Abstract: A power and thermal management system thermal bus vapor cycle architecture including a bypass duct formed between a gas turbine inner fan duct and a gas turbine outer fan duct; a condenser coil located within the bypass duct, the condenser coil being configured as an annulus disk shape heat exchanger; at least one evaporator coil fluidly coupled to the condenser coil and located within the bypass duct upstream of the condenser coil relative to a bypass air flow direction; a gearbox in operative communication with the gas turbine engine; and a variable speed compressor in operative communication with the gearbox, the variable speed compressor being fluidly coupled with the at least one evaporator coil and the condenser coil.

Abstract: A gas turbine engine assembly includes a core engine that includes a core flow path where a core airflow is compressed in a compressor section, communicated to a combustor section, mixed with fuel and ignited to generate an exhaust gas flow that is expanded through a turbine section. The turbine section is coupled to drive the compressor section through an engine drive shaft. A tap is at a location up stream of the combustor section for drawing a bleed airflow. A bleed air heat exchanger places the bleed airflow in thermal communication with an auxiliary flow for heating the bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow.

Abstract: A gas turbine engine assembly including a tap that is at a location up stream of the combustor section for drawing a bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow and communicate the heated bleed airflow into the turbine section where it is expanded to drive the turbine section.

Abstract: A gear carrier drives a propulsor. Two ring gears engage with respective tooth locations on planet gears. A first gear tooth set on a ring lock selectively engage with a first ring gear and a second gear tooth set on a ring lock selectively engage with a second ring gear. When the first gear tooth set on the ring lock engages the first ring gear, rotation of the first ring gear stops, and the planet gears drive the carrier at a first speed. When the second ring lock gear tooth set engages the second ring gear, rotation of the second ring gear stops and the planet gear drives the carrier at a second speed which is different than the first speed. An aircraft and a method are also disclosed.

Abstract: A gas turbine engine assembly including a tap that is at a location up stream of the combustor section for drawing a bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow and communicate the heated bleed airflow into the turbine section where it is expanded to drive the turbine section.

Abstract: A gas turbine engine assembly including a tap that is at a location up stream of the combustor section for drawing a bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow and communicate the heated bleed airflow into the turbine section where it is expanded to drive the turbine section.

Abstract: A propulsion system for an aircraft includes a core engine that includes a core flow path where a core flow is compressed in a compressor section, communicated to a combustor section, mixed with a hydrogen-based fuel, and ignited to generate a gas flow that is expanded through a turbine section. A fuel system is configured to supply a hydrogen based fuel to the combustor through a fuel flow path. A condenser is arranged along the core flow path and configured to extract water from the gas flow. An intercooling system receives a portion of water from the condenser for cooling a portion of the core flow at a first location within the compressor section. Heated water from the intercooling system is exhausted to a second location within the core flow path downstream of the first location.

Abstract: A method can comprise dividing a heat exchanger design into a plurality of modules, the plurality of modules arranged in a grid, each module in the plurality of modules including: a first fluid conduit defining an inlet, an outlet, and a heat-transfer surface, and a first flow direction, and a second fluid conduit defining a second inlet, a second outlet, a second heat-transfer surface, and a second flow direction, the second flow direction different from the first flow direction; and determining a heat-transfer augmenter arrangement for the first fluid conduit and the second fluid conduit of each module in the plurality of modules based on a stress threshold of the module in the plurality of modules.

Abstract: A gas turbine engine assembly including a tap that is at a location up stream of the combustor section for drawing a bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow and communicate the heated bleed airflow into the turbine section where it is expanded to drive the turbine section.

Abstract: A gas turbine engine assembly includes a core engine that includes a core flow path where a core airflow is compressed in a compressor section, communicated to a combustor section, mixed with fuel and ignited to generate an exhaust gas flow that is expanded through a turbine section. The turbine section is coupled to drive the compressor section through an engine drive shaft. A tap is at a location up stream of the combustor section for drawing a bleed airflow. A bleed air heat exchanger places the bleed airflow in thermal communication with an auxiliary flow for heating the bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow.

Abstract: A propulsion system for an aircraft includes a core engine that includes a core flow path where a core flow is compressed in a compressor section, communicated to a combustor section, mixed with a hydrogen-based fuel and ignited to generate a gas flow that is expanded through a turbine section. A fuel system is configured to supply a hydrogen based fuel to the combustor through a fuel flow path. A condenser is arranged along the core flow path and configured to extract water from the gas flow. An intercooling system receives a portion of water from the condenser for cooling a portion of the core flow at a first location within the compressor section. Heated water from the intercooling system is exhausted to a second location within the core flow path downstream of the first location.

Abstract: A propulsion system for an aircraft includes a core engine that includes a core flow path where a core flow is compressed in a compressor section, communicated to a combustor section, mixed with a hydrogen-based fuel and ignited to generate a gas flow that is expanded through a turbine section. A fuel system is configured to supply a hydrogen based fuel to the combustor through a fuel flow path. A condenser is arranged along the core flow path and configured to extract water from the gas flow. An intercooling system receives a portion of water from the condenser for cooling a portion of the core flow at a first location within the compressor section. Heated water from the intercooling system is exhausted to a second location within the core flow path downstream of the first location.

Abstract: A gas turbine engine has a compressor section and a turbine section. A secondary cooling air includes a first fluid connection to tap cooling air and pass the cooling air through a plurality of tubes, and a second fluid connection for returning air from the tubes back to at least one of the compressor and turbine for cooling. A sensor senses a condition of the cooling air downstream of the tubes and a control compares the sensed condition of the cooling air to an expected condition, and to identify a potential concern in the cooling air system should the sensed condition differ from the expected condition by more than a predetermined amount.

Abstract: A gas turbine engine has a compressor section and a turbine section. A secondary cooling air includes a first fluid connection to tap cooling air and pass the cooling air through a plurality of tubes, and a second fluid connection for returning air from the tubes back to at least one of the compressor and turbine for cooling. A sensor senses a condition of the cooling air downstream of the tubes and a control compares the sensed condition of the cooling air to an expected condition, and to identify a potential concern in the cooling air system should the sensed condition differ from the expected condition by more than a predetermined amount.