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EP3536974B1 - Compresseur de turbine à gaz - Google Patents

Compresseur de turbine à gaz Download PDF

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Publication number
EP3536974B1
EP3536974B1 EP19159823.4A EP19159823A EP3536974B1 EP 3536974 B1 EP3536974 B1 EP 3536974B1 EP 19159823 A EP19159823 A EP 19159823A EP 3536974 B1 EP3536974 B1 EP 3536974B1
Authority
EP
European Patent Office
Prior art keywords
groove
upstream
edge
blade tip
downstream
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
EP19159823.4A
Other languages
German (de)
English (en)
Other versions
EP3536974A1 (fr
Inventor
Giovanni Brignole
Tobias Mayenberger
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 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 MTU Aero Engines AG filed Critical MTU Aero Engines AG
Publication of EP3536974A1 publication Critical patent/EP3536974A1/fr
Application granted granted Critical
Publication of EP3536974B1 publication Critical patent/EP3536974B1/fr
Active 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
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/126—Baffles or ribs
    • 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
    • F05D2240/55—Seals

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 blade tips each having an upstream leading edge and a downstream trailing edge, and a flow channel wall radially opposite these blade tips, in which a circumferential groove is arranged, which has an upstream and a downstream groove edge, wherein webs are arranged in the circumferential groove, each having a radial cutback.
  • An object of an embodiment of the present invention is to improve a gas turbine compressor.
  • a gas turbine compressor in particular an axial one, has one or more blades arranged next to one another in the circumferential direction with blade tips, in particular without a shroud, and a flow channel wall radially opposite these.
  • the gas turbine compressor is a gas turbine compressor for an aircraft engine or an aircraft engine, it can in particular be a low-pressure compressor arranged in a gas turbine upstream of another gas turbine compressor or a high-pressure compressor arranged downstream of another gas turbine compressor.
  • the blades are A rotatably mounted rotor has rotating blades arranged on it, the radially outer blade tips of which are opposite the flow channel wall fixed to the housing on the outside.
  • the blades are guide vanes fixed to the housing, the rotating, rotatably mounted flow channel wall being opposite the radially inside.
  • an axial direction is, in the usual way, parallel to the rotation axis of the compressor, a circumferential direction is a direction of rotation around this rotation axis and a radial direction is perpendicular to the axial and circumferential directions.
  • upstream or downstream refers, in the usual way, to a (normal) flow (direction) through the compressor, so that in one embodiment, upstream is closer to an inlet and downstream is closer to an outlet of the compressor.
  • a circumferential groove is arranged in the flow channel wall.
  • this has an upstream groove flank that merges into the flow channel wall in an upstream groove edge, a downstream groove flank that 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 square or rounded or have a radius, whereby its center or intersection of its two outermost tangents can then define the groove edge for dimensional information.
  • the upstream groove flank and/or the downstream groove flank has an axial undercut, the cross-sectional area of which in at least one meridian section in a further development is less than 10% of a cross-sectional area of the circumferential groove between its upstream and downstream groove edge.
  • a meridian section in the sense of the present invention is a plane section that contains the rotation axis of the compressor.
  • An axial undercut of the upstream groove flank is a region of this groove flank that is arranged upstream in the axial direction in front of the upstream groove edge.
  • a axial undercut of the downstream groove flank 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 accordingly the area which is limited in the meridional section by the groove base, a straight connecting line between the upstream and downstream groove edge and perpendiculars through the upstream and downstream groove edge.
  • the circumferential groove extends, in particular continuously or without interruption, over the entire circumference of the flow channel wall or over 360°.
  • the upstream and downstream groove edges are each a continuous edge that extends uninterrupted over 360°. In one embodiment, this can improve the production and/or aerodynamics of the circumferential groove.
  • One or more webs are arranged in the circumferential groove.
  • 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. This can improve the production and/or aerodynamics of the circumferential groove in one embodiment.
  • adjacent webs can be designed in different ways in one embodiment, in particular have different dimensions and/or contours. This can specifically represent or compensate for asymmetries in one embodiment.
  • Three or more, in particular all, webs can be equidistantly spaced in the circumferential direction.
  • three or more, in particular all, webs can have different distances from one another in the circumferential direction in pairs.
  • a radial cutback is understood to mean in particular an empty space between a blade-side front side of the web and its projection into a reference surface which extends from the upstream groove edge to the downstream groove edge, wherein the curvature of the reference surface in the meridian sections through the front side is equal to infinity or at the upstream and downstream groove edge is equal to the curvature of the flow channel wall and is continuously linear in the axial direction therebetween.
  • the radial cutback is understood to be the free area between a blade tip-side upper edge of the cross-section of the web and a reference curve which 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 edge being equal to the curvature of the flow channel wall and is continuously linear in the axial direction therebetween.
  • a radial cutback in one embodiment is understood to mean the empty space or the free area between the blade-side front side or upper edge of the web and a flow channel contour that is virtually continued over the circumferential groove, whereby this virtually continued contour can be a straight connecting plane or line or can connect the groove edges with a curvature that corresponds to the curvature of the flow channel contour at the groove edges and interpolates linearly therebetween.
  • an axial distance between the upstream leading edge of the blade tip and the downstream groove edge is at least 5%, in particular at least 7.5%, in one embodiment at least 10%, and/or at most 40%, in particular at most 35%, in one embodiment at most 30%, of the chord length between the upstream leading edge and the downstream trailing edge of the blade tip or the gas turbine compressor is designed in such a way or this axial distance is selected in such a way.
  • an axial distance between the upstream leading edge of the blade tip and a kink of a blade tip-side upper edge of the web in the cutback is at most 10%, in particular at most 7.5%, in one embodiment at most 5%, of the chord length between the upstream leading edge and the downstream trailing edge of the blade tip, wherein in one embodiment the kink is arranged downstream, in another embodiment the kink is arranged upstream of the upstream leading edge of the blade tip, or the gas turbine compressor is designed in such a way or this axial distance is selected in such a way.
  • the kink of the blade tip-side upper edge can be sharp-edged or angular or rounded or have a radius, wherein for dimensional specifications its center or intersection of its two outermost tangents can then define the kink.
  • a kink is defined in particular as a discontinuity (point) of the tangent to the upper edge of the web.
  • the front side of the blade tip or the upper edge of the web in the cutback can also be kink-free.
  • a, in particular minimum, maximum and/or average, distance in the radial direction (“(minimum/maximum/average) radial distance") between the blade tip, in particular its upstream leading edge, and a blade tip-side upper edge of the web in the cutback is at least 50%, in particular at least 75%, in one embodiment at least 100%, and/or at most 1500%, in particular at most 1250%, in one embodiment at most 1000%, of a radial distance between the blade tip and the downstream groove edge radially opposite it, or the gas turbine compressor is designed in this way or this radial distance is selected in this way.
  • chord length refers in the usual way to the length of the profile chord or centerline of the blade tip or its projection in the axial direction or the axial distance between the leading and trailing edges of the blade tip.
  • an upstream beginning of the cutback is arranged axially downstream of the upstream groove edge between this groove edge and the upstream leading edge of the blade tip and/or a downstream end of the cutback is arranged in a half of a radial height of the circumferential groove closer to the blade tip.
  • an upstream start of the cutback is understood to mean the axial position from which the blade-side front side or upper edge of the web deviates from the virtually continued flow channel contour or the reference surface or curve away from the blade tip towards the groove base.
  • an upstream start of the cutback is understood to mean the axial position from which the blade-side front side 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 arranged axially downstream of the upstream groove edge and upstream of the upstream leading edge of the blade tip.
  • the blade-side front side or in one or more, preferably all, meridional sections through the blade-tip-side front side of the web, the upper edge) of the web in one embodiment continues the flow channel contour with a constant curvature or without a sudden change in the curvature.
  • a downstream end of the cutback is understood to mean the axial position at which the blade-side front face or upper edge of the web again flows into the reference surface or curve or into the downstream groove flank.
  • a downstream end of the cutback is understood to mean the axial position from which the blade-side front face or upper edge of the web deviates from the straight reference surface or curve towards 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 is understood to mean in particular a maximum distance between the groove base and the reference surface or curve, in particular a maximum distance between the groove base and the groove edge closer to the blade tip, in the radial direction or a direction perpendicular to the connecting line of the upstream and downstream groove edges, whereby such a distance perpendicular to the connecting line is also generally referred to as the radial height of the circumferential groove.
  • the radial cutback ends in the reference surface or curve, in a further development axially upstream in front of or downstream behind the upstream leading edge of the blade tip.
  • the blade-side front face or in one or more, preferably all, meridional sections through the blade-tip-side front face of the web, the upper edge of the web in one embodiment continues the flow channel contour with a constant curvature or without an abrupt change in the curvature from the downstream groove edge upstream.
  • the radially upper half is generally referred to as the part of the downstream groove flank that extends in the radial direction or a direction perpendicular to the connecting line of 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/or downstream groove flank of the circumferential groove, and can thus extend in particular axially through the groove or its maximum axial length.
  • a blade tip-side upper edge of the web at the upstream groove edge can have the same curvature as the flow channel contour, i.e. have a continuous curvature at the upstream groove edge, and continue this continuously up to the beginning of the cutback.
  • the web can be straight or curved.
  • the blade-side front side of the web can open, at least essentially, axially into the upstream groove edge. Additionally or alternatively, the blade-side front side can open curved into the downstream groove flank in or against a direction of rotation of the blade tip.
  • the area of the cutback in at least one meridian section is limited to a maximum of 30%, in particular a maximum of 25% of the cross-sectional area of the circumferential groove.
  • the web in one or more, in particular all, meridian sections through the blade tip-side front side of the web has a cross-sectional area that 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 limited 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 in one or more, in particular all, meridional sections through the blade tip-side face of the web at the upstream groove edge forms an angle of between 60° and 90° with the flow channel wall. This can in particular provide an advantageous axial undercut.
  • 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 thereof.
  • the leading edge of the blade tip is arranged 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 the downstream trailing edge of the blade tip.
  • the web In a section perpendicular to a rotation axis of the compressor, the web can be straight or curved, whereby it or its tangents can run radially or be inclined against the radial direction. Accordingly, in one embodiment, in one or more, in particular all, sections perpendicular to the rotation axis of the compressor through the blade tip-side face of the web, 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.
  • Dimensions in one version refer to a component temperature of 20°C and/or components without elastic deformation.
  • Fig.1 shows a meridian section of a part of a gas turbine compressor according to an embodiment of the present invention or of a gas turbine compressor designed according to an embodiment of the present invention.
  • the meridian section contains the axis of rotation of the compressor (horizontal in Fig.1 ), in the Fig.1 vertical direction is a radial direction.
  • the gas turbine compressor has a circumferential direction (perpendicular to the plane of the Fig.1 ) side-by-side arranged rotor blades with shroudless blade tips, of which in the meridional section of the Fig.1 a rotor blade tip 10 is partially shown, and a flow channel wall 20 fixed to the housing radially outwardly opposite these.
  • 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 base 33 connecting these groove flanks.
  • the upstream groove flank has an axial undercut whose cross-sectional area in the meridional section is less than 10% of a cross-sectional area of the circumferential groove between its upstream and downstream groove edge.
  • This cross-sectional area of the circumferential groove between its upstream and downstream groove edge is the area which in the meridional section of the Fig.1 from the groove base, a straight connecting line 24 between the upstream and downstream groove edges and perpendiculars through the upstream and downstream groove edges, which in Fig.1 are indicated by dash-dotted lines, the cross-sectional area of the undercut corresponding to the area between the upstream groove flank 31 and the Fig.1 left dash-dotted perpendicular to the connecting line 24 .
  • Fig.1 denotes a straight connecting line 24 between the upstream and downstream groove edges 21, 22. This thus represents a reference curve which extends from the upstream groove edge to the downstream groove edge, with its curvature equal to infinity.
  • Fig.1 another reference curve is designated, 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 the axial direction between them, ie the curvature of the flow channel wall 20 between the groove edges 21, 22 is linearly interpolated.
  • This reference curve 23 thus virtually continues the flow channel contour 20 across the circumferential groove.
  • the reference curves 23, 24 each represent a corresponding reference surface 23, 24 extending in the circumferential direction in the meridian section of the Fig.1 by a blade tip-side front face or upper edge 43 of the web 40.
  • the blade tip-side front surface or upper edge 43 deviates from a point or a circumferential line 41 to a 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 groove base radially (upwards in Fig.1 ) away.
  • the blade-side front side 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 circumferential line 41 thus defines an upstream beginning of a radial cutback 44 of the web.
  • the blade-side front side or upper edge of the web continues the flow channel contour 20 with a continuous curvature.
  • the point or circumferential line 42 defines a downstream end of the radial cutback 44, at which the blade-side front side or upper edge 43 of the web opens into the downstream groove flank 32.
  • the blade-side front face or upper edge 43 of the web flows back into the reference surface or curve 23. Then the point or circumferential line at which the blade-side front face or upper edge 43 of the web flows back into the reference surface or curve 23, or the point or circumferential line from which the blade-side front face or upper edge of the web deviates from the straight reference surface or curve 24 towards the groove base 33 by less than 1% of the maximum radial distance between the groove edge 22 closer to the blade tip and the groove base 33, represents the downstream end of the radial cutback.
  • the blade-side front side or upper edge of the web can form the flow channel contour with a continuous curvature from the downstream groove edge 22 upstream (to the left in Fig.1 ) to this end of the cutback, as is shown or explained analogously for the area between the upstream groove edge 21 and the upstream beginning 41 of the cutback.
  • the empty space or the free area between the blade-side front side or upper 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 cut 44 is axially downstream (right in Fig.1 ) from the upstream groove edge 21 between this groove edge 21 and the upstream leading 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 radial height can be defined as the maximum distance between the groove base 33 and the groove edge 22 closer to the blade tip in the radial direction (vertically in Fig.1 ) or, as in Fig.1 As 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 radial cutback ends in the radially upper half 34 of the downstream groove flank 32, the web is radially cutback continuously from the beginning 41.
  • the radially upper half is the part or area of the downstream groove flank 32 which extends in the radial direction or the direction perpendicular to the connecting line 24 of the upstream and downstream groove edges over 50% of the maximum distance of the downstream groove edge 22 from the groove base 33 in this direction.
  • the web 40 opens into the upstream and downstream groove flanks 31, 32 of the circumferential groove, thus extending axially through the groove.
  • the blade tip-side front surface or upper edge of the web at the upstream groove edge 21 has the same curvature as the flow channel contour 20 and continues this continuously up to the beginning 41 of the cutback 44.
  • the web 40 has a Fig.1 hatched cross-sectional area 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.
  • the circumferential groove at the upstream groove edge 21 forms an angle ⁇ with the flow channel wall 20 which is between 60° and 90°.
  • Fig.1 In the execution of the Fig.1 is an axial distance between the upstream groove edge 21 and the downstream groove edge 22 (right in Fig.1 ) arranged leading edge 11 of the blade tip 10 is greater than an axial distance between the downstream groove edge 22 and the leading edge 11 arranged upstream thereof.
  • 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 12 of the blade tip 10.
  • S AX schematically indicates an axial chord length of the blade tip 10, which can equally correspond to the axial distance between the leading and trailing edges 11, 12 or to the length of the profile chord or centerline of the blade tip 10.
  • An axial distance L KOZ between the upstream beginning 41 of the cutback 44 and the upstream leading edge 11 of the blade tip is between 1% and 40%, preferably between 2% and 15%, of this chord length S AX defined in this way.
  • An axial distance L OL between the upstream leading edge 11 of the blade tip and the downstream groove edge 22 is between 5% and 40%, preferably between 10% and 30%, of the chord length S AX .
  • An axial distance ⁇ 45 between the upstream leading edge 11 of the blade tip and a kink 45 of the blade tip-side front face or upper edge 43 of the web in the cutback amounts to a maximum of 10%, preferably a maximum of 5%, of the chord length S AX .
  • a radial distance between the blade tip 10 and the blade tip-side front side or upper edge 43 of the web in the recess 44 is between 50% and 1500%, preferably between 100% and 1000%, of a radial distance H GAP between the blade tip 10 and the radially opposite downstream groove edge 22.
  • H GAP a radial distance between the blade tip 10 and the radially opposite downstream groove edge 22.

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Claims (13)

  1. Compresseur de turbine à gaz, comportant au moins une pointe d'aube (10) qui présente un bord d'attaque amont (11) et un bord de fuite aval (12), et une paroi de canal d'écoulement (20) opposée radialement à cette pointe d'aube, paroi dans laquelle est disposée une rainure périphérique (31-33) qui présente un bord de rainure amont (21) et un bord de rainure aval (22),
    au moins une nervure (40) étant disposée dans la rainure périphérique, laquelle nervure présente une découpe radiale (44) ;
    le bord d'attaque amont (11) de la pointe d'aube (10) étant disposé entre un début amont (41) de la découpe et le bord de rainure aval (22) dans la direction axiale dans au moins une section méridienne à travers le côté frontal, côté pointe d'aube, de la nervure,
    caractérisé en ce
    que, dans au moins la section méridienne à travers un côté frontal, côté pointe d'aube, de la nervure, une distance axiale (LOL) entre le bord d'attaque amont (11) de la pointe d'aube et le bord de rainure aval (22) vaut au moins 5 % et au plus 40 % de la longueur de corde (SAX) entre le bord d'attaque amont (11) et le bord de fuite aval (12) de la pointe d'aube et/ou
    une distance axiale (Δ45) entre le bord d'attaque amont (11) de la pointe d'aube et un coude (45) d'un bord supérieur (43), côté pointe d'aube, de la nervure dans la découpe vaut au plus 10 % de la longueur de corde (SAX) entre le bord d'attaque amont (11) et le bord de fuite aval (12) de la pointe d'aube.
  2. Compresseur de turbine à gaz selon la revendication 1,
    caractérisé en ce qu'une distance radiale (HKOZ) entre la pointe d'aube (10) et un bord supérieur (43), côté pointe d'aube, de la nervure dans la découpe vaut au moins 50 % et/ou au plus 1500 % d'une distance radiale (HGAP) entre la pointe d'aube (10) et le bord de rainure aval (22) opposé radialement à celle-ci.
  3. Compresseur de turbine à gaz selon la revendication 1 ou 2,
    caractérisé en ce que le début amont (41) de la découpe est disposé axialement en aval du bord de rainure amont (21) entre ce bord de rainure et le bord d'attaque amont (11) de la pointe d'aube et/ou une extrémité aval (42) de la découpe est disposée dans une moitié (34) proche de la pointe d'aube d'une hauteur radiale (35) de la rainure périphérique.
  4. Compresseur de turbine à gaz selon l'une des revendications précédentes, caractérisé en ce que, dans au moins une section méridienne, le bord supérieur (43), côté pointe d'aube, de la nervure présente, au niveau du bord de rainure amont, en particulier jusqu'au début de la découpe, une courbure continue ; et/ou en ce qu'un côté frontal (43), côté aube, 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.
  5. Compresseur de turbine à gaz selon l'une des revendications précédentes, caractérisé en ce que la nervure débouche dans un flanc de rainure (31, 32) amont et/ou aval de la rainure périphérique et/ou présente, dans au moins une section méridienne, une surface de section transversale qui vaut au moins 70 %, en particulier au moins 75 %, d'une surface de section transversale de la rainure périphérique.
  6. 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 et/ou 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.
  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 d'attaque, disposé en aval de celui-ci, de la pointe d'aube est supérieure à la distance axiale entre le bord de rainure aval et le bord d'attaque, disposé en amont de celui-ci, de la pointe d'aube ; et/ou en ce qu'une distance axiale entre le bord de rainure amont et le bord de rainure aval vaut au moins 25 % d'une distance axiale entre le bord d'attaque amont et le 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° par rapport à une direction radiale ; et/ou 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.
  9. 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 extérieure d'une aube mobile (10) à laquelle la paroi de canal d'écoulement est opposée radialement vers l'extérieur, ou une pointe d'aube radialement interne d'une aube directrice à laquelle la paroi de canal d'écoulement est opposée radialement vers l'intérieur.
  10. Compresseur de turbine à gaz selon l'une des revendications précédentes, 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 au moins 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.
  11. Moteur d'aéronef comportant un compresseur de turbine à gaz selon l'une des revendications précédentes.
  12. Procédé destiné à la conception d'un compresseur de turbine à gaz selon l'une des revendications précédentes, caractérisé en ce que,
    dans au moins une section méridienne, une distance axiale (LOL) entre le bord d'attaque amont (11) de la pointe d'aube et le bord de rainure aval (22) est sélectionnée de telle sorte qu'elle vaut au moins 5 % et au plus 40 % de la longueur de corde (SAX) entre le bord d'attaque amont (11) et le bord de fuite aval (12) de la pointe d'aube et/ou une distance axiale (Δ45) entre le bord d'attaque amont (11) de la pointe d'aube et un coude d'un bord supérieur (43), côté pointe d'aube, de la nervure dans la découpe est sélectionnée de telle sorte qu'elle vaut au plus 10 % de la longueur de corde (SAX) entre le bord d'attaque amont (11) et le bord de fuite aval (12) de la pointe d'aube.
  13. Procédé destiné à la conception d'un compresseur de turbine à gaz selon la revendication 12, caractérisé en ce qu'une distance radiale (HKOZ) entre la pointe d'aube (10) et un bord supérieur (43), côté pointe d'aube, de la nervure dans la découpe est sélectionnée de telle sorte qu'elle vaut au moins 50 % et/ou au plus 1500 % d'une distance radiale (HGAP) entre la pointe d'aube (10) et le bord de rainure aval (22) opposé radialement à celle-ci.
EP19159823.4A 2018-03-06 2019-02-27 Compresseur de turbine à gaz Active EP3536974B1 (fr)

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DE102018203304.8A DE102018203304A1 (de) 2018-03-06 2018-03-06 Gasturbinenverdichter

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CN112685829B (zh) * 2020-12-22 2021-11-02 中国船舶重工集团公司第七0三研究所 一种船舶燃气轮机压气机带槽环式处理机匣设计方法

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2003072910A1 (fr) * 2002-02-28 2003-09-04 Mtu Aero Engines Gmbh Structure de recirculation de turbocompresseurs
WO2004018844A1 (fr) * 2002-08-23 2004-03-04 Mtu Aero Engines Gmbh Structure de recirculation d'un turbocompresseur

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Publication number Priority date Publication date Assignee Title
EP1478857B1 (fr) 2002-02-28 2008-04-23 MTU Aero Engines GmbH Compresseur avec moyens de traitement antiblocage d'extremites
DE10330084B4 (de) * 2002-08-23 2010-06-10 Mtu Aero Engines Gmbh Rezirkulationsstruktur für Turboverdichter
DE102007056953B4 (de) * 2007-11-27 2015-10-22 Rolls-Royce Deutschland Ltd & Co Kg Strömungsarbeitsmaschine mit Ringkanalwandausnehmung
DE102008011644A1 (de) * 2008-02-28 2009-09-03 Rolls-Royce Deutschland Ltd & Co Kg Gehäusestrukturierung für Axialverdichter im Nabenbereich
DE102008031982A1 (de) * 2008-07-07 2010-01-14 Rolls-Royce Deutschland Ltd & Co Kg Strömungsarbeitsmaschine mit Nut an einem Laufspalt eines Schaufelendes
EP2927503B1 (fr) * 2014-04-03 2023-05-17 MTU Aero Engines AG Compresseur de turbine à gaz, moteur d'avion et méthode de dimensionnement

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2003072910A1 (fr) * 2002-02-28 2003-09-04 Mtu Aero Engines Gmbh Structure de recirculation de turbocompresseurs
WO2004018844A1 (fr) * 2002-08-23 2004-03-04 Mtu Aero Engines Gmbh Structure de recirculation d'un turbocompresseur

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EP3536974A1 (fr) 2019-09-11
DE102018203304A1 (de) 2019-09-12
US11686207B2 (en) 2023-06-27

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