Cooled gas turbine transition duct

• US 7,310,938 B2
• Filed: 12/16/2004
• Issued: 12/25/2007
• Est. Priority Date: 12/16/2004
• Status: Active Grant

• Current Assignee: Siemens Energy Incorporated (Siemens AG)
• Original Assignee: Siemens Power Generation, Inc.
• Inventors: Slentz, Kenneth, Gill, David Alan, Marcum, Steven
• Primary Examiner(s): KIM, TAE JUN
• Application Number:
  US11/014,294
• Publication Number:
  US 20060130484A1
• Time in Patent Office: 1,104 Days
• Field of Search: 607/52, 60/393.7
• US Class Current: 60/39.37
• 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 2260/20 Heat transfer, e.g. cooling
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: F05D 2260/203 by transpiration cooling
• CPC Class Codes: F23R 2900/00005 Preventing fatigue failures or reducing mechanical stress in gas turbine components
• CPC Class Codes: F23R 2900/03041 Effusion cooled combustion chamber walls or domes
• CPC Class Codes: F23R 3/002 Wall structures F23R3/02 and F23R3/007 take precedence

Abstract

A transition duct (40) for a gas turbine engine (10) incorporating a combination of cooling structures that provide active cooling in selected regions of the duct while avoiding cooling of highly stressed regions of the duct. In one embodiment, a panel (74) formed as part of the transition duct includes some subsurface cooling holes (92) that extend under a central portion of a stiffening rib (90) attached to the panel and some subsurface cooling holes (94) that have a truncated length so as to avoid extending under a rib end (45). Effusion cooling holes (88) used to cool a side subpanel (48) of the panel may have a distribution that reduces to zero approaching a double bend region (48) of the panel. An upstream subpanel (76) of the panel may be actively cooled only when the panel is located on an extrados of the transition duct.

Claims

1. A panel of a transition duct for a gas turbine engine, the panel comprising:

an upstream subpanel joined to a downstream subpanel;

side subpanels joined along respective opposed sides of the upstream panel and the downstream panel, each side subpanel comprising a double bend region of the transition duct; and

cooling structures formed in each of the side subpanels in only regions remote from the respective double bend regions.

2. The panel of claim 1, further comprising a distribution of effusion cooling holes reduced from a first value to zero in a direction approaching the respective double bend regions.

3. The panel of claim 1, further comprising:

a stiffening rib comprising opposed rib ends attached to the downstream panel;

a first subsurface cooling passage formed in the downstream subpanel and extending under the stiffening rib remote from the rib ends; and

a second subsurface cooling passage formed in the downstream subpanel extending toward one of the rib ends and being truncated so as not to extend under the one of the rib ends.

4. The panel of claim 1 disposed on an extrados of the transition duct, further comprising a plurality of subsurface cooling channels formed in the upstream subpanel.

5. A transition duct for conveying hot combustion gas from a combustor to a turbine in a gas turbine engine, the transition duct comprising:

a plurality of panels joined together to form a duct comprising a generally cylindrical inlet end and a generally rectangular outlet end disposed radially inwardly of the inlet end when installed in the gas turbine engine;

a double bend region formed in a first of the panels;

a stiffening rib end region in a second of the panels proximate an end of a stiffening rib joined to an outside surface of the second of the panels;

a plurality of cooling structures formed in the panels for passing respective flows of cooling air through the panels; and

wherein the cooling structures are formed to avoid both the double bend region and the stiffening rib end region.

6. The transition duct of claim 5, wherein the cooling structures comprise:

a plurality of subsurface cooling passages formed through respective ones of the plurality of the panels, each subsurface cooling passage having an inlet opening to an outside surface of the duct and an outlet opening to an inside surface of the duct; and

a plurality of effusion cooling holes formed through a plurality of the panels in regions remote from the subsurface cooling passages.

7. A transition duct for conveying hot combustion gas from a combustor to a turbine in a gas turbine engine, the transition duct comprising:

a plurality of panels joined together to form a duct comprising a generally cylindrical inlet end and a generally rectangular outlet end disposed radially inwardly of the inlet end when installed in the gas turbine engine;

the outlet end comprising an outlet mouth formed to extend across at least approximately a 45°

arc of a turbine inlet;

a stiffening rib end region in one of the panels proximate an end of a stiffening rib joined to an outside surface of the one of the panels;

a plurality of subsurface cooling passages formed through the one of the panels, each subsurface cooling passage having an inlet opening to an outside surface of the duct and an outlet opening to an inside surface of the duct; and

wherein the cooling passages are formed to avoid the stiffening rib end region.

8. The transition duct of claim 7, further comprising:

a first portion of the subsurface cooling passages extending through the one of the panels directly under the stiffening rib remote from the stiffening rib end region; and

a second portion of the subsurface cooling passages extending through the one of the panels in a direction toward the stiffening rib end region but having an axial length truncated so as not to extend proximate the stiffening rib end region.