５１、５２
にすることで、ボルト部の締付けが強固になり、孔４の
内部に挿入した棒６に遠心力に伴う力が生じてもボルト
が抜けることのないような構造となっている。また図５
に示す例では、棒６を挿入後ボルト５１を挿入し、その
ボルト５１の上面と孔とが接する部分(図中ｄ点)をかし
めることにより棒６およびボルトが抜けることのないよ
うな構造になっている。さらに図４、図５に示すエアフ
ォイル２先端とボルト５２の接触面(図中ｃ点)をかしめ
ることでさらに棒６およびボルトの飛散を防止すること
が可能となる。本実施例では孔の封止をボルトにて実施
しているが、同様の効果を持つ他の手法、たとえば、溶
接、ピン圧入、接着などで実現してもよい。FIG. 4 is an enlarged view of the bolt portion shown in FIGS. As shown in FIG.
By doing so, the tightening of the bolt portion is strengthened, so that the bolt does not come off even if a force due to centrifugal force is generated in the rod 6 inserted into the hole 4. FIG.
In the example shown in FIG. 7, the rod 6 is inserted, the bolt 51 is inserted, and the portion where the upper surface of the bolt 51 contacts the hole (point d in the drawing) is crimped to prevent the rod 6 and the bolt from coming off. It has become. Further, by caulking the contact surface (point c in the figure) between the tip of the airfoil 2 and the bolt 52 shown in FIGS. 4 and 5, it is possible to further prevent the rod 6 and the bolt from scattering. In this embodiment, the holes are sealed with bolts, but may be realized by other methods having the same effect, for example, welding, press-fitting, bonding, or the like.
【００２４】図６に示す実施例では、図１１に示す高次
の固有値の節ａの部分のエアフォイル２板厚方向に孔７
を開けその孔７の中に圧縮機翼の材料よりも剛性の低い
材料を挿入することにより、圧縮機翼１の固有値を高次
の運転時の励振力の周波数より低い方向に定めることが
できる。また、この実施例ではき裂が進展しても孔７の
部分でき裂を停留させることも可能である。さらに、こ
の実施例では据え付け後に圧縮機翼１の固有値を測定
し、運転時の励振力の周波数と圧縮機翼１の固有値とが
近接していてかつ運転時の励振力の周波数よりも圧縮機
翼１の固有値が高い場合には、エアフォイル２側面より
孔７を開けエアフォイル２の材料よりも剛性の低い材料
を挿入することができるため、据え付けた後でも補修が
可能である。In the embodiment shown in FIG. 6, a hole 7 is formed in the thickness direction of the airfoil 2 at the node a of the higher eigenvalue shown in FIG.
And by inserting a material having a lower rigidity than the material of the compressor blade into the hole 7, the eigenvalue of the compressor blade 1 can be determined in a direction lower than the frequency of the excitation force during higher-order operation. . Further, in this embodiment, even if the crack propagates, it is possible to stop the crack in the hole 7 only. Further, in this embodiment, the eigenvalue of the compressor blade 1 is measured after installation, and the frequency of the excitation force during operation and the eigenvalue of the compressor blade 1 are close to each other and the frequency of the compressor force is lower than the frequency of the excitation force during operation. When the characteristic value of the wing 1 is high, a hole 7 can be opened from the side surface of the airfoil 2 and a material having a lower rigidity than the material of the airfoil 2 can be inserted, so that repair can be performed even after installation.
【００２５】図６に示す実施例では、エアフォイル２に
開けた孔７を１つにしたが、本方法は図７に示すよう
に、高次の固有値の節ａに複数個の孔７１を開け、その
内部にエアフォイル２材よりも剛性の低い材料を入れて
やることで圧縮機翼１の固有値を下げてやることも可能
である。また、図８のように高次の固有値の節ａに沿っ
て長孔７２を開けてやり、その内部にエアフォイル２材
よりも剛性の低い材料を入れてやることで圧縮機翼１の
固有値を下げてやることも可能である。In the embodiment shown in FIG. 6, a single hole 7 is formed in the airfoil 2. However, as shown in FIG. 7, the present method includes a plurality of holes 71 in a node e having a higher eigenvalue. It is also possible to lower the eigenvalue of the compressor blade 1 by opening it and putting a material having a lower rigidity than the two airfoil members inside. Also, as shown in FIG. 8, a long hole 72 is formed along a node a having a higher eigenvalue, and a material having a rigidity lower than that of the two air foils is put into the elongate hole 72 to thereby obtain an eigenvalue of the compressor blade 1. It is also possible to lower.
【００２６】図９に示す本発明の実施例では、圧縮機翼
１のダブテイル３とコンプレッサディスク８の接続面に
隙間９を設けることにより、圧縮機翼１の減衰を高くす
ることができ、運転時の励振力の周波数に対する圧縮機
翼１の応答を下げることが可能となる。また、図１０に
示すように圧縮機翼１のダブテイル３とコンプレッサデ
ィスク８の接触面９１の面にローレット加工等の加工を
施し、接触面の面粗さを粗くすることで、この接触面９
１での摩擦エネルギーを大きくすることができ、圧縮機
翼１の減衰が高くなり、運転時の励振力の周波数に対す
る圧縮機翼１の応答を下げることが可能となる。In the embodiment of the present invention shown in FIG. 9, a gap 9 is provided on the connecting surface between the dovetail 3 of the compressor blade 1 and the compressor disk 8, so that the damping of the compressor blade 1 can be increased, and The response of the compressor blade 1 to the frequency of the exciting force at the time can be reduced. Further, as shown in FIG. 10, the surface of the contact surface 91 between the dovetail 3 of the compressor blade 1 and the compressor disk 8 is subjected to a process such as knurling to make the contact surface rough.
In this case, the frictional energy of the compressor blade 1 can be increased, the damping of the compressor blade 1 increases, and the response of the compressor blade 1 to the frequency of the exciting force during operation can be reduced.
【００２７】図１４に本発明を適用したガスタービン断
面図を示す。ガスタービン２０においては、空気を圧縮
して燃焼器２２へ送るための圧縮機２１が設けられてい
る。圧縮機２１内部には、ガスタービン２０の中心軸周
りに回転する圧縮機ロータが設けられ、このロータに固
定されたコンプレッサディスクに圧縮機翼１が埋め込ま
れている。FIG. 14 is a sectional view of a gas turbine to which the present invention is applied. In the gas turbine 20, a compressor 21 for compressing air and sending the compressed air to a combustor 22 is provided. A compressor rotor that rotates around the central axis of the gas turbine 20 is provided inside the compressor 21, and the compressor blade 1 is embedded in a compressor disk fixed to the rotor.
【００２８】[0028]
【発明の効果】本発明によると圧縮機翼に孔をあけ、圧
縮機翼の固有値を運転時の励振力の周波数より下げてや
ることで、経年劣化、き裂の進展により圧縮機翼の固有
値が下がったとしても圧縮機翼の応答を下げることが可
能である。また、該孔にものを挿入し、孔とものの摩擦
により圧縮機翼の減衰を大きくすることで、運転時の励
振力の周波数での圧縮機翼の応答を下げることも可能と
なる。According to the present invention, a hole is formed in the compressor blade, and the characteristic value of the compressor blade is lowered below the frequency of the excitation force during operation, whereby the characteristic value of the compressor blade is deteriorated due to aging and crack propagation. It is possible to reduce the response of the compressor blade even if the pressure decreases. In addition, by inserting an object into the hole and increasing the damping of the compressor blade due to friction between the hole and the hole, the response of the compressor blade at the frequency of the excitation force during operation can be reduced.
【図１】本発明の一実施例で圧縮機翼エアフォイルの端
部より孔を設けた図。FIG. 1 is a view showing a compressor blade airfoil according to an embodiment of the present invention, in which a hole is provided from an end of the airfoil.
【図２】発明の一実施例で圧縮機翼エアフォイルの端部
より孔を開け、孔の内部に剛性の低い材料を挿入した
図。FIG. 2 is a view showing a hole from the end of a compressor blade airfoil according to an embodiment of the present invention, and a material having low rigidity inserted into the hole.
【図３】本発明の一実施例で圧縮機翼エアフォイルたん
部より穴を開け、その孔の内部にものを挿入した図。FIG. 3 is a view showing a state in which a hole is formed in a compressor blade airfoil portion and an object is inserted into the hole in an embodiment of the present invention.
【図４】図２、３に示す実施例のボルト部拡大図。FIG. 4 is an enlarged view of a bolt portion of the embodiment shown in FIGS.
【図５】図２、３に示す実施例のボルト部拡大図。FIG. 5 is an enlarged view of a bolt portion of the embodiment shown in FIGS.
【図６】本発明の一実施例で圧縮機翼エアフォイル板厚
方向に孔を開けた図。FIG. 6 is a view showing a hole in a thickness direction of a compressor blade airfoil according to an embodiment of the present invention.
【図７】図６に示す実施例の孔の数を増やした図。FIG. 7 is a diagram of the embodiment shown in FIG. 6 in which the number of holes is increased.
【図８】図６に示す実施例の孔の形を変えた図。FIG. 8 is a view of the embodiment shown in FIG. 6 in which the shape of the hole is changed.
【図９】本発明の一実施例でダブテイルとコンプレッサ
ディスクの接触面に隙間を設けた図。FIG. 9 is a view showing a gap between contact surfaces of a dovetail and a compressor disk according to an embodiment of the present invention.
【図１０】本発明の一実施例でダブテイルとコンプレッ
サディスクの接触面の面粗さを粗くした図。FIG. 10 is a diagram showing a roughened surface of a contact surface between a dovetail and a compressor disk in one embodiment of the present invention.
【図１１】圧縮機翼の高次の高応力部の節を示す図。FIG. 11 is a view showing a node of a high-order high-stress portion of a compressor blade.
【図１２】運転時周波数と圧縮機翼の固有値の関係を説
明する図。FIG. 12 is a view for explaining the relationship between the operating frequency and the eigenvalue of the compressor blade.
【図１３】き裂長さと固有値の関係を説明する図。FIG. 13 is a view for explaining a relationship between a crack length and an eigenvalue.
【図１４】ガスタービン概要図。FIG. 14 is a schematic diagram of a gas turbine.
【符号の説明】 １…圧縮機翼、２…エアフォイル、３…ダブテイル、４
…孔部、５・・・ボルト部、６…孔部挿入物、７…孔部、
８…コンプレッサディスク、９・・・隙間部、２０…ガス
タービン、２１…圧縮機、２２…燃焼器、２３・・・ター
ビン。[Explanation of Signs] 1 ... Compressor blade, 2 ... Air foil, 3 ... Dovetail, 4
... Hole, 5 ... Bolt, 6 ... Hole insert, 7 ... Hole,
8 Compressor disk, 9 Crevice, 20 Gas turbine, 21 Compressor, 22 Combustor, 23 Turbine.
───────────────────────────────────────────────────── フロントページの続き (51)Int.Cl.7 識別記号 ＦＩ テーマコート゛(参考） Ｆ０４Ｄ 29/34 Ｆ０４Ｄ 29/34 Ｔ Ｆ (72)発明者 桜井 茂雄 茨城県土浦市神立町502番地 株式会社日 立製作所機械研究所内 Ｆターム(参考） 3H032 AA01 3H033 AA02 AA16 BB03 BB08 CC01 DD06 DD18 DD29 EE06 EE18──────────────────────────────────────────────────続 き Continued on the front page (51) Int.Cl. 7 Identification symbol FI Theme coat ゛ (Reference) F04D 29/34 F04D 29/34 TF (72) Inventor Shigeo Sakurai 502, Kandachicho, Tsuchiura-shi, Ibaraki, Inc. F-term in the Hitachi Machinery Laboratory (reference) 3H032 AA01 3H033 AA02 AA16 BB03 BB08 CC01 DD06 DD18 DD29 EE06 EE18
Claims (5)
成される圧縮機翼において、エアフォイルの板厚中心部
に翼長方向にエアフォイル先端から孔を設けたことを特
徴とするガスタービンの圧縮機翼。1. A compressor blade for a gas turbine, comprising: a compressor blade comprising an airfoil and a dovetail portion, wherein a hole is provided from a tip of the airfoil in a blade length direction at a center portion of a thickness of the airfoil. .
において、穿孔した孔の内に減衰材を有することを特徴
とするガスタービンの圧縮機翼。2. The gas turbine compressor blade according to claim 1, further comprising a damping material in the perforated hole.
において、穿孔した穴の中に、該孔に対し、相対的にす
べることのできる挿入物を有することを特徴としたガス
タービンの圧縮機翼。3. The gas turbine compressor blade according to claim 1, further comprising an insert which can be slid relative to said hole in said hole. Machine wing.
成される圧縮機翼において、エアフォイルの板厚方向に
孔を開け、その孔の内部に剛性の低い材料を挿入したこ
とを特徴とするガスタービンの圧縮機翼。4. A compressor blade comprising an airfoil and a dovetail portion, wherein a hole is formed in a thickness direction of the airfoil, and a material having low rigidity is inserted into the hole. Compressor wings.
成される圧縮機翼において、ダブテイルとコンプレッサ
ディスクの側面の接触部に隙間を設け、さらにその接触
面の面粗さを粗くしたことを特徴とするガスタービンの
圧縮機翼。5. A compressor blade comprising an airfoil and a dovetail portion, wherein a gap is provided at a contact portion between the dovetail and a side surface of the compressor disk, and the contact surface has a roughened surface. Turbine compressor blades.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP11177781A JP2001012390A (en) | 1999-06-24 | 1999-06-24 | Compressor blade of gas turbine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP11177781A JP2001012390A (en) | 1999-06-24 | 1999-06-24 | Compressor blade of gas turbine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JP2001012390A true JP2001012390A (en) | 2001-01-16 |
Family
ID=16037005
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP11177781A Pending JP2001012390A (en) | 1999-06-24 | 1999-06-24 | Compressor blade of gas turbine |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JP2001012390A (en) |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2005291211A (en) * | 2004-04-01 | 2005-10-20 | General Electric Co <Ge> | Frequency-tuned compressor stator blade and related method |
| JP2017137826A (en) * | 2016-02-04 | 2017-08-10 | 三菱日立パワーシステムズ株式会社 | Blade, rotary machine and blade frequency adjustment method |
| JP2018114565A (en) * | 2017-01-16 | 2018-07-26 | 三菱マテリアル株式会社 | Cutting tool |
-
1999
- 1999-06-24 JP JP11177781A patent/JP2001012390A/en active Pending
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2005291211A (en) * | 2004-04-01 | 2005-10-20 | General Electric Co <Ge> | Frequency-tuned compressor stator blade and related method |
| JP4711717B2 (en) * | 2004-04-01 | 2011-06-29 | ゼネラル・エレクトリック・カンパニイ | Frequency-adjusting compressor stator blades and related methods |
| JP2017137826A (en) * | 2016-02-04 | 2017-08-10 | 三菱日立パワーシステムズ株式会社 | Blade, rotary machine and blade frequency adjustment method |
| JP2018114565A (en) * | 2017-01-16 | 2018-07-26 | 三菱マテリアル株式会社 | Cutting tool |
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