The invention relates to a monitoring system for a gas turbine, in particular for an aircraft engine. The monitoring system comprises at least one borescope device that is able to be mounted in a borescope opening of a gas turbine housing and has a housing, in which at least one optical sensor device for acquiring images of at least one inner region of the gas turbine is arranged, and an evaluation device that is able to be connected to the at least one borescope device in order to exchange data and is designed to inspect the at least one inner region for the presence of a fault on the basis of the at least one image acquired by way of the sensor device. The invention furthermore relates to a borescope device to an evaluation device and to a gas turbine.

An instrumentation assembly configured to measure properties of an engine exhaust stream is disclosed in this paper. The instrumentation assembly may include an outer support ring that extends around a central axis, an inner support ring arranged radially inward of the outer support ring around the central axis, and a plurality of instrumentation rake assemblies. The plurality of instrumentation rake assemblies extends from the outer support ring to the inner support ring across an annular passageway defined between the outer support ring and the inner support ring configured to carry the engine exhaust stream.

A blade damage evaluation apparatus includes: a registration unit for design information of a turbine and maintenance information; an acquisition processor for detection data of sensors; a first discrimination processor for first facility states of the turbine at a plurality of past time points; a classification processor for classes C.sub.n of a plurality of first facility states; a first determination processor for first operating state values of the turbine; another registration unit for first damage rates at past time points; a setting processor of a characteristic function for each of the classes C.sub.n; a second discrimination processor for a second facility state of the turbine at the current time point; a second determination processor for a second operating state value of the turbine at the current time point; and an analyzer for a second damage rate at the current time point.

A device for detecting particles in a lubricating oil of a machine, comprising a particle separator; at least one particle detector; a bypass conduit for the particle-concentrating oil, fluidly connected to an oil outlet of the particle separator, concentrating the particles; and wherein the at least one particle detector is operatively mounted on the bypass conduit so as to be able to detect particles in the bypass conduit.

A damage detection system includes one or more emitters, one or more receivers, and a controller. Emitters are arranged to transmit continuous beam or intermittent light pulses toward rotor blades during operation of a turbomachine. Light returns collected at the receivers define a light return amplitude profile. The controller analyzes the light return amplitude profile to identify light amplitude changes indicative of one or more damaged blades. When the light return profile satisfies one or more damage criteria, the controller outputs an indication of blade damage to another turbomachine controller, system, or display.

A power-receiving-module-type telemeter transmitter of a telemeter measuring system that is provided in a rotor and that performs communication using power received from a power transmitting antenna formed in an arc shape on a stator side, includes: a plurality of power receiving modules disposed in the rotor at intervals in a circumferential direction; an OR circuit that outputs a highest voltage among voltages input from the plurality of power receiving modules; a sensor that measures a state of a monitoring target and that outputs measurement data indicating the state; and a transmitting unit that is driven by the voltage input from the OR circuit as a power source and that transmits a radio signal including the measurement data of the sensor.

Disclosed herein is a method of automatically determining an operating condition of at least part of a gas turbine engine 10 for an aircraft, the method comprising: measuring one or more gas pressure waves by a gas pressure detector 401, wherein the gas pressure detector 401 is located in the gas turbine engine 10; and automatically determining, by a computing system, an operating condition of at least part of a gas turbine engine 10 in dependence on an output signal of the gas pressure detector 401.

A gas turbine engine, and an aircraft including the gas turbine engine. The gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a shaft system connecting the turbine to the compressor. The shaft system is axially located by a thrust bearing located forward of the turbine, and the engine is configured such that, in the event of a shaft break which divides the shaft system into a front portion located by the thrust bearing and a rear portion unlocated by the thrust bearing, the rear portion is free to move axially rearwardly under a gas load; and the engine further comprises an axial movement sensor configured to register a shaft break when it detects the axial movement of the rear portion of the shaft system.

A heat management system for turbomachinery is provided. The heat management system includes a thermal delivery system configured to providing heating, cooling, or a combination thereof, to a turbomachinery. The thermal delivery system includes a first heat transfer conduit and a second heat transfer conduit. The second heat transfer conduit is disposed on the turbomachinery adjacent to the first heat transfer conduit. The heat management system additionally includes a controller operatively coupled to the thermal delivery system and configured to control the heating, the cooling, or the combination thereof, of the turbomachinery via the thermal delivery system.

A heat management system for turbomachinery is provided. The heat management system includes a thermal delivery system configured to providing heating, cooling, or a combination thereof, to a turbomachinery. The thermal delivery system includes a first heat transfer conduit and a second heat transfer conduit. The second heat transfer conduit is disposed on the turbomachinery adjacent to the first heat transfer conduit. The heat management system additionally includes a controller operatively coupled to the thermal delivery system and configured to control the heating, the cooling, or the combination thereof, of the turbomachinery via the thermal delivery system.