A gas turbine engine includes a rotatable first shaft, a first disk connected to the first shaft, a second disk connected to the first shaft, a combustor radially outward from the first disk and the second disk, and a heat exchanger connected to the combustor aft of the second disk. The first disk includes a row of low pressure compressor blades and a row of high pressure turbine blades connected to a radially outer end of the row of low pressure compressor blades. The second disk includes a row of high pressure compressor blades and a row of low pressure turbine blades connected to a radially outer end of the row of high pressure compressor blades.

The present disclosure provides an axial flow compressor comprising a high pressure compressor section having a core flow path, an aft stage and a forward stage; a diffuser in fluid communication with the core flow path and coupled to the aft stage; a plenum coupled to the diffuser; a pre-swirl nozzle coupled to the plenum, an exit of the pre swirl nozzle being directed at an aft stage rotor disk and configured to impart a swirl to a cooling fluid. The axial flow compressor further may further comprise an aft stage rotor cavity defined by a portion of the aft stage rotor disk and having an aft stage axial overlap seal, wherein a portion of the cooling fluid returns to the core flow path though the aft stage labyrinth seal. The present disclosure provides a method of high pressure compressor aft stage cooling.

The present disclosure provides an axial flow compressor comprising a high pressure compressor section having a core flow path, an aft stage and a forward stage; a diffuser in fluid communication with the core flow path and coupled to the aft stage; a plenum coupled to the diffuser; a pre-swirl nozzle coupled to the plenum, an exit of the pre swirl nozzle being directed at an aft stage rotor disk and configured to impart a swirl to a cooling fluid. The axial flow compressor further may further comprise an aft stage rotor cavity defined by a portion of the aft stage rotor disk and having an aft stage axial overlap seal, wherein a portion of the cooling fluid returns to the core flow path though the aft stage labyrinth seal. The present disclosure provides a method of high pressure compressor aft stage cooling.

A rotating detonation combustion (RDC) system is provided. The RDC includes a first outer wall and a second outer wall each extended around a centerline axis, and a detonation chamber formed radially inward of the second outer wall. A fuel passage extended between the first outer wall and the second outer wall, the fuel passage including a first inlet opening proximate to the aft end through which a flow of fuel is received into the fuel passage. The flow of fuel is provided through the fuel passage from the aft end to the forward end of the RDC system and to the detonation chamber.

A combustor basket assembly for a gas turbine engine that includes a combustor basket having a basket liner including an input end and an output end. A double-wall exit cone is mounted to the output end of the basket liner, where the exit cone includes an inner wall and an outer wall defining an exit cone channel therebetween. A splash plate is mounted to the outer wall to define a splash plate channel between the splash plate and the basket liner. A series of pairs of cooling feed holes are provided through the basket liner, where one of the feed holes in each pair provides cooling air to the cone channel and the other feed hole provides cooling air to the splash plate channel. The outer surface of the outer wall and the inner surface of the inner wall are coated with a thermal barrier coating.

A vehicle propulsion system comprises at least two motors. Combustion occurs upstream of a first motor, and a second motor is downstream of said first motor. The first motor is a turbine that drives a primary propulsion element to effect propulsion and a compressor to effect compression. The second motor is an expansion device whose incoming gases arrive from said first motor. The first motor and the second motor intercommunicate energy via electrical, electromagnetic, and/or mechanical means. Pressurized gases that result from said compression, combustion, or both are rendered or wastegated for auxiliary usage such as aerial thrust, vertical takeoff and/or vertical landing, near-vertical takeoff and/or near-vertical landing, pneumatic storage for hybrid drive, pneumatic lift and/or drive for towing and/or raising another vehicle, aerial vehicle steering, aerial vehicle pitch stabilization or manipulation, aerial vehicle roll stabilization or manipulation, and/or aerial vehicle yaw stabilization or manipulation.

A vehicle propulsion system comprises at least two motors. Combustion occurs upstream of a first motor, and a second motor is downstream of said first motor. The first motor is a turbine that drives a primary propulsion element to effect propulsion and a compressor to effect compression. The second motor is an expansion device whose incoming gases arrive from said first motor. The first motor and the second motor intercommunicate energy via electrical, electromagnetic, and/or mechanical means. Pressurized gases that result from said compression, combustion, or both are rendered or wastegated for auxiliary usage such as aerial thrust, vertical takeoff and/or vertical landing, near-vertical takeoff and/or near-vertical landing, pneumatic storage for hybrid drive, pneumatic lift and/or drive for towing and/or raising another vehicle, aerial vehicle steering, aerial vehicle pitch stabilization or manipulation, aerial vehicle roll stabilization or manipulation, and/or aerial vehicle yaw stabilization or manipulation.

Systems and methods for adjusting airflow distortion in a gas turbine engine using a secondary airflow passage assembly are disclosed. A gas turbine engine can include a compressor section, a combustion section, and a turbine section in series flow and defining at least in part an engine airflow path. A casing can enclose the gas turbine engine and be at least partially exposed to a bypass airflow. The gas turbine engine can further include a secondary airflow passage assembly comprising a door and a duct, the duct defining an inlet located on the casing, the duct defining an outlet in airflow communication with the engine airflow path, the duct comprising an airflow passage extending between the inlet and outlet. The door can be moveable between an open and closed position to allow a portion of the bypass airflow to flow through the airflow passage to adjust airflow distortion.

Systems and methods for adjusting airflow distortion in a gas turbine engine using a secondary airflow passage assembly are disclosed. A gas turbine engine can include a compressor section, a combustion section, and a turbine section in series flow and defining at least in part an engine airflow path. A casing can enclose the gas turbine engine and be at least partially exposed to a bypass airflow. The gas turbine engine can further include a secondary airflow passage assembly comprising a door and a duct, the duct defining an inlet located on the casing, the duct defining an outlet in airflow communication with the engine airflow path, the duct comprising an airflow passage extending between the inlet and outlet. The door can be moveable between an open and closed position to allow a portion of the bypass airflow to flow through the airflow passage to adjust airflow distortion.

A gas turbine having a combustion chamber with exhaust section through which combustion gas is exhaustable, plenum chamber and compressor are provided. The plenum chamber is coupled to the compressor wherein a first quantity of compressed fluid is injectable therein at a radially inner wall of the plenum chamber. A guide vane section with at least one airfoil is coupled to the exhaust section so combustion gas is flowable against the airfoil. The exhaust section and guide vane section are housed inside the plenum chamber. The airfoil has a first flow chamber where a second quantity of compressed fluid is flowable through the guide vane section from the compressor in the direction from the inner wall to a outer wall of the plenum chamber before being discharged. The second quantity of compressed fluid streamable through the guide vane section is larger than the first quantity of the compressed fluid.