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US 11,655,757 B2Granted
Modular multistage compressor system for gas turbine engines
AssigneeRolls-Royce North American Technologies Inc.
InventorBenjamin M. Iwrey
Life of this patent
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Abstract
A method of assembling a gas turbine engine is disclosed herein. The method comprises providing a set of standard axial compressor stages. Each axial compressor stage included in the set of standard axial compressor stages includes a rotor having a plurality of blades configured to rotate about an axis and a stator having a plurality of stator vanes.
Claims20
A method of assembling a gas turbine engine, the method comprising:
- providing a set of standard axial compressor stages that each include a rotor having a plurality of blades configured to rotate about an axis and a stator having a plurality of stator vanes, wherein the set of standard axial compressor stages ranges from a first compressor stage to an Nth compressor stage where N is a natural number greater than 1 and whereby a radial length of the plurality of blades and a radial length of the stator vanes on each compressor stage included in the standard compressor stages gradually decreases in size from the first compressor stage to the Nth compressor stage and each compressor stage of the set of standard axial compressor stages is set in size and the size does not change,
- determining an engine performance capability for the gas turbine engine, the engine performance capability including a predetermined inlet corrected flow and a predetermined pressure ratio
- selecting an initial axial compressor stage from the set of standard axial compressor stages for the gas turbine engine based on the predetermined inlet corrected flow, the initial axial compressor stage being the furthest axially upstream compressor stage in a compressor of the compressor for the gas turbine engine, and
- adding two or more sequential axial compressor stages from the set of standard axial compressor stages downstream of the initial axial compressor stage based on the predetermined pressure ratio to provide the gas turbine engine.
1 more element
A method comprising:
- providing a set of standard axial compressor stages that each include a rotor having a plurality of blades configured to rotate about an axis, wherein the set of standard axial compressor stages ranges from a first compressor stage to an Nth compressor stage where N is a natural number greater than 1 and whereby a radial length of the plurality of blades on each compressor stage included in the standard compressor stages gradually decreases in size from the first compressor stage to the Nth compressor stage and each compressor stage of the set of standard axial compressor stages is set in size and the size does not change,
- determining an engine performance capability for a first gas turbine engine, the engine performance capability including a first predetermined inlet corrected flow and a first predetermined pressure ratio,
- selecting an initial axial compressor stage from the set of standard axial compressor stages for the first gas turbine engine based on the first predetermined inlet corrected flow, the initial axial compressor stage being the furthest axially upstream compressor stage in a compressor of the first gas turbine engine, and
- adding two or more sequential axial compressor stages from the set of standard axial compressor stages downstream of the initial axial compressor stage for the first gas turbine engine based on the first predetermined pressure ratio to provide the compressor for the first gas turbine engine.
1 more element
Dependencies are inferred from the claim text.