System and method for improved engine cooling in conjunction with an improved gas bearing face seal assembly
AssigneeGeneral Electric Company
InventorsAdam N. Pope, Ming-Fong Hwang
Life of this patent
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Abstract
A method and system for cooling a gas turbine engine in conjunction with an improved gas bearing face seal assembly. The gas turbine engine has a multi-stage compressor which discharges compressor discharge air for expansion by a combustor. A cavity which is aerodynamically linked to a turbine stage on its downstream end is formed by a metal structure the upstream end of which is connected to a seal which substantially inhibits the flow of compressor discharge air into the cavity. A conduit links an intermediate stage of the multi-stage compressor to the cavity for purposes of cooling the cavity and the turbine stage components located downstream of the cavity.
Claims5
A method for improving efficiency of a gas turbine engine by reducing cooling air requirements, the gas turbine engine having a combustor and a multi-stage air compressor having at least an intermediate pressure stage and a high pressure compressor discharge stage, discharge air being directed into the combustor, a cavity defined at least adjacent one surface of the combustor for reducing heat transfer to adjacent structures by air flow through the cavity, a conduit coupled from the intermediate pressure stage to the cavity, a rotor structure supported generally centrally of said combustor and spaced therefrom, wherein said rotor structure forms a portion of said cavity and includes a rotating member having a primary rotating sealing surface, the gas turbine engine further including a gas bearing face seal assembly connected between said rotor structure and said combustor structure, said gas bearing face seal assembly including a face seal ring member having a primary flow restrictor dam and a gas bearing surface, said seal assembly further including means for movably supporting said face seal ring member on said combustor structure, said method comprising the steps of:
- a) extracting a preselected flow of air from the intermediate stage of the multi-stage compressor via the conduit;
- b) subjecting the structure in the cavity to the flow of air from the intermediate compressor stage;
- and
- c) substantially inhibiting the flow of air from the compressor discharge stage into the cavity, wherein the step of substantially inhibiting comprises the steps of
- i) forming a primary gas bearing face clearance between said primary rotating sealing surface and each of said primary flow restrictor dam and said gas bearing surface, and
- ii) utilizing gas bearing forces to maintain said clearance at a predetermined size.
A gas turbine engine comprising:
- a) a multi-stage air compressor having a compressor discharge stage and at least one lower pressure stage;
- b) an annular combustor structure having air inlet means coupled for receiving air from said compressor discharge stage;
- c) a rotor structure supported generally centrally of said combustor and spaced therefrom whereby an annular cavity is defined between said rotor structure and said combustor structure;
- d) sealing means connected between said rotor structure and said combustor structure adjacent said air inlet means for inhibiting compressor discharge air from entering said cavity;
- e) means for extracting a preselected flow of air from said lower pressure stage of said compressor and for directing said extracted air onto said rotor structure;
- f) wherein said sealing means comprises a gas bearing face seal assembly, said gas bearing face seal assembly comprising
- i) a face seal ring member including
- A) a primary flow restrictor dam, said dam being a radially and circumferentially extending surface, and
- B) a gas bearing face surface, and
- ii) means for supporting said face seal ring member on said combustor structure for translation of said primary flow restrictor dam and said gas bearing face surface relative to a primary rotating sealing surface on a rotating member which comprises a portion of said rotating structure to establish an axially and circumferentially extending primary gas bearing face clearance between said primary rotating sealing surface and each of said primary flow restrictor dam and said gas bearing face surface.
Dependencies are inferred from the claim text.