A gas turbine engine (10) includes a compressor section, a turbine section, and an accessory gearbox (100). A turning unit (200) for the gas turbine engine includes an output assembly (204) configured to be mechanically coupled to the gas turbine engine, and an electric motor (202). The electric motor is operable to rotate, through the output assembly, one or more components of the compressor section or the turbine section at a rotational speed less than about fifty revolutions per minute during a shut-down condition of the gas turbine engine.

In accordance with at least one aspect of this disclosure, a fluid system of an aircraft includes a primary fluid conduit that conveys a primary fluid, and a leak detection system disposed around at least a portion of the primary fluid conduit and forming one or more detection volumes. The leak detection system determines whether there is a primary fluid leak into the one or more detection volumes by sensing a pressure change in the one or more detection volumes.

Examples described herein provide a computer-implemented method that includes monitoring a hybrid electric turbine engine of an aircraft, the hybrid electric turbine engine including a first electric machine associated with a high speed spool and a second electric machine associated with a low speed spool. The method further includes receiving an indication of a failed electric machine, the failed electric machine being an electric machine on another hybrid electric turbine engine of the aircraft. The method further includes, responsive to detecting the failed electric machine, distributing power from one or more of the first electric machine or the second electric machine to a spool associated with the failed electric machine.

Systems and methods for optimizing clearances within an engine include an adjustable coupling configured to couple a thrust link to the aircraft engine, an actuator coupled to the adjustable coupling, where motion produced by the actuator adjusts a hinge point of the adjustable coupling, sensors configured to capture real time flight data, and an electronic control unit. The electronic control unit receives flight data from the sensors, implements a machine learning model trained to predict clearance values within the engine based on the received flight data, predicts, with the machine learning model, the clearance values within the engine based on the received flight data, determines an actuator position based on the clearance values, and causes the actuator to adjust to the determined actuator position.

A measurement system for an aircraft gas turbine engine includes a probe and a heated-gas source in fluid communication with the pressure probe. The probe includes a probe body defining an internal cavity of the probe. The probe further includes a plurality of sensor inlet ports extending through the probe body and configured to receive a sensed fluid flow. The probe further includes a plurality of probe conduits. Each probe conduit of the plurality of probe conduits is coupled to a respective sensor inlet port of the plurality of sensor inlet ports and extending from the respective sensor inlet port to an exterior of the probe body. The heated-gas source is configured to supply a heated gas flow to one or both of: the plurality of sensor inlet ports via the plurality of probe conduits and an interior of the probe body outside of the plurality of probe conduits.

Methods and systems for detecting a shaft event of a gas turbine engine are described. The method comprises monitoring at least one engine parameter and comparing the at least one engine parameter to a schedule for the at least one parameter defining a first threshold and a second threshold greater than the first threshold; applying a limit to the at least one engine parameter when the at least one engine parameter is inside a parameter limiting region between the first threshold and the second threshold, the first threshold separating the parameter limiting region from a normal operating region, the second threshold separating the parameter limiting region from a hazardous operating region; and detecting the shaft event when the at least one engine parameter crosses the second threshold and issuing a signal in response to the detecting.

Methods and systems for detecting a shaft event of a gas turbine engine are described. The method comprises monitoring at least one engine parameter and comparing the at least one engine parameter to a schedule for the at least one parameter defining a first threshold and a second threshold greater than the first threshold; applying a limit to the at least one engine parameter when the at least one engine parameter is inside a parameter limiting region between the first threshold and the second threshold, the first threshold separating the parameter limiting region from a normal operating region, the second threshold separating the parameter limiting region from a hazardous operating region; and detecting the shaft event when the at least one engine parameter crosses the second threshold and issuing a signal in response to the detecting.

An operation method of a turbine fracturing device and a turbine fracturing device are provided. The turbine fracturing device includes a turbine engine, a speed reducer, a brake mechanism, and a fracturing pump, the method includes: driving, by the turbine engine, the fracturing pump to perform a fracturing operation through the speed reducer so as to keep the fracturing pump in an operating state, the fracturing pump being configured to suck fluid of a first pressure and discharge fluid of a second pressure, the second pressure being greater than the first pressure; and in response to an idling instruction, the turbine engine entering an idling state and triggering a brake operation so as to keep the fracturing pump in a non-operating state.

An aircraft power plant comprising: a high-pressure spool including a high-pressure compressor, a high-pressure turbine, and a high-pressure shaft drivingly engaging the high-pressure turbine to the high-pressure compressor; a low-pressure spool including a low-pressure compressor, a low-pressure turbine, and a low-pressure shaft drivingly engaging the low-pressure turbine to the low-pressure compressor; an electrical machine operable as a generator; and a transmission having a first input drivingly engaged by the high-pressure shaft, a second input drivingly engaged by the low-pressure shaft, and an output drivingly engaging the electrical machine, the transmission having a coupling system selectively interconnecting the output with one of: the first input, with the second input disconnected from the output; the second input, with the first input disconnected from the output; and both of the first input and the second input.

Automatic testing for control valves is provided for diagnosing of actuators, including actuators not equipped with analog or discrete position transmitters. A valve controller confirms steady-state conditions for a turbo-compressor system that includes a control valve in a first position and sends, to an actuator for the control valve, a signal to initiate a partial valve stroke to move the control valve away from the first position. The valve controller receives feedback signals from sensors in the turbo-compressor system and monitors the feedback signals for a change from the steady-state conditions. When the monitoring detects a change from the steady-state conditions within a defined time period, the valve controller sends, to the actuator, a signal to return the control valve to the first position. When the monitoring does not detect a change from the steady-state conditions within the defined time period, the valve controller generates an alarm signal.