Gas turbine engine shaft bearing configuration
AssigneeUnited Technologies Corporation
InventorsBrian D. Merry, Gabriel L. Suciu
Life of this patent
Estimating the term…
Abstract
A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supporting the shaft relative to the inlet case. The second bearing is arranged radially outward from the shaft.
Claims26
A gas turbine engine comprising:
- a core housing including an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path;
- a shaft supporting a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path,
- a geared architecture coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture, wherein the geared architecture includes a sun gear supported on the shaft;
- a first bearing supporting the shaft relative to the intermediate case;
- and
- a second bearing supporting the shaft relative to the inlet case, and the second bearing is arranged radially outward from the shaft.
A gas turbine engine comprising:
- a core housing providing a core flow path;
- a fan;
- a shaft supporting a compressor section arranged within the core flow path, wherein the compressor section is fluidly connected to the fan, the compressor section comprising a first pressure compressor and a second pressure compressor upstream from the first pressure compressor, the second pressure compressor including multiple compressor stages;
- and
- first and second bearings supporting the shaft relative to the core housing and are arranged radially inward of and axially overlapping with at least some of the multiple compressor stages;
- and
- wherein the gas turbine engine is a high bypass geared aircraft engine having a bypass ratio of greater than about six (6).
A gas turbine engine comprising:
- a core housing providing a core flow path;
- a fan;
- a shaft supporting a compressor section arranged within the core flow path, wherein the compressor section is fluidly connected to the fan, the compressor section comprising a first pressure compressor and a second pressure compressor upstream from the first pressure compressor, the second pressure compressor including multiple compressor stages;
- and
- first and second bearings supporting the shaft relative to the core housing and being arranged radially inward of and axially overlapping with at least some of the multiple compressor stages;
- a combustor fluidly connected to the compressor section;
- a turbine section fluidly connected to the combustor, the turbine section comprising:
- a high pressure turbine;
- a low pressure turbine;
- and
- wherein the gas turbine engine includes at least one of a low Fan Pressure Ratio of less than about 1.45 and a low pressure turbine pressure ratio that is greater than about 5.
Dependencies are inferred from the claim text.