Abstract: An aircraft propulsion system 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 a gas flow that is expanded through a turbine section for powering a primary propulsor. The aircraft propulsion system further includes a tap that is at a location upstream of the combustor section where a bleed airflow is drawn, a heat exchanger where the bleed airflow is heated by the gas flow, a power turbine through which heated bleed airflow is expanded to generate a work output, and a secondary propulsor that is driven by the work output that is generated by the power turbine.

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 turbine engine assembly includes a tap is at a location upstream of the combustor section where a bleed airflow is drawn. The bleed air is pressurized in an auxiliary compressor section, heated in an exhaust heat exchanger, and expanded through a power turbine that is coupled to drive the auxiliary compressor section.

Abstract: In some embodiments, a computer system displays a virtual surface for containing one or more virtual objects in a three-dimensional environment. In some embodiments, a computer system automatically resizes virtual surfaces that contain objects. In some embodiments, a computer system displays feedback related to removal and/or addition of objects to virtual surfaces.

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 that is in communication with a compressor section, combustor section and a turbine section, the core engine is configured to generate a high energy gas flow, a first cycle turbine that is configured to drive a first cycle compressor at a cycle speed in response to expansion of a heated working fluid flow between a first inlet and a first outlet of the first cycle turbine, and a first power turbine that is configured to drive a first output shaft at a power speed that is different than the cycle speed in response to expansion of a working fluid flow received from the first outlet of the first cycle turbine.

Abstract: A transformer bobbin includes a first bobbin member including a first end defining a first opening, a second end defining a second opening, an inner surface defining the first opening and an outer surface. The inner surface includes a first plurality of fins that extend between the first end and the second end. A second bobbin member is disposed about and spaced from the first bobbin member. The second bobbin member includes a first end portion defining a first opening portion, a second end portion, an inner surface portion and an outer surface portion. One of the outer surface of the first bobbin member and the outer surface portion of the second bobbin member includes a second plurality of fins that extend between the first end portion and the second end portion.

Abstract: A propulsion system for an aircraft includes a core flow path in communication with a compressor section, combustor section and a turbine section. A first bottoming cycle system includes a bottoming working fluid flow in thermal communication with a high energy exhaust gas flow that is generated by the core engine. The first bottoming cycle is configured to recover power from the high energy exhaust gas flow in a first engine operating condition and in a second engine operating condition. A second bottoming cycle system is configured to recover power from the high energy exhaust gas flow in the first engine operating condition and not to recover power in a second engine operating condition.

Abstract: A gas turbine engine includes a core engine that includes a core flow path that connects a compressor section, combustor section and a turbine section. The gas turbine engine further includes a bottoming cycle system that includes a supercritical CO2 (sCO2) working fluid flow. A first recuperator is disposed in the core flow path downstream of the turbine section, the first recuperator is configured to transfer thermal energy from a core flow aft of the turbine section to the sCO2 working fluid flow. A second recuperator is disposed in the compressor section, the second recuperator is configured to transfer thermal energy from the sCO2 working fluid flow to a location forward of the combustor section.

Abstract: A turbine engine assembly includes a tap is at a location upstream of the combustor section where a bleed airflow is drawn. The bleed air is pressurized in an auxiliary compressor section, heated in an exhaust heat exchanger and expanded through a power turbine that is coupled to drive the auxiliary compressor 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 a high energy combusted gas flow that is expanded through a turbine section, a first tap at a location up stream of the combustor section for communicating a portion of the core airflow as a bleed airflow downstream of the combustor section, a heat exchanger that places the bleed airflow that is communicated from the first tap in thermal communication with the high energy combusted gas flow downstream of the combustor section, and a power turbine that is configured to generate shaft power from expansion of the heated bleed airflow, the power turbine includes an inlet that is configured to receive the heated bleed airflow from the heat exchanger.

Abstract: A CO2 bottoming cycle system includes a first compressor operatively connected to a first turbine through a first shaft. A first generator is operatively connected to the first shaft. A second compressor is fluidically connected to the first compressor. The second compressor is operatively connected to a second turbine through a second shaft. A second generator is operatively connected to the second shaft. The first turbine is fluidically connected to the second turbine.

Abstract: An ejector assembly for a cooling system of a gas turbine engine may comprise: a tail cone having a tail cone outlet in fluid communication with a cooling air flow of the cooling system; an ejector body defining a mixing section, a constant area section, and a diffuser section; and a nozzle section in fluid communication with an exhaust air flow of the gas turbine engine, the ejector assembly configured to entrain the cooling air flow via the exhaust air flow.

Abstract: A closed-loop Brayton cycle system utilizes supercritical carbon dioxide as the working fluid for the system to achieve higher efficiencies than can be achieved with traditional open-loop gas turbine engines. A bleed channel is used to direct a flow of cooling fluid to cool the turbine blades during operation of the system, preventing damage to the turbine blades during operation of the system. The bleed channel includes a bleed inlet fluidly coupled between a first recuperator and a second recuperator and a bleed outlet fluidly coupled to the turbine blades. The bleed channel is configured to direct the flow of cooling fluid to the turbine blades at a desired temperature and pressure.

Abstract: An engine driven environmental control system (ECS) air circuit includes a gas turbine engine having a compressor section. The compressor section includes a plurality of compressor bleeds. A selection valve selectively connects each of said bleeds to an input of an intercooler. A second valve is configured to selectively connect an output of said intercooler to at least one auxiliary compressor. The output of each of the at least one auxiliary compressors is connected to an ECS air input.

Abstract: An ejector assembly for a cooling system of a gas turbine engine may comprise: a tail cone having a tail cone outlet in fluid communication with a cooling air flow of the cooling system; an ejector body defining a mixing section, a constant area section, and a diffuser section; and a nozzle section in fluid communication with an exhaust air flow of the gas turbine engine, the ejector assembly configured to entrain the cooling air flow via the exhaust air flow.

Abstract: An engine driven environmental control system (ECS) air circuit includes a gas turbine engine having a compressor section. The compressor section includes a plurality of compressor bleeds. A selection valve selectively connects each of said bleeds to an input of an intercooler. A second valve is configured to selectively connect an output of said intercooler to at least one auxiliary compressor. The output of each of the at least one auxiliary compressors is connected to an ECS air input.

Abstract: An engine driven environmental control system (ECS) air circuit includes a gas turbine engine having a compressor section. The compressor section includes a plurality of compressor bleeds. A selection valve selectively connects each of said bleeds to an input of an intercooler. A second valve is configured to selectively connect an output of said intercooler to at least one auxiliary compressor. The output of each of the at least one auxiliary compressors is connected to an ECS air input.

Abstract: A gas turbine engine includes a main engine compressor section. A booster compressor changing a pressure of airflow received from the main engine compressor section to a pressure desired for a pneumatic system. The booster compressor operates at airflow conditions greater than a demand by the pneumatic system. A speed change system driving the booster compressor at speeds corresponding to a demand of the pneumatic system. A bleed air system for a gas turbine engine and a method of controlling engine bleed airflow are also disclosed.

Abstract: A gas turbine engine includes a main compressor section. A booster compressor includes an inlet and an outlet. The inlet receives airflow from the main compressor section and the outlet provides airflow to a pneumatic system. A recirculation passage is between the inlet and the outlet. A flow splitter valve controls airflow between the outlet and the inlet through the recirculation passage for controlling airflow to the pneumatic system based on airflow output from the booster compressor. A bleed air system for a gas turbine engine and a method of controlling engine bleed airflow are also disclosed.