Patent ID: 11879386
Assignee: ROLLS-ROYCE NORTH AMERICAN TECHNOLOGIES, INC.
Field: Engines, pumps, turbines (Mechanical engineering)
Classification: CPC F | IPC F

Claim 12:
13. A method comprising:
providing a first axial compressor including a first predetermined number of axial compressor stages based on an engine performance capability for a first gas turbine engine, the first axial compressor having a first exit corrected flow output,
providing a set of standard turbine stages whereby each standard turbine stage includes a single turbine rotor having a plurality of turbine blades and a subset of standard turbine vane rings associated with the single turbine rotor and including a plurality of turbine vanes, wherein the set of standard turbine stages ranges from a first turbine stage to an Mth turbine stage where M is a natural number greater than 1, and wherein each of the subset of standard turbine vane rings ranges from a first turbine vane ring to a Zth turbine vane ring where Z is a natural number greater than 1 and whereby a throat area of the plurality of turbine vanes of each turbine vane ring included in the subset of standard turbine vane rings gradually increases from the first turbine vane ring to the Zth turbine vane ring,
selecting an initial turbine stage from the set of standard turbine stages for the first gas turbine engine based on the first exit corrected flow output of the first axial compressor,
selecting an initial turbine vane ring from the subset of standard turbine vane rings included in the initial turbine stage for the first gas turbine engine based on the first exit corrected flow output of the first axial compressor of the first gas turbine engine,
locating the initial turbine vane ring for the first gas turbine engine downstream of the first axial compressor of the first gas turbine engine and locating an initial turbine rotor associated with the initial turbine stage for the first gas turbine engine directly downstream of the initial turbine vane ring of the first gas turbine engine to provide a turbine for the first gas turbine engine,
providing a second axial compressor including a second predetermined number of axial compressor stages based on an engine performance capability for a second gas turbine engine, the second axial compressor having a second exit corrected flow output,
selecting an initial turbine stage from the set of standard turbine stages for the second gas turbine engine based on the second exit corrected flow output of the second axial compressor,
selecting an initial turbine vane ring from the subset of standard turbine vane rings included in the initial turbine stage for the second gas turbine engine based on the second exit corrected flow output of the second axial compressor, and
locating the initial turbine vane ring for the second gas turbine engine downstream of the second axial compressor of the second gas turbine engine and locating an initial turbine rotor associated with the initial turbine stage for the second gas turbine engine directly downstream of the initial turbine vane ring of the second gas turbine engine to provide a turbine for the second gas turbine engine.