Gas turbine engine having an improved transition duct support

• US 4,016,718 A
• Filed: 07/21/1975
• Issued: 04/12/1977
• Est. Priority Date: 07/21/1975
• Status: Expired due to Term

• Current Assignee: United Technologies Corporation (Rtx Corporation)
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: Lauck, Lawrence J.
• Primary Examiner(s): Croyle, Carlton R.
• Assistant Examiner(s): Casaregola, L. J.
• Application Number:
  US05/597,876
• Time in Patent Office: 631 Days
• Field of Search: 60/39.31, 60/39.32, 60/39.37, 60/39.65, 60/39.66, 415/134, 415/138
• 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: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

An axial flow gas turbine engine is provided with a plurality of combustion chambers distributed circumferentially about the engine axis and a plurality of transition ducts which carry propulsive gases between the burners and an annular turbine inlet at a bulkhead between the combustion and turbine sections of the engine. Support for the transition ducts at the turbine inlet is provided by an outer support ring fixedly secured to the engine bulkhead and an inner support ring which is held by a plurality of struts from the outer support ring. Flexible sealing means extend between the inner support ring and the bulkhead adjacent the turbine inlet to permit the inner ring to expand or contract radially with the transition ducts and at the same time to prevent compressed air from flowing directly into the turbine section. The outer ring and the flexible sealing means are apertured to permit a limited flow of relatively cool compressor air to enter the turbine along with the propulsive gases and thereby cool the support structure and the inlet guide vanes of the turbine.

Claims

1. In combination in an axial flow gas turbine engine having a combustion section including a plurality of burners disposed circumferentially about the engine axis for receiving and mixing compressed air and fuel in a combustion process to produce propulsive gases which are transmitted from the plurality of burners through a plurality of circumferentially disposed transition ducts between the respective burners and a turbine inlet at the turbine section, the improvement comprising:

an outer support ring connected to the plurality of transition ducts and positioned in the engine coaxially of the engine axis, the outer ring circumscribing the transition ducts radially outside of the ducts;

an inner support ring connected to a plurality of transition ducts and positioned in the engine coaxially of the engine axis, the inner ring extending around the engine axis radially within the circumaxially disposed transition ducts;

a plurality of struts extending from the outer support ring to the inner support ring, the struts being interdigitated between the transition ducts; and

one of the support rings being fixedly secured in the engine between the combustion and turbine sections for holding the transition ducts in the engine, and the other of the support rings being suspended in the engine by the struts from said one of the support rings, said other of the support rings thereby being free to expand and contract radially in the engine with the transition ducts.

2. In an axial flow gas turbine, the improvement of claim 1 further including

flexible sealing means circumscribing the engine axis and extending between said other of the support rings and the turbine inlet to impede the leakage of compressed air from the combustion section directly into the propulsive gases entering the turbine inlet.

3. The improvement of claim 2 wherein the sealing means is provided with leakage apertures through which a limited flow of compressed air enters the turbine inlet for cooling.

4. The improvement of claim 1 in an axial flow gas turbine having a turbine inlet defined by a series of inlet guide vanes bearing radially inner and radially outer vane platforms which platforms are serially interconnected to form inner and outer walls of an annular turbine inlet adjacent the corresponding inner and outer support rings, wherein:

sealing means is interposed between said other of the support rings free to expand or contract and the corresponding wall of the turbine inlet and includes leakage apertures admitting a limited flow of compressed air into the inlet and over the vane platforms defining the corresponding turbine wall for cooling the platforms and vanes.

5. The improvement of claim 4 wherein:

said one of the support rings fixedly secured in the engine defines a wall of a cooling chamber circumscribing the engine axis between the circumferentially disposed transition ducts and the vane platforms which form one wall of the turbine inlet, the support ring being apertured to admit compressed air to the chamber from the combustion section for cooling the vane platforms.

6. The improvement of claim 5 in an axial flow gas turbine having the transition ducts projecting into the turbine inlet with finite spacing between the vane platforms and the ducts wherein:

said one of the support rings defining a wall of the cooling chamber extends axially of the engine between the vane platforms defining the turbine inlet wall and the transition duct to allow the compressed air from the cooling chamber to flow over the surface of the vane platforms.

7. The improvement of claim 6 in an axial flow gas turbine having interior cooling ducts in the inlet guide vanes of the turbine inlet wherein the cooling chamber also communicates with the cooling ducts of the inlet guide vanes for cooling the vanes internally.

8. The improvement of claim 1 in a gas turbine engine having the transition ducts projecting axially rearward in the engine to the turbine inlet wherein:

said one of the support rings fixedly secured in the engine projects axially forward of the turbine inlet to a given axial station in the engine; and

sleeves attached to each of the respective transition ducts adjacent the turbine inlet also project axially forward of the turbine inlet and are connected to said one of the support rings at the given axial station.

9. The improvement of claim 8 wherein:

said other of the support rings free to expand and contract radially in the engine has at least one portion also located at the given axial station in the engine; and

the sleeves circumscribe the respective transition ducts and attach to said other of the support rings at the given axial station.

10. The improvement of claim 1 wherein:

shrouds extend over the respective transition ducts and also connect with the inner and outer support rings for support within the engine.

11. The improvement of claim 10 wherein:

the shrouds extend forward in the engine over the respective transition ducts and associated burners and define an annular space about the respective burners and transition ducts; and

spacing means are interposed between the shrouds and the burners and ducts for maintaining the annular space as defined.

12. The improvement of claim 11 wherein:

further means are provided for connecting the forward ends of the shrouds to the engine in the combustion section.