Turbine nozzle guide vane assembly in a turbomachine

Abstract

A sectorized nozzle for a turbine engine turbine including an inner sectorized annular platform and an outer sectorized annular platform connected together by radial airfoils, at least one of the platforms including a plurality of orifices for passing air in a neighborhood of its upstream end, the orifices being distributed over the circumference of the platform and opening out at their ends remote from the airfoils into a circumferential annular cavity of the sector of the platform, which cavity is closed by a metal sheet fastened to the platform sector and pierced by orifices for feeding cooling air.

Claims

1. A sectorized nozzle for a turbine engine turbine comprising: an inner sectorized annular platform and an outer sectorized annular platform connected together by substantially radial airfoils, at least one of the platforms including a plurality of through orifices for passing air upstream of a leading edge of the airfoils, the through orifices being distributed over the circumference of the platform, wherein the through orifices for passing air in each sector of a platform open out upstream from the airfoils into a circumferential annular cavity of the sector of the platform, the cavity being closed and further comprising cooling air feed orifices that are offset tangentially in a staggered configuration relative to the through orifices in the platform, wherein the cavity is formed by a groove arranged in a thickness of the at least one of the platforms of each nozzle sector and closed by a metal sheet including the cooling air feed orifices, wherein the cavity is disposed downstream of a radial wall of the at least one of the platforms, and a sealing device is fastened to the radial wall, and wherein the metal sheet being inserted radially between the radial wall and a radial tab of the sealing device.

2. A nozzle according to claim 1, wherein axes of the through orifices and cooling air feed orifices in the platform and in the annular cavity of each nozzle sector lie substantially in a common plane perpendicular to a longitudinal axis of the nozzle.

3. A nozzle according to claim 1, wherein the through orifices in the platform and in the annular cavity of each nozzle sector are regularly spaced apart from one another.

4. A nozzle according to claim 1, wherein a number of orifices in the platform of each nozzle sector is equal to a number of orifices in the annular cavity of the nozzle sector.

5. A nozzle according to claim 1, wherein the through orifices in the platform of each nozzle sector are of a diameter substantially identical to a diameter of the cooling air feed orifices in the cavity of the sector.

6. A nozzle according to claim 1, wherein the metal sheet is fastened by brazing or welding on two cylindrical bearing surfaces of the platform of each nozzle sector, which surfaces are situated respectively upstream and downstream relative to the cavity.

7. A turbine engine, or an airplane turboprop or turbojet, comprising a turbine nozzle according to claim 1, arranged at an outlet from an annular combustion chamber.

Description

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

(1) Other advantages and characteristics of the invention appear on reading the following description made by way of non-limiting example and with reference to the accompanying drawings, in which:

(2) FIG. 1 is a diagrammatic axial section view of a combustion chamber and of the first nozzle of a prior art high-pressure turbine;

(3) FIG. 1A is a diagrammatic view on a larger scale of the zone in a chain-dotted box in FIG. 1;

(4) FIG. 2 is a diagrammatic view in perspective from above of two sectorized inner platforms of the FIG. 1 nozzle;

(5) FIG. 3 is a diagrammatic axial section view of the upstream end of an inner platform of a nozzle of a high-pressure turbine of the invention;

(6) FIG. 4 is a diagrammatic perspective view from the inside of an inner platform of a nozzle of the invention; and

(7) FIG. 5 is a diagrammatic perspective view and a section view on a cross-section plane showing the upstream end of the inner platform of FIG. 4.

DETAILED DESCRIPTION OF THE INVENTION

(8) Reference is made initially to FIG. 1, which shows a prior art annular combustion chamber 10 in which an upstream high pressure compressor (not shown) feeds air under pressure into an annular space 12 defined between two coaxial casings, a radially outer casing 14 and a radially inner casing 18, which casings contain the combustion chamber 10.

(9) The combustion chamber 10 has two coaxial walls forming inner and outer bodies of revolution 20 and 22, and an upstream annular chamber end wall 24 having an annular fairing 26 fastened thereto and extending upstream.

(10) The annular fairing 26 and the annular chamber end wall 24 have a plurality of openings for passing air and for mounting injectors 28 that are fastened externally onto the outer casing 14.

(11) The outlet from the combustion chamber 10 leads to a high-pressure turbine 30 having at least one nozzle of stationary vanes and at least one rotor wheel of blades. In FIG. 1, only the inlet nozzle 32 of the high-pressure turbine is shown. This nozzle 32 is fastened to an outer casing 34 of the turbine, which is fastened in turn at its upstream end to the downstream end of the outer casing 14 of the combustion chamber 10. The nozzle 32 has inner and outer annular platforms 36 and 38 extending one inside the other and connected together by substantially radial airfoils 40. These inner and outer platforms 36 and 38 thus define between them an annular passage for the flow of burnt gas coming from the combustion chamber 10.

(12) The nozzle 32 is sectorized and is made up of sectors arranged one beside another around a circumference centered on the axis of revolution 142 of the combustion chamber 10.

(13) The downstream ends of the inner and outer cylindrical walls 20 and 22 of the chamber are in alignment and they are connected by sealing means to the upstream ends of the sectors of the inner and outer platforms 36 and 38 of the nozzle.

(14) FIG. 1A is a view on a larger scale of the inner sealing means, which means are similar to the outer sealing means. These sealing means comprise blades 42 that are arranged circumferentially beside one another around the axis of revolution 142 of the combustion chamber 10, and having joint covers (not shown) mounted thereon. Each blade 42 is formed by a plane plate in a circumferential orientation that extends substantially upstream and inwards in the mounted position. It is fastened by rivets in its middle portion to a nozzle sector and it bears via its inner peripheral edge 44 against a radial face of a cylindrical rim 46 at the downstream end of the inner wall 20.

(15) At its upstream end, the inner platform 36 of the nozzle sector has a radial wall 48 and a plurality of radial tabs 50 that are spaced apart axially downstream relative to the radial wall and that are regularly distributed around the circumference of the inner platform 36. The radial tabs 50 have respective orifices 52 for passing rivets 54 with upstream ends that pass through the blades 42 and that are housed in notches in the radially inner periphery of the radial wall 48. In this way, the blades 42 are fastened to the radial wall 48.

(16) The blades 42 are mounted between the radial wall and the radial tabs and they bear against the downstream radial face of the radial wall, being urged upstream by a spring 56 mounted between the blades 42 and the upstream faces of the radial tabs 50 of the inner platform 36.

(17) Upstream from the leading edges of the airfoils, each sector of an inner platform 36 has a circumferential row of substantially radial orifices 58 passing through the platform 36 (FIG. 2). These orifices 58 open out into the annular passage defined by the inner and outer platforms 36 and 38. On the side of the inner platform 36 remote from the airfoils 40, these orifices 58 open out axially between the radial wall 48 and the radial tabs 50 of the inner platform 36.

(18) In similar manner, each sector of the outer platform 38 has orifices opening out in the outer face of the platform 36 between a radial wall and radial tabs, and in the inner face of the outer platform 38 upstream from the leading edges of the airfoils 40.

(19) In operation, air (arrows A) flowing around the combustion chamber 10 feeds the orifices 58 in the inner and outer platforms 36 and 38 respectively, thereby enabling the platforms to be cooled. Nevertheless, the air streams passing through the inner and outer platforms 36 and 38 are flowing at a relatively high speed, which leads to a cooling air penetrating a long way into the annular passage, thereby failing to achieve optimum cooling of the platforms of the nozzle.

(20) The invention provides a simple solution to the problems of the prior art by forming a circumferential annular cavity 60 over at least the sectors of the inner and/or outer platform, the orifices 64 of each platform opening out at one end beside the airfoils upstream from the leading edges 63 of the airfoils, and at the opposite end in a closed cavity having cooling air feed orifices.

(21) FIG. 3 shows the upstream inner end of a nozzle sector of the invention, connected to the downstream end of the outer wall 20 via a seal of the type described above with reference to FIGS. 1 and 1A.

(22) In FIG. 3, the cavity is formed axially between an inner radial wall 70 of the upstream end of the inner platform 62 and inner radial fastener tabs 72 of the sealing means.

(23) This cavity 60 is formed by machining an annular groove in the thickness of each sector of the inner platform 62, and it is closed by a metal sheet 66 that is inserted radially between the radial wall 70 and the radial tabs 72 and that is fastened by brazing or welding to cylindrical bearing surfaces 74 formed upstream and downstream on either side of the inner opening of the groove 60.

(24) The circumferential ends of the groove 60 are closed and do not open out at the circumferential ends of the sector of the platform 62 (FIG. 4). In this way, it is possible to conserve a sealed junction between two facing platform sectors by known means of the type comprising blades each having one half inserted in a slot 76 in a circumferential edge of a platform of a nozzle sector, with the other half inserted in the slot 76 of a facing circumferential edge of a platform of an adjacent nozzle sector.

(25) Each metal sheet 66 extends over the same angular distance as the nozzle sector. The orifices 68, 64 in the sheet 66 and the platform are regularly spaced apart from one another and their respective axes 80, 82 lie in a common plane perpendicular to the longitudinal axis 78 of the nozzle.

(26) As shown in FIG. 5, the axes 80 of the orifices 68 in the sheet 66 are each situated in a plane containing the longitudinal axis 78 of the nozzle and lies substantially between two consecutive orifices 64 of the platform. This staggered arrangement of the orifices 68 in the sheet 66 relative to the orifices 64 in the platform 62 enables the platform 62 to be cooled by air impact.

(27) The number of orifices in the platform 62 of each nozzle sector may be equal to the number of orifices in the sheet 66 of the nozzle sector, and may for example lie in the range 20 to 25.

(28) The orifices 68 in the sheets 66 and the orifices 64 in the platforms 62 of the nozzle sectors may be identical in diameter, e.g. having a diameter of about 0.5 millimeters (mm) to 0.6 mm.

(29) In a particular embodiment of the invention, each sheet 66 may be about 1 mm thick.

(30) The invention is described above with reference to an inner platform 62 of a nozzle. Nevertheless, it applies equally to the outer platforms of nozzle sectors.

(31) It can also be understood that it is possible to form closed circumferential cavities solely in the inner platforms, or else solely in the outer platforms, or indeed in both the inner and the outer annular platforms of the nozzle.

(32) As mentioned above, incorporating a cavity 60 into which the orifices 64 of the platforms 62 open out and that is closed by a sheet 66 pierced by orifices 68, makes it possible to reduce the penetration speed of the fraction of the air that flows around the chamber and that is reintroduced via the orifices in the platforms, thus enabling the air that is reintroduced to flow closer to the faces of the platforms that face the airfoils.

(33) In addition, because the air that has flowed around the chamber and that has been reintroduced into the annular passage presents a speed that is slower than in the prior art, this air becomes mixed less quickly with the hot gas from the combustion chamber, thereby further reducing circumferential temperature non-uniformities and thus achieving better circumferential uniformity in the temperature of the nozzle.