EP2665896B1 - Intermediate casing of a gas turbine engine comprising an outer boundary wall wich comprises upstream of a support strut a variable contour in circumferential direction in order to reduce secondary flow losses - Google Patents

Intermediate casing of a gas turbine engine comprising an outer boundary wall wich comprises upstream of a support strut a variable contour in circumferential direction in order to reduce secondary flow losses Download PDF

Info

Publication number
EP2665896B1
EP2665896B1 EP12716196.6A EP12716196A EP2665896B1 EP 2665896 B1 EP2665896 B1 EP 2665896B1 EP 12716196 A EP12716196 A EP 12716196A EP 2665896 B1 EP2665896 B1 EP 2665896B1
Authority
EP
European Patent Office
Prior art keywords
boundary
wall
housing
boundary wall
radially outer
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.)
Not-in-force
Application number
EP12716196.6A
Other languages
German (de)
French (fr)
Other versions
EP2665896A1 (en
Inventor
Martin Hoeger
Inga Mahle
Jochen Gier
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 EP2665896A1 publication Critical patent/EP2665896A1/en
Application granted granted Critical
Publication of EP2665896B1 publication Critical patent/EP2665896B1/en
Not-in-force legal-status Critical Current
Anticipated expiration legal-status Critical

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
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/14—Casings modified therefor
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/11—Shroud seal segments
    • 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
    • F05D2270/00—Control
    • F05D2270/01—Purpose of the control system
    • F05D2270/17—Purpose of the control system to control boundary layer

Definitions

  • the invention relates to a housing of a gas engine according to the preamble of patent claim 1, as this from the US 2008/0276621 A1 and WO 2009/019282 A2 is known.
  • a multi-shaft turbomachine such as a multi-shaft gas engine has a plurality of compressor components, at least one combustion chamber and a plurality of turbine components.
  • a two-shaft gas engine has a low-pressure compressor, a high-pressure compressor, at least one combustion chamber, a high-pressure turbine and a low-pressure turbine.
  • a three-shaft gas engine has a low-pressure compressor, a medium-pressure compressor, a high-pressure compressor, at least one combustion chamber, a high-pressure turbine, a medium-pressure turbine and a low-pressure turbine.
  • Fig. 1 shows a highly schematic section of a multi-shaft gas engine in the region of a rotor 10 of a high-pressure turbine 11 and a rotor 12 of a low-pressure turbine 13. Between the high-pressure turbine 11 and the low-pressure turbine 13 extends between an intermediate housing 14 with a transitional flow channel 33 to the flow, which the high-pressure turbine 11 leaves to supply the low-pressure turbine 13, wherein in the transitional flow channel 33 at least one support rib 15 is positioned.
  • the support rib 15 is a stator-side component, which carries the flow flowing through the transition flow channel 33.
  • a flow-guiding support rib 15 has a front edge 16, which is also referred to as a flow inlet edge, via a trailing edge 17, which is also referred to as a flow outlet edge, and via side walls 18.
  • transitional flow channel 33 can (see Fig. 1 ) open upstream of the support ribs 15 in the region of an entry into the transitional flow channel 33 and in the region of a front edge 34 of the intermediate housing 14 radially outwardly in the same a cavity 19 through which a small amount of cooling air 21 a can emerge, which deals with the high-pressure turbine 11 leaving gas flow 20 mixed.
  • This cavity 19 is located between the NDT housing and the intermediate housing 14, which is sealed with a seal 21c. Only a weak leakage flow 21b flows through this seal 21c, since the NDT housing and the intermediate housing 14 can not be firmly connected to each other.
  • the static pressure of the gas flow 20 in the region of entry into the cavity 19 is below the pressure of the cooling air 21b in the secondary air region 21d outside the annulus.
  • FIG. 2 can be removed, arises in the from the prior art according to Fig. 1 known turbomachine upstream of the leading edges 16 of the support ribs 15 due to a blocking of the gas flow flowing through the transitional flow channel 33 at circumferential positions on which the support ribs are positioned, a pressure increase + .DELTA.p of the static pressure, whereas according Fig. 2 on circumferential positions between adjacent support ribs 15 sets a pressure drop - ⁇ p of the static pressure.
  • Fig. 2 is shown a dimensionless circumferential direction u / t, where t corresponds to the support rib pitch in the circumferential direction u.
  • the present invention is based on the problem to provide an intermediate housing, by means of which the efficiency can be increased.
  • the radially outer boundary wall has a contour which changes in the circumferential direction at least in a section upstream of the support rib.
  • the present invention relates to the field of multi-shaft turbomachinery, in particular multi-shaft gas engines, with several compressor components and several turbine components.
  • the basic structure of such a turbomachine is familiar to the person mentioned here and has already been in connection with Fig. 1 described.
  • the present invention now relates to details of an intermediate housing 14 of such a turbomachine, by means of which the entry of a guided in a cooling air flow passage 19 cooling air flow can be improved in the outflow passage 33 of the intermediate housing 14 guided gas flow, namely in an inlet region of the transitional flow channel 33 upstream of the transitional flow channel 33 positioned support ribs 15th
  • the invention is applicable both to an intermediate housing 14 of a twin-shaft turbomachine which extends between a high-pressure turbine 11 and a low-pressure turbine 13, and to an intermediate housing of a three-shaft turbomachine which extends between a high-pressure turbine and a medium-pressure turbine or between a medium-pressure turbine and a low-pressure turbine, used.
  • Fig. 3 shows a section of a turbomachine in the region of an intermediate housing 14, a transitional flow channel 33 of this intermediate housing 14 and an upstream of the transitional flow channel 33 positioned, formed in the illustrated embodiment as a high-pressure turbine 11 turbine component, according to Fig. 3 the cooling air flow passage 19 opens radially outward into the transition flow passage 33, namely upstream of support ribs 15 positioned in the transition flow passage 33.
  • the cooling air flow channel 19 is thereby limited in sections by the front edge 34 of the intermediate housing 14.
  • the transitional flow channel 33 is bounded radially inwardly by a stator-side boundary wall 23 and also radially on the outside by a stator-side boundary wall 24.
  • a boundary wall 25 of the high-pressure turbine 11 adjoins the rotor 10 of the high-pressure turbine 11 radially on the outside.
  • the radially outer boundary wall 24 of the transition flow channel 33 at least in a transition section between the front edge 34 of the intermediate housing 14 and the transition flow channel 33 has a circumferentially changing contour.
  • This circumferentially changing contour of the radially outer boundary wall 24 of the transition flow channel 33 may according to Fig. 3 extend to a region downstream of the leading edges 16 of the support ribs 15, wherein Fig. 3 two contours 24 and 24 'formed at different circumferential positions u / t for the radially outer boundary wall of the transitional flow channel 33.
  • the radially outer boundary wall 24 of the transition flow channel 33 has in the inlet region of the transition flow channel 33 upstream of the leading edges 16 of the support ribs 15 via a boundary wall portion or boundary wall 26 with a minimum radius of curvature and thus maximum curvature.
  • the contour of the radially outer boundary wall 24 of the transition flow channel 33 changes in the circumferential direction u or u / t such that an axial position (axial direction x) and / or a radial position (radial direction r) of the boundary wall section or boundary wall point 26 with a minimum radius of curvature in the circumferential direction u or u / t changed.
  • both the axial position and the radial position of the boundary wall point 26 change with a minimum radius of curvature.
  • the axial position of the boundary wall 26 with minimal radius of curvature changes in the circumferential direction u or u / t such that approximately at the circumferential position of the leading edges 16 of the support ribs 15 of this boundary wall 26 in the axial direction x maximum upstream and approximately in a circumferential position half pitch between two adjacent supporting ribs in the axial direction x offset or positioned downstream of maximum. Between these maximum upstream and downstream axial positions, the axial position of the boundary wall 26 gradually changes in the circumferential direction.
  • the radial position of the boundary wall 26 with minimal radius of curvature changes in the circumferential direction u or u / t such that approximately at the circumferential position of the leading edges 16 of the support ribs 15 of this boundary wall 26 in the radial direction r maximum radially outward and approximately in a circumferential position half pitch between two adjacent support ribs 15 in the radial direction r is offset or positioned maximally radially inward. Between these maximum radially inner and radially outer radial positions, the radial position of the boundary wall point 26 changes continuously or continuously in the circumferential direction.
  • contour 24 of the radially outer boundary wall of the transition flow channel 33 corresponds to the contour thereof approximately at the circumferential position of a front edge 16 of a support rib 15, whereas the in Fig. 3 shown contour 24 'of the same contour approximately in a circumferential position half pitch between two adjacent support ribs 15 corresponds.
  • Fig. 4 is on the horizontally extending axis an absolute value ratio ⁇ x / x KS between the axial distance ⁇ x (see Fig. 3 ) Downstream of the axial position and the axial position of the maximum of the upstream boundary wall point 26 with a minimum radius of curvature and the axial distance ⁇ KS (see Fig. 3 ) of a downstream end 27 of the radially outer boundary wall 25 of the upstream of the transition channel 33 positioned turbine component 11 and the front edge 16 of the support ribs 15 applied. Furthermore, in Fig. 4 on the horizontally extending axis an absolute value ratio ⁇ r / x KS between the radial distance ⁇ r (see Fig.
  • x corresponds to KS (see Fig. 3 ) the distance between the downstream end 27 of the radially outer boundary wall 25 of the high-pressure turbine 11 and the front edge 16 of the support ribs 15th
  • the area 28 of the Fig. 4 visualizes a preferred scope for extending u and u / t changing in the circumferential direction ratio Ax / x KS and / or AR / x KS and thus the to u and u / t changing in the circumferential direction of offset of the axial position and / or the radial position of the Boundary wall point 26 with a minimum radius of curvature.
  • the ratios ⁇ x / x KS and ⁇ r / x KS are up to 40%.
  • Curve 29 within region 28 visualizes the preferred circumferentially varying ratio ⁇ x / x KS, and hence the circumferentially varying offset of the axial position of the minimum wall radius limiting wall 26, where, according to curve 29, the axial position offset is in the half pitch range between two adjacent support ribs is largest and the ratio ⁇ x / x KS is about 20%.
  • Curve 30 within region 28 illustrates the preferred circumferentially varying ratio ⁇ r / x KS, and thus the circumferentially varying offset of the radial position of boundary wall 26 with minimum radius of curvature, and at approximately half pitch between adjacent support ribs, the ratio ⁇ r / x KS is about 2.5% and the offset of the radial position in the area of half pitch between two adjacent support ribs is the largest.
  • the offset of the axial position of the boundary wall 26 having the minimum radius of curvature and the offset of the radial position of the boundary wall 26 having the minimum radius of curvature and the above ratios ⁇ x / x KS and ⁇ r / x KS are respectively continuous and continuous, and preferably nonlinear.
  • Fig. 5 visualizes the effect of the contouring according to the invention of the radially outer boundary wall 24 of the transitional flow channel 33 wherein Fig. 5 is plotted on the horizontal axis a ratio (pp m ) / p m between the difference (pp m ) of the static pressure p of the gas flow in the transitional flow channel 14 and the mean value p m of this static pressure and the mean value p m ; vertically extending axis the dimensionless circumferential direction u / t is plotted.
  • the curve 31 of the Fig. 5 corresponds to a state of the art adjusting course of the ratio (pp m ) / p m and the curve 32 according to the invention adjusting the course of the ratio (pp m ) / p m .
  • Fig. 5 It can be seen that with the invention, an improved, uniform pressure distribution of the static pressure in the circumferential direction can be provided, whereby the formation of a secondary flow in the mouth portion of the cooling air flow channel 19 in the transitional flow channel 33 can be effectively counteracted. Thereby, an unhindered entry of the cooling air flow into the transitional flow channel 33 can be ensured, whereby the efficiency of the turbomachine can be improved. Furthermore, the flow in the transitional flow passage 33 between adjacent support ribs 15 can be improved.

Landscapes

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

Description

Die Erfindung betrifft ein Gehäuse eines Gastriebwerks nach dem Oberbegriff des Patentanspruchs 1, wie dieses aus der US 2008/0276621 A1 und WO 2009/019282 A2 bekannt ist.The invention relates to a housing of a gas engine according to the preamble of patent claim 1, as this from the US 2008/0276621 A1 and WO 2009/019282 A2 is known.

Eine mehrwellige Strömungsmaschine wie zum Beispiel ein mehrwelliges Gastriebwerk verfügt über mehrere Verdichterkomponenten, mindestens eine Brennkammer und mehrere Turbinenkomponenten. So verfügt ein zweiwelliges Gastriebwerk über einen Niederdruck-verdichter, einen Hochdruckverdichter, mindestens eine Brennkammer, eine Hochdruckturbine sowie eine Niederdruckturbine. Ein dreiwelliges Gastriebwerk verfügt über einen Niederdruckverdichter, einen Mitteldruckverdichter, einen Hochdruckverdichter, mindestens eine Brennkammer, eine Hochdruckturbine, eine Mitteldruckturbine und eine Niederdruckturbine.A multi-shaft turbomachine such as a multi-shaft gas engine has a plurality of compressor components, at least one combustion chamber and a plurality of turbine components. For example, a two-shaft gas engine has a low-pressure compressor, a high-pressure compressor, at least one combustion chamber, a high-pressure turbine and a low-pressure turbine. A three-shaft gas engine has a low-pressure compressor, a medium-pressure compressor, a high-pressure compressor, at least one combustion chamber, a high-pressure turbine, a medium-pressure turbine and a low-pressure turbine.

Fig. 1 zeigt einen stark schematisierten Ausschnitt aus einem mehrwelligen Gastriebwerk im Bereich eines Rotors 10 einer Hochdruckturbine 11 sowie eines Rotors 12 einer Niederdruckturbine 13. Zwischen der Hochdruckturbine 11 und der Niederdruckturbine 13 erstreckt sich ein Zwischengehäuse 14 mit einem Übergangströmungskanal 33, um die Strömung, welche die Hochdruckturbine 11 verlässt, der Niederdruckturbine 13 zuzuführen, wobei im Übergangsströmungskanal 33 mindestens eine Stützrippe 15 positioniert ist. Fig. 1 shows a highly schematic section of a multi-shaft gas engine in the region of a rotor 10 of a high-pressure turbine 11 and a rotor 12 of a low-pressure turbine 13. Between the high-pressure turbine 11 and the low-pressure turbine 13 extends between an intermediate housing 14 with a transitional flow channel 33 to the flow, which the high-pressure turbine 11 leaves to supply the low-pressure turbine 13, wherein in the transitional flow channel 33 at least one support rib 15 is positioned.

Bei der Stützrippe 15 handelt es sich um ein statorseitiges Bauteil, welches die den Übergangsströmungskanal 33 durchströmende Strömung führt. Eine solche strömungsführende Stützrippe 15 verfügt über eine Vorderkante 16, die auch als Strömungseintrittskante bezeichnet wird, über eine Hinterkante 17, die auch als Strömungsaustrittskante bezeichnet wird, und über Seitenwände 18.The support rib 15 is a stator-side component, which carries the flow flowing through the transition flow channel 33. Such a flow-guiding support rib 15 has a front edge 16, which is also referred to as a flow inlet edge, via a trailing edge 17, which is also referred to as a flow outlet edge, and via side walls 18.

In den Übergangströmungskanal 33 kann (siehe Fig. 1) stromaufwärts der Stützrippen 15 im Bereich eines Eintritts in den Übergangsströmungskanal 33 bzw. im Bereich einer Vorderkante 34 des Zwischengehäuses 14 radial außen in denselben eine Kavität 19 münden, durch die in geringem Maße Kühlluft 21 a austreten kann, die sich mit der die Hochdruckturbine 11 verlassenden Gasströmung 20 vermischt. Diese Kavität 19 befindet sich zwischen den NDT-Gehäuse und dem Zwischengehäuse 14, die mit einer Dichtung 21c abgedichtet wird. Durch diese Dichtung 21c fließt nur eine schwache Leckageströmung 21b, da das NDT-Gehäuse und das Zwischengehäuse 14 nicht fest miteinander verbunden werden können.Into the transitional flow channel 33 can (see Fig. 1 ) open upstream of the support ribs 15 in the region of an entry into the transitional flow channel 33 and in the region of a front edge 34 of the intermediate housing 14 radially outwardly in the same a cavity 19 through which a small amount of cooling air 21 a can emerge, which deals with the high-pressure turbine 11 leaving gas flow 20 mixed. This cavity 19 is located between the NDT housing and the intermediate housing 14, which is sealed with a seal 21c. Only a weak leakage flow 21b flows through this seal 21c, since the NDT housing and the intermediate housing 14 can not be firmly connected to each other.

Um den Eintritt der Leckage 21 a in den Übergangsströmungskanal 33 zu ermöglichen und ein Einströmen der Gasströmung 20 über die Kavität 19 zu verhindern, liegt der statische Druck der Gasströmung 20 im Bereich des Eintritts in die Kavität 19 unterhalb des Drucks der Kühlluft 21b im Sekundärluftbereich 21d außerhalb des Ringraumes.In order to allow the entry of the leakage 21 a in the transitional flow channel 33 and to prevent an influx of the gas flow 20 through the cavity 19, the static pressure of the gas flow 20 in the region of entry into the cavity 19 is below the pressure of the cooling air 21b in the secondary air region 21d outside the annulus.

Wie Fig. 2 entnommen werden kann, stellt sich bei der aus dem Stand der Technik gemäß Fig. 1 bekannten Strömungsmaschine stromaufwärts der Vorderkanten 16 der Stützrippen 15 in Folge einer Verblockung der den Übergangsströmungskanal 33 durchströmenden Gasströmung auf Umfangspositionen, auf welchen die Stützrippen positioniert sind, ein Druckanstieg +Δp des statischen Drucks ein, wohingegen sich gemäß Fig. 2 auf Umfangspositionen zwischen benachbarten Stützrippen 15 ein Druckabfall -Δp des statischen Drucks einstellt. In Fig. 2 ist eine dimensionslose Umfangsrichtung u/t gezeigt, wobei t der Stützrippenteilung in Umfangsrichtung u entspricht.As Fig. 2 can be removed, arises in the from the prior art according to Fig. 1 known turbomachine upstream of the leading edges 16 of the support ribs 15 due to a blocking of the gas flow flowing through the transitional flow channel 33 at circumferential positions on which the support ribs are positioned, a pressure increase + .DELTA.p of the static pressure, whereas according Fig. 2 on circumferential positions between adjacent support ribs 15 sets a pressure drop -Δp of the static pressure. In Fig. 2 is shown a dimensionless circumferential direction u / t, where t corresponds to the support rib pitch in the circumferential direction u.

Die in Fig. 2 durch gestrichelte Linien dargestellten Druckfelder des Druckanstiegs +Δp auf der Umfangsposition der Stützrippen 15 und des Druckabfalls -Δp auf der Umfangsposition zwischen benachbarten Stützrippen 15 jeweils stromaufwärts der Vorderkanten 16 der Stützrippen 15 reicht in die Kavität 19 hinein, sodass sich im Mündungsbereich der Kavität 19 und im Übergangsströmungskanal 33 eine verlustbehaftete Sekundärströmung 22 ausbildet. Weiter führt die Druckschwankung gemäß Fig. 2 in der Kavität zu einem höheren Druckgefälle zwischen der Gasströmung 20 und der Kühlluftströmung 21b, was letztendlich die Leckage erhöht und zu einem verschlechterten Wirkungsgrad der Strömungsmaschine führt.In the Fig. 2 indicated by dashed lines pressure fields of the pressure increase + Δp on the circumferential position of the support ribs 15 and the pressure drop -Δp on the circumferential position between adjacent support ribs 15 respectively upstream of the leading edges 16 of the support ribs 15 extends into the cavity 19, so that in the mouth region of the cavity 19 and in the transitional flow passage 33, a lossy secondary flow 22 is formed. Next leads the pressure fluctuation according to Fig. 2 in the cavity to a higher pressure gradient between the gas flow 20 and the cooling air flow 21b, which ultimately increases the leakage and leads to a deteriorated efficiency of the turbomachine.

Hiervon ausgehend liegt der vorliegenden Erfindung das Problem zu Grunde, ein Zwischengehäuse zu schaffen, mit Hilfe dessen der Wirkungsgrad gesteigert werden kann.On this basis, the present invention is based on the problem to provide an intermediate housing, by means of which the efficiency can be increased.

Dieses Problem wird durch ein Zwischengehäuse gemäß Anspruch 1 gelöst.This problem is solved by an intermediate housing according to claim 1.

Erfindungsgemäß weist die radial außen liegende Begrenzungswand zumindest in einem Abschnitt stromaufwärts der Stützrippe eine sich in Umfangsrichtung verändernde Kontur auf.According to the invention, the radially outer boundary wall has a contour which changes in the circumferential direction at least in a section upstream of the support rib.

Mit der Erfindung ist es möglich, der Ausbildung der sich nach dem Stand der Technik im Kühlluftströmungskanal einstellenden, verlustbehafteten Sekundärströmung effizient entgegen zu wirken. Da mit einem geringeren Druckgefälle zwischen der Gasströmung und der Kühlluftströmung gearbeitet werden kann, kann der Wirkungsgrad gegenüber dem Stand der Technik verbessert werden.With the invention, it is possible to effectively counteract the formation of the state of the art in the cooling air flow channel adjusting, lossy secondary flow. Since it is possible to work with a lower pressure gradient between the gas flow and the cooling air flow, the efficiency can be improved compared to the prior art.

Bevorzugte Weiterbildungen der Erfindung ergeben sich aus den Unteransprüchen und der nachfolgenden Beschreibung. Ausführungsbeispiele der Erfindung werden, ohne hierauf beschränkt zu sein, an Hand der Zeichnung näher erläutert. Dabei zeigt:

Fig. 1
einen stark schematisierten, ausschnittsweisen Längsschnitt durch eine aus dem Stand der Technik bekannte Strömungsmaschine im Bereich eines Zwischengehäuses und damit Strömungskanals zwischen zwei Turbinenkomponenten;
Fig. 2
einen Ausschnitt aus der Anordnung der Fig. 1 in radialer Blickrichtung;
Fig. 3
einen stark schematisierten, ausschnittsweisen Längsschnitt durch eine Strömungsmaschine im Bereich eines erfindungsgemäßen Zwischengehäuses, das zwischen zwei Turbinenkomponenten positioniert ist;
Fig. 4
ein Diagramm zur Verdeutlichung der Erfindung; und
Fig. 5
ein weiteres Diagramm zur Verdeutlichung der Erfindung.
Preferred embodiments of the invention will become apparent from the dependent claims and the description below. Embodiments of the invention will be described, without being limited thereto, with reference to the drawings. Showing:
Fig. 1
a highly schematic, partial longitudinal section through a known from the prior art turbomachine in the region of an intermediate housing and thus flow channel between two turbine components;
Fig. 2
a section of the arrangement of Fig. 1 in the radial direction;
Fig. 3
a highly schematic, fragmentary longitudinal section through a turbomachine in the region of an intermediate housing according to the invention, which is positioned between two turbine components;
Fig. 4
a diagram for illustrating the invention; and
Fig. 5
another diagram to illustrate the invention.

Die hier vorliegende Erfindung betrifft den Bereich mehrwelliger Strömungsmaschinen, insbesondere mehrwelliger Gastriebwerke, mit mehreren Verdichterkomponenten sowie mehreren Turbinenkomponenten. Der grundsätzliche Aufbau einer solchen Strömungsmaschine ist dem hier angesprochenen Fachmann geläufig und wurde bereits im Zusammenhang mit Fig. 1 beschrieben.The present invention relates to the field of multi-shaft turbomachinery, in particular multi-shaft gas engines, with several compressor components and several turbine components. The basic structure of such a turbomachine is familiar to the person mentioned here and has already been in connection with Fig. 1 described.

Die hier vorliegende Erfindung betrifft nun Details eines Zwischengehäuses 14 einer derartigen Strömungsmaschine, mithilfe derer der Eintritt einer in einem Kühlluftströmungskanal 19 geführten Kühlluftströmung in die vom Übergangsströmungskanal 33 des Zwischengehäuses 14 geführte Gasströmung verbessert werden kann, nämlich in einem Eintrittsbereich des Übergangsströmungskanals 33 stromaufwärts von im Übergangsströmungskanal 33 positionierten Stützrippen 15.The present invention now relates to details of an intermediate housing 14 of such a turbomachine, by means of which the entry of a guided in a cooling air flow passage 19 cooling air flow can be improved in the outflow passage 33 of the intermediate housing 14 guided gas flow, namely in an inlet region of the transitional flow channel 33 upstream of the transitional flow channel 33 positioned support ribs 15th

Die Erfindung ist sowohl bei einem Zwischengehäuse 14 einer zweiwelligen Strömungsmaschine, das sich zwischen einer Hochdruckturbine 11 sowie eine Niederdruckturbine 13 erstreckt, als auch bei einem Zwischengehäuse einer dreiwelligen Strömungsmaschine, das sich zwischen einer Hochdruckturbine und einer Mitteldruckturbine oder zwischen einer Mitteldruckturbine und einer Niederdruckturbine erstreckt, einsetzbar.The invention is applicable both to an intermediate housing 14 of a twin-shaft turbomachine which extends between a high-pressure turbine 11 and a low-pressure turbine 13, and to an intermediate housing of a three-shaft turbomachine which extends between a high-pressure turbine and a medium-pressure turbine or between a medium-pressure turbine and a low-pressure turbine, used.

Fig. 3 zeigt einen Ausschnitt aus einer Strömungsmaschine im Bereich eines Zwischengehäuses 14, eines Übergangsströmungskanals 33 dieses Zwischengehäuses 14 und einer stromaufwärts des Übergangsströmungskanals 33 positionierten, im gezeigten Ausführungsbeispiel als Hochdruckturbine 11 ausgebildeten Turbinenkomponente, wobei gemäß Fig. 3 der Kühlluftströmungskanal 19 von radial außen in den Übergangsströmungskanal 33 mündet, nämlich stromaufwärts von Stützrippen 15, die im Übergangsströmungskanal 33 positioniert sind. Der Kühlluftströmungskanal 19 wird dabei von der Vorderkante 34 des Zwischengehäuses 14 abschnittsweise begrenzt. Fig. 3 shows a section of a turbomachine in the region of an intermediate housing 14, a transitional flow channel 33 of this intermediate housing 14 and an upstream of the transitional flow channel 33 positioned, formed in the illustrated embodiment as a high-pressure turbine 11 turbine component, according to Fig. 3 the cooling air flow passage 19 opens radially outward into the transition flow passage 33, namely upstream of support ribs 15 positioned in the transition flow passage 33. The cooling air flow channel 19 is thereby limited in sections by the front edge 34 of the intermediate housing 14.

Der Übergangsströmungskanal 33 wird radial innen von einer statorseitigen Begrenzungswand 23 und radial außen ebenfalls von einer statorseitigen Begrenzungswand 24 begrenzt.The transitional flow channel 33 is bounded radially inwardly by a stator-side boundary wall 23 and also radially on the outside by a stator-side boundary wall 24.

An den Rotor 10 der Hochdruckturbine 11 grenzt radial außen eine Begrenzungswand 25 der Hochdruckturbine 11 an.A boundary wall 25 of the high-pressure turbine 11 adjoins the rotor 10 of the high-pressure turbine 11 radially on the outside.

Um nun einen ungehinderten Eintritt der vom Kühlluftströmungskanal 19 geführten Kühlluft in die die Hochdruckturbine 11 verlassende und vom Übergangsströmungskanal 33 des Zwischengehäuses 14 geführte Gasströmung zu ermöglichen, kann die radial außen liegende Begrenzungswand 24 des Übergangsströmungskanals 33 zumindest in einem Abschnitt stromaufwärts der Stützrippen 15 mit einer sich in Umfangsrichtung verändernden Kontur versehen sein.In order now to allow an unhindered entry of the cooling air flow channel 19 guided cooling air in the high pressure turbine 11 leaving and guided by the transition flow channel 33 of the intermediate housing 14 gas flow, the radially outer boundary wall 24 of the transitional flow channel 33 at least in a portion upstream of the support ribs 15 with a be provided in the circumferential direction changing contour.

Vorzugsweise weist die radial außen liegende Begrenzungswand 24 des Übergangsströmungskanals 33 zumindest in einem Übergangsabschnitt zwischen der Vorderkante 34 des Zwischengehäuses 14 und dem Übergangsströmungskanals 33 eine sich in Umfangsrichtung verändernde Kontur auf.Preferably, the radially outer boundary wall 24 of the transition flow channel 33 at least in a transition section between the front edge 34 of the intermediate housing 14 and the transition flow channel 33 has a circumferentially changing contour.

Diese sich in Umfangsrichtung verändernde Kontur der radial außen liegenden Begrenzungswand 24 des Übergangsströmungskanals 33 kann sich gemäß Fig. 3 auch bis in einen Bereich stromabwärts der Vorderkanten 16 der Stützrippen 15 erstrecken, wobei Fig. 3 zwei an unterschiedlichen Umfangpositionen u/t ausgebildete Konturen 24 und 24' für die radial äußere Begrenzungswand des Übergangsströmungskanals 33 zeigt.This circumferentially changing contour of the radially outer boundary wall 24 of the transition flow channel 33 may according to Fig. 3 extend to a region downstream of the leading edges 16 of the support ribs 15, wherein Fig. 3 two contours 24 and 24 'formed at different circumferential positions u / t for the radially outer boundary wall of the transitional flow channel 33.

Die radial außen liegende Begrenzungswand 24 des Übergangsströmungskanals 33 verfügt im Eintrittsbereich des Übergangsströmungskanals 33 stromaufwärts der Vorderkanten 16 der Stützrippen 15 über einen Begrenzungswandabschnitt bzw. Begrenzungswandpunkt 26 mit minimalem Krümmungsradius und demnach maximaler Krümmung.The radially outer boundary wall 24 of the transition flow channel 33 has in the inlet region of the transition flow channel 33 upstream of the leading edges 16 of the support ribs 15 via a boundary wall portion or boundary wall 26 with a minimum radius of curvature and thus maximum curvature.

Die Kontur der radial außen liegenden Begrenzungswand 24 des Übergangsströmungskanals 33 verändert sich dabei in Umfangsrichtung u bzw. u/t derart, dass sich eine Axialposition (Axialrichtung x) und/oder einer Radialposition (Radialrichtung r) des Begrenzungswandabschnitts bzw. Begrenzungswandpunkts 26 mit minimalem Krümmungsradius in Umfangsrichtung u bzw. u/t verändert.The contour of the radially outer boundary wall 24 of the transition flow channel 33 changes in the circumferential direction u or u / t such that an axial position (axial direction x) and / or a radial position (radial direction r) of the boundary wall section or boundary wall point 26 with a minimum radius of curvature in the circumferential direction u or u / t changed.

Vorzugsweise verändert sich in Umfangsrichtung sowohl die Axialposition als auch die Radialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius. In einer vereinfachten Ausführung der Erfindung ist es jedoch auch möglich, dass sich ausschließlich die Axialposition oder ausschließlich die Radialposition dieses Begrenzungswandpunkts 26 in Umfangsrichtung verändert.Preferably, in the circumferential direction, both the axial position and the radial position of the boundary wall point 26 change with a minimum radius of curvature. In a simplified embodiment of the invention, however, it is also possible that only the axial position or exclusively the radial position of this boundary wall point 26 changes in the circumferential direction.

Die Axialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius verändert sich in Umfangsrichtung u bzw. u/t derart, dass in etwa auf der Umfangsposition der Vorderkanten 16 der Stützrippen 15 dieser Begrenzungswandpunkt 26 in Axialrichtung x maximal stromaufwärts und in etwa auf einer Umfangsposition halber Teilung zwischen zwei benachbarten Stützrippen in Axialrichtung x maximal stromabwärts versetzt bzw. positioniert ist. Zwischen diesen maximalen stromaufwärtigen und stromabwärtigen Axialpositionen verändert sich die Axialposition des Begrenzungswandpunkts 26 in Umfangsrichtung kontinuierlich bzw. stetig.The axial position of the boundary wall 26 with minimal radius of curvature changes in the circumferential direction u or u / t such that approximately at the circumferential position of the leading edges 16 of the support ribs 15 of this boundary wall 26 in the axial direction x maximum upstream and approximately in a circumferential position half pitch between two adjacent supporting ribs in the axial direction x offset or positioned downstream of maximum. Between these maximum upstream and downstream axial positions, the axial position of the boundary wall 26 gradually changes in the circumferential direction.

Die Radialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius verändert sich in Umfangsrichtung u bzw. u/t derart, dass in etwa auf der Umfangsposition der Vorderkanten 16 der Stützrippen 15 dieser Begrenzungswandpunkt 26 in Radialrichtung r maximal nach radial außen und in etwa auf einer Umfangsposition halber Teilung zwischen zwei benachbarten Stützrippen 15 in Radialrichtung r maximal nach radial innen versetzt bzw. positioniert ist. Zwischen diesen maximalen radial inneren und radial äußeren Radialpositionen verändert sich die Radialposition des Begrenzungswandpunkts 26 in Umfangsrichtung kontinuierlich bzw. stetig.The radial position of the boundary wall 26 with minimal radius of curvature changes in the circumferential direction u or u / t such that approximately at the circumferential position of the leading edges 16 of the support ribs 15 of this boundary wall 26 in the radial direction r maximum radially outward and approximately in a circumferential position half pitch between two adjacent support ribs 15 in the radial direction r is offset or positioned maximally radially inward. Between these maximum radially inner and radially outer radial positions, the radial position of the boundary wall point 26 changes continuously or continuously in the circumferential direction.

Die in Fig. 3 gezeigte Kontur 24 der radial äußeren Begrenzungswand des Übergangsströmungskanals 33 entspricht der Kontur derselben in etwa auf der Umfangsposition einer Vorderkante 16 einer Stützrippe 15, wohingegen die in Fig. 3 gezeigte Kontur 24' der Kontur derselben in etwa auf einer Umfangsposition halber Teilung zwischen zwei benachbarten Stützrippen 15 entspricht.In the Fig. 3 shown contour 24 of the radially outer boundary wall of the transition flow channel 33 corresponds to the contour thereof approximately at the circumferential position of a front edge 16 of a support rib 15, whereas the in Fig. 3 shown contour 24 'of the same contour approximately in a circumferential position half pitch between two adjacent support ribs 15 corresponds.

Weitere Details hinsichtlich des Versatzes der Axialposition sowie Radialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius in Umfangsrichtung u bzw. u/t werden nachfolgend unter Bezugnahme auf Fig. 4 beschrieben.Further details regarding the offset of the axial position as well as the radial position of the boundary wall point 26 with the minimum radius of curvature in the circumferential direction u or u / t will be described below with reference to FIG Fig. 4 described.

In Fig. 4 ist auf der horizontal verlaufenden Achse ein betragsmäßiges Verhältnis Δx/xKS zwischen dem Axialabstand Δx (siehe Fig. 3) der stromabwärtigen Axialposition und der maximal stromaufwärtigen Axialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius und dem Axialabstand ΔKS (siehe Fig. 3) eines stromabwärtigen Endes 27 der radial außenliegenden Begrenzungswand 25 der stromaufwärts des Übergangskanals 33 positionierten Turbinenkomponente 11 und der Vorderkante 16 der Stützrippen 15 aufgetragen. Weiterhin ist in Fig. 4 auf der horizontal verlaufenden Achse ein betragsmäßiges Verhältnis Δr/xKS zwischen dem Radialabstand Δr (siehe Fig. 3) der maximal radial äußeren Radialposition und der radial inneren Radialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius und diesem Axialabstand xKS aufgetragen. Wie bereits erwähnt, entspricht xKS (siehe Fig. 3) dem Abstand zwischen dem stromabwärtigen Ende 27 der radial außenliegenden Begrenzungswand 25 der Hochdruckturbine 11 und der Vorderkante 16 der Stützrippen 15.In Fig. 4 is on the horizontally extending axis an absolute value ratio Δx / x KS between the axial distance Δx (see Fig. 3 ) Downstream of the axial position and the axial position of the maximum of the upstream boundary wall point 26 with a minimum radius of curvature and the axial distance Δ KS (see Fig. 3 ) of a downstream end 27 of the radially outer boundary wall 25 of the upstream of the transition channel 33 positioned turbine component 11 and the front edge 16 of the support ribs 15 applied. Furthermore, in Fig. 4 on the horizontally extending axis an absolute value ratio Δr / x KS between the radial distance Δr (see Fig. 3 ) is applied to the maximum radially outer radial position and the radially inner radial position of the boundary wall point 26 with a minimum radius of curvature and this axial distance x KS . As already mentioned, x corresponds to KS (see Fig. 3 ) the distance between the downstream end 27 of the radially outer boundary wall 25 of the high-pressure turbine 11 and the front edge 16 of the support ribs 15th

Auf der vertikal verlaufenden Achse ist in Fig. 4 die dimensionslose Umfangsrichtung u/t aufgetragen, wobei auf den Umfangspositionen u/t=0 und u/t=1 jeweils eine Vorderkante 16 einer Stützrippe 15 positioniert ist, und wobei eine Umfangsposition u/t=0.5 einer Umfangsposition in der Mitte zwischen zwei benachbarten Stützrippen 15 entspricht.On the vertical axis is in Fig. 4 plotted on the circumferential positions u / t = 0 and u / t = 1 respectively a leading edge 16 of a support rib 15, and a circumferential position u / t = 0.5 of a circumferential position in the middle between two adjacent Support ribs 15 corresponds.

So kann Fig. 4 entnommen werden, dass sich die Verhältnisse Δx/xKS und Δr/xKS in dimensionsloser Umfangsrichtung u/t gesehen zwischen zwei benachbarten Stützrippen 15 kontinuierlich verändern, wobei auf der Umfangsposition u/t=0.5 von in etwa halber Teilung zwischen zwei benachbarten Stützrippen 15 das Verhältnis Δx/xKS und damit der Versatz der Axialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius nach stromabwärts sowie das Verhältnis Δr/xKS und damit der Versatz der Radialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius nach radial innen am größten sind, und in etwa auf den Umfangspositionen u/t=0 und u/t=1, auf denen die Vorderkanten 16 der Stützrippen 15 positioniert sind, diese Verhältnisse und damit Versatze am Kleinsten sind.So can Fig. 4 It can be seen that the ratios .DELTA.x / x KS and .DELTA.r / x KS in the dimensionless circumferential direction u / t continuously change between two adjacent support ribs 15, wherein on the circumferential position u / t = 0.5 of approximately half the pitch between two adjacent support ribs 15 the ratio .DELTA.x / x KS and thus the offset of the axial position of the boundary wall 26 with minimum radius of curvature to the downstream and the ratio .DELTA.r / x KS and thus the offset of the radial position of the boundary wall 26 with the minimum radius of curvature radially inward are greatest, and in about the circumferential positions u / t = 0 and u / t = 1, on which the leading edges 16 of the support ribs 15 are positioned, these ratios and thus offsets are the smallest.

Der Bereich 28 der Fig. 4 visualisiert einen bevorzugten Gültigkeitsbereich für das sich in Umfangsrichtung u bzw. u/t verändernde Verhältnis Δx/xKS und/oder Δr/xKS und damit den sich in Umfangsrichtung u bzw. u/t verändernden Versatz der Axialposition und/oder der Radialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius.The area 28 of the Fig. 4 visualizes a preferred scope for extending u and u / t changing in the circumferential direction ratio Ax / x KS and / or AR / x KS and thus the to u and u / t changing in the circumferential direction of offset of the axial position and / or the radial position of the Boundary wall point 26 with a minimum radius of curvature.

Die Verhältnisse Δx/xKS und Δr/xKS betragen bis zu 40%.The ratios Δx / x KS and Δr / x KS are up to 40%.

Die Verhältnisse Δx/xKS und Δr/xKS betragen auf der Umfangsposition u/t=0.5 von in etwa halber Teilung zwischen zwei Stützrippen 15 maximal 40% und minimal 2%. Die Verhältnisse Δx/XKS und Δr/xKS betragen auf den Umfangspositionen u/t=0 und u/t=1 0%. Dazwischen verändern sich diese Verhältnisse Δx/xKS und Δr/xKS kontinuierlich, steig und vorzugsweise nicht linear.The ratios .DELTA.x / x KS and .DELTA.r / x KS are at the circumferential position u / t = 0.5 of about half pitch between two support ribs 15 maximum 40% and a minimum of 2%. The ratios Δx / X KS and Δr / x KS are on the circumferential positions u / t = 0 and u / t = 1 0%. In between, these ratios Δx / x KS and Δr / x KS change continuously, increasing and preferably not linearly.

Insbesondere beträgt das sich in Umfangsrichtung u bzw. u/t verändernde Verhältnis Δx/sKS auf der Umfangsposition u/t=0.5 von in etwa halber Teilung zwischen zwei Stützrippen 15 insbesondere zwischen 2% und 25%.In particular, the ratio Δx / s KS changing in the circumferential direction u or u / t at the circumferential position u / t = 0.5 of approximately half the pitch between two support ribs 15 is in particular between 2% and 25%.

Das sich in Umfangsrichtung u bzw. u/t verändernde Verhältnis Δr/xKS beträgt auf der Umfangsposition u/t=0.5 von in etwa halber Teilung zwischen zwei Stützrippen 15 insbesondere zwischen 2% und 5%.The ratio Δr / x KS which changes in the circumferential direction u or u / t is in the circumferential position u / t = 0.5 of approximately half the pitch between two support ribs 15, in particular between 2% and 5%.

Die Kurve 29 innerhalb des Bereichs 28 visualisiert das bevorzugte, sich in Umfangsrichtung verändernde Verhältnis Δx/xKS und damit den sich in Umfangsrichtung verändernden Versatz der Axialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius, wobei gemäß der Kurve 29 der Versatz der Axialposition im Bereich halber Teilung zwischen zwei benachbarten Stützrippen am Größten ist und das Verhältnis Δx/xKS in etwa 20% beträgt.Curve 29 within region 28 visualizes the preferred circumferentially varying ratio Δx / x KS, and hence the circumferentially varying offset of the axial position of the minimum wall radius limiting wall 26, where, according to curve 29, the axial position offset is in the half pitch range between two adjacent support ribs is largest and the ratio Δx / x KS is about 20%.

Die Kurve 30 innerhalb des Bereichs 28 verdeutlicht das bevorzugte, sich in Umfangsrichtung verändernde Verhältnis Δr/xKS und damit den sich in Umfangsrichtung verändernden Versatz der Radialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius, wobei bei in etwa halber Teilung zwischen benachbarten Stützrippen das Verhältnis Δr/xKS in etwa 2.5% beträgt und der Versatz der Radialposition im Bereich halber Teilung zwischen zwei benachbarten Stützrippen am Größten ist.Curve 30 within region 28 illustrates the preferred circumferentially varying ratio Δr / x KS, and thus the circumferentially varying offset of the radial position of boundary wall 26 with minimum radius of curvature, and at approximately half pitch between adjacent support ribs, the ratio Δr / x KS is about 2.5% and the offset of the radial position in the area of half pitch between two adjacent support ribs is the largest.

In Umfangsrichtung gesehen verändern sich der Versatz der Axialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius und der Versatz der Radialposition des Begrenzungswandpunkts 26 mit minimalem Krümmungsradius bzw. die obigen Verhältnisse Δx/xKS und Δr/xKS jeweils kontinuierlich bzw. stetig und vorzugsweise nicht linear.In the circumferential direction, the offset of the axial position of the boundary wall 26 having the minimum radius of curvature and the offset of the radial position of the boundary wall 26 having the minimum radius of curvature and the above ratios Δx / x KS and Δr / x KS are respectively continuous and continuous, and preferably nonlinear.

Fig. 5 visualisiert den Effekt der erfindungsgemäßen Konturierung der radial außen liegenden Begrenzungswand 24 des Übergangsströmungskanals 33 wobei in Fig. 5 auf der horizontal verlaufenden Achse ein Verhältnis (p-pm)/pm zwischen der Differenz (p-pm) des statischen Drucks p der Gasströmung im Übergangsströmungskanal 14 und dem Mittelwert pm dieses statischen Drucks und dem Mittelwert pm aufgetragen ist, und wobei auf der vertikal verlaufenden Achse die dimensionslose Umfangsrichtung u/t aufgetragen ist. Fig. 5 visualizes the effect of the contouring according to the invention of the radially outer boundary wall 24 of the transitional flow channel 33 wherein Fig. 5 is plotted on the horizontal axis a ratio (pp m ) / p m between the difference (pp m ) of the static pressure p of the gas flow in the transitional flow channel 14 and the mean value p m of this static pressure and the mean value p m ; vertically extending axis the dimensionless circumferential direction u / t is plotted.

Die Kurve 31 der Fig. 5 entspricht einem sich nach dem Stand der Technik einstellenden Verlauf des Verhältnisses (p-pm)/pm und die Kurve 32 dem sich nach der Erfindung einstellenden Verlauf des Verhältnisses (p-pm)/pm.The curve 31 of the Fig. 5 corresponds to a state of the art adjusting course of the ratio (pp m ) / p m and the curve 32 according to the invention adjusting the course of the ratio (pp m ) / p m .

Fig. 5 kann entnommen werden, dass mit der Erfindung ein verbesserter, gleichförmiger Druckverlauf des statischen Drucks in Umfangsrichtung bereitgestellt werden kann, wodurch der Ausbildung einer Sekundärströmung im Mündungsabschnitt des Kühlluftströmungskanals 19 in den Übergangsströmungskanal 33 effektiv entgegengewirkt werden kann. Dadurch kann ein ungehinderter Eintritt der Kühlluftströmung in den Übergangsströmungskanal 33 gewährleistet werden, wodurch der Wirkungsgrad der Strömungsmaschine verbessert werden kann. Weiterhin kann die Strömung im Übergangsströmungskanal 33 zwischen benachbarten Stützrippen 15 verbessert werden. Fig. 5 It can be seen that with the invention, an improved, uniform pressure distribution of the static pressure in the circumferential direction can be provided, whereby the formation of a secondary flow in the mouth portion of the cooling air flow channel 19 in the transitional flow channel 33 can be effectively counteracted. Thereby, an unhindered entry of the cooling air flow into the transitional flow channel 33 can be ensured, whereby the efficiency of the turbomachine can be improved. Furthermore, the flow in the transitional flow passage 33 between adjacent support ribs 15 can be improved.

Claims (6)

  1. A housing (14) of a gas engine having a radially inner boundary wall (23) and having a radially outer boundary wall (24, 24'), having a crossflow channel (33) which is formed by the boundary walls (23, 24, 24') and in which there is positioned at least one supporting rib (15) that has a leading edge (16), a trailing edge (17) and also side walls (18) extending between the leading edge (16) and the trailing edge (18) and guiding a gas flow flowing through the crossflow channel (33), wherein the radially outer boundary wall (24) has at least in a section upstream of the supporting rib (15) a contour that changes in the peripheral direction, and wherein the contour of the radially outer boundary wall (24) changes in such a way that an axial and/or radial position of a boundary-wall section or a boundary-wall point (26) with a minimum radius of curvature changes in the peripheral direction,
    characterised in that the housing (14) is an intermediate housing (14) of turbines (11, 13) of the gas engine, wherein
    - the axial position of the boundary-wall section or boundary-wall point (26) with a minimum radius of curvature changes in the peripheral direction in such a way that this boundary-wall point (26) is positioned maximally upstream at the peripheral position of leading edges (16) of the supporting ribs (15) and maximally downstream substantially at a peripheral position of a half pitch between two adjacent supporting ribs; and/or
    - the radial position of the boundary-wall section or boundary-wall point (26) with a minimum radius of curvature changes in the peripheral direction in such a way that this boundary-wall point (26) is positioned maximally radially outwards substantially at the peripheral position of leading edges (16) of the supporting ribs (15) and maximally radially inwards substantially at a peripheral position of a half pitch between two adjacent supporting ribs.
  2. A housing (14) according to claim 1, characterised in that the radially outer boundary wall (24) of the crossflow channel (33) has, at least in a crossover section between a leading edge (34) of the intermediate housing (14) and the crossflow channel (33), a contour that changes in the peripheral direction.
  3. A housing (14) according to one of the preceding claims, characterised in that an absolute-value ratio between the axial spacing of the downstream and the maximally upstream axial position of the boundary-wall point (26) with a minimum radius of curvature and the axial spacing of a downstream end (27) of a radially outer boundary wall (25) of a turbine component (11), positioned upstream of the crossflow channel (33), and the leading edge (16) of the supporting ribs (15) amounts to up to 40%.
  4. A housing (14) according to claim 3, characterised in that the ratio amounts to up to 25%.
  5. A housing (14) according to one of the preceding claims, characterised in that an absolute-value ratio between the radial spacing of the maximally radially outer and the radially inner radial position of the boundary-wall point (26) with a minimum radius of curvature and the axial spacing between a downstream end (27) of a radially outer housing wall (25) of a turbine component (11), positioned upstream of the crossflow channel (33), and the leading edge (16) of the supporting ribs (15) amounts to up to 40%.
  6. A housing (14) according to claim 5, characterised in that the ratio amounts to up to 5%.
EP12716196.6A 2011-01-19 2012-01-16 Intermediate casing of a gas turbine engine comprising an outer boundary wall wich comprises upstream of a support strut a variable contour in circumferential direction in order to reduce secondary flow losses Not-in-force EP2665896B1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102011008812A DE102011008812A1 (en) 2011-01-19 2011-01-19 intermediate housing
PCT/DE2012/000032 WO2012097798A1 (en) 2011-01-19 2012-01-16 Intermediate housing of a gas turbine with an outer bounding wall, having upstream of a supporting rib a contour that changes in the circumferential direction, for reducing secondary flow losses

Publications (2)

Publication Number Publication Date
EP2665896A1 EP2665896A1 (en) 2013-11-27
EP2665896B1 true EP2665896B1 (en) 2015-06-10

Family

ID=45999502

Family Applications (1)

Application Number Title Priority Date Filing Date
EP12716196.6A Not-in-force EP2665896B1 (en) 2011-01-19 2012-01-16 Intermediate casing of a gas turbine engine comprising an outer boundary wall wich comprises upstream of a support strut a variable contour in circumferential direction in order to reduce secondary flow losses

Country Status (4)

Country Link
US (1) US9382806B2 (en)
EP (1) EP2665896B1 (en)
DE (1) DE102011008812A1 (en)
WO (1) WO2012097798A1 (en)

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9598981B2 (en) * 2013-11-22 2017-03-21 Siemens Energy, Inc. Industrial gas turbine exhaust system diffuser inlet lip
ES2632613T3 (en) 2014-08-29 2017-09-14 MTU Aero Engines AG Gas turbine construction group
DE102017222193A1 (en) 2017-12-07 2019-06-13 MTU Aero Engines AG Turbomachinery flow channel

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009019282A2 (en) * 2007-08-06 2009-02-12 Alstom Technology Ltd Gap cooling between a combustion chamber wall and a turbine wall of a gas turbine installation

Family Cites Families (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19650656C1 (en) 1996-12-06 1998-06-10 Mtu Muenchen Gmbh Turbo machine with transonic compressor stage
EP1515000B1 (en) * 2003-09-09 2016-03-09 Alstom Technology Ltd Blading of a turbomachine with contoured shrouds
EP1760257B1 (en) * 2004-09-24 2012-12-26 IHI Corporation Wall shape of axial flow machine and gas turbine engine
US7179049B2 (en) 2004-12-10 2007-02-20 Pratt & Whitney Canada Corp. Gas turbine gas path contour
US8511978B2 (en) 2006-05-02 2013-08-20 United Technologies Corporation Airfoil array with an endwall depression and components of the array
US7594405B2 (en) * 2006-07-27 2009-09-29 United Technologies Corporation Catenary mid-turbine frame design
JP5283855B2 (en) * 2007-03-29 2013-09-04 株式会社Ｉｈｉ Turbomachine wall and turbomachine
DE102008021053A1 (en) * 2008-04-26 2009-10-29 Mtu Aero Engines Gmbh Reformed flow path of an axial flow machine to reduce secondary flow
DE102008031789A1 (en) * 2008-07-04 2010-01-07 Deutsches Zentrum für Luft- und Raumfahrt e.V. Method and device for influencing secondary flows in a turbomachine
DE102008060847B4 (en) 2008-12-06 2020-03-19 MTU Aero Engines AG Fluid machine
EP2248996B1 (en) * 2009-05-04 2014-01-01 Alstom Technology Ltd Gas turbine
EP2261462A1 (en) 2009-06-02 2010-12-15 Alstom Technology Ltd End wall structure for a turbine stage

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009019282A2 (en) * 2007-08-06 2009-02-12 Alstom Technology Ltd Gap cooling between a combustion chamber wall and a turbine wall of a gas turbine installation

Also Published As

Publication number Publication date
US9382806B2 (en) 2016-07-05
DE102011008812A1 (en) 2012-07-19
US20130064657A1 (en) 2013-03-14
EP2665896A1 (en) 2013-11-27
WO2012097798A1 (en) 2012-07-26

Similar Documents

Publication Publication Date Title
EP2356321B1 (en) Turbo engine
DE10210866C5 (en) Guide vane mounting in a flow channel of an aircraft gas turbine
EP0799973B1 (en) Wall contour for an axial turbomachine
DE112015001237B4 (en) Exhaust gas turbocharger
EP2179143B1 (en) Gap cooling between combustion chamber wall and turbine wall of a gas turbine installation
EP1621733A2 (en) Flow device for a gas turbine
EP2639411B1 (en) Casing of a turbomachine with a by-passing system
DE102006048933A1 (en) Arrangement for influencing the flow
EP1632648B1 (en) Gas turbine with transition duct
DE102015219556A1 (en) Diffuser for radial compressor, centrifugal compressor and turbo machine with centrifugal compressor
EP2140111B1 (en) Turbomachine
DE102007050916A1 (en) Stator arrangement for compressor of fluid conveying arrangement in gas turbine engine, has radial passage conduit formed in part of stator ring segment, where radial passage conduit is arranged adjacent to stator blade passage conduit
EP2665896B1 (en) Intermediate casing of a gas turbine engine comprising an outer boundary wall wich comprises upstream of a support strut a variable contour in circumferential direction in order to reduce secondary flow losses
WO1999013199A1 (en) Blade for a turbo-machine and steam turbine
EP2526263A2 (en) Housing system for an axial turbomachine
EP3034944A1 (en) Gas turbine combustor with altered wall thickness
DE102013210171A1 (en) Structural assembly for a turbomachine
EP3686396B1 (en) Lead frame for a flow engine
WO2018157957A1 (en) Turbine casing and method for assembling a turbine having a turbine casing
EP1731711A1 (en) Transition from combustion chamber to turbine, heat shield, and turbine vane in a gas turbine
EP3670845A1 (en) Static seal assembly and turbomachine
EP3581760A1 (en) Turbine blade for a flow engine
EP1895105A1 (en) Cooling method for stator vanes and sector of a stator with at least two airfoils
WO2021110191A1 (en) Seal carrier for a turbomachine, having slot-like openings in the seal body
DE102018201389A1 (en) Turbine with axially extending wave-like annular space contour

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

17P Request for examination filed

Effective date: 20120823

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

DAX Request for extension of the european patent (deleted)
17Q First examination report despatched

Effective date: 20140623

REG Reference to a national code

Ref country code: DE

Ref legal event code: R079

Ref document number: 502012003417

Country of ref document: DE

Free format text: PREVIOUS MAIN CLASS: F01D0005140000

Ipc: F16B0037040000

RIC1 Information provided on ipc code assigned before grant

Ipc: F16B 37/04 20060101AFI20141128BHEP

Ipc: F01D 9/02 20060101ALI20141128BHEP

Ipc: F01D 5/14 20060101ALI20141128BHEP

GRAP Despatch of communication of intention to grant a patent

Free format text: ORIGINAL CODE: EPIDOSNIGR1

INTG Intention to grant announced

Effective date: 20150108

GRAS Grant fee paid

Free format text: ORIGINAL CODE: EPIDOSNIGR3

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

REG Reference to a national code

Ref country code: GB

Ref legal event code: FG4D

Free format text: NOT ENGLISH

REG Reference to a national code

Ref country code: CH

Ref legal event code: EP

REG Reference to a national code

Ref country code: AT

Ref legal event code: REF

Ref document number: 731038

Country of ref document: AT

Kind code of ref document: T

Effective date: 20150715

REG Reference to a national code

Ref country code: DE

Ref legal event code: R096

Ref document number: 502012003417

Country of ref document: DE

REG Reference to a national code

Ref country code: IE

Ref legal event code: FG4D

Free format text: LANGUAGE OF EP DOCUMENT: GERMAN

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: ES

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: FI

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: NO

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150910

Ref country code: LT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

REG Reference to a national code

Ref country code: NL

Ref legal event code: MP

Effective date: 20150610

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: GR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150911

Ref country code: LV

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: RS

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: BG

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150910

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 5

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: EE

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: PT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20151012

Ref country code: SK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: PL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: CZ

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: RO

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20150610

Ref country code: IS

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20151010

REG Reference to a national code

Ref country code: DE

Ref legal event code: R097

Ref document number: 502012003417

Country of ref document: DE

PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: DK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: IT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

26N No opposition filed

Effective date: 20160311

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: BE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20160131

Ref country code: SI

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: LU

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20160116

REG Reference to a national code

Ref country code: CH

Ref legal event code: PL

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: MC

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: CH

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20160131

Ref country code: LI

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20160131

REG Reference to a national code

Ref country code: IE

Ref legal event code: MM4A

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 6

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: IE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20160116

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: SE

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: NL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: MT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 7

REG Reference to a national code

Ref country code: AT

Ref legal event code: MM01

Ref document number: 731038

Country of ref document: AT

Kind code of ref document: T

Effective date: 20170116

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: CY

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: SM

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: AT

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20170116

Ref country code: HU

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT; INVALID AB INITIO

Effective date: 20120116

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: HR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: MK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

Ref country code: TR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: AL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20150610

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: FR

Payment date: 20230123

Year of fee payment: 12

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: GB

Payment date: 20230124

Year of fee payment: 12

Ref country code: DE

Payment date: 20230119

Year of fee payment: 12

REG Reference to a national code

Ref country code: DE

Ref legal event code: R119

Ref document number: 502012003417

Country of ref document: DE

GBPC Gb: european patent ceased through non-payment of renewal fee

Effective date: 20240116

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: DE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20240801

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: GB

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20240116

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: FR

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20240131