Mounting a high pressure turbine nozzle in leaktight manner to one end of a combustion chamber in a gas turbine

• US 7,237,387 B2
• Filed: 06/16/2005
• Issued: 07/03/2007
• Est. Priority Date: 06/17/2004
• Status: Active Grant

• Current Assignee: Safran Aircraft Engine Company
• Original Assignee: Snecma (Safran SA)
• Inventors: Conete, Eric, De Sousa, Mario, Aumont, Caroline, Hernandez, Didier
• Primary Examiner(s): CASAREGOLA, LOUIS J
• Application Number:
  US11/153,354
• Publication Number:
  US 20060032236A1
• Time in Patent Office: 747 Days
• Field of Search: 607/52, 607/53, 607/96, 607/98, 608/00, 415/191
• US Class Current: 60/796
• CPC Class Codes: F01D 9/023 Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
• CPC Class Codes: F01D 9/042 fixing blades to stators fixing stator-rings in the casing or to each other F01D25/246
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

The downstream end portions of the inner and outer walls of the combustion chamber of the gas turbine are pressed against the inner and outer platforms respectively of the vanes of the high pressure turbine nozzle and they extend thereover substantially to the downstream end of the platforms, and the means for mechanically connecting the turbine nozzle to the downstream end portions of the combustion chamber walls are formed by connection screws implanted in the airfoils of the vanes of the turbine nozzle, passing through holes formed in the end portions of the combustion chamber walls. Thus, the downstream end portions of the combustion chamber walls contribute to preventing leakage from the gas stream at the interface between the chamber and the nozzle, and also along the nozzle.

Claims

1. A gas turbine comprising:

an annular combustion chamber having an inner wall and an outer wall;

a high pressure turbine nozzle having a plurality of stationary vanes distributed around an axis that coincides with the axis of the combustion chamber, the vanes having airfoils secured to inner and outer platforms in the form of juxtaposed ring sectors with inside faces defining between them a flow path through the turbine nozzle for the gas stream coming from the combustion chamber; and

connection means for mechanically connecting the turbine nozzle to downstream end portions of the inner and outer walls of the combustion chamber;

wherein;

the downstream end portions of the inner and outer walls of the combustion chamber are pressed against the inner and outer platforms respectively of the vanes of the turbine nozzle substantially all the way to the downstream ends of the platforms; and

the connection means are formed by substantially radial connection screws implanted in vane airfoils of the turbine nozzle, passing through holes formed in the end portions of the inner and outer walls of the combustion chamber.

2. A gas turbine according to claim 1, in which the assembly constituted by the combustion chamber and the turbine nozzle is held between inner and outer metal shrouds by means of respective inner and outer ferrules, wherein the inner and outer ferrules have end portions connected to the assembly formed by the combustion chamber and the turbine nozzle by means of said connection screws.

3. A gas turbine according to claim 2, wherein the downstream end portion of the inner wall of the combustion chamber is clamped between the inner platforms of the vanes and an inner ring pressing against the outside face of the end portion of the inner wall of the combustion chamber.

4. A gas turbine according to claim 3, wherein said end portion of the inner ferrule comes to press against an outside face of the inner ring.

5. A gas turbine according to claim 2, wherein the downstream end portion of the outer wall of the combustion chamber is clamped between the outer platforms of the vanes and an outer ring pressing against the outside face of the downstream end portion of the outer wall of the combustion chamber.

6. A gas turbine according to claim 5, wherein said end portion of the outer ferrule comes to press against an outside face of the outer ring.

7. A gas turbine according to claim 3, wherein the inner ring is secured to an annular radial flange that co-operates with an annular radial flange of the inner metal shroud to hold between them an annular sealing gasket closing the space extending between the inner wall of the combustion chamber and the inner metal shroud in the vicinity of the downstream end of the turbine nozzle.

8. A gas turbine according to claim 5, wherein the outer ring is secured to an annular radial flange which co-operates with an annular radial flange integral with the outer metal shroud to hold between them an annular sealing gasket closing the space extending between the outer wall of the combustion chamber and the outer metal shroud in the vicinity of the downstream end of the turbine nozzle.

9. A gas turbine according to claim 2, wherein locking means are also provided to prevent the turbine nozzle from turning so as to avoid turning forces induced by the interaction between the stationary vanes of the turbine nozzle and the gas stream coming from the combustion chamber being transmitted to the inner and outer ferrules.

10. A gas turbine according to claim 9, wherein the locking means comprise elements providing mutual locking in turning acting on at least one of the pairs of radial flanges between which an annular sealing gasket is held.