WO2017195782A1 - Pale de stator de turbine et turbine comprenant celle-ci - Google Patents

Pale de stator de turbine et turbine comprenant celle-ci Download PDF

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Publication number
WO2017195782A1
WO2017195782A1 PCT/JP2017/017557 JP2017017557W WO2017195782A1 WO 2017195782 A1 WO2017195782 A1 WO 2017195782A1 JP 2017017557 W JP2017017557 W JP 2017017557W WO 2017195782 A1 WO2017195782 A1 WO 2017195782A1
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WO
WIPO (PCT)
Prior art keywords
turbine
inner end
stationary blade
outer end
radially inner
Prior art date
Application number
PCT/JP2017/017557
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English (en)
Japanese (ja)
Inventor
浩史 渡邊
Original Assignee
三菱重工業株式会社
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Filing date
Publication date
Application filed by 三菱重工業株式会社 filed Critical 三菱重工業株式会社
Publication of WO2017195782A1 publication Critical patent/WO2017195782A1/fr

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles

Definitions

  • the present invention relates to a turbine vane configured to suppress pressure loss.
  • a gas turbine is well known as one that obtains power by changing the kinetic energy of a fluid into rotational energy.
  • it is an important issue to improve the turbine efficiency, and in order to improve the turbine efficiency, it is most important to reduce the pressure loss generated in the moving blades and the stationary blades. It is considered effective.
  • Patent Document 1 discloses a turbine blade for the purpose of suppressing pressure loss.
  • Patent Document 1 The turbine blade disclosed in Patent Document 1 is intended for a moving blade. However, when improving the turbine efficiency, not only the pressure loss generated in the moving blade but also the pressure loss generated in the stationary blade. It is necessary to take some measures.
  • some of the moving blades installed in the gas turbine have a gap between the tip surface of the moving blade and the casing on the fixed side. Such a gap is referred to as a tip clearance.
  • a tip clearance Has been.
  • combustion gas leaks from the tip clearance, and pressure loss may occur in the stationary blade located at the subsequent stage (next stage) of the moving blade.
  • the tip surface 13a of the rotor blade 13 is not provided with a shroud (not shown). For this reason, a tip clearance 15 serving as a gap is formed between the inner peripheral surface 11 a of the casing 11 and the tip surface 13 a of the rotor blade 13. As described above, when the tip clearance 15 is installed, the combustion gas G leaked from the tip clearance 15 may flow toward the stationary blade 50 located at the rear stage of the moving blade 13. Large pressure loss may occur.
  • the main flow F ⁇ b> 1 of the combustion gas G flows along the back surface 13 c and the abdominal surface 13 d of the moving blade 13, and then the stationary blade located at the subsequent stage of the moving blade 13. It flows toward 50.
  • the main flow F1 of the combustion gas G flowing into the stationary blade 50 flows along the back surface 50c and the abdominal surface 50d of the stationary blade 50.
  • the leakage flow F2 of the combustion gas G leaking from the tip clearance 15 flows into the stationary blade 50 at an inflow angle different from the inflow angle of the main flow F1.
  • the stagnation line Po is in the blade height direction. It is preferably formed so as to extend linearly in the (radial direction). However, when the leakage flow F2 of the combustion gas G leaking from the tip clearance 15 is generated, the generation position of the stagnation line Po changes.
  • the leakage flow F2 of the combustion gas G flows into the stationary blade 50 from a position near the rear surface 50c in the vicinity of the front edge 50e of the stationary blade 50.
  • the stagnation line Po is shifted toward the back surface 50c in the vicinity of the chip surface 50a of the front edge 50e. That is, in the stagnation line Po generated in the stationary blade 50, only the vicinity of the tip surface 50a moves to the back surface 50c side due to the inflow of the leakage flow F2.
  • the pressure distribution of the combustion gas G flowing along the back surface 50c of the stationary blade 50 changes following the inclination of the stagnation line Po, and the isopressure line P indicating the pressure level difference approaches the tip surface 50a. It will distort toward the back 50c side.
  • a flow that is orthogonal to the distorted portion of the constant pressure line P and is directed from the tip surface 50a side to the hub surface 50b side is induced.
  • the branch flow F3 of the combustion gas G flowing along the back surface 50c flows from the front edge 50e toward the rear edge 50f with almost no change in the blade height direction position.
  • the branch flow F3 passing through the distorted portion of the isobaric line P flows as the secondary flow F4 from the tip surface 50a side to the hub surface 50b side while moving from the front edge 50e to the rear edge 50f.
  • Such a secondary flow F4 of the combustion gas G is a strong flow directed radially inward.
  • An object of the present invention is to provide a turbine vane that can be improved and a turbine including the same.
  • a turbine vane according to a first invention for solving the above-described problem is Turbine vanes provided at predetermined intervals in the circumferential direction on the inner peripheral surface of a casing that rotatably supports the rotor, A chip surface supported by the inner peripheral surface of the casing and serving as a radially outer end surface; A hub surface facing the outer peripheral surface of the rotor in the radial direction and serving as a radially inner end surface; Connecting the tip surface and the hub surface in the radial direction, a front edge serving as a vertical edge disposed on the upstream side; The tip surface and the hub surface are connected in the radial direction, and includes a rear edge serving as a vertical edge disposed on the downstream side, The leading edge is A radially inner end connected to the hub surface; A retreat start point set at a predetermined radial height from the radially inner end, A radial outer end that is connected to the chip surface and retreated toward the downstream side in the axial direction from the retreat start point; It has
  • the turbine vane according to the second invention for solving the above-described problem is
  • the retraction starting point and the radially inner end are connected so as to form a concave shape or a straight shape toward the upstream side in the axial direction.
  • a turbine vane according to a third invention for solving the above-described problem is
  • the trailing edge has a radially inner end connected to the hub surface;
  • the axial distance between the radially inner end of the leading edge and the radially inner end of the trailing edge is divided into two equal parts, and from the axial center line extending in the radial direction,
  • the axial distance to the radially inner end is Lo
  • the axial distance from the axial center line to the radially outer end of the leading edge is L
  • the receding rate of the radially outer end at the front edge is (Lo-L) / Lo
  • the range of the receding rate (Lo-L) / Lo is 0 ⁇ (Lo ⁇ L) /Lo ⁇ 0.5.
  • a turbine vane according to a fourth invention for solving the above-described problem is
  • the radial height from the radially inner end of the leading edge to the retraction start point is H
  • the range of the radial height H is 0 ⁇ H ⁇ 0.8Ho.
  • a turbine vane according to a fifth invention for solving the above-described problem is When the angle between the tangential direction at the radially outer end of the front edge and the radial direction is ⁇ , The range of the angle ⁇ is 0 ° ⁇ ⁇ 45 °.
  • a turbine vane according to a sixth invention for solving the above-described problem is
  • the trailing edge has a radially outer end connected to the chip surface, The radially outer end of the trailing edge is retracted toward the downstream side in the axial direction from the radially inner end of the trailing edge.
  • a turbine according to a seventh invention for solving the above-described problem is as follows.
  • the stationary blades provided at predetermined intervals in the circumferential direction on the inner peripheral surface of the casing that rotatably supports the rotor and the moving blades provided at predetermined intervals in the circumferential direction on the outer peripheral surface of the rotor are alternately arranged in the axial direction.
  • the turbine vane is It is arrange
  • the radially outer end portion of the front edge is located on the downstream side in the axial direction from the retreat start point set at the intermediate portion in the height direction of the front edge.
  • the front edge curved portion connecting the radially outer end and the retraction starting point is convex toward the upstream side in the axial direction.
  • the stationary blade leading edge may be configured so as to form a concave shape or a straight shape between the retreat start point and the radially inner end portion toward the upstream side in the axial direction.
  • FIG. 2 is a cross-sectional view taken along the line II-II in FIG. It is the front perspective view which looked at the stationary blade from the ventral surface side. It is the front perspective view which looked at the stationary blade from the back side. It is a meridional sectional view of a stationary blade. It is the figure which showed the relationship between the retreat rate of a retreat end part, and the turbine whole stage efficiency. It is the longitudinal cross-sectional view which selected a part of blade cascade structure in the conventional gas turbine.
  • FIG. 8 is a cross-sectional view taken along arrow VIII-VIII in FIG. 7.
  • a rotor 12 is rotatably supported in a casing 11 of the gas turbine 1. Furthermore, in the casing 11, a plurality of moving blades 13 and stationary blades 14 are arranged in an annular shape around the rotor 12.
  • the rotor blades 13 are supported on the outer peripheral surface of the rotor 12 so as to extend in the radial direction, and are arranged at regular intervals in the circumferential direction. That is, the hub surface 13b that is the radially inner end surface of the rotor blade 13 is supported by the outer peripheral surface of the rotor 12, while the tip surface 13a that is the radially outer end surface of the rotor blade 13 is the inner peripheral surface 11a of the casing 11. Is spaced radially inward.
  • a tip clearance (gap) 15 having a predetermined gap amount in the radial direction is formed between the inner peripheral surface 11 a of the casing 11 and the tip surface 13 a of the rotor blade 13. That is, a shroud for filling the chip clearance 15 is not attached to the chip surface 13a.
  • the stationary blades 14 are supported on the inner peripheral surface 11a of the casing 11 so as to extend in the radial direction, and are arranged at regular intervals in the circumferential direction. That is, the tip surface 14 a serving as the radially outer end surface of the stationary blade 14 is supported by the inner peripheral surface 11 a of the casing 11, while the hub surface 14 b serving as the radially inner end surface of the stationary blade 14 is the outer peripheral surface of the rotor 12. Is spaced radially outward.
  • the moving blades 13 and the stationary blades 14 are alternately arranged in the axial direction, thereby forming a cascade structure composed of a plurality of stages.
  • a combustion gas (working fluid) G is supplied from the axial direction between the inner peripheral surface 11 a of the casing 11 and the outer peripheral surface of the rotor 12.
  • the combustion gas G flows from the upstream side, which is upstream in the gas flow direction, toward the downstream side, which is downstream in the gas flow direction, so that the moving blades 13 and the stationary blades 14 constituting each stage are moved to the upstream side. Passes in order toward the rear side.
  • the rotor blades 13 and the stationary blades 14 have shapes (blade shapes) viewed from the outside in the radial direction that are both curved in the circumferential direction.
  • the rotor blade 13 is curved so as to be convex toward the downstream side in the rotational direction of the rotor 12, and the convex shape is formed on the side surfaces provided on both sides in the circumferential direction.
  • the formed surface is a back surface 13c that becomes a negative pressure surface
  • the concave surface is an abdominal surface 13d that becomes a positive pressure surface.
  • the stationary blade 14 is curved so as to be convex toward the upstream side in the rotational direction of the rotor 12, and is convex among the side surfaces provided on both sides in the circumferential direction.
  • a back surface 14c that is a negative pressure surface
  • a concave surface is an abdominal surface 14d that is a positive pressure surface.
  • the leading edge 21 and the trailing edge 22 are formed on the stationary blade 14.
  • the front edge 21 is a vertical edge disposed on the upstream side of the stationary blade 14 and forms a top where the front end of the back surface 14c and the front end of the abdominal surface 14d intersect, and the front end of the tip surface 14a and the hub It connects between the front end of the surface 14b.
  • the rear edge 22 is a vertical edge disposed on the downstream side of the stationary blade 14 and forms a top portion where the rear end portion of the back surface 14c and the rear end portion of the abdominal surface 14d intersect with each other. The rear end portion is connected to the rear end portion of the hub surface 14b.
  • the secondary flow F4 caused by the leakage flow F2 of the combustion gas G leaked from the tip clearance 15 (see FIGS. 10 and 11).
  • the radially outer portion of the front edge 21 is retracted toward the downstream side in the axial direction.
  • FIG. 5 shows a meridional section of the stationary blade 14.
  • This meridional section refers to a line (camber line) connecting a middle point in the blade thickness direction between the back surface 14c and the abdominal surface 14d. It is a longitudinal section of the stationary blade 14 when penetrating in the height direction (radial direction).
  • the front edge 21 connects the front end portion of the chip surface 14a and the front end portion of the hub surface 14b, and is connected to (coincides with) the front end portion of the chip surface 14a.
  • Part 21a and a radially inner end 21b connected to (matching with) the front end of hub surface 14b.
  • a retreat starting point 21c is set at an intermediate portion between the radially outer end 21a and the radially inner end 21b, and among the front edges 21, from the retreat starting point 21c to the radially outer end 21a. This portion constitutes the leading edge curved portion 21d.
  • the rear edge 22 connects the rear end portion of the chip surface 14a and the rear end portion of the hub surface 14b, and is connected to (coincides with) the rear end portion of the chip surface 14a in the radially outer end portion. 22a and a radially inner end 22b connected (coincident with) the rear end of the hub surface 14b. Further, a retreat starting point 22c is set at an intermediate portion between the radially outer end 22a and the radially inner end 22b. Among the rear edges 22, the retreat starting point 22c to the radially outer end 22a. This portion constitutes the trailing edge curved portion 22d.
  • the axial distance between the radially inner end 21b and the radially inner end 22b is divided into two equal parts, and a straight line extending in the radial direction (blade height direction) is defined as the axial center line O. Then, the axial distance from the axial center line O to the radially inner end 21b is Lo, and the axial distance from the axial center line O to the radially outer end 21a is L. As a result, the retraction rate of the radially outer end portion 21a (the front end portion of the chip surface 14a) with respect to the radially inner end portion 21b is (Lo-L) / Lo.
  • the receding rate (Lo-L) / Lo is set in a range of 0 ⁇ (Lo ⁇ L) /Lo ⁇ 0.5.
  • the leading edge curved portion 21d becomes a curve that is convex toward the upstream side in the axial direction.
  • the leading edge of the stationary blade connecting between the retreat start point 21c and the radially inner end 21b is linear in FIG. 5, but is not limited thereto, and is concave toward the upstream side in the axial direction. It may be a leading edge curved portion.
  • the radial height H of the receding start point 21c is the radial height from the radial inner end 21b, and the radial height Ho of the radial outer end 21a is from the radial inner end 21b. It is the height in the radial direction.
  • the radial height H of the retreat starting point 21c is set in a range of 0 ⁇ H ⁇ 0.8Ho.
  • the angle formed by the tangential direction at the radially outer end 21a of the leading edge curved portion 21d and the radial direction is defined as ⁇ .
  • the angle ⁇ is set in a range of 0 ° ⁇ ⁇ 45 °. Note that the shape (fillet) of the connecting portion between the stationary blade and the casing is not considered.
  • the radially outer end 22a is retracted toward the downstream side in the axial direction, so that the trailing edge curved portion 22d becomes a curve that is convex toward the upstream side in the axial direction. Yes.
  • the trailing edge curved portion 22d is formed on the trailing edge 22
  • at least the radially outer end portion 22a only needs to recede toward the downstream side in the axial direction from the radially inner end portion 22b.
  • the backward starting point 22c is set at the same position as the backward starting point 21c in the radial direction. That is, the radial height from the radially inner end 21b at the retreat start point 22c is H.
  • the main flow F1 of the combustion gas G flows toward the rotor blade 13
  • the main flow F1 of the combustion gas G flows from the side of the abdominal surface 13d to the vicinity of the front edge, and then flows along the back surface 13c and the abdominal surface 13d.
  • the blade shape is convex toward the downstream side in the rotational direction, and the leading edge is disposed downstream of the trailing edge in the rotational direction.
  • the rotor blades 13 are caused to flow toward the upstream side in the rotational direction while flowing from the front edge side toward the rear edge side.
  • the main flow F1 of the combustion gas G is changed in the flow direction by the moving blade 13 and gives a force in the rotating direction to the moving blade 13.
  • the main flow F1 of the combustion gas G flows toward the stationary blade 14
  • the main flow F1 of the combustion gas G flows from the abdominal surface 14d side to the vicinity of the front edge 21, and then flows along the back surface 14c and the abdominal surface 14d.
  • the blade shape has a convex shape toward the upstream side in the rotation direction, and the rear edge 22 is disposed downstream in the rotation direction from the front edge 21.
  • F1 is caused to flow toward the downstream side in the rotational direction while being caused to flow from the front edge 21 side toward the rear edge 22 side by the stationary blade 14.
  • the main flow F1 of the combustion gas G is rectified by the stationary blade 14 and the flow direction is changed.
  • the main flow F1 of the combustion gas G flows into the moving blade 13, as described above, the main flow F1 of the combustion gas G flows from the abdominal surface 13d side, so that the main flow of fuel flowing along the surface of the moving blade 13 In the pressure of F1, the pressure on the abdominal surface 13d side becomes higher than the pressure on the back surface 13c side. Further, since the moving blade 13 has the tip clearance 15, a part of the main flow F1 of the combustion gas G flowing on the abdominal surface 13d side is caused by the pressure difference between the pressure on the back surface 13c side and the pressure on the abdominal surface 13d side. Flows from the side of the abdominal surface 13d where the pressure is high to the side of the back surface 13c where the pressure is low via the chip clearance 15.
  • the leakage flow F2 which is the flow of the combustion gas G leaking from the tip clearance 15, flows into the downstream stationary blade 14 while going downstream in the rotational direction.
  • the leakage flow F2 of the combustion gas G flows into the stationary blade 14
  • the leakage flow F2 of the combustion gas G flows from the back surface 14c side in the vicinity of the radially outer end 21a of the front edge 21, that is, the tip. It flows toward the vicinity of the front end of the surface 14a.
  • the stagnation line Po in the vicinity where the leakage flow F2 collides tends to be inclined toward the back surface 14c because the leakage flow F2 of the combustion gas G slightly collides with the back surface 14c from the front edge 21 and collides. .
  • the radially outer end 21a which is a portion where the leakage flow F2 of the combustion gas G collides, is retracted axially downstream from the retreat starting point 21c.
  • the leading edge curved portion 21d connecting the retracted radially outer end 21a and the retracting start point 21c is convex toward the upstream side in the axial direction.
  • the portion where the leakage flow F2 of the combustion gas G collides with the stationary blade 14 can be brought close to the portion where the main flow F1 of the combustion gas G collides with the stationary blade 14, so that the stagnation line Po is set in the radial direction (blade height). Can be formed substantially linearly.
  • the pressure of the combustion gas G flowing along the back surface 14c of the stationary blade 14 becomes substantially equal in the radial direction, and the isobaric line P indicating the pressure level difference of the combustion gas G is also almost in the radial direction. It is formed linearly.
  • the branch flow F3 that branches from the vicinity of the front edge curved portion 21d to the back surface 14c side is along the back surface 14c. Even when flowing from the front edge 21 side toward the rear edge 22 side, there is no secondary flow F4 that is a strong flow toward the radially inner side on the back surface 14c.
  • the distance (distance in the axial direction) becomes long, and the combustion gas G easily comes into contact with the inner peripheral surface 11a between them.
  • the low speed region of the combustion gas G is formed along the inner peripheral surface 11a corresponding to the portion where the interstage distance is increased, and there is a possibility that the turbine efficiency is reduced.
  • the retracting rate (Lo-L) / L of the radially outer end 21a is set in a range of 0 ⁇ (Lo ⁇ L) /L ⁇ 0.5.
  • the retreat rate (Lo-L) / L is set to 0 ⁇ (Lo-L).
  • the overall turbine efficiency can be improved as compared with a conventional gas turbine including a stationary blade 50 that does not have the leading edge curved portion 21d.
  • the turbine full-stage efficiency in the gas turbine 1 according to the present invention is indicated by a solid line
  • the turbine full-stage efficiency in a conventional gas turbine is indicated by ⁇ .
  • the stationary blade 14 by setting the radial height H of the receding start point 21c in the range of 0 ⁇ H ⁇ 0.8Ho, the combustion gas is generated near the leading edge curved portion 21d. Since the leakage flow F2 of G easily collides, the generation of the secondary flow F4 due to the leakage flow F2 can be suppressed.
  • the angle ⁇ at the radially outer end 21a of the leading edge curved portion 21d becomes a large angle exceeding 45 °, the space between the leading edge curved portion 21d and the inner peripheral surface 11a of the casing 11 is increased. (Gap) becomes very narrow. As a result, the combustion gas G stagnates in the space, which may cause a decrease in turbine efficiency.
  • the angle ⁇ is set in a range of 0 ° ⁇ ⁇ 45 °, so that the space between the leading edge curved portion 21d and the inner peripheral surface 11a of the casing 11 is set. Since the space at is not narrowed, a decrease in turbine efficiency can be suppressed.
  • the shaft starts from the retreat starting point 21c set at the radially outer end 21a of the front edge 21 at the intermediate portion in the height direction of the front edge 21.
  • the front edge curved portion 21d connecting the radially outer end portion 21a and the retreat start point 21c is convex toward the upstream side in the axial direction.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

La présente invention concerne une pale de stator (14) qui est disposée sur une surface périphérique interne (11a) d'un boîtier (11) qui soutient de façon rotative un rotor (12), et comprend un bord avant (21) qui relie une surface de pointe (14a), qui est soutenu sur la surface périphérique interne (11a) du boîtier (11), avec une surface de moyeu (14b), qui fait face à une surface périphérique externe du rotor (12) dans une direction radiale. Le bord avant (21) comporte : une partie de bord interne radial (21b) qui est reliée à la surface de moyeu (14b) ; une origine de rétraction (21c) qui est définie à une hauteur prédéterminée (H) depuis la partie de bord interne radial (21b) dans la direction radiale ; une partie de bord externe radial (21a) qui est reliée à la surface de pointe (14a) et est rétractée depuis l'origine de rétraction (21c) vers le côté aval dans la direction axiale ; et une partie de bord avant incurvée (21d) qui est raccordée entre la partie de bord externe radial (21a) et l'origine de rétraction (21c) de façon à former une forme convexe vers le côté amont dans la direction axiale.
PCT/JP2017/017557 2016-05-09 2017-05-09 Pale de stator de turbine et turbine comprenant celle-ci WO2017195782A1 (fr)

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JP2016093616 2016-05-09
JP2016-093616 2016-05-09

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WO2017195782A1 true WO2017195782A1 (fr) 2017-11-16

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2019187435A1 (fr) 2018-03-30 2019-10-03 三菱重工航空エンジン株式会社 Turbine à gaz pour aéronef

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPH0681603A (ja) * 1992-09-03 1994-03-22 Hitachi Ltd 軸流形ターボ機械の静翼構造
JP2002213206A (ja) * 2001-01-12 2002-07-31 Mitsubishi Heavy Ind Ltd ガスタービンにおける翼構造
EP1612372A1 (fr) * 2004-07-01 2006-01-04 Alstom Technology Ltd Aube de turbine avec une partie découpée au pied ou à l'extremité de l'aube
JP2009121468A (ja) * 2007-11-09 2009-06-04 Alstom Technology Ltd 蒸気タービン

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPH0681603A (ja) * 1992-09-03 1994-03-22 Hitachi Ltd 軸流形ターボ機械の静翼構造
JP2002213206A (ja) * 2001-01-12 2002-07-31 Mitsubishi Heavy Ind Ltd ガスタービンにおける翼構造
EP1612372A1 (fr) * 2004-07-01 2006-01-04 Alstom Technology Ltd Aube de turbine avec une partie découpée au pied ou à l'extremité de l'aube
JP2009121468A (ja) * 2007-11-09 2009-06-04 Alstom Technology Ltd 蒸気タービン

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2019187435A1 (fr) 2018-03-30 2019-10-03 三菱重工航空エンジン株式会社 Turbine à gaz pour aéronef
JP2019178636A (ja) * 2018-03-30 2019-10-17 三菱重工航空エンジン株式会社 航空機用ガスタービン
US11111820B2 (en) 2018-03-30 2021-09-07 Mitsubishi Heavy Industries Aero Engines, Ltd. Gas turbine for aircraft
JP7061497B2 (ja) 2018-03-30 2022-04-28 三菱重工航空エンジン株式会社 航空機用ガスタービン

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