Modular Transvane Assembly

• US 20100077719A1
• Filed: 04/08/2009
• Published: 04/01/2010
• Est. Priority Date: 09/29/2008
• Status: Active Grant

• Current Assignee: Siemens Energy Incorporated (Siemens AG)
• Original Assignee: Siemens Energy Incorporated (Siemens AG)
• Inventors: Charron, Richard C., Wilson, Jody W., Nordlund, Raymond S.
• Granted Patent: US 8,230,688 B2
• Time in Patent Office: Days
• US Class Current: 60/39.370
• 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: F05D 2240/40 Use of a multiplicity of similar components
• CPC Class Codes: F05D 2250/121 square
• CPC Class Codes: F05D 2250/141 circular
• CPC Class Codes: F05D 2250/322 tangential
• CPC Class Codes: F05D 2250/323 convergent
• CPC Class Codes: F05D 2250/324 divergent
• CPC Class Codes: F23R 3/425 Combustion chambers comprising a tangential or helicoidal arrangement of the flame tubes
• CPC Class Codes: F23R 3/46 Combustion chambers comprising an annular arrangement of several essentially tubular flame tubes within a common annular casing or within individual casings

Abstract

An arrangement for delivering gasses from can combustors of a can annular gas turbine combustion engine to a turbine first stage section including a first row of turbine blades, the arrangement including a flow-directing structure for each combustor, wherein each flow-directing structure includes a straight path and an annular chamber end, wherein the annular chamber ends together define an annular chamber for delivering the gas flow to the turbine first stage section, wherein gasses flow from respective combustors, through respective straight paths, and into the annular chamber as respective straight gas flows, and wherein the annular chamber is configured to unite the respective straight gas flows along respective shear planes to form a singular annular gas flow, and wherein the annular chamber is configured to impart circumferential motion to the singular annular gas flow before the singular annular gas flow exits the annular chamber to the first row of blades.

Claims

1. An arrangement for delivering gasses from a plurality of combustors of a can annular gas turbine combustion engine to a turbine first stage section comprising a first row of turbine blades, the arrangement comprising a flow-directing structure for each combustor, wherein each flow-directing structure comprises a straight path portion for receiving a gas flow from a respective combustor, and an annular chamber end, wherein the respective annular chamber ends together define an annular chamber that is oriented concentric to a gas turbine engine longitudinal axis, for delivering the gas flow to the turbine first stage section;

wherein gasses flow from respective combustors, through respective straight paths, and into the annular chamber as respective straight gas flows, in respective substantially uniform flow directions;

wherein the annular chamber is configured to unite the respective straight gas flows to form a singular annular gas flow and wherein the annular chamber is configured to impart circumferential motion to the singular annular gas flow before the singular annular gas flow exits the annular chamber.

2. The arrangement of claim 1, wherein each flow-directing structure comprises a downstream cross sectional area that is less than an entry cross sectional area, to define a throat effective to insulate the respective combustor from pressure pulsations originating downstream of the throat.

3. The arrangement of claim 1, wherein an axial length of the singular annular gas flow within the annular chamber is greater than or equal to 0.1 inches, and less than or equal to a greatest width of an active airfoil portion of a blade of the first row of turbine blades.

4. The arrangement of claim 1, wherein an axial length of the singular annular gas flow within the annular chamber is greater than or equal to 0.25 inches, and less than or equal to 1.0 inch.

5. The arrangement of claim 1, wherein an axial length of the singular annular gas flow within the annular chamber is greater than or equal to 0.5 inches, and less than or equal to 1.0 inch.

6. The arrangement of claim 1, wherein a longitudinal axis of a respective straight path portion intersects a plane of the first row of turbine blades at between 5 and 50 degrees.

7. The arrangement of claim 1, wherein a longitudinal axis of a respective straight path portion intersects a plane of the first row of turbine blades at 17 degrees.

8. A can annular gas turbine combustor comprising the arrangement of claim 1.

9. In a can annular gas turbine engine, an improvement comprising orienting each combustor can and a straight path portion of a transition module to share a longitudinal axis with a gas flow of a gas flowing from the respective combustor can, wherein the gas flow travels through respective straight path portions, into a common annular chamber defined by adjoined downstream annular chamber ends of the transition module, wherein the common annular chamber is configured to combine individual gas flows into a singular annular gas flow, and the common annular chamber is further configured to impart circumferential rotation to the singular annular gas flow upstream of a first stage section of the gas turbine engine.

10. The improvement of claim 9, the transition module further comprising a downstream cross sectional area that is less than an upstream cross sectional area, to define a throat effective to insulate the respective combustor cans from pressure pulsations originating downstream of the throat.

11. The improvement of claim 9, wherein the common annular chamber is shaped to define an area of shear between adjacent gas flows, and wherein the areas of shear constrain the respective adjacent gas flows as the gas flows in the annular chamber.

12. The improvement of claim 10, wherein the transition module comprises interchangeable modules each comprising a straight path portion and an annular chamber end.

13. The improvement of claim 12, wherein a straight path component forms the straight path portion, and an annular chamber component forms the annular chamber end, and the straight path component and the annular chamber component are adapted to sealingly engage each other.

14. The improvement of claim 9, wherein an axial length of the singular annular gas flow within the annular chamber is greater than or equal to 0.1 inches, and less than or equal to a height of an active airfoil portion of a blade of the first stage section of the gas turbine engine.

15. The improvement of claim 9, wherein an axial length of the singular annular gas flow within the annular chamber is greater than or equal to 0.25 inches, and less than or equal to 1.0 inch.

16. The improvement of claim 9, wherein an axial length of the singular annular gas flow within the annular chamber is greater than or equal to 0.5 inches, and less than or equal to 1.0 inch.

17. A can annular gas turbine combustor comprising the improvement of claim 9.