Abstract: A system for generating combustion gases in a turbine engine may comprise a combustion chamber and a fuel injection system configured to inject a metal-based fuel into the combustion chamber. A metal-based fuel source may be configured to deliver the metal-based fuel to the fuel injection system. The metal-based fuel may comprise a metal powder.

Abstract: Hybrid electric propulsion systems are described. The systems include a gas turbine engine having a low speed spool and a high speed spool. The low speed spool includes a low pressure compressor and a low pressure turbine and the high speed spool includes a high pressure compressor and a high pressure turbine. An electric machine is configured to augment rotational power of at least one of the high speed spool and the low speed spool. A controller is configured to control the electric machine to one of add or subtract rotational energy to or from at least one of the high speed spool and the low speed spool during a transition to or from an idle state of operation of the gas turbine engine.

Abstract: A system for providing real time aircraft engine sensor analysis includes a computer system configured to receive an engine operation data set in real time. The computer system includes a machine learning based analysis tool and a user interface configured to display a real time analysis of the engine operation data set. The user interface includes at least one portion configured to identify a plurality of anomalies in the engine operation data set.

Abstract: A system for providing real time aircraft engine sensor analysis includes a computer system configured to receive an engine operation data set in real time. The computer system includes a machine learning based analysis tool and a user interface configured to display a real time analysis of the engine operation data set. The user interface includes at least one portion configured to identify a plurality of anomalies in the engine operation data set.

Abstract: A gas turbine engine includes a rotor section proximate to a combustor section, where the rotor section is subject to bowing effects due to thermal differences and heat transfer at engine shutdown. The gas turbine engine also includes a heat pipe system. The heat pipe system includes one or more heat pipes installed between an upper portion of the rotor section and a lower portion of the rotor section. The heat pipe system is operable to accept heat at a hot side of the heat pipe system at the upper portion, flow heat from the hot side to a cold side of the heat pipe system, and reject heat from the cold side of the heat pipe system at the lower portion to reduce a thermal differential between the upper portion and the lower portion at engine shutdown.

Abstract: A deflection measurement assembly according to an example of the present disclosure includes, among other things, a nacelle arranged about an axis to define a flow path, a cable assembly arranged at least partially about the axis, and a transducer coupled to the cable assembly.

Abstract: A gas turbine engine includes a rotor section proximate to a combustor section, where the rotor section is subject to bowing effects due to thermal differences and heat transfer at engine shutdown. The gas turbine engine also includes a heat pipe system. The heat pipe system includes one or more heat pipes installed between an upper portion of the rotor section and a lower portion of the rotor section. The heat pipe system is operable to accept heat at a hot side of the heat pipe system at the upper portion, flow heat from the hot side to a cold side of the heat pipe system, and reject heat from the cold side of the heat pipe system at the lower portion to reduce a thermal differential between the upper portion and the lower portion at engine shutdown.

Abstract: A system is provided for a turbine engine. The system includes a rotor assembly and a monitoring system. The rotor assembly includes a plurality of rotor blades connected to a shaft. The monitoring system includes a processing system. This processing system is adapted to receive blade data indicative of a rotational position of at least a first of the rotor blades. The processing system is also adapted to process shaft data with the blade data to provide torque data indicative of a torque to which at least a portion of the rotor assembly is being subjected, where the shaft data is indicative of a rotational position of the shaft.

Abstract: A gas turbine engine includes a bleed structure which includes a forward wall and a rear structural wall to define a deposit space downstream of the bleed structure for a hail event of a predetermined duration.

Abstract: A deflection measurement assembly according to an example of the present disclosure includes, among other things, a nacelle arranged about an axis to define a flow path, a cable assembly arranged at least partially about the axis, and a transducer coupled to the cable assembly.

Abstract: A system is provided for a turbine engine. The system includes a rotor assembly and a monitoring system. The rotor assembly includes a plurality of rotor blades connected to a shaft. The monitoring system includes a processing system. This processing system is adapted to receive blade data indicative of a rotational position of at least a first of the rotor blades. The processing system is also adapted to process shaft data with the blade data to provide torque data indicative of a torque to which at least a portion of the rotor assembly is being subjected, where the shaft data is indicative of a rotational position of the shaft.

Abstract: A gas turbine engine includes a bleed structure which includes a forward wall and a rear structural wall to define a deposit space downstream of the bleed structure for a hail event of a predetermined duration.

Abstract: A gas turbine engine includes a bleed structure which includes a forward wall and a rear structural wall to define a deposit space downstream of the bleed structure for a hail event of a predetermined duration.

Abstract: A method of operating a gas turbine engine for testing, comprising: providing an aircraft on a tarmac, the aircraft having a gas turbine engine with an inlet; selecting a power setting for the engine that is capable of producing a vortex between the inlet and the tarmac; and inhibiting formation of the vortex. A suppressor for preventing a vortex between an inlet of a gas turbine engine on an aircraft and a tarmac. The suppressor comprises: a base facing the tarmac; and an inclined surface extending in a direction from the tarmac towards the inlet at an angle to the base. The suppressor prevents formation of the vortex.

Abstract: A method of operating a gas turbine engine for testing, comprising: providing an aircraft on a tarmac, the aircraft having a gas turbine engine with an inlet; selecting a power setting for the engine that is capable of producing a vortex between the inlet and the tarmac; and inhibiting formation of the vortex. A suppressor for preventing a vortex between an inlet of a gas turbine engine on an aircraft and a tarmac. The suppressor comprises: a base facing the tarmac; and an inclined surface extending in a direction from the tarmac towards the inlet at an angle to the base. The suppressor prevents formation of the vortex.