Abstract: A hybrid energy storage and control system for a clearance control system for a gas turbine engine may comprise a hybrid electric power source, a first converter, a second converter configured to receive electric power from the hybrid electric power source via the first converter and configured to send the electric power to a heating element for controlling a blade tip clearance between a rotor blade and an outer structure of the gas turbine engine, and a controller in electronic communication with the second converter. The hybrid electric power source may comprise a battery, a supercapacitor, and/or an ultracapacitor.

Abstract: An aspect includes a dirt mitigation system for a gas turbine engine. The dirt mitigation system includes a plurality of bleeds of the gas turbine engine and a control system configured to determine a particulate ingestion estimate indicative of dirt ingested in the gas turbine engine. The control system is further configured to determine one or more operating parameters of the gas turbine engine and alter a bleed control schedule of the gas turbine engine to purge at least a portion of the dirt ingested in the gas turbine engine through one or more of the bleeds of the gas turbine engine based on the particulate ingestion estimate and the one or more operating parameters.

Abstract: An aspect includes a system for a gas turbine engine. The system includes one or more bleeds of the gas turbine engine and a control system configured to check one or more activation conditions of a dirt rejection mode in the gas turbine engine. A bleed control schedule of the gas turbine engine is adjusted to extend a time to hold the one or more bleeds of the gas turbine engine partially open at a power setting above a threshold based on the one or more activation conditions. One or more deactivation conditions of the dirt rejection mode in the gas turbine engine are checked. The dirt rejection mode is deactivated to fully close the one or more bleeds based on the one or more deactivation conditions.

Abstract: An aircraft gas turbine engine system comprises first and second gas turbine engines. The first gas turbine engine has first and second spools. A first power linkage connects the second gas turbine engine to the first spool of the gas turbine engine, and a second power linkage connects the second gas turbine engine to the second spool of the first gas turbine engine.

Abstract: An aircraft gas turbine engine system comprises first and second gas turbine engines connected by an inter-engine gas path. The first gas turbine engine has a first spool with a first compressor section, and a second spool with a second compressor section downstream of and rotationally independent from the first compressor section. The second gas turbine engine is configured to provide power to at least one of the first and second spools of the first gas turbine engine. The inter-engine gas path is disposed to receive gas flow bled from a bleed location in the first gas turbine engine downstream of the first compressor section, and to supply this gas flow to an inlet of the second gas turbine engine.

Abstract: A gas turbine engine according to an exemplary embodiment of this disclosure includes among other possible things, a fan section including a plurality of fan blades, a first electric motor assembly that provides a first drive input for driving the fan blades about an axis, a turbine section, and a geared architecture driven by the turbine section and coupled to the fan section to provide a second drive input for driving the fan blades, and second electric motor assembly is coupled to rotate the geared architecture relative to a fixed structure controls a speed of the fan blades provided by a combination of the first drive input and the second drive input.

Abstract: A clearance control system for a gas turbine engine may comprise a rotor blade, an outer structure disposed radially outward from the rotor blade, and a heating element configured to cause the outer structure to be heated in response to electric current being supplied to the heating element, wherein a gap between the rotor blade and the outer structure is at least one of increased, decreased, and maintained in response to the outer structure being heated.

Abstract: A hybrid energy storage and control system for a clearance control system for a gas turbine engine may comprise a hybrid electric power source, a first converter, a second converter configured to receive electric power from the hybrid electric power source via the first converter and configured to send the electric power to a heating element for controlling a blade tip clearance between a rotor blade and an outer structure of the gas turbine engine, and a controller in electronic communication with the second converter. The hybrid electric power source may comprise a battery, a supercapacitor, and/or an ultracapacitor.

Abstract: A method of assembling a gas turbine engine includes setting a build clearance at assembly in response to a running tip clearance defined with a cooled cooling air. A method of operating a gas turbine engine includes supplying a cooled cooling air to a high pressure turbine in response to an engine rotor speed.

Abstract: Aspects of the disclosure are directed to an engine of an aircraft. The engine may include a first fan configured to output a first air flow, a second fan configured to receive a first portion of the first air flow and output a second air flow, a core configured to receive a first portion of the second air flow and generate a first stream, and at least one valve configured to assume one of at least three states in association with a generation of a second stream and a third stream based on at least one of the first air flow and the second air flow.

Abstract: A gas turbine engine includes low and high spools constructed and arranged to rotate about an engine axis. The low spool drives at least one leading stage of a fan section and the high spool drives an aft stage of the fan section. The aft stage may generally include a bypass duct for controllably flowing a bypass stream directly from the leading stage and controllably and/or selectively into an auxiliary flowpath and/or into a second flowpath both located radially outward from a core flowpath.

Abstract: A gas turbine engine includes an auxiliary airflow control system to provide selective communication of a portion of a second stream from a second stream flow path into a third stream flow path, the auxiliary airflow control system in communication with a high temperature flow sourced from the turbine section.

Abstract: A gas turbine engine includes a compressor section having a first compressor and a second compressor, the second compressor having a higher pressure than the first compressor, and a turbine section having a first turbine and a second turbine, the second turbine having a higher pressure than the first turbine. The first compressor is connected to the first turbine via a first shaft. The second compressor is connected to the second turbine via a second shaft. An electric motor is connected to the first shaft such that rotational energy generated by the electric motor is translated to the first shaft. A fan is connected to the first shaft via a gear system. The gas turbine engine includes at least a takeoff mode of operations, a top of climb mode of operations and a maximum cruise mode of operations. The gas turbine engine is sized to operate at peak efficiency in the maximum cruise mode of operations.

Abstract: A method of assembling a gas turbine engine includes setting a build clearance at assembly in response to a running tip clearance defined with a cooled cooling air. A method of operating a gas turbine engine includes supplying a cooled cooling air to a high pressure turbine in response to an engine rotor speed.

Abstract: A gas generator has at least one compressor rotor, at least one gas generator turbine rotor and a combustion section. A fan drive turbine is positioned downstream of a path of the products of combustion having passed over the at least one gas generator turbine rotor. The fan drive turbine drives a shaft and the shaft engages gears to drive at least three fan rotors.

Abstract: A gas generator has at least one compressor rotor, at least one gas generator turbine rotor and a combustion section. A fan drive turbine is positioned downstream of a path of the products of combustion having passed over the at least one gas generator turbine rotor. The fan drive turbine drives a shaft and the shaft engages gears to drive at least three fan rotors.

Abstract: A gas turbine engine includes an auxiliary airflow control system to provide selective communication of a portion of a second stream from a second stream flow path into a third stream flow path, the auxiliary airflow control system in communication with a high temperature flow sourced from the turbine section.

Abstract: Aspects of the disclosure are directed to an engine of an aircraft. The engine may include a first fan configured to output a first air flow, a second fan configured to receive a first portion of the first air flow and output a second air flow, a core configured to receive a first portion of the second air flow and generate a first stream, and at least one valve configured to assume one of at least three states in association with a generation of a second stream and a third stream based on at least one of the first air flow and the second air flow.

Abstract: A gas turbine engine includes a core flowpath for flowing a core stream, a second flowpath located radially outward from the core flowpath for flowing a second stream, and an auxiliary flowpath located radially outward from the second flowpath for flowing an auxiliary stream. A heat exchanging device is constructed and arranged to divert a portion of the second stream into the auxiliary flowpath. A turbine exhaust case is constructed and arranged to flow the auxiliary stream into the core flowpath for mixing with the core stream.

Abstract: A gas turbine engine includes low and high spools constructed and arranged to rotate about an engine axis. The low spool drives at least one leading stage of a fan section and the high spool drives an aft stage of the fan section. The aft stage may generally include a bypass duct for controllably flowing a bypass stream directly from the leading stage and controllably and/or selectively into an auxiliary flowpath and/or into a second flowpath both located radially outward from a core flowpath.