EP2927503B1 - Compresseur de turbine à gaz, moteur d'avion et méthode de dimensionnement - Google Patents

Compresseur de turbine à gaz, moteur d'avion et méthode de dimensionnement Download PDF

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Publication number
EP2927503B1
EP2927503B1 EP14163465.9A EP14163465A EP2927503B1 EP 2927503 B1 EP2927503 B1 EP 2927503B1 EP 14163465 A EP14163465 A EP 14163465A EP 2927503 B1 EP2927503 B1 EP 2927503B1
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EP
European Patent Office
Prior art keywords
groove
upstream
edge
downstream
blade tip
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.)
Active
Application number
EP14163465.9A
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German (de)
English (en)
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EP2927503A1 (fr
Inventor
Giovanni Brignole
Tobias Mayenberger
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MTU Aero Engines AG
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MTU Aero Engines AG
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Publication date
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Priority to EP14163465.9A priority Critical patent/EP2927503B1/fr
Priority to US14/672,959 priority patent/US10450869B2/en
Publication of EP2927503A1 publication Critical patent/EP2927503A1/fr
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Publication of EP2927503B1 publication Critical patent/EP2927503B1/fr
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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
    • F01D5/141—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
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • 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
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/522—Casings; Connections of working fluid for axial pumps especially adapted for elastic fluid pumps
    • F04D29/526—Details of the casing section radially opposing blade tips
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • F04D29/685—Inducing localised fluid recirculation in the stator-rotor interface
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • 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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor

Definitions

  • the present invention relates to a gas turbine compressor and an aircraft engine with such a gas turbine compressor and a method for designing such a gas turbine compressor.
  • a gas turbine compressor with casing structuring (“casing treatment” CT) is known.
  • This comprises a circumferential or interrupted circumferential groove which is arranged axially in the area of a circumferential blade tip and in which deflection means in the form of webs are arranged.
  • the lands have random radial cutbacks.
  • the WO 2004/018844 A1 discloses a recirculation structure for turbo compressors, with an annular chamber arranged in the area of the free blade ends of a blade ring, radially adjoining the main flow channel, and with a large number of guide elements arranged in the annular chamber and distributed over its circumference, with the annular chamber in the front and/or rear area allows a flow passage in the circumferential direction, and the guide elements are firmly connected to at least one wall of the annular chamber and are otherwise designed to be free-standing.
  • the US 6,290,458 B1 relates to a turbomachine having a casing on the inner airfoil of which a plurality of first grooves are formed to connect an inlet side of blades of an impeller and a portion of the airfoil where the blades of the impeller lie.
  • An object of an embodiment of the present invention is to improve a gas turbine compressor.
  • a gas turbine compressor in particular an axial compressor, has one or more blades arranged next to one another in the circumferential direction with blade tips, in particular without shrouds, and a flow channel wall lying radially opposite thereto.
  • the gas turbine compressor is a gas turbine compressor for an aircraft engine or aircraft engine; in particular, it can be a low-pressure compressor arranged in a gas turbine upstream of a further gas turbine compressor or a high-pressure compressor arranged downstream of a further gas turbine compressor.
  • the blades are available in one version rotating blades arranged on a rotatably mounted rotor and rotating during operation, the radially outer blade tips of which are opposite the flow channel wall fixed to the housing radially on the outside.
  • the vanes are guide vanes fixed to the housing, which the rotatably mounted flow channel wall, which revolves during operation, is opposite radially on the inside.
  • a circumferential groove is arranged in the flow channel wall.
  • this has an upstream groove flank which merges into the flow channel wall in an upstream groove edge, a downstream groove flank which merges into the flow channel wall in a downstream groove edge, and a groove base connecting these groove flanks.
  • a groove edge can be sharp-edged or angular or also rounded or have a radius, in which case its center point or intersection point of its two outermost tangents can then define the groove edge for dimensions.
  • the upstream groove flank and/or the downstream groove flank has an axial undercut whose cross-sectional area in a meridian section is less than 10% of a cross-sectional area of the circumferential groove between its upstream and downstream groove edge.
  • a meridional section within the meaning of the present invention is a planar section that contains the axis of rotation of the compressor.
  • An axial undercut of the upstream groove flank is a region of this groove flank which is arranged in the axial direction upstream in front of the upstream groove edge.
  • an axial undercut of the downstream groove flank is a region of this groove flank which is arranged downstream in the axial direction behind the downstream groove edge.
  • a cross-sectional area of the circumferential groove between its upstream and downstream groove edge is correspondingly the area bounded in the meridian section by the groove bottom, a straight line connecting the upstream and downstream groove edge and perpendicular through the upstream and downstream groove edge.
  • the circumferential groove extends, in particular continuously or without interruption, over the full circumference of the flow channel wall or over 360°.
  • the upstream and downstream groove edges are each a continuous edge that extends over 360° without interruption. In this way, in one embodiment, the manufacture and/or aerodynamics of the circumferential groove can be improved.
  • One or more webs are arranged in the circumferential groove.
  • Several adjacent webs, in particular all webs can be designed in the same way in one embodiment, in particular have at least essentially identical dimensions and contours. In this way, in one embodiment, the manufacture and/or aerodynamics of the circumferential groove can be improved.
  • adjacent webs can be designed in different ways in one embodiment, in particular have different dimensions and/or contours.
  • asymmetries can be specifically represented or compensated for in one embodiment.
  • Three or more, in particular all, webs can be spaced equidistantly in the circumferential direction. Equally, three or more, in particular all, webs can have different distances from one another in pairs in the circumferential direction.
  • a radial cutback is understood to mean in particular an empty space between a blade-side end face of the web and its projection into a reference surface that extends from the upstream groove edge to the downstream groove edge, the curvature of the reference surface in the meridian sections through the end face being equal to infinity or to of the upstream and downstream groove edges is equal to the curvature of the flow passage wall and continuously linear therebetween in the axial direction.
  • radial cutback is understood to correspond to the free area between a blade tip-side upper edge of the cross section of the web and a reference curve that extends from the upstream groove edge to the downstream groove edge, the curvature of the reference curve being equal to infinity or at the upstream and downstream groove edges equal to the curvature of the flow channel wall and continuously linear therebetween in the axial direction.
  • a radial cutback is understood in one embodiment to be the empty space or the free area between the blade-side end face or upper edge of the web and a flow channel contour that is virtually continued beyond the circumferential groove, with this virtually continued contour connecting the groove edges with a curvature that corresponds to the groove edges of the curvature of the flow channel contour and is linearly interpolated between them.
  • an upstream beginning of the cutback is axially downstream of the upstream groove edge between this groove edge and the upstream leading edge of the blade tip and a downstream end of the Cutback arranged in a blade tip closer half of a radial height of the circumferential groove.
  • a cutback according to the invention which begins downstream after the upstream groove edge and upstream before the upstream leading edge of the blade tip and ends in the half of the circumferential groove closer to the blade tip, in one embodiment has the advantages of the housing structuring in non-design operation ("off-design "), at least essentially, can be maintained, while at the same time undesired flow phenomena can be reduced in design operation or under nominal operating conditions.
  • an upstream start of the cutback is understood to mean that axial position from which the blade-side end face or upper edge of the web deviates from the virtually continued flow channel contour or the reference surface or curve from the blade tip away to the groove base.
  • an upstream start of the cutback is understood to mean that axial position from which the blade-side end face or upper edge of the web deviates from the straight reference surface or curve in the radial direction towards the groove base by at least 1%, in particular at least 5% of a maximum radial distance between a groove edge closer to the blade tip and the groove base.
  • the upstream beginning of the cutback is located axially downstream of the upstream groove edge and upstream of the upstream leading edge of the blade tip.
  • the blade-side face (or in one or more, preferably all, meridian sections through the blade-tip-side face of the web the top edge) of the web continues the flow channel contour with a constant curvature or without an abrupt change in curvature.
  • a downstream end of the cutback is understood to mean that axial position at which the blade-side end face or upper edge of the web again opens into the reference surface or reference curve or into the downstream groove flank.
  • a downstream end of the cutback is understood to mean that axial position from which the blade-side front side or upper edge of the web deviates from the straight reference surface or curve to the groove base in the radial direction by less than 5%, in particular less than 1 % of the maximum radial distance between the groove edge closest to the blade tip and the groove base.
  • the downstream end of the cutback is arranged in a half of a radial height of the circumferential groove that is closer to the blade tip.
  • a radial height of the circumferential groove means in particular a maximum distance between the groove base and the reference surface or reference curve, i.e. in particular a maximum distance between the groove base and the groove edge closer to the blade tip, in the radial direction or in a direction perpendicular to the connecting line of the understood upstream and downstream groove edge, such a distance perpendicular to the connecting line is generally referred to as the radial height of the circumferential groove.
  • the radial cut-back ends in the reference surface or curve, in a further development axially downstream behind the upstream one leading edge of the blade tip.
  • the blade-side face (or in one or more, preferably all, meridian sections through the blade-tip-side face of the web the top edge) of the web sets the flow channel contour with a continuous curvature or without an abrupt change in curvature from the downstream one groove edge upstream.
  • the radial cut-back ends in the radially upper half of the downstream groove flank, and the web is radially cut-back continuously from the start of the cut-back.
  • the radially upper half is generally referred to as the part of the downstream groove flank that extends in the radial direction or in a direction perpendicular to the line connecting the upstream and downstream groove edges over 50% of the maximum distance of the downstream groove edge from the groove base in this direction.
  • the web opens into the upstream and downstream groove flank of the circumferential groove, it thus extends axially through the groove or its maximum axial length.
  • a blade-tip-side top edge of the web at the upstream groove edge has the same curvature as the flow channel contour, i.e. a constant curvature at the upstream groove edge, and this continue until the start of the cut back.
  • the web can be or run straight or curved.
  • the end face of the web on the blade side can open out, at least essentially, axially into the upstream groove edge.
  • the end face on the blade side can open into the downstream groove flank in a curved manner in or counter to a direction of rotation of the blade tip.
  • the area of the cutback in at least one meridian section is preferably limited to at most 30%, in particular at most 25% of the cross-sectional area of the circumferential groove.
  • the web has a cross-sectional area in one or more, in particular all meridian sections through the end face of the web on the blade tip side, which is at least 70%, in particular at least 75%, of the cross-sectional area of the circumferential groove in this meridian section.
  • a cross-sectional area of the circumferential groove is the area that is delimited in the meridian section by the groove base, the groove flanks and a straight connecting line between the upstream and downstream groove edges.
  • the circumferential groove encloses an angle of between 60° and 90° with the flow channel wall in one or more, in particular all meridian sections through the end face of the web on the blade tip side at the upstream groove edge. In this way, in particular, an advantageous axial undercut can be produced.
  • an axial distance between the upstream groove edge and the leading edge of the blade tip arranged downstream thereof is greater than an axial distance between the downstream groove edge and the leading edge of the blade tip arranged upstream therefrom.
  • the leading edge of the blade tip is located between the upstream and downstream groove edges and closer to the downstream groove edge.
  • an axial distance between the upstream and downstream groove edges is at least 25% of an axial distance between the upstream leading edge and a downstream trailing edge of the blade tip.
  • the web in a section perpendicular to an axis of rotation of the compressor, can be straight or curved, in which case it or its tangents can run radially or be inclined against the radial direction.
  • one or more, in particular all, sections are perpendicular to the axis of rotation of the compressor through the end face of the web on the blade tip side, the web is inclined towards the groove base of the circumferential groove in the direction of rotation of the blade tip, in particular by at least 25° and/or at most 65° against the radial direction.
  • FIG. 1 shows a part of a gas turbine compressor according to an embodiment of the present invention or a gas turbine compressor designed according to an embodiment of the present invention in a meridian section.
  • the meridian section contains the axis of rotation of the compressor (horizontally in 1 ), in the 1 vertical direction is a radial direction.
  • the gas turbine compressor has in the circumferential direction (perpendicular to the plane of the 1 ) Side-by-side moving blades with blade tips without shrouds, of which in the meridian section the 1 a rotor blade tip 10 is shown in part, and a flow channel wall 20 fixed to the housing radially on the outside opposite this.
  • a circumferential groove is arranged in the flow channel wall, which has an upstream groove flank 31, which merges into the flow channel wall in an upstream groove edge 21, a downstream groove flank 32, which merges into the flow channel wall in a downstream groove edge 22, and a groove bottom 33 connecting these groove flanks.
  • the upstream groove flank has an axial undercut whose cross-sectional area in the meridian section is less than 10% of a cross-sectional area of the circumferential groove between its upstream and downstream groove edges.
  • This cross-sectional area of the circumferential groove between its upstream and downstream groove edge is the area in the meridian section of the 1 from the bottom of the groove, a straight connecting line 24 between the upstream and downstream groove edges and perpendicular through the upstream and downstream groove edges, which is shown in FIG 1 are indicated by dot-dash lines, the cross-sectional area of the undercut corresponds to the area between the upstream groove flank 31 and the in 1 left dash-dotted perpendicular to the connecting line 24 .
  • circumferential groove In the circumferential groove are several ridges in the circumferential direction (perpendicular to the plane of the drawing 1 ) spaced apart from those in the meridian section of the 1 a web 40 is shown in section.
  • At 24 is in 1 , as already explained above, designates a straight connecting line 24 between the upstream and downstream groove edges 21, 22. This thus represents a reference curve extending from the upstream groove edge to the downstream groove edge, with its curvature equal to infinity.
  • At 23 is in 1 designates another reference curve which also extends from the upstream groove edge to the downstream groove edge, the curvature of this reference curve at the upstream and downstream groove edges being equal to the curvature of the flow channel wall and continuously linear in between in the axial direction, i.e. the curvature of the flow channel wall 20 between the groove edges 21, 22 linearly interpolated.
  • This reference curve 23 thus continues the flow channel contour 20 virtually beyond the circumferential groove.
  • the reference curves 23, 24 each represent a circumferentially extending corresponding reference surface 23, 24 in the meridian section of FIG 1 by an end face or upper edge 43 of the web 40 on the blade tip side.
  • the end face or top edge 43 on the blade tip side deviates from a point or a peripheral line 41 to a point further point or a further circumferential line 42 from the reference curve or surface 23 or the virtually continued flow channel contour from the blade tip away to the bottom of the groove radially (upward in 1 ) away.
  • the blade-side end face or upper edge 43 also deviates from the straight reference surface or curve 24 towards the groove base by at least 1% of a maximum radial distance between the groove edge 22 closer to the blade tip and the groove base 33.
  • the point or perimeter 41 thus defines an upstream beginning of a radial cutback 44 of the web.
  • the blade-side end face or upper edge of the web continues the flow channel contour 20 with a constant curvature.
  • the point or circumferential line 42 defines a downstream end of the radial cutback 44 at which the blade-side face or top edge 43 of the land opens into the downstream groove flank 32 .
  • the blade-side end face or upper edge 43 of the web leads back into the reference surface or curve 23. Then the point or the circumferential line at which the blade-side end face or upper edge 43 of the web again opens into the reference surface or curve 23, or the point or the circumferential line from which the blade-side face or upper edge of the web deviates from the straight reference surface or curve 24 to the groove base 33 again by less than 1%. of the maximum radial distance between the groove edge 22 closer to the blade tip and the groove bottom 33, represents the downstream end of the radial cutback.
  • the blade-side end face or upper edge of the web can follow the flow channel contour with a continuous curvature from the downstream groove edge 22 upstream (to the left in 1 ) to this end of the cutback, as shown or explained analogously for the area between the upstream groove edge 21 and the upstream beginning 41 of the cutback.
  • the void or free area between the blade-side face or top edge 43 of the web and the reference surface or curve 23 thus defines the radial cutback 44 with its upstream beginning 41 and its downstream end 42.
  • this upstream beginning 41 of the cutback 44 becomes or is axially downstream (on the right in 1 ) from the upstream groove edge 21 between this groove edge 21 and the upstream front edge 11 of the blade tip 10 and the downstream end 42 of the cutback 44 in a half 34 of a radial height 35 of the circumferential groove closer to the blade tip.
  • the maximum distance between the groove base 33 and the groove edge 22 closer to the blade tip in the radial direction (vertically in 1 ) or, as in 1 indicated, the maximum distance 35 between the groove base 33 and the groove edge 22 closer to the blade tip can be defined in a direction perpendicular to the straight connecting line 24 of the upstream and downstream groove edges.
  • the part or region of the downstream groove flank 32 that extends in the radial direction or the direction perpendicular to the connecting line 24 of the upstream and downstream groove edge over 50% of the maximum distance of the downstream groove edge 22 from the groove bottom 33 in this direction is referred to as the radially upper half .
  • the web 40 opens into the upstream and downstream groove flank 31, 32 of the circumferential groove, and it thus extends axially through the groove.
  • the end face or upper edge of the web on the blade tip side has the same curvature at the upstream groove edge 21 as the flow channel contour 20 and continues this continuously up to the start 41 of the cutback 44 .
  • the web 40 has an in 1 cross-sectional area indicated by hatching, which is at least 75% of the cross-sectional area of the circumferential groove in this meridian section, which is defined by the groove flanks 31, 32, the groove base 33 and the connecting line 24 between the two groove edges 21, 22.
  • an axial distance between the upstream groove edge 21 and that downstream thereof (right in 1 ) arranged front edge 11 of the blade tip 10 is greater than an axial distance between the downstream groove edge 22 and the front edge 11 arranged upstream thereof 1
  • an axial distance between the upstream and downstream groove edges 21, 22 is at least 25% of an axial distance between the upstream leading edge 11 and a downstream trailing edge (not shown) of the blade tip 10.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Claims (14)

  1. Compresseur de turbine à gaz, comportant au moins une pointe d'aube (10) et une paroi de canal d'écoulement (20) radialement opposée à celle-ci et dans laquelle est disposée une rainure périphérique (31-33), dans laquelle rainure est disposée au moins une nervure (40) présentant une découpe radiale (44) ;
    dans lequel
    un début (41) amont de la découpe est disposé axialement en aval d'un bord de rainure amont (21) entre ledit bord de rainure et un bord d'attaque amont (11) de la pointe d'aube et une extrémité aval (42) de la découpe est disposée dans une moitié proche de la pointe d'aube (34) d'une hauteur radiale (35) de la rainure périphérique, la nervure débouchant dans un flanc de rainure amont et un flanc de rainure aval (31, 32) de la rainure périphérique,
    le flanc de rainure amont (31) de la rainure périphérique présentant une contre-dépouille axiale
    caractérisé en ce que, dans au moins une section méridienne, un bord supérieur (43) côté pointe d'aube de la nervure présente, au niveau du bord de rainure amont, la même courbure qu'un contour de canal d'écoulement de la paroi de canal d'écoulement, c'est-à-dire qu'il présente une courbure continue au niveau du bord de rainure amont et celle-ci se poursuit en continu jusqu'au début (41) de la découpe.
  2. Compresseur de turbine à gaz selon la revendication précédente, caractérisé en ce qu'une face frontale côté aube (43) de la nervure débouche, au moins sensiblement, axialement dans le bord de rainure amont et/ou de manière incurvée dans le flanc de rainure aval dans un sens de rotation de la pointe d'aube ou dans le sens opposé à celui-ci.
  3. Compresseur de turbine à gaz selon l'une des revendications précédentes, caractérisé en ce que la nervure présente, dans au moins une section méridienne, une surface de section transversale qui est égale à au moins 70 %, en particulier à au moins 75 %, d'une surface de section transversale de la rainure périphérique.
  4. Compresseur de turbine à gaz selon l'une des revendications précédentes, caractérisé en ce que la rainure périphérique s'étend sur toute la périphérie de la paroi de canal d'écoulement.
  5. Compresseur de turbine à gaz selon l'une des revendications précédentes, caractérisé en ce que la rainure périphérique forme un angle (α) compris entre 60° et 90° avec la paroi de canal d'écoulement dans au moins une section méridienne au niveau du bord de rainure amont.
  6. Compresseur de turbine à gaz selon l'une des revendications précédentes, caractérisé en ce qu'une distance axiale entre le bord de rainure amont et le bord d'attaque disposé en aval de celui-ci de la pointe d'aube est supérieure à une distance axiale entre le bord de rainure aval et le bord d'attaque disposé en amont de celui-ci de la pointe d'aube.
  7. Compresseur de turbine à gaz selon l'une des revendications précédentes, caractérisé en ce qu'une distance axiale entre le bord de rainure amont et le bord de rainure aval est égale à au moins 25 % d'une distance axiale entre le bord d'attaque amont et un bord de fuite aval de la pointe d'aube.
  8. Compresseur de turbine à gaz selon l'une des revendications précédentes, caractérisé en ce que la nervure est inclinée, dans au moins une section perpendiculaire à un axe de rotation du compresseur, vers un fond de rainure de la rainure périphérique dans le sens de rotation de la pointe d'aube, en particulier d'au moins 25° et/ou d'au plus 65° contre une direction radiale.
  9. Compresseur de turbine à gaz selon l'une des revendications précédentes, caractérisé en ce qu'au moins trois nervures identiques ou différentes sont disposées dans la rainure périphérique, de manière équidistante ou à des distances différentes les unes des autres dans la direction périphérique.
  10. Compresseur de turbine à gaz selon l'une des revendications précédentes, caractérisé en ce que la pointe d'aube est une pointe d'aube (11) radialement externe d'une aube mobile (10) à laquelle la paroi de canal d'écoulement est opposée radialement vers l'extérieur.
  11. Compresseur de turbine à gaz selon l'une des revendications précédentes 1 à 10, caractérisé en ce que la pointe d'aube est une pointe d'aube radialement interne d'une aube directrice à laquelle la paroi de canal d'écoulement est opposée radialement vers l'intérieur.
  12. Compresseur de turbine à gaz selon l'une des revendications précédentes 1, caractérisé en ce qu'un flanc de rainure amont (31) et/ou un flanc de rainure aval (32) de la rainure périphérique présentent une contre-dépouille axiale dont la surface de section transversale, dans une section méridienne, est inférieure à 10 % d'une surface de section transversale de la rainure périphérique entre son bord de rainure amont et son bord de rainure aval.
  13. Moteur d'aéronef comportant un compresseur de turbine à gaz selon l'une des revendications précédentes.
  14. Procédé destiné à la conception d'un compresseur de turbine à gaz selon l'une des revendications précédentes, dans lequel un début (41) amont de la découpe est disposé axialement en aval d'un bord de rainure amont (21) entre ledit bord de rainure et un bord d'attaque amont (11) de la pointe d'aube et une extrémité aval (42) de la découpe est disposée dans une moitié proche de la pointe d'aube (34) d'une hauteur radiale (35) de la rainure périphérique.
EP14163465.9A 2014-04-03 2014-04-03 Compresseur de turbine à gaz, moteur d'avion et méthode de dimensionnement Active EP2927503B1 (fr)

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EP14163465.9A EP2927503B1 (fr) 2014-04-03 2014-04-03 Compresseur de turbine à gaz, moteur d'avion et méthode de dimensionnement
US14/672,959 US10450869B2 (en) 2014-04-03 2015-03-30 Gas turbine compressor

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DE102018203304A1 (de) 2018-03-06 2019-09-12 MTU Aero Engines AG Gasturbinenverdichter
US12018621B1 (en) 2023-08-16 2024-06-25 Rolls-Royce North American Technologies Inc. Adjustable depth tip treatment with rotatable ring with pockets for a fan of a gas turbine engine
US11965528B1 (en) 2023-08-16 2024-04-23 Rolls-Royce North American Technologies Inc. Adjustable air flow plenum with circumferential movable closure for a fan of a gas turbine engine
US11970985B1 (en) 2023-08-16 2024-04-30 Rolls-Royce North American Technologies Inc. Adjustable air flow plenum with pivoting vanes for a fan of a gas turbine engine

Citations (1)

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WO2009103278A1 (fr) * 2008-02-21 2009-08-27 Mtu Aero Engines Gmbh Structure d'écoulement pour turbocompresseur

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Publication number Priority date Publication date Assignee Title
US6290458B1 (en) * 1999-09-20 2001-09-18 Hitachi, Ltd. Turbo machines
DE10390754D2 (de) * 2002-02-28 2005-05-12 Mtu Aero Engines Gmbh Rezirkulationsstruktur für Turboverdichter
EP1530670B1 (fr) * 2002-08-23 2006-05-10 MTU Aero Engines GmbH Structure de recirculation d'un turbocompresseur
GB2408546B (en) * 2003-11-25 2006-02-22 Rolls Royce Plc A compressor having casing treatment slots
DE102008031982A1 (de) 2008-07-07 2010-01-14 Rolls-Royce Deutschland Ltd & Co Kg Strömungsarbeitsmaschine mit Nut an einem Laufspalt eines Schaufelendes
FR2989743B1 (fr) * 2012-04-19 2015-08-14 Snecma Carter de compresseur a cavites de longueurs variees

Patent Citations (1)

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Publication number Priority date Publication date Assignee Title
WO2009103278A1 (fr) * 2008-02-21 2009-08-27 Mtu Aero Engines Gmbh Structure d'écoulement pour turbocompresseur

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EP2927503A1 (fr) 2015-10-07
US20150285079A1 (en) 2015-10-08

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