US20130156562A1 - Turbomachine and turbomachine stage - Google Patents

Turbomachine and turbomachine stage Download PDF

Info

Publication number
US20130156562A1
US20130156562A1 US13/687,789 US201213687789A US2013156562A1 US 20130156562 A1 US20130156562 A1 US 20130156562A1 US 201213687789 A US201213687789 A US 201213687789A US 2013156562 A1 US2013156562 A1 US 2013156562A1
Authority
US
United States
Prior art keywords
cascade
contour
recited
guide vane
turbomachine
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.)
Abandoned
Application number
US13/687,789
Other languages
English (en)
Inventor
Inga Mahle
Jochen Gier
Kai Koerber
Michael ENGELKE
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.)
MTU Aero Engines AG
Original Assignee
MTU Aero Engines GmbH
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 MTU Aero Engines GmbH filed Critical MTU Aero Engines GmbH
Publication of US20130156562A1 publication Critical patent/US20130156562A1/en
Assigned to MTU AERO ENGINES GMBH reassignment MTU AERO ENGINES GMBH ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: ENGEL, KARL, GIER, JOCHEN, KOERBER, KAI, MAHLE, INGA
Abandoned legal-status Critical Current

Links

Images

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
    • F01D9/00—Stators
    • 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
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/142—Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • F01D5/143—Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/184—Two-dimensional patterned sinusoidal
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/60—Structure; Surface texture
    • F05D2250/61—Structure; Surface texture corrugated
    • F05D2250/611—Structure; Surface texture corrugated undulated
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to a turbomachine, in particular a gas turbine, preferably an aircraft engine gas turbine, having at least one turbomachine stage, in particular a compressor stage or a turbine stage, including a guide vane cascade and a rotor blade cascade, and to such a turbomachine stage.
  • a turbomachine stage has a cascade of rotating rotor blades and a cascade of guide vanes disposed adjacent to the rotor blade cascade on the upstream or downstream side.
  • the rotor blades terminate in a radially outer airfoil platform at the root end.
  • guide vanes may be provided at the tip end with a radially outer airfoil platform, for example, in the form of a shroud.
  • An axial gap is formed between the guide vane cascade and the rotor blade cascade.
  • pressure gradients are formed therein, the pressure gradients varying around the circumference and causing secondary flows.
  • a rotating cascade of turbine rotor blades may force working fluid into the axial gap on its pressure side and, conversely, draw working fluid from the gap on its suction side.
  • a compensating flow is generated, which degrades the efficiency of the turbomachine.
  • a gas turbine having shroudless rotor blades is disclosed in EP 2 372 102 A2, which proposes that the radially inner platforms of guide vanes and rotor blades have a non-axisymmetric contour, in particular a radially and/or axially undulated contour.
  • European Patent Publication EP 2 136 033 A1 discloses a turbomachine stage having guide vanes which, together with radially inner and/or radially outer airfoil platforms, form a guide vane cascade.
  • the turbomachine further has rotor blades which, together with radially inner and/or radially outer airfoil platforms, form a rotor blade cascade.
  • an axial gap which is bounded by gap regions of the airfoil platforms of the rotor blade cascade and the guide vane cascade.
  • European Patent Publication EP 1 067 273 A1 describes a turbomachine stage having a rotor blade.
  • the radially outer airfoil platform associated with the rotor blade has a contour which varies in the axial direction, while the radially outer airfoil platform associated with a guide vane has a contour which does not vary in the axial direction.
  • a turbomachine stage according to the present invention includes a plurality of rotor blades which are preferably equidistantly distributed around the circumference and, at their root or rotor end, are connected to, in particularly integrally formed with, radially inner airfoil platforms. At their tip or casing ends, the rotor blades may be connected to, in particularly integrally formed with, radially outer airfoil platforms. Rotor blades may be removably or non-removably attached to, in particular integrally formed with, a rotor (member) of the turbomachine, either individually or in groups.
  • a plurality of guide vanes are preferably equidistantly distributed around the circumference and removably or non-removably attached to, in particular integrally formed with, a casing (member) of the turbomachine.
  • the guide vanes are connected to, in particular integrally formed with, radially outer airfoil platforms.
  • the guide vanes may be connected to, particularly integrally formed with, radially inner airfoil platforms.
  • the guide vanes may be connected to, in particular integrally formed with, radially inner airfoil platforms.
  • the airfoil platforms may extend axially beyond these cascade regions on the upstream and/or downstream side(s); i.e., beyond the airfoil leading and/or trailing edges.
  • These regions of the airfoil platforms bound an axial gap extending axially between the guide vane cascade and the rotor blade cascade and, therefore, are hereinafter collectively referred to as gap regions of the airfoil platforms.
  • An airfoil platform may have radially outer gap regions including a plurality of sections.
  • the radially outer airfoil platforms of a rotor blade cascade or a guide vane cascade may have one or more radial shoulders whose circumferential surfaces bound the axial gap radially and whose end faces bound the axial gap axially.
  • the following explanations may refer to one or more, in particular to all of the sections of a gap region.
  • the contour(s) of one or more circumferential surfaces may vary in the radial direction and/or the contour(s) of one or more end faces may vary in the axial direction.
  • the casing member in which the recess is formed may form a radially outer gap region of the guide vane platforms according to the present invention.
  • a radially outer rotor blade platform which may be disposed in particular in a recess of the casing may form a gap region according to the present invention.
  • a component which is radially outwardly connected to, or integrally formed with, at least one guide vane or rotor blade and whose contour, possibly together with additional contours, radially and/or axially bounds the axial gap between the rotor blade cascade and the guide vane cascade, may constitute a radially outer gap region of an airfoil platform according to the present invention.
  • a contour of one or more of these gap regions varies in the radial and/or axial direction around the circumference.
  • a variation in the radial direction is understood, in particular, to be an outside radius R of the contour which, in polar coordinates, varies with the circumferential angle ⁇ around the axis of rotation of the turbomachine stage, and analogously, a variation in the axial direction is understood, in particular, to be an axial coordinate X of the contour which varies with the circumferential angle.
  • the contour varies periodically, in particular sinusoidally:
  • R ( ⁇ ) R 0 + ⁇ R ⁇ sin( ⁇ R ⁇ + ⁇ R )and/or
  • this variation may be formed solely in the radial direction, solely in the axial direction, or in both the axial and radial directions.
  • the contour of a cylindrical gap region having a smooth end face and an undulated circumferential surface varies solely in the radial direction, that of a cylindrical gap region having an undulated end face and a smooth circumferential surface varies solely in the axial direction, while that of a cylindrical gap region having an undulated end face and an undulated circumferential surface and that of conical gap region having an undulated circumferential surface vary in both the axial and radial directions.
  • the undulation may be formed solely on one or more gap regions of radially outer guide vane platforms, solely on one or more gap regions of radially outer rotor blade platforms, or also on one or more gap regions of radially outer platforms of both guide vanes and rotor blades.
  • an undulation may additionally be provided on one or more gap regions of radially inner guide vane platforms and/or on one or more gap regions of radially inner rotor blade platforms.
  • a contour of a gap region of an airfoil platform of one of the guide vane and rotor blade cascades and an axially and/or radially opposite contour of a gap region of an airfoil platform of the other of the guide vane and rotor blade cascades may vary around the circumference, preferably identically, in particular in parallel, or with a phase offset of preferably at least 45°, in particular at least 90°, preferably at least 135° and/or preferably of no more than 270°, in particular no more than 210°, and preferably no more than 180°.
  • a gap region has two opposite contours, such as an inner and an outer circumferential surface of an annular flange such as, in particular, a shroud extension, or on a casing, then these two opposite contours may vary around the circumference, preferably differently or identically, in particular in parallel, or with a phase offset of preferably at least 45°, in particular at least 90°, preferably at least 135° and/or preferably of no more than 270°, in particular no more than 210°, and preferably no more than 180°. If the two contours vary in parallel, the wall thickness of this gap region of the airfoil platform remains constant. It may equally be provided that only one of such opposite contours, in the case of an annular-flange-like shroud extension preferably the radially inner contour, varies while the other remains constant around the circumference.
  • an entire contour of a gap region may vary around the circumference. It is equally possible that only a section of the contour has an undulation.
  • the inner circumferential surface of an annular flange may vary in the radial direction only in one or more axial sections, or an end face may vary in the axial direction only in one or more radial sections.
  • a radial variation of a contour of a gap region of an airfoil platform of a cascade may be constant in the axial direction, so that troughs and crests are oriented parallel to the axis of rotation of the turbomachine stage.
  • a radial variation of a contour of a gap region of an airfoil platform of a cascade may also vary in the axial direction, so that troughs and crests extend at an angle to the axis of rotation.
  • a phase offset may be provided which varies with the axial position x, preferably linearly:
  • R ( ⁇ , x ) R 0 + ⁇ R ⁇ sin( ⁇ R ⁇ + ⁇ R ⁇ x )
  • an axial variation of a contour of a gap region of an airfoil platform of a cascade may be constant in the radial direction, so that troughs and crests are oriented perpendicularly to the axis of rotation of the turbomachine stage.
  • an axial variation of a contour of a gap region of an airfoil platform of a cascade may also vary in the radial direction, so that troughs and crests are inclined at an angle to the axis of rotation.
  • a phase offset may be provided which varies with the radial position r, preferably linearly:
  • the cascade region of the airfoil platform varies as well, at least partially, around the circumference in one of the ways described above.
  • a gap region whose contour varies around the circumference merges smoothly into this cascade region, especially in such a way that a trough of the gap region contour merges into a trough of the cascade region, and a crest of the gap region contour merges into a crest of the cascade region.
  • smooth transition is used, in particular, to refer to a transition which has no sharp edges or bends, but which preferably has a continuous curvature.
  • an extreme extent i.e., a maximum or minimum extent of a radially varying contour of a gap region of an airfoil platform of a cascade is circumferentially located in the pressure-side half of the segment between two adjacent airfoil leading edges or in the suction-side half of the segment between two adjacent airfoil trailing edges of the cascade in order to compensate the pressure increases and decreases induced there.
  • two airfoil leading or trailing edges of a guide vane or rotor blade cascade define a segment therebetween which extends in the circumferential direction and is divided into two halves by the channel center.
  • the segment half adjoining the pressure side of the airfoil is referred to as “pressure-side half”, the other one as “suction-side half” accordingly.
  • These halves define a circumferential angular range in which, according to the present invention, an extreme extent of a varying contour is disposed. Since the varying contour does not lie at the axial height of this segment itself, this segment may be imagined as being displaced parallel to an extension of the mean camber line of the airfoil for the positioning of the extreme extent.
  • a maximum variation in the radial direction of gap region of an airfoil platform of a cascade is no more than 50%, particularly no more than 40% of the pitch of the cascade.
  • an extreme i.e., a maximum or minimum extent of an axially varying contour of a gap region may be circumferentially located in the region of an airfoil edge, in particular circumferentially spaced from the airfoil edge by no more than 25% of the cascade pitch.
  • a maximum extent of a guide vane platform in the axial direction extends preferably axially toward the rotor blade cascade and, analogously, a maximum extent of a rotor blade platform in the axial direction extends preferably axially toward the guide vane cascade.
  • a maximum variation in the axial direction of gap region of an airfoil platform of a cascade is no more than 50%, particularly no more than 40% of the pitch of the cascade.
  • FIG. 1 a developed view of a portion of a gas turbine stage according to the present invention including a guide vane cascade and a rotor blade cascade having radially outer airfoil platforms whose gap region contours vary in the axial direction around the circumference;
  • FIGS. 2A , 2 B meridional sections at different circumferential positions through a gas turbine stage according to the present invention including a guide vane cascade and a rotor blade cascade having radially outer airfoil platforms whose gap region contour varies in the radial direction around the circumference; and
  • FIG. 3 a meridional section, similar to those of FIG. 2 , through a gas turbine stage according to the present invention, in which a gap region of a radially outer guide vane platform has a recess formed therein for receiving radially outer rotor blade platforms.
  • FIG. 1 shows a developed view of a portion of a gas turbine stage according to the present invention, as seen from an axis of rotation; i.e., viewed from radially inside to radially outside, showing a stationary cascade of guide vanes 1 and, opposite thereto, a rotating cascade of rotor blades 2 .
  • the rotation is indicated by a filled vertical arrow, the flow of working fluid is indicated by an empty arrow in the region of the guide vane cascade.
  • This configuration is merely exemplary for purposes of illustration.
  • the present invention may be used equally in turbine and compressor stages, where the guide vane cascade is disposed upstream and/or downstream of the rotor blade cascade.
  • Integrally formed with airfoils 1 , 2 are radially outer airfoil platforms, which are shown from above in FIG. 1 ; i.e., as viewed from the axis of rotation of the gas turbine stage.
  • Each airfoil may either have a separate airfoil platform, or several or all of the airfoils of a cascade may be connected to, in particular integrally formed with, the same airfoil platform which, in accordance with the present invention, may then be imagined as being divided into separate airfoil platforms associated with the individual airfoils. Therefore, FIG. 1 does not show any airfoil platform boundaries in the circumferential direction (vertically in FIG. 1 ).
  • the radially outer platforms of guide vanes 1 may be, for example, a part, in particular an integral part, of a casing of a gas turbine (stage), or be attached to such a casing.
  • the radially outer platforms of rotor blades 2 may be, for example, shrouds, in particular interconnected shrouds.
  • a cascade region 10 . 1 of the guide vane platforms and a cascade region 20 . 1 of the rotor blade platforms extend axially between the respective leading edge (left in FIG. 1 ) and the respective trailing edge (right in FIG. 1 ), said cascade regions being hatched from top left to bottom right in FIG. 1 .
  • Gap regions 10 . 2 T and 20 . 2 L each have substantially the shape of a radial shoulder whose circumferential surface facing toward the spoke-like pattern of the respective cascade and whose end face facing toward the respective other airfoil cascade radially and axially bound a radially inner axial gap A between the rotor blade cascade and the guide vane cascade.
  • this gap region 10 . 2 T, respectively 20 . 2 L, and more particularly its end face facing the respective other airfoil cascade varies in the axial direction around the circumference; i.e., in the vertical direction in FIG. 1 . That is, the generating lines of the end face, which extend from the axis of rotation of the turbomachine to the peripheral edge of the radial shoulder, have different axial positions, so that the end face has a maximum axial extent A max 10 , respectively A max 20 , toward the respective other airfoil cascade at selected circumferential positions, as measured from a generating line which is axially farthest away from the respective other airfoil cascade.
  • the generating lines may be perpendicular to the axis of rotation of the turbomachine, or inclined thereto at the same angle or at an angle that varies in the circumferential direction.
  • the generating lines are perpendicular to the axis of rotation.
  • Their axial position varies sinusoidally around the circumference, so that maximum axial extents A max 20 of gap region 20 . 2 L of the rotor blade platforms are disposed near respective leading edges of rotor blades 2 , and maximum axial extents A max 10 of gap region 10 . 2 T of the guide vanes are disposed near respective trailing edges of guide vanes 1 , as viewed in the circumferential direction.
  • Maximum extents A max 10 and A max 20 are each 50% of the respective cascade pitch.
  • FIGS. 2A , 2 B show meridional sections at different circumferential positions through a gas turbine stage according to the present invention including a guide vane cascade and a rotor blade cascade having radially outer airfoil platforms whose gap region contour varies in the radial direction around the circumference.
  • this gas turbine stage may be the one described hereinabove with reference to FIG. 1 , so that a radial undulation is combined with an axial undulation. Therefore, in the following, reference is made to the above description and only the aspects of the radial undulation will be described. It is equally possible to provide only an axial undulation, as described hereinabove with reference to FIG. 1 , or only a radial undulation, such as will be described hereinafter.
  • the maximum positive amplitude i.e., the maximum radial extent R max 20 in the radially outward direction, is circumferentially located in the pressure-side half of the segment between two successive rotor blade leading edges. Again, the positions in the circumferential direction may be imagined as being displaced parallel to the extension of the mean camber line to the respective axial positions.
  • FIG. 2A shows the trough R min 20 sloping upwardly in the direction of the flow
  • FIG. 2B shows the crests R max 20 sloping downwardly in the direction of the flow.
  • this radial undulation of gap region 20 . 2 merges smoothly into a corresponding undulation of cascade region 20 . 1 between rotor blades 2 .
  • the gap region 10 . 2 T of the guide vane platforms which faces the rotor blade cascade, may also have a radial undulation, such as described hereinabove with reference to gap region 20 . 2 of the rotor blade platforms.
  • this trailing-edge gap region is shaped like an annular flange, and therefore has two radially opposite surfaces (upper and lower in FIG. 2 ).
  • a radial undulation may in particular be provided on the radially inner surface (lower in FIG. 2 ) or on both surfaces. In the latter case, preferably, they vary identically, so that the wall thickness of the annular flange remains constant.
  • FIG. 3 shows, in a view similar to that of FIG. 2 , a portion of a gas turbine stage according to a modified embodiment of the present invention.
  • Corresponding elements are identified by the same reference numerals, so that reference is made to the above explanations in their entirety, and only the differences in the modified embodiment will be discussed below.
  • FIG. 3 shows the downstream trailing edge of a rotor blade 2 and the upstream leading edge of a following guide vane 1 .
  • the radially outer gap regions are designated 20 . 2 T (for trailing edge) and 10 . 2 L (for leading edge, and further with an M or S designation as discussed below) to illustrate by way of example that the explanations are equally applicable to leading and trailing edges of rotor blade platforms and guide vane platforms, respectively.
  • the radially outer airfoil platforms are inclined at an angle to the turbine axis to illustrate a divergent flow channel.
  • the explanations apply equally to convergent flow channels (not shown), in particular in compressor stages.
  • the radially outer rotor blade platform which is in the form of a shroud, has an annular flange formed in its trailing-edge gap region 20 . 2 T.
  • This annular flange has radially opposite circumferential surfaces, such as described hereinbefore with reference to trailing-edge gap region 10 . 2 T of guide vane 1 of FIG. 1 .
  • Trailing-edge gap region 20 . 2 T may also have an undulation, in particular the same undulation, on its radially inner (lower in FIG. 3 ) and/or outer circumferential surface.
  • the leading-edge gap region includes a radially inner annular flange in a radially outer groove-like recess of the gas turbine casing. Accordingly, the leading-edge gap region has three circumferential surfaces 10 . 2 LM, namely the radially inner and radially outer circumferential surfaces of the annular flange and the circumferential surface of the recess itself, as well as two end faces 10 . 2 LS, namely that of the annular flange and that of the recess itself.
  • each of these sections 10 . 2 LM, 10 . 2 LS may have an undulation in the radial direction ( 10 . 2 LM) and in the axial direction ( 10 . 2 LS), respectively.
  • the modification shown in FIG. 3 is intended to illustrate in one view different variants where a contour of a gap region may vary in the radial and/or axial direction around the circumference.

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US13/687,789 2011-12-20 2012-11-28 Turbomachine and turbomachine stage Abandoned US20130156562A1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EP11194433.6A EP2607625B1 (de) 2011-12-20 2011-12-20 Turbomaschine und turbomaschinenstufe
EPEP11194433.6 2011-12-20

Publications (1)

Publication Number Publication Date
US20130156562A1 true US20130156562A1 (en) 2013-06-20

Family

ID=45442918

Family Applications (1)

Application Number Title Priority Date Filing Date
US13/687,789 Abandoned US20130156562A1 (en) 2011-12-20 2012-11-28 Turbomachine and turbomachine stage

Country Status (3)

Country Link
US (1) US20130156562A1 (de)
EP (1) EP2607625B1 (de)
ES (1) ES2891562T3 (de)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20140140822A1 (en) * 2012-11-16 2014-05-22 General Electric Company Contoured Stator Shroud
US20160061111A1 (en) * 2014-08-29 2016-03-03 MTU Aero Engines AG Gas turbine subassembly
US20180252107A1 (en) * 2017-03-03 2018-09-06 MTU Aero Engines AG Contouring a blade/vane cascade stage
US20190106995A1 (en) * 2017-10-11 2019-04-11 Doosan Heavy Industries & Construction Co., Ltd. Compressor and gas turbine including the same
WO2021148607A1 (fr) * 2020-01-24 2021-07-29 Safran Aircraft Engines Basculement ondulé de plateformes aux entrefers rotor-stator dans un compresseur de turbomachine
EP3913192A1 (de) * 2018-12-14 2021-11-24 Rolls-Royce plc Eiskristallschutz für einen gasturbinenmotor

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7320574B2 (en) * 2003-09-09 2008-01-22 Alstom Technology Ltd Turbomachine
US20110243749A1 (en) * 2010-04-02 2011-10-06 Praisner Thomas J Gas turbine engine with non-axisymmetric surface contoured rotor blade platform

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB9915648D0 (en) * 1999-07-06 1999-09-01 Rolls Royce Plc Improvement in or relating to turbine blades
DE10327977A1 (de) * 2003-06-21 2005-01-05 Alstom Technology Ltd Seitenwandgestaltung eines umlenkenden Strömungskanals
US7249928B2 (en) * 2005-04-01 2007-07-31 General Electric Company Turbine nozzle with purge cavity blend
JP5283855B2 (ja) * 2007-03-29 2013-09-04 株式会社Ｉｈｉ ターボ機械の壁、及びターボ機械
DE102007027427A1 (de) * 2007-06-14 2008-12-18 Rolls-Royce Deutschland Ltd & Co Kg Schaufeldeckband mit Überstand
US8356975B2 (en) * 2010-03-23 2013-01-22 United Technologies Corporation Gas turbine engine with non-axisymmetric surface contoured vane platform

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7320574B2 (en) * 2003-09-09 2008-01-22 Alstom Technology Ltd Turbomachine
US20110243749A1 (en) * 2010-04-02 2011-10-06 Praisner Thomas J Gas turbine engine with non-axisymmetric surface contoured rotor blade platform

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20140140822A1 (en) * 2012-11-16 2014-05-22 General Electric Company Contoured Stator Shroud
US20160061111A1 (en) * 2014-08-29 2016-03-03 MTU Aero Engines AG Gas turbine subassembly
US9822706B2 (en) * 2014-08-29 2017-11-21 MTU Aero Engines AG Gas turbine subassembly
US20180252107A1 (en) * 2017-03-03 2018-09-06 MTU Aero Engines AG Contouring a blade/vane cascade stage
US10648339B2 (en) * 2017-03-03 2020-05-12 MTU Aero Engines AG Contouring a blade/vane cascade stage
US20190106995A1 (en) * 2017-10-11 2019-04-11 Doosan Heavy Industries & Construction Co., Ltd. Compressor and gas turbine including the same
US11162373B2 (en) * 2017-10-11 2021-11-02 Doosan Heavy Industries & Construction Co., Ltd. Compressor and gas turbine including the same
EP3913192A1 (de) * 2018-12-14 2021-11-24 Rolls-Royce plc Eiskristallschutz für einen gasturbinenmotor
WO2021148607A1 (fr) * 2020-01-24 2021-07-29 Safran Aircraft Engines Basculement ondulé de plateformes aux entrefers rotor-stator dans un compresseur de turbomachine
US11846194B2 (en) 2020-01-24 2023-12-19 Safran Aircraft Engines Wavy tilting of platforms at the rotor-stator gaps in a turbine engine compressor

Also Published As

Publication number Publication date
EP2607625B1 (de) 2021-09-08
ES2891562T3 (es) 2022-01-28
EP2607625A1 (de) 2013-06-26

Similar Documents

Publication Publication Date Title
EP2820279B1 (de) Turbomaschinenschaufel
JP5154165B2 (ja) エアフォイル形状を有するブレードエアフォイルを備えるタービンブレードシステム、および該タービンブレードシステム用のリングプラットフォーム
US7690890B2 (en) Wall configuration of axial-flow machine, and gas turbine engine
US20130156562A1 (en) Turbomachine and turbomachine stage
EP2187000B1 (de) Kaskadenendwand für eine turbinenschaufel
US8894364B2 (en) Aerofoil blade for an axial flow turbomachine
US9638040B2 (en) Blade of a row of rotor blades or stator blades for use in a turbomachine
US10934849B2 (en) Endwall contouring for a turbomachine
US9512727B2 (en) Rotor of an axial compressor stage of a turbomachine
US9863251B2 (en) Turbomachine and turbomachine stage
US10458248B2 (en) Blade channel, blade cascade and turbomachine
US9822795B2 (en) Stator of an axial compressor stage of a turbomachine
US9745859B2 (en) Radial-inflow type axial flow turbine and turbocharger
US20170218773A1 (en) Blade cascade and turbomachine
US9945232B2 (en) Gas turbine blade configuration
CN110873075B (zh) 用于涡轮机的压缩机的具有突起的叶片
US20110097205A1 (en) Turbine airfoil-sidewall integration
US10704393B2 (en) Contouring of an airfoil array platform
US8777564B2 (en) Hybrid flow blade design
US8721289B2 (en) Flow balancing slot
US11293289B2 (en) Shrouded blades with improved flutter resistance
WO2019160547A1 (en) Assembly of turbine blades and corresponding article of manufacture
US10450869B2 (en) Gas turbine compressor
US11686207B2 (en) Gas turbine compressor
US20230383662A1 (en) Annulus contouring

Legal Events

Date Code Title Description
AS Assignment

Owner name: MTU AERO ENGINES GMBH, GERMANY

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:MAHLE, INGA;GIER, JOCHEN;KOERBER, KAI;AND OTHERS;SIGNING DATES FROM 20130703 TO 20130807;REEL/FRAME:031063/0049

STCB Information on status: application discontinuation

Free format text: ABANDONED -- AFTER EXAMINER'S ANSWER OR BOARD OF APPEALS DECISION