A motor rotor includes an inner sleeve, a cylindrical permanent magnet disposed around the inner sleeve, and a resin portion formed by charging a gap between the inner sleeve and the permanent magnet with a resin. The inner sleeve includes a small diameter portion. The small diameter portion is located to face an inner peripheral surface of an end portion of the permanent magnet in an axial direction. The small diameter portion is formed to have a diameter smaller than a diameter of a portion facing an inner peripheral surface of a center portion of the permanent magnet in the axial direction.

A motor rotor includes an inner sleeve, a cylindrical permanent magnet disposed around the inner sleeve, and a resin portion formed by charging a gap between the inner sleeve and the permanent magnet with a resin. The inner sleeve includes a small diameter portion. The small diameter portion is located to face an inner peripheral surface of an end portion of the permanent magnet in an axial direction. The small diameter portion is formed to have a diameter smaller than a diameter of a portion facing an inner peripheral surface of a center portion of the permanent magnet in the axial direction.

An improved power generation system and methods of its operation are provided, wherein the system combines: (i) a wave reformer, (ii) an optional second wave rotor, and (iii) a gas turbine in a flexible range of novel designs. Such a hybrid power generation system can thermally crack or decompose hydrocarbon fuel to produce a high-pressure fuel product, including mainly hydrogen and lead to a higher thermal efficiency than existing engines with low to no direct emission footprint.

An improved power generation system for aircraft and methods of its operation are provided, wherein the system combines a wave reformer providing a contiguous fuel supply to a jet engine, and use of ammonia as the fuel source from which hydrogen and/or a duel supply of ammonia and hydrogen will be supplied to aircraft jet engines leading to a higher thermal efficiency than existing engines with low to no direct emission footprint.

A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

An eductor for an auxiliary power unit with a gas turbine engine and a load compressor incorporates an eductor housing having an cooling airflow inlet and a turbine exhaust inlet opening fluidly connected through a primary plenum to an outlet opening. A splitter plate is positioned in the primary plenum fixing a flow stagnation point with respect to a wall of the primary plenum.

An eductor for an auxiliary power unit with a gas turbine engine and a load compressor incorporates an eductor housing having an cooling airflow inlet and a turbine exhaust inlet opening fluidly connected through a primary plenum to an outlet opening. A splitter plate is positioned in the primary plenum fixing a flow stagnation point with respect to a wall of the primary plenum.