EP2824279B1 - Turbomachine with sealing arrangement - Google Patents

Description

The invention relates to a turbomachine according to the preamble of patent claim 1.

In order to prevent a hot gas from a radially outer hot gas channel in a radially inner cooling channel through an axial gap between a series of airfoils and a downstream blade row in a turbomachine such as a gas turbine, so-called Fischmauldichtungen be used in which a rotating rotor ring of the blade row in an annular sealing space runs, which is bounded by a radially inner stator ring and a radially outer stator ring of the row of guide vanes. In the sealing space vortex form, which should prevent the ingress of hot gas from the hot gas channel into the cooling channel. The inner stator ring can as in the US 2010172749 A1 shown formed as an integral component with the outer stator ring or as in the DE 2847012 A1 shown as a separate component.

From the US 7540709 B1 is a turbomachine with a sealing device between a row of vanes and a blade row is known, which has a radially inner hook-like stator ring and a radially outer hook-like rotor ring. The two rings are in axial overlap, with a sealing space formed between them due to their hook-like design.

In the US 8075256 B2 a turbomachine is shown with a sealing device which is formed as a labyrinth seal between a radially outer stator ring and a radially inner rotor ring. The rings are in axial overlap, wherein the rotor ring in addition to its radially outwardly directed end portion to approximately its half axial extent has a circumferential radially outwardly directed sealing ridge. The stator ring has a radially inwardly directed end portion.

From the EP 2573329 A2 FIG. 2 shows a turbomachine with a sealing device between a stationary airfoil row arranged in a turbine intermediate housing and a subsequent blade row.

Other turbomachinery with sealing devices are from the US2010183426A1 . EP1380726A2 and US2010074733A1 known.

The object of the invention is to provide a turbomachine having an alternative sealing means for sealing an axial gap between a series of airfoils and a downstream blade row.

This object is achieved by a turbomachine having the features of patent claim 1.

A turbomachine according to the invention with a series of airfoils in the form of a guide vane row and a row of blades downstream of the airfoil, between which a radially inner cooling channel is formed, the airfoil row and the blade row each having a platform ring defining a radially outer hot gas channel, has a sealing device for preventing a Heißgaseinzuges through an axial gap between the rows of the hot gas duct in the cooling channel. According to the invention, the sealing device has a downstream-oriented guide vane-side axial ring, which extends axially beyond a platform ring trailing edge of the airfoil row and has a radially outwardly directed end portion. In addition, the sealing device has an upstream platform ring overhang, which is arranged on the blade row and having a radially inwardly directed end portion, wherein the axial ring is arranged radially inward to the platform ring overhang and forms an axial overlap with this. In this case, the front platform ring overhang follows, at least in sections, an ideal flow pattern of the hot gas over the rotor blade-side platform ring.

The formation of the airfoil series without a downstream or rear platform ring overhang in conjunction with the platform ring radially inwardly running rear hook-like axial ring and with it in axial overlap with the axial ring located radially outer front hook-like platform ring overhang causes a sealing space or a vortex zone in which / is applied to the incoming hot gas with a high tangential velocity component. The inflowing hot gas is swirled and thereby made much more difficult or prevented at its entry into the cooling channel. In addition, the sealing space acts as a damping area between the hot gas duct and the inner cooling channel, whereby pressure differences between the hot gas duct and the inner cooling channel are weakened, whereby also a hot gas from the hot gas duct into the inner cooling channel is difficult. Furthermore, the sealing device according to the invention allows large axial and radial relative movements of the rotor and the stator to each other in comparison to conventional Fischmauldichtungen at a small axial gap. Due to the fact that the front platform ring overhang follows, at least in sections, an ideal flow course of the hot gas over the blade row side platform ring, in the region of the platform overhang in the hot gas channel, eddies are prevented or substantially reduced. The sealing device is preferably arranged on the turbine side of the turbomachine such as a gas turbine and in particular an aircraft engine. For example, the airfoil series is a vane row in a low-pressure turbine of the turbomachine. However, the airfoil series can also consist of a plurality of blade profiles arranged in a turbine intermediate housing between a high-pressure turbine and a low-pressure turbine.

In one embodiment, the end portions are spaced apart in the radial direction. As a result, a radial gap is essentially created, can be blown through the cooling air from the cooling passage in the sealing chamber.

In another embodiment, the end portions terminate at the same radial height. As a result, an axial gap is substantially created, can be blown through the cooling air from the cooling passage in the sealing chamber.

Preferably, the end portion of the axial ring has a greater radial extent than the end portion of the front platform ring overhang. The end portion of the rotating platform ring overhang is shortened in this embodiment compared to the end portion of the fixed axial ring. The end portions of the platform ring overhang are through the stabilizes short radial extent, whereby, an introduction of disturbing vibrations in the rotor by the rotating end portions is prevented.

In one embodiment, the end portions are parallel and oriented radially offset from one another. For example, the end portions are orthogonal to the axial direction of the turbomachine. Due to the parallelism, an axial component of the radial gap is kept constant despite different thermal radial expansion behavior of the axial ring and the platform ring overhang or of the stator and of the rotor.

In another embodiment, the end portions are set to each other. For example, the end portion of the axial ring is inclined downstream in the axial direction and the end portion of the front platform ring overhang is orthogonal to the axial direction. Due to the inclination or inclination of the gap between the cooling chamber and the sealing space can be opened.

A vane for a turbomachine according to the invention may have a downstream axial overhang extending axially beyond a platform trailing edge and having a radially outwardly directed end portion. Such a guide vane allows the formation of a row of guide vanes into which the preferred sealing device for preventing fuel gas is integrated in sections.

A bucket for a turbomachine according to the invention may have an upstream front platform overhang having a radially inwardly directed end portion. Such a blade allows the formation of a blade row, in which the preferred sealing means for preventing Heizgaseizugs is partially integrated.

Figur 1

einen Längsschnitt durch eine erfindungsgemäße Strömungsmaschine im Bereich einer ersten beispielhaften Dichteinrichtung,

Figur 2

einen Längsschnitt durch eine erfindungsgemäße Strömungsmaschine im Bereich einer zweiten beispielhaften Dichteinrichtung, und

Figur 3

eine Seitendarstellung einer Laufschaufel.

In the following, preferred embodiments of the invention will be explained in more detail with reference to highly simplified schematic illustrations. Show it:

FIG. 1

a longitudinal section through a flow machine according to the invention in the region of a first exemplary sealing device,

FIG. 2

a longitudinal section through a flow machine according to the invention in the region of a second exemplary sealing device, and

FIG. 3

a page illustration of a blade.

In FIG. 1 is a longitudinal section through a turbomachine in the region of a guide blade row 1 and a row of blades 1 downstream blade row 2 shown. The guide blade row 1 and the blade row 2 have a plurality of blades 35, 36, which are arranged side by side in the circumferential direction of the turbomachine and their respective blade 4, 6 is arranged in the hot gas duct 8 of the turbomachine. The turbomachine is a gas turbine and in particular an aircraft engine. The guide vane row 1 and the blade row 2 are arranged in particular on the turbine side and, for example, in the low-pressure turbine of the turbomachine.

The hot gas channel 8 is from a hot gas stream as shown in FIG FIG. 2 flows through from left to right. It is bounded radially inwardly by a guide blade row side platform ring 10 and by a rotor blade side platform ring 12. Radially inside the hot gas duct 8, a cooling space 14 is formed, which is flowed through by a cooling air flow.

In order to prevent a hot gas inlet or an intake of a hot gas partial flow 16 through an axial gap between the guide vane row 1 and the blade row 2 from the hot gas channel 8 into the cooling space 14, a sealing device 18 is arranged in the axial gap between the guide vane row 1 and the blade row 2.

This in FIG. 1 shown first embodiment of the sealing device 18 forms a ring-like sealing space 20. The sealing space 20 is in the in FIG. 1 shown embodiment via an annular axial outer gap 21 to the hot gas channel 8 and via an annular axial inner gap 23 to the cooling channel 14. Through the outer gap 21 occurs the Hot gas partial flow 16 in the sealing chamber 20 and is swirled in this. For vortex guidance a cooling air partial stream 22 is simultaneously injected through the inner gap 23 into the sealing chamber 20, which lays on a hot gas partial flow vortex 25 in the sealing chamber 20 and exits into the hot gas channel 8.

The sealing device 18 has an opposite to a guide vane row side front platform ring trailing edge 24 radially inwardly disposed axial ring 26 which is oriented downstream in the direction of the blade row 2 and has a radially outwardly directed end portion 28. In addition, the sealing device 18 has a blade row-side front platform ring overhang 30, which is oriented downstream in the direction of the guide blade row 1 and has a radially inwardly directed end portion 32. The circumferential end portions 28, 32 thus have in opposite radial directions, wherein they face each other.

To form the sealing space 20, the axial ring 26 and the platform ring overhang 30 are in axial overlap. They are in the in FIG. 1 shown embodiment in axial overlap that the sealing space 20 is not open over its entire axial extent to the hot gas channel 8, but only in the area near the platform ring trailing edge 24th

The axial ring 26 extends with a cylindrical annular portion 34 in the axial direction of the turbomachine and is arranged radially inward to the platform ring overhang 30. Its end portion 28 extends from the ring portion 34 and is made inclined downstream, extending radially outwardly. The end portion 28 thus points in the direction of the front platform ring overhang 30.

The front platform ring overhang 30 follows in the illustrated embodiment with a ring portion 37 an ideal flow of the hot gas over the blade row side platform ring 12. He goes for this purpose on the one steplessly in the platform 12 and has viewed in the flow direction a funnel-shaped opening contour.

The end portion 32 of the front platform ring overhang 30 extends from the ring portion 37. It is orthogonal to the axial direction and thus orthogonal to the ring portion 34 of the axial ring 26, extending radially inwardly. The end portion 32 has a shorter radial extent than the end portion 28 of the axial ring 26. In the illustrated embodiment, it has a radial extent such that it ends at or about the same radial height as the end portion 28 of the axial ring 26. Downstream forms the end portion 32 of the platform ring overhang 30 together with the end portion 28 of the axial ring 26, the inner gap 23, the here in FIG. 1 almost an axial gap. Due to the inclination of the end portion 28 of the inner gap 23 is thereby axially expanded with respect to a non-inclined employment of the end portion 28. Upstream, the end portion 32 of the front platform ring overhang 30 forms the outer nip 21, together with the vane side platform ring trailing edge 24.

At the in FIG. 2 shown embodiment of a sealing device 18 between a stator blade 1 and a downstream blade row 2 of a turbomachine to prevent a hot gas inlet 16 from a radially outer hot gas channel 8 in a radially inner cooling channel 14 to form a sealing space 20 are essentially different from the embodiment according to FIG. 1 both end portions 28, 32 oriented orthogonal to the axial ring portion 34 and to the axial direction of the turbomachine. The end sections 28, 32 are aligned almost parallel to one another. However, in the radial direction of the turbomachine, they are objected to one another in such a way that a ring-like inner gap 23 for injecting a cooling air partial flow 22 into the sealing space 20 has a greater axial component than a radial component and thus is virtually a radial gap.

In addition, the end sections 28, 32 are substantially different from a guide blade-side platform ring trailing edge 24 in that a ring-like axial outer gap 21 extends virtually over the entire axial extent of the axial ring 26 and the sealing space 20 thus opens virtually over its entire axial extent to the hot gas channel 8 is.

As in the Figures 1 and 2 to recognize a respective vane 35 of the vane row 1 is formed without a rear platform overhang. The vane 35 has in each case one to a rear platform edge radially inwardly positioned axial projection with a radially to outwardly directed end region. In the state mounted to the guide vane row 1, the adjacent axial projections form the axial ring 26, wherein the end regions form the peripheral end section 28.

In FIG. 3 For example, a blade 36 usable for the invention is shown for forming an aforementioned blade row 2. The blade 36 has seen in the direction of flow of a hot gas, a front platform overhang 38 with a radially inwardly directed end portion 40. The front platform overhang 38 follows an ideal flow of the hot gas through a paddle platform 42. The front platform overhang 38 steplessly into the paddle platform 42 and over is in accordance with the blade platform 42 ascending inclined in the flow direction. In addition, the in FIG. 3 shown blade 36 a rear platform overhang 44 which is aligned almost axially to the axial direction. In the assembled state of the blade row 2 state, the adjacent front platform overhangs 38 form the front platform ring overhang 30, wherein the end portions 40 form the peripheral end portion 32.

LIST OF REFERENCE NUMBERS

1

vane row

2

Blade row

4

airfoil

6

airfoil

8th

Hot gas duct

10

platform ring

12

platform ring

14

cooling channel

16

Hot gas partial flow / hot gas intake

18

sealing device

20

tight space

21

outer gap

22

Cooling air partial stream

23

interior gap

24

Platform ring trailing edge

25

Hot gas partial flow vortices

26

axial ring

28

end

30

Platform ring overhang

32

end

34

ring section

35

vane

36

blade

37

ring section

38

Deck overhang

40

end

42

blade platform

44

front platform overhang

Claims (6)

1. Turbomachine comprising a stationary flow profile row (1), which is a guide vane row (1), and a downstream rotor blade row (2), between which a radially inner cooling channel (14) is formed, the guide vane row (1) and the rotor blade row (2) each comprising a platform ring (10, 12) which delimits a radially outer hot-gas channel (8), and comprising a sealing means (18) for preventing an intake of hot gas (16) into the cooling channel (14) from the hot-gas channel (8) through an axial gap between the rows (1, 2), the sealing means (18) having an axial ring (26) on the guide vane side, which ring is oriented downstream, extends axially over a platform ring rear edge (24) of the guide vane row (1) and has an end portion (28) that is directed radially outward, and the sealing means (18) comprising a platform ring overhang (30) on the rotor blade side, which overhang is directed upstream and comprises an end portion (32) that is directed radially inward, the axial ring (26) being arranged radially inward of the platform ring overhang (30) and forming an axial overlap therewith, the front platform ring overhang (30) seamlessly transitioning into the platform ring (12) and thus following, at least in portions, an ideal flow path of the hot gas over the platform ring (12) on the rotor blade row side, characterized in that the guide vanes (35) of the guide vane row (1) are formed without a rear platform ring overhang, and each comprise an axial projection that is positioned radially inward of a rear platform edge and has an end region that is directed radially outward, such that, when mounted on the guide vane row (1), the adjacent axial projections form the axial ring (26) and the end regions form the circumferential end portion (28).
2. Turbomachine according to claim 1, wherein the end portions (28, 32) are spaced apart in the radial direction.
3. Turbomachine according to claim 1, wherein the end portions (28, 32) end at the same radial height.
4. Turbomachine according to any of the preceding claims, wherein the end portion (28) of the axial ring (26) has a larger radial extension than the end portion (32) of the front platform ring overhang (30).
5. Turbomachine according to any of claims 1 to 4, wherein the end portions (28, 32) extend in parallel with one another and so as to be offset from one another in the radial direction.
6. Turbomachine according to any of claims 1 to 4, wherein the end portions (28, 32) are inclined toward one another.

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