A gas turbine engine includes a compressor section, a combustor section, and a turbine section operably coupled to the compressor section. A primary flow path is defined through the compressor section, the combustor section, and the turbine section. An engine case surrounds the compressor section, the combustor section, and the turbine section. The gas turbine engine also includes a means for providing an active variable cooling flow through a bypass duct external to the engine case to a secondary flow cavity of the turbine section.

A gas turbine engine includes a compressor section, a combustor section, and a turbine section operably coupled to the compressor section. A primary flow path is defined through the compressor section, the combustor section, and the turbine section. An engine case surrounds the compressor section, the combustor section, and the turbine section. The gas turbine engine also includes a means for providing an active variable cooling flow through a bypass duct external to the engine case to a secondary flow cavity of the turbine section.

Fault-tolerant electric drive systems including a machine having a rotor and a stator having coils arranged in pairs. Each coil in each pair separated by 180 degrees, a first phase (ϕA) having one of the coil pairs and a phase drive circuit connected therewith, a second phase (ϕB) having a second one of the coil pairs and a second phase drive circuit connected therewith, a third phase (ϕC) having a third one of the coil pairs and a third phase drive circuit connected therewith, and a fourth phase having a fourth one of the coil pairs and a fourth phase drive circuit connected therewith. Further included is a controller connected with the first, second, third and fourth phase drive circuits to control operation thereof.

Aircraft power plants and associated methods are provided. A method for driving a load on an aircraft includes: transferring motive power from an internal combustion (IC) engine to the load; discharging a flow of first exhaust gas from the IC engine when transferring motive power from the IC engine to the load; receiving the flow of first exhaust gas from the IC engine into a combustor; mixing fuel with the first exhaust gas in the combustor and igniting the fuel to generate a flow of second exhaust gas; receiving the flow of second exhaust gas at a turbine and driving the turbine with the flow of second exhaust gas from the combustor; and transferring motive power from the turbine to the load.

A power management system for a multi engine rotorcraft having a main rotor system with a main rotor speed. The power management system includes a first engine that provides a first power input to the main rotor system. A second engine selectively provides a second power input to the main rotor system. The second engine has at least a zero power input state and a positive power input state. A power anticipation system is configured to provide the first engine with a power adjustment signal in anticipation of a power input state change of the second engine during flight. The power adjustment signal causes the first engine to adjust the first power input to maintain the main rotor speed within a predetermined rotor speed threshold range during the power input state change of the second engine.

A combined cycle power generation system for an aircraft includes fuel cell and supercritical CO.sub.2 cycles. The fuel cell cycle includes a compressor and turbine disposed on a first shaft, a fuel cell in fluid communication with the compressor and a fuel source, and a combustor in fluid communication with the fuel cell and the turbine. The combustor is configured to combust partially spent fuel from the fuel cell and produce combustion exhaust gas for delivery to the turbine. The supercritical CO.sub.2 cycle includes a compressor and turbine disposed on a second shaft, a supercritical CO.sub.2 fluid circuit in thermal communication with the combustor and configured to deliver CO.sub.2 to the turbine and compressor, and a heat exchanger in thermal communication with the supercritical CO.sub.2 fluid circuit and a source of cooling fluid. A mechanical linkage is configured to transfer power from the second shaft to the first shaft.

A combined cycle power generation system for an aircraft includes fuel cell and supercritical CO.sub.2 cycles. The fuel cell cycle includes a compressor and turbine disposed on a first shaft, a fuel cell in fluid communication with the compressor and a fuel source, and a combustor in fluid communication with the fuel cell and the turbine. The combustor is configured to combust partially spent fuel from the fuel cell and produce combustion exhaust gas for delivery to the turbine. The supercritical CO.sub.2 cycle includes a compressor and turbine disposed on a second shaft, a supercritical CO.sub.2 fluid circuit in thermal communication with the combustor and configured to deliver CO.sub.2 to the turbine and compressor, and a heat exchanger in thermal communication with the supercritical CO.sub.2 fluid circuit and a source of cooling fluid. A mechanical linkage is configured to transfer power from the second shaft to the first shaft.

A propulsion system is provided including an unducted rotating fan defining a fan axis; and a turbomachine disposed downstream from the unducted rotating fan, wherein the turbomachine defines a working gas flowpath flowing therethrough; wherein the propulsion system defines a third stream flowpath and an inlet passage having an inlet that is offset from the fan axis, wherein the inlet passage is configured to provide an inlet airflow to the working gas flowpath, and wherein the third stream flowpath bypasses at least a portion of the turbomachine.

An apparatus is provided for a gas turbine engine. This gas turbine engine apparatus includes a shaft base, a flange and a plurality of lobes. The shaft base extends axially along an axis between a shaft first end and a shaft second end. The flange is connected to the shaft base at the shaft first end. The flange projects radially out from the shaft base. The flange includes a plurality of fastener apertures, and the fastener apertures include a first fastener aperture. The lobes are arranged circumferentially about the axis. Each of the lobes is connected to and projects radially away from the shaft base. Each of the lobes is connected to and projects axially out from the flange. The lobes include a first lobe and a second lobe. The first fastener aperture is arranged circumferentially between the first lobe and the second lobe. The second lobe radially overlaps the first fastener aperture.

In one aspect, there is provided a planetary gearbox, comprising a sun gear, a plurality of planet gear assemblies, each planet gear assembly having a main gear meshed with the sun gear, a fore lateral gear and an aft lateral gear disposed on opposite sides of the main gear and rotating therewith, a diameter of the main gear being different than a diameter of the fore and aft lateral gears, a planet carrier rotatably supporting at least some of the planet gear assemblies, and at least one fore ring gear meshed with the fore lateral gears, at least one aft ring gear meshed with the aft lateral gears, wherein one of the sun gear, the planet carrier, and the ring gears is configured to be operatively connected to an input, one is configured to be operatively connected to an output, and rotation of a remaining one is limited.