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EP2309097A1 - Profil und zugehörige Leitschaufel, Laufschaufel, Gasturbine und Strömungsmaschine - Google Patents

Profil und zugehörige Leitschaufel, Laufschaufel, Gasturbine und Strömungsmaschine Download PDF

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Publication number
EP2309097A1
EP2309097A1 EP09012404A EP09012404A EP2309097A1 EP 2309097 A1 EP2309097 A1 EP 2309097A1 EP 09012404 A EP09012404 A EP 09012404A EP 09012404 A EP09012404 A EP 09012404A EP 2309097 A1 EP2309097 A1 EP 2309097A1
Authority
EP
European Patent Office
Prior art keywords
airfoil
platform
gas turbine
ridge
distance
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Withdrawn
Application number
EP09012404A
Other languages
English (en)
French (fr)
Inventor
Andrew Shepherd
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Siemens AG
Original Assignee
Siemens AG
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Siemens AG filed Critical Siemens AG
Priority to EP09012404A priority Critical patent/EP2309097A1/de
Priority to PCT/EP2010/061625 priority patent/WO2011038971A1/en
Publication of EP2309097A1 publication Critical patent/EP2309097A1/de
Withdrawn legal-status Critical Current

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Classifications

    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/127—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with a deformable or crushable structure, e.g. honeycomb

Definitions

  • the invention relates to an airfoil, a guide vane, and a blade, particularly of a compressor or a turbine section of a turbomachine. Furthermore the invention relates to a turbomachine, especially a gas turbine, comprising at least one airfoil.
  • a conventional gas turbine engine comprises a compressor which is used for pressurising ambient air.
  • the compressed air is mixed with fuel and burned in a combustor.
  • Energy is extracted from the hot combustion gas by passing it through a turbine section of the gas turbine engine.
  • the rotatable sections of an engine typically comprise annular arrays of compressor or turbine rotor blades, the blades normally being intersected with annular arrays of static aerodynamic guide vanes, also called stator vanes.
  • stator vanes also called stator vanes.
  • Each disc of guide vanes and each disc of rotor blades are arranged axially alternately.
  • One disc of rotor blades and one disc of guide vanes are referred to as a stage.
  • the guide vanes ensure the gas impinges on the rotor blades at the correct angle.
  • the first stages of a compressor may be equipped with variable guide vanes that are pivotable to adapt the stagger angle during different operating modes of the compressor.
  • a rotor comprises a plurality of rotor blades arranged at even intervals in a circumferential direction, each of which is obliquely positioned to both the front direction and the rotating direction so as to compress air downstream by rotation thereof.
  • a first face - e.g. directed forward - is to suck air and therefore referred to as a suction side or suction surface
  • a second face - e.g. directed aftward - is to compress air and referred to as a pressure side or pressure surface within this application.
  • the suction side is typically made convex and the pressure side is made concave.
  • This clearance may be defined as the distance between these mentioned elements after their assembly but prior to initial operation of the turbine (cold clearance measurement).
  • each blade - i.e. a tip section of the blade - may be in frictional contact with an inner face of a case of the compressor and may be coated with a hard coating to abrade the inner face of the case of the compressor. Because of the abrasiveness of the hard coating the distal end itself is protected from deterioration by frictional contact.
  • the opposite member preferentially wears in comparison between the tip section and the opposite member. Such a configuration allows a perfect match of the tip of the blade and the opposing compressor casing, even with manufacturing or assembly tolerances.
  • US patent US 4,526,509 A1 discloses a further type of seal that is in rubbing contact with a tip of a rotor blade, but not being abradable but resilient.
  • the present invention seeks to mitigate drawbacks of the prior art. Specifically it is an object of the invention to provide an airfoil or a component comprising such an airfoil that leakage over the tip of the airfoil is reduced, by also considering material costs and/or durability of the airfoil.
  • an airfoil of a turbomachine particularly a gas turbine
  • the airfoil particularly being arranged in a compressor section of the gas turbine, the airfoil being in particular a vane or a blade.
  • the gas turbine may be an axial flow turbine and the compressor section an axial flow compressor.
  • the invention may also apply to blades or vanes within a turbine section.
  • the airfoil comprises a platform, a pressure surface, a suction surface, and a tip region. The pressure surface and a suction surface both extend substantially perpendicular to the platform along a first direction, the first direction being defined as a direction substantially perpendicular to the platform and progressing with a distance to the platform.
  • the tip region is located between the pressure and the suction surface opposite to the platform in respect of the pressure and the suction surfaces.
  • the tip region is defined as a distant end of the airfoil in direction along the first direction and is configured such that the tip region comprises a first and a second surface.
  • the first surface is in form of a ridge and has a first surface expanse less than a cross sectional plane of the airfoil perpendicular to the first direction and has a first distance to the platform along the first direction.
  • the second surface has a second surface expanse and a second distance to the platform along the first direction. Additionally the first distance is greater than the second distance.
  • tip region the part of the airfoil is meant, when being assembled within the gas turbine, that is directed to an opposing surface of a component of the gas turbine - e.g. a compressor casing or a rotor.
  • the "ridge” defines a sharp edge, substantially straight.
  • the ridge may extend from a leading edge of the airfoil - upstream during operation of the gas turbine - to a trailing edge of the airfoil - downstream during operation of the gas turbine.
  • the first surface expanse is compared to the cross sectional plane of the airfoil perpendicular to the first direction, i.e. substantially parallel to the platform or parallel to the spread of the tip section.
  • a cross sectional plane of a blade differs in size based on the position where the measurement is taken. Particularly a smallest possible cross sectional plane, an average cross sectional plane, a largest cross section, or a cross section of a plane close to the tip region could be taken for the to be compared cross sectional plane.
  • the first and second distance may be average values, if the distance of the first surface (or the second surface) in relation to the platform is different based on at which actual position within the first surface (or the second surface) the measurement is taken.
  • the inventive airfoil may be advantageous - especially when also considering the subject matter of the dependent claims - , because this form of the airfoil provides only a small contact surface - the first surface - to an opposing surface, once assembled. This may lead to reduced force on blades to abrade the tip region. A reduced operating clearance between the tip region and the opposing surface may be provided which leads to reduced leakage and eventually to an improved compressor efficiency.
  • the ridge may be an extension of the suction side surface or of the pressure side surface. Therefore the ridge spans over the whole length of the tip region of the airfoil. This may reduce leakages.
  • the second surface may be at least one of a tilted surface between the pressure surface and the ridge and a tilted surface between the suction surface and the ridge. This guarantees that the first surface is highest elevation of the tip region.
  • the tip region - particularly the first surface and possibly only the first surface - may be composed of a deformable material. This would allow that only the tip region adapts to the form of an opposing surface.
  • the tip region - particularly the first surface and possibly only the first surface - may be either coated with an abrasive coating for abrading an opposing surface or coated with an abradable coating, so that the tip region gets abraded.
  • This has the advantage that the tip region and the opposing surface would become a perfect fit due to the intended abrasion. This is possible by using only less abradable or abrasive coatings on the tip region, because only the ridge may be coated. This again may reduce costs.
  • the invention is also directed to a blade and/or a guide vane, particularly a variable guide vane of a gas turbine, comprising an airfoil as explained.
  • the blade and/or vane may be located in a compressor or a turbine section of the gas turbine.
  • the invention is also directed to a gas turbine comprising at least one of a compressor section or a turbine section, the compressor section or the turbine section comprising at least one airfoil configured as explained.
  • a gas turbine comprising at least one of a compressor section or a turbine section, the compressor section or the turbine section comprising at least one airfoil configured as explained.
  • the invention is directed to all kinds of turbomachinery equipment, like turbines, compressors, pumps that comprise rotating parts.
  • Fig. 1 shows a gas turbine 1 in a longitudinal partial section. In the interior, it has a rotor 3 which is rotatably mounted about a rotation axis 2 and is also referred to as turbine rotor or rotor shaft. Following one another along the rotor 3 are an intake casing 4, a compressor 5, torus-like annular combustion chambers 6 (only one can be seen in the figure) each having a burner 7, a turbine unit 8 and an exhaust-gas casing 9. As an alternative to the annular combustion chambers 6, also separate can combustion chambers may be present, each with one or more burners, of which there may be a plurality of them arranged around the axis of the gas turbine engine.
  • annular compressor duct 10 as a flow path for air which narrows in cross section in the direction of the annular combustion chamber 6, i.e. downstream of a to be compressed fluid.
  • a diffuser 11 Arranged at the downstream end of the compressor 5 is a diffuser 11, which is fluidically connected to the annular combustion chambers 6.
  • Each of the annular combustion chambers 6 form a combustion space 12 for a mixture of liquid and/or gaseous fuel and compressed air from the compressor 5.
  • a hot-gas duct 13 arranged in the turbine unit 8 is fluidically connected to the combustion spaces 12, the exhaust-gas casing 9 being arranged downstream of the hot-gas duct 13.
  • Sections with blades and vanes are arranged in the compressor duct 10 and in the hot-gas duct 13.
  • a turnable blade section 17 of blades 16 alternately follows a vane section 15 of non-rotatable guide vanes 14.
  • These fixed guide vanes 14 are in this case connected to one or more guide vane carriers 18, whereas the moving blades 16 are fastened to the rotor 3 by means of a disc 19.
  • the turbine unit 8 has a conically widening hot-gas duct 13, the outer guide surface 21 of which widens concentrically in the direction of flow of the working fluid 20.
  • the inner guide surface 22, on the other hand, is oriented essentially parallel to the rotation axis 2 of the rotor 3.
  • the moving blades 16 each have an edge as a tip region 29 (see Fig. 2 ) - possibly grazing edges -, which form a radial gap 23 with the outer guide surfaces 21 opposite them.
  • air L is drawn in from the compressor 5 through the intake casing 4 and is compressed in the compressor duct 10.
  • the air L provided at the burner-side end of the compressor 5 is directed through the diffuser 11 to the burners 7 and is mixed there with a fuel.
  • the mixture is then burned, with a working fluid 20 being formed in the combustion space 12.
  • the working fluid 20 flows from there into the hot-gas duct 13.
  • the working fluid expands in an impulse-transmitting manner, so that the rotor 3 is driven.
  • An inlet-side compressor bearing 32 may serve, in addition to the axial and radial mounting, as an adjusting device for a displacement of the rotor 3.
  • the rotor 3, in the steady state, may be displaced, to the left in fig. 1 , from an initial position into a steady operating position against the direction of flow of the working fluid 20.
  • the radial gap 23 formed in the turbine unit 8 by moving blades 16 and the outer guide surface 21 is reduced. This leads to a reduction in the flow losses in the turbine unit 8 and therefore to an increase in the efficiency of the gas turbine 1.
  • the radial gap 23 may be different for each of the blades 16 due to tolerances.
  • a section of the annular compressor duct 10 of the compressor 5 with two sections 17 of blades 16 and with a guide-vane section 15 arranged in between is shown in fig. 2 .
  • the annular compressor duct 10 is in this case designed as a flow duct for air L as the flow medium.
  • each blade 16 has a respective platform 25, the surfaces of which define the compressor duct 10 on the radial inside.
  • each guide vane 14, at its fixed end, has a platform 25, which defines the compressor duct 10 on the radial outside.
  • Extending from the platform 25 of the blade 16 (or of the guide vane 14) into the compressor duct 10 is a rotatable airfoil 27 (or respectively a fixed guide airfoil 28) which compress the air L during operation of the compressor 5.
  • the free ends of the airfoils 27, 28 opposite to the platformside ends form a tip region 29 of the airfoils 27, 28 and are opposite respective guide rings 30, with the radial gap 23 being formed.
  • the radial gap 23 is in each case oriented substantially parallel to the rotation axis 2 in one section.
  • the platforms 25 arranged in the section may be each inclined relative to the rotation axis 2 of the rotor 3, so that the flow duct 24 narrows as viewed in the axial direction.
  • a cylindrical contour of the flow duct 10 is obtained in the regions of the radially opposite fixed and rotating components, which as viewed in the axial direction lie in sections and in the radial direction lie inside and respectively outside the guide profiles and moving profiles, respectively.
  • both the outer guide surface 21 and the inner guide surface 22 alternately run cylindrically and in such a way as to be inclined relative to the rotation axis 2 of the rotor 3, the cylindrical guide surface 21, 22 in each case being opposite an inclined guide surface 21, 22 as viewed in the radial direction of the rotor 3.
  • the blade 16 comprises an airfoil 27.
  • the airfoil 27 has an outer wall 41 comprising a pressure sidewall or pressure surface 42 and a suction sidewall or suction surface 43.
  • the pressure and suction surfaces 42, 43 are joined together along an upstream leading edge 44 and a downstream trailing edge 45, where the leading and trailing edges 44, 45 are spaced axially or chordally from each other with respect to a chordal direction C.
  • the airfoil 27 extends radially along a longitudinal direction of the blade 16 - a radial direction of the compressor 4 -, from a radially inner airfoil platform 25 to a radially outer tip region 29.
  • the tip region 29 includes a blade tip surface 46 having an airfoil shape, and pressure tip side 51 and suction tip side 50 which are joined together at spaced apart leading tip edge 55 and trailing tip edge 56 of the tip region 29.
  • the pressure and suction tip sides 51, 50 form the radial ends of the pressure and suction surfaces 42, 43, respectively, of the airfoil 27.
  • cooling holes 52 may be provided along the trailing edge of the airfoil 27, as indicated in the figure 3 .
  • first direction D is defined in figure 3 .
  • This direction D should also apply to the following figures and defines a direction substantially perpendicular to the platform 25 and progressing with a distance to the platform 25.
  • the first direction is perpendicular to inner guide surface 22 of the platform 25.
  • FIG 4 several profiles of the tip region 29 are depicted, as it can be seen in a sectional view from the direction indicated as A-A in figure 3 .
  • FIG. 4 All embodiments of figure 4 show on the left hand side the suction surface 43 and the pressure surface 42 of the airfoil 27.
  • the platform 25 is shown in figure 4 only partially.
  • a top surface of the platform 25 - the inner guide surface 22 - defines the basis for the first direction D, which is indicated by an arrow with the accompanying reference sign D.
  • a tip region 29 is shown at a distant end of the airfoil 27 in respect to the platform 25, having a greater distance to the platform 25 in direction of the first direction D than any other part of the airfoil 27.
  • the tip region 29 is comprised of sections, a ridge surface 60 as a first surface and a tilted surface 61 as a second surface according to the claims.
  • the tilted surface 61 may be a continuous surface as shown in figures 4A and 4B or a discontinuous surface as shown in figure 4C .
  • the ridge surface 60 forms a ridge, i.e. a sharp edge. This ridge continues from the leading tip edge 55 to the trailing tip edge 56.
  • the ridge preferably is of a substantially same distance to the platform 25, independently at which location of the ridge between the leading tip edge 55 and the trailing tip edge 56 the measurement is taken.
  • the ridge preferably may be of a continuous elevation.
  • the ridge has a ridge surface expanse 62 as a first surface expanse which should be seen as the surface expanse of the edge itself, considering that the edge will not be a mathematical perfect line but being slightly round and having a little expansion in the direction of the suction and the pressure surface 43, 42.
  • the surface expanse is shown in figure 5 , which shows embodiments of a top view as seen from the direction of B-B as indicated in figure 3 , in which FIG. 5A corresponds to FIG. 4A, FIG. 5B to FIG. 4B, and FIG. 5C to FIG. 4C . With surface expanse not the projection from the top of the blade is meant even though this is shown in figure 5 , but the actual expanse of the surface shell.
  • the tilted surface 61 is the remaining surface of the tip region 29 without the ridge surface 60.
  • the tilted surface 61 has the tilted surface expanse 63 as second surface expanse, as indicated in figure 5 .
  • the ridge surface expanse 62 is less than a cross sectional plane of the airfoil 27 perpendicular to the first direction D, i.e. parallel to the platform 25. This means that the ridge surface expanse 62 has a lesser expanse than compared to an airfoil with a flat surface at the tip region 29, as it can be seen for example in figure 3 .
  • any possible plane is meant that could be defined parallel to the platform 25, e.g. a plane close to the tip region 29, a plane close to the platform 25, or any plane in between, especially a plane with the smallest possible cross section.
  • a first distance D1 between the ridge to the platform 25 is defined along the first direction D.
  • the distance may be taken from the most distant location or from an intermediate location representing the distance of the whole ridge to the platform 25.
  • a second distance D2 between the tilted surface 61 to the platform 25 is defined along the first direction D.
  • the distance may be taken from a location that may be seen representative for the tilted surface 61, so preferably an intermediate location representing the distance of the whole tilted surface 61 to the platform 25.
  • the first distance D1 is greater than the second distance D2.
  • Figures 4A and 5A show an embodiment in which the ridge surface 60 is located at a radial continuation of the pressure surface 42.
  • the tilted surface 61 extends from the suction tip side 50 to the ridge surface 60 as a slant.
  • An end of the tilted surface 61 at the suction tip side 50 has a distance less than a distance of another end of the tilted surface 61 at the ridge surface 60.
  • Figures 4B and 5B show an alternative embodiment in which the ridge surface 60 is located at a radial continuation of the suction surface 43.
  • the tilted surface 61 extends from the pressure tip side 51 to the ridge surface 60 as a slant.
  • An end of the tilted surface 61 at the pressure tip side 50 has a distance less than a distance of another end of the tilted surface 61 at the ridge surface 60.
  • FIGS 4C and 5C show an alternative embodiment in which the ridge surface 60 is located at a medium location between the suction surface 43 and the pressure surface 42. It may be right in the middle as shown in figure 4C but may also be off the centre.
  • the tilted surface 61 comprises two tilted surfaces 61A and 61B.
  • the tilted surface 61A extends from the suction tip side 50 to the ridge surface 60 as a slant.
  • An end of the tilted surface 61A at the suction tip side 50 has a distance less than a distance of another end of the tilted surface 61A at the ridge surface 60.
  • the tilted surface 61B extends from the pressure tip side 51 to the ridge surface 60 as a slant.
  • An end of the tilted surface 61B at the pressure tip side 50 has a distance less than a distance of another end of the tilted surface 61B at the ridge surface 60.
  • the slant of the tilted surface 61 may form in the cross sectional view substantially a straight line, as indicated in figure 4 , but may also be convex or concave.
  • the given profiles may be advantageous because only a fraction of the tip region 29 may be in contact with an opposing surface like the outer guide surface 21. Particularly, only the ridge surface 60 may be in contact.
  • the tip region 29 is designed to abrade.
  • it is sufficient to only coat the ridge surface 60 with an abradable or abrasive coating because this is the only surface which is in contact with the opposing surface (which itself may be coated with the "opposite" coating, i.e. with abrasive coating if the ridge surface 60 comprises an abradable coating and with abradable coating if the ridge surface 60 comprises an abrasive coating).
  • the amount of coating material may be reduced. Furthermore the force on the blades to abrade is reduced.
  • the radial gap 23 - which may be affected by manufacturing or assembly tolerances - during operation may be reduced so that also leakage of air will be reduced. This again improves the efficiency of the compressor.
  • the invention reduces the need to grind the fully assembled compressors, particularly the need to grind the rotor blades to size, because the tip region of the blade will adapt to the opposing surface simply during operation of the gas turbine.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP09012404A 2009-09-30 2009-09-30 Profil und zugehörige Leitschaufel, Laufschaufel, Gasturbine und Strömungsmaschine Withdrawn EP2309097A1 (de)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP09012404A EP2309097A1 (de) 2009-09-30 2009-09-30 Profil und zugehörige Leitschaufel, Laufschaufel, Gasturbine und Strömungsmaschine
PCT/EP2010/061625 WO2011038971A1 (en) 2009-09-30 2010-08-10 Airfoil and corresponding guide vane, blade, gas turbine and turbomaschine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP09012404A EP2309097A1 (de) 2009-09-30 2009-09-30 Profil und zugehörige Leitschaufel, Laufschaufel, Gasturbine und Strömungsmaschine

Publications (1)

Publication Number Publication Date
EP2309097A1 true EP2309097A1 (de) 2011-04-13

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EP09012404A Withdrawn EP2309097A1 (de) 2009-09-30 2009-09-30 Profil und zugehörige Leitschaufel, Laufschaufel, Gasturbine und Strömungsmaschine

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EP (1) EP2309097A1 (de)
WO (1) WO2011038971A1 (de)

Cited By (11)

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Publication number Priority date Publication date Assignee Title
WO2012168114A1 (de) * 2011-06-07 2012-12-13 Bosch Mahle Turbo Systems Gmbh & Co. Kg Leitschaufeln für eine variable turbinen- oder verdichtergeometrie
EP2573326A1 (de) * 2011-09-23 2013-03-27 United Technologies Corporation Schaufelspitzendichtungsanordnung
EP2604798A1 (de) * 2011-12-13 2013-06-19 United Technologies Corporation Bauteil eines Turbinentriebwerks und zugehöriges Herstellungsverfahren
EP2458156A3 (de) * 2010-11-24 2013-08-28 United Technologies Corporation Turbinenmotorstator, z.B. Kompressorstator
EP2514922A3 (de) * 2011-04-20 2014-08-13 General Electric Company Verdichter mit Schaufelspitzengeometrie zur Reduzierung der Spannungen
EP2952685A1 (de) * 2014-06-04 2015-12-09 United Technologies Corporation Schaufel, zugehöriges Gasturbinentriebwerk und Verfahren zur Reduzierung von Erhitzung durch Reibung
EP2952686A1 (de) * 2014-06-04 2015-12-09 United Technologies Corporation Laufschaufel, zugehöriges Gasturbinentriebwerk und Herstellungsverfahren
EP2935790B1 (de) * 2012-12-19 2019-05-01 Rolls-Royce PLC Verbundschaufelstruktur mit einem schneidekantenspitzenteil
US10697311B2 (en) 2014-11-20 2020-06-30 Mitsubishi Heavy Industries, Ltd. Turbine blade and gas turbine
US11066937B2 (en) 2014-06-04 2021-07-20 Raytheon Technologies Corporation Cutting blade tips
EP3882437A1 (de) * 2020-03-20 2021-09-22 Raytheon Technologies Corporation Integral beschaufelter rotor, gasturbinentriebwerk und verfahren zur herstellung eines integral beschaufelten rotors

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Publication number Priority date Publication date Assignee Title
EP3477059A1 (de) * 2017-10-26 2019-05-01 Siemens Aktiengesellschaft Kompressorschaufel
GB201900961D0 (en) 2019-01-24 2019-03-13 Rolls Royce Plc Fan blade
WO2023242949A1 (ja) * 2022-06-14 2023-12-21 三菱重工業株式会社 圧縮機の動翼及び圧縮機

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US4671735A (en) 1984-01-19 1987-06-09 Mtu-Motoren-Und Turbinen-Union Munchen Gmbh Rotor of a compressor, more particularly of an axial-flow compressor
EP0291407A1 (de) * 1987-05-13 1988-11-17 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Bewegbare Gebläseschaufel mit einer Schneidkante am Ende
JPH03122202U (de) * 1990-03-27 1991-12-13
GB2282856A (en) * 1993-10-15 1995-04-19 United Technologies Corp Reducing stress on the tips of turbine or compressor blades
US5456576A (en) * 1994-08-31 1995-10-10 United Technologies Corporation Dynamic control of tip clearance
US6086328A (en) * 1998-12-21 2000-07-11 General Electric Company Tapered tip turbine blade
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EP1770244A1 (de) * 2005-09-30 2007-04-04 Snecma Verdichterschaufel mit angefaster Spitze
EP1930547A2 (de) 2006-11-24 2008-06-11 IHI Corporation Kompressorschaufel für eine Gasturbine

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Publication number Priority date Publication date Assignee Title
GB1107024A (en) * 1965-11-04 1968-03-20 Parsons C A & Co Ltd Improvements in and relating to blades for turbo-machines
US4295786A (en) 1978-08-04 1981-10-20 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Composite seal for turbomachinery
US4526509A (en) 1983-08-26 1985-07-02 General Electric Company Rub tolerant shroud
US4671735A (en) 1984-01-19 1987-06-09 Mtu-Motoren-Und Turbinen-Union Munchen Gmbh Rotor of a compressor, more particularly of an axial-flow compressor
EP0291407A1 (de) * 1987-05-13 1988-11-17 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Bewegbare Gebläseschaufel mit einer Schneidkante am Ende
JPH03122202U (de) * 1990-03-27 1991-12-13
GB2282856A (en) * 1993-10-15 1995-04-19 United Technologies Corp Reducing stress on the tips of turbine or compressor blades
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