Abstract: An aircraft propulsion system includes a core engine and a heat exchanger assembly that is operable in a heat exchanger configuration where the exhaust gas flow is in thermal communication with a working fluid flow and a bypass configuration where the exhaust gas flow passes through the heat exchanger assembly without thermal communication with the working fluid flow.

Abstract: An aircraft propulsion system includes a primary energy conversion device that uses a cryogenic fuel and air to generate power and thermal energy, a pylon that is configured to support the primary energy conversion device, the pylon includes an interior space, a bottoming cycle that is supported within the interior space of the pylon, the bottoming cycle includes a working fluid that is circulated within a closed circuit that includes a bottoming compressor section and a bottoming turbine section, the working fluid is compressed in the bottoming compressor section and expanded through the bottoming turbine section to generate shaft power, a primary heat exchanger that is configured to communicate thermal energy that is generated by the primary energy conversion device into the working fluid, and a fuel/working fluid heat exchanger that is configured to transfer thermal energy from the working fluid into the cryogenic fuel.

Abstract: A turbine engine is provided that includes a propulsor rotor and an engine core. The propulsor rotor is rotatable about a propulsor axis. The engine core is configured to power operation of the propulsor rotor. The engine core includes a core compressor section, a core combustor section, a core turbine section, a first rotating structure and a second rotating structure. The first rotating structure includes a first compressor rotor arranged within the core compressor section. The first rotating structure is rotatable about a first structure axis which is offset from the propulsor rotor axis. The second rotating structure includes a second compressor rotor arranged within the core compressor section. The second rotating structure is rotatable about a second structure axis which is offset from the propulsor rotor axis and the first structure axis.

Abstract: An airfoil for a gas turbine engine including an airfoil body extending between a leading edge and a trailing edge and between a pressure side and a suction side. The airfoil body includes a strut portion extending from the leading edge and a flap portion extending from the trailing edge. The flap portion is pivotable relative to the strut portion. A flexible skin surrounds both the strut portion and the flap portion on both the pressure side and the suction side.

Abstract: A propulsion system for an aircraft includes an intercooling system configured to inject water into a compressor section to reduce a temperature of a core airflow within the core flow path. The intercooling system includes a plurality of injectors for injecting water at a water injection angle relative to an airflow to generate a defined swirl of a combined airflow and water to the compressor section.

Abstract: A turbine engine is provided that includes a propulsor rotor and an engine core. The propulsor rotor is rotatable about a propulsor axis. The engine core is configured to power operation of the propulsor rotor. The engine core includes a core compressor section, a core combustor section, a core turbine section, a first rotating structure and a second rotating structure. The first rotating structure includes a first compressor rotor arranged within the core compressor section. The first rotating structure is rotatable about a first structure axis which is offset from the propulsor rotor axis. The second rotating structure includes a second compressor rotor arranged within the core compressor section. The second rotating structure is rotatable about a second structure axis which is offset from the propulsor rotor axis and the first structure axis.

Abstract: A propulsion system is provided for an aircraft. This propulsion system includes a turbine engine, and the turbine engine includes a propulsor rotor, an engine core and a heat exchanger. The propulsor rotor is configured to rotate about an axis and direct a first air stream along a primary flowpath that bypasses the engine core. The propulsor rotor is configured to direct a second air stream along a secondary flowpath with a geometry extending at least eighty degrees about the axis. The turbine engine is configured such that the first air stream enters the primary flowpath at a first pressure and the second air stream enters the secondary flowpath at a second pressure that is different than the first pressure. The engine core is configured to drive rotation of the propulsor rotor. The heat exchanger is disposed within the secondary flowpath.

Abstract: A propulsion system is provided for an aircraft. This propulsion system includes a turbine engine, and the turbine engine includes a propulsor rotor, an engine core and a heat exchanger. The propulsor rotor is configured to rotate about an axis and direct a first air stream along a primary flowpath that bypasses the engine core. The propulsor rotor is configured to direct a second air stream along a secondary flowpath with a geometry extending at least eighty degrees about the axis. The turbine engine is configured such that the first air stream enters the primary flowpath at a first pressure and the second air stream enters the secondary flowpath at a second pressure that is different than the first pressure. The engine core is configured to drive rotation of the propulsor rotor. The heat exchanger is disposed within the secondary flowpath.

Abstract: A propulsion system for an aircraft includes an intercooling system with a plurality of injectors for injecting an intercooling waterflow into a portion of a compressor passage. The plurality of injectors are spaced apart about the compressor passage and the injection of the intercooling water flow is varied between the plurality of injectors to influence flow that is communicated through the compressor passage.

Abstract: A compressor section for a gas turbine engine includes a blade including a platform, a tip and an airfoil extending between the platform and tip. The airfoil includes a root portion adjacent to the platform, a midspan portion and a tip portion. Each of the root portion, midspan portion and tip portion define a meridional velocity at stage exit with the tip portion including a first meridional velocity greater than a second meridional velocity of the midspan portion. A blade for an axial compressor of a gas turbine engine and a method of operating a compressor section of a gas turbine engine are also disclosed.

Abstract: An airfoil for a gas turbine engine including an airfoil body extending between a leading edge and a trailing edge and between a pressure side and a suction side. The airfoil body includes a strut portion extending from the leading edge and a flap portion extending from the trailing edge. The flap portion is pivotable relative to the strut portion. A flexible skin surrounds both the strut portion and the flap portion on both the pressure side and the suction side.

Abstract: A gas turbine engine includes a fan section, a compressor section, and a turbine section. The fan section has a plurality of vane assemblies spaced circumferentially about an engine axis. The vane assemblies each include an airfoil extending between a leading edge and a trailing edge, a control rod extending through the airfoil, and a mechanism driven by the control rod to change the shape of the airfoil. A vane system for a gas turbine engine is also disclosed.

Abstract: A gas turbine engine includes a fan section, a compressor section, and a turbine section. The fan section has a plurality of vane assemblies spaced circumferentially about an engine axis. The vane assemblies each include an airfoil extending between a leading edge and a trailing edge, a control rod extending through the airfoil, and a mechanism driven by the control rod to change the shape of the airfoil. A vane system for a gas turbine engine is also disclosed.

Abstract: A compressor section for a gas turbine engine includes a blade including a platform, a tip and an airfoil extending between the platform and tip. The airfoil includes a root portion adjacent to the platform, a midspan portion and a tip portion. Each of the root portion, midspan portion and tip portion define a meridional velocity at stage exit with the tip portion including a first meridional velocity greater than a second meridional velocity of the midspan portion. A blade for an axial compressor of a gas turbine engine and a method of operating a compressor section of a gas turbine engine are also disclosed.

Abstract: A compressor section for a gas turbine engine includes a blade including a platform, a tip and an airfoil extending between the platform and tip. The airfoil includes a root portion adjacent to the platform, a midspan portion and a tip portion. Each of the root portion, midspan portion and tip portion define a meridional velocity at stage exit with the tip portion including a first meridional velocity greater than a second meridional velocity of the midspan portion. A blade for an axial compressor of a gas turbine engine and a method of operating a compressor section of a gas turbine engine are also disclosed.

Abstract: A stator vane for a gas turbine engine section includes a stator vane having an airfoil extending between a leading edge and a trailing edge. The airfoil has a suction side and a pressure side. There is at least one piezoelectric actuator for changing a shape of at least one of the leading edge and the trailing edge. A gas turbine engine is also disclosed.

Abstract: A compressor for use in a gas turbine engine including: an airfoil configured to rotate about an engine central longitudinal axis of the gas turbine engine; a casing, the casing including a radially inward surface; and a groove located within the casing opposite the airfoil, the groove is recessed in a radially outward direction from the radially inward surface of the casing, wherein the groove is defined by a forward groove wall, an aft groove wall opposite the forward groove wall, and a base groove interposed between the forward groove wall and the aft groove wall, and wherein the forward groove wall is operably shaped to generate an aft directed jet.

Abstract: A rotor blade of a gas turbine engine includes a pressure side, and a suction side opposite the pressure side and defining a rotor blade profile therebetween, the pressure side and the suction side each extending from a blade root to a blade tip. The rotor blade defines a cross-sectional median line midway between the pressure side and the suction side. The cross-sectional median line extends in a generally radial direction from the blade root to a lean point between the blade root and the blade tip. The cross-sectional median line extends off of radial from the lean point to the blade tip, defining a lean of the rotor blade between the lean point and the blade tip.

Abstract: A stator vane for a gas turbine engine section includes a stator vane having an airfoil extending between a leading edge and a trailing edge. The airfoil has a suction side and a pressure side. There is at least one piezoelectric actuator for changing a shape of at least one of the leading edge and the trailing edge. A gas turbine engine is also disclosed.

Abstract: A compressor for use in a gas turbine engine including: an airfoil configured to rotate about an engine central longitudinal axis of the gas turbine engine; a casing, the casing including a radially inward surface; and a groove located within the casing opposite the airfoil, the groove is recessed in a radially outward direction from the radially inward surface of the casing, wherein the groove is defined by a forward groove wall, an aft groove wall opposite the forward groove wall, and a base groove interposed between the forward groove wall and the aft groove wall, and wherein the forward groove wall is operably shaped to generate an aft directed jet.