Abstract: In some examples, a gas turbine engine including a high-pressure (HP) spool assembly including a HP shaft, a HP compressor and HP turbine; a lower pressure (LP) spool assembly including a LP shaft and LP turbine; a motor-generator coupled to the LP shaft; and a controller. The controller is configured to control the motor-generator to operate in a motor mode to apply torque the LP shaft during a starting of the HP spool assembly, and control the motor-generator to operate in a generator mode for a least a period of time following the starting of the HP spool assembly.

Abstract: A method for determining the flow rate of combustible fluid injected into a combustion chamber (120) of a turbine (100) includes determining the cross section of the orifice of the at least one injector (112, 113, 114, 115) through which the combustible fluid is injected into the combustion chamber (120). The pressure of the combustible fluid upstream of the orifice of the injector (112, 113, 114, 115) is determined. The pressure downstream of the orifice of the injector (112, 113, 114, 115) is determined. The flow rate of combustible fluid flowing through the orifice of the at least one injector (112, 113, 114, 115) is determined.

Abstract: An anti-icing system includes a nozzle assembly, a bleed air supply, a bleed control valve assembly, and a controller. The bleed air supply is configured to direct pressurized bleed air to the nozzle assembly. The bleed air supply includes a first pressure sensor configured to measure a first pressure of the pressurized bleed air. The bleed control valve assembly includes a control valve and a valve actuator. The control valve is positionable to control a flow rate of the pressurized bleed air. The valve actuator is configured to control a position of the control valve. The controller is configured to identify a power condition of the bleed air supply as a first power condition or a second power condition and control the valve actuator to position the control valve in a fully opened position for the first power condition and in a predetermined position based on the first pressure for the second power condition.

Abstract: A method for detecting blowout precursors in at least one gas turbine combustor comprising: receiving combustion dynamics acoustic data measured by an acoustic measuring device associated with the combustor in real time; performing wavelet analysis on the acoustic data using simplified Mexican Hat wavelet transform analysis; and determining the existence of a blowout precursor based at least in part on the wavelet analysis. Provided also is a system and a non-transitory computer readable medium configured to perform the method.

Abstract: A gas turbine engine for an aircraft that includes a nacelle, a fan, an engine core, a bypass duct extending between the engine core and the nacelle and guiding a bypass airflow through the bypass duct, and at least one non-structural strut extending in the radial direction within the bypass duct, wherein the non-structural strut includes an outside wall acting as a heat exchanger, and wherein the outside wall includes first transport means configured to transport in the outside wall at least one fluid to be cooled. It is provided that the non-structural strut further includes second transport means configured to transport a fluid to be heated, wherein the first transport means and the second transport means are configured such that the fluid to be heated is heated by the at least one fluid to be cooled and the at least one fluid to be cooled is cooled both by the bypass airflow and the fluid to be heated.

Abstract: Methods and systems for controlling a bleed-off valve of a gas turbine engine are described. The method comprises maintaining a first bleed-off valve associated with a first compressor of the gas turbine engine at least partially open upon detection of an unintended engine disturbance causing a drop in pressure of a combustion chamber of the engine; monitoring a rotor acceleration of the first compressor; and controlling closure of the first bleed-off valve when the rotor acceleration of the first compressor reaches a first threshold for a first duration.

Abstract: An oil delivery system is provided that includes a tank for a gas turbine engine, where the tank is positioned radially outward from a compressor section, a combustor section, and/or a turbine section of the gas turbine engine. The tank is configured to store oil for the gas turbine engine. The oil delivery system further includes a primary lubrication system including a sump of a power gearbox, a pump, and an oil feed line. The oil feed line extends from the tank to the primary lubrication system. The oil feed line is configured to allow a flow of oil to pass from the tank to the pump and from the pump through the power gearbox to the sump of the power gearbox.

Abstract: The invention regards a gas turbine engine control system and a method for limiting turbine overspeed in case of a shaft failure. The control system includes: an overspeed protection system that activates an activation member in case a shaft failure is detected; a fuel limiting mechanism coupled with the activation member, wherein the fuel limiting mechanism is configured to limit the fuel supply to the gas turbine engine combustor if the activation member is activated; a variable stator vane mechanism which is configured to adjust variable stator vanes of a compressor of the gas turbine engine in their rotational position, the variable stator vanes having a closed position which blocks air flow through the compressor.

Abstract: A fuel staging system for a gas turbine engine has a plurality of fuel injectors each having a mains burner. The system has a mains manifold connected to a mains delivery line and configured to distribute fuel from the mains delivery line to the mains burner of each of the plurality of fuel injectors, and a check valve disposed in the mains delivery line upstream of the mains manifold. The check valve is configured to permit flow of fuel from the mains delivery line to the mains manifold when the pressure of fuel in the mains delivery line exceeds a threshold pressure.

Abstract: A system includes an engine case. A heat exchanger is included inside the engine case. The heat exchanger includes an air passage and a fuel passage. The air passage and fuel passage are in fluid isolation from one another, but are in thermal communication with one another for exchange of heat.

Abstract: A method of cooling a gas turbine engine case assembly includes moving a fan air valve that is operatively connected to a pre-cooler having a bypass inlet that is configured to receive bypass air that bypasses a gas turbine engine core to facilitate a provision of bypass air through a fan air valve inlet to the bypass inlet to a first open position, in response to a core compartment temperature being greater than a target core compartment temperature. The method further includes bleeding the bypass air through a bypass outlet of the pre-cooler into a core compartment.

Abstract: An inlet body for a fluid injector for a turbomachine. The inlet body includes a casing defined by an internal surface and a seal valve housed inside the casing. The valve includes a sealing member including an intake duct, and internal duct into which the intake duct opens, and a seat for the sealing member. The seat defines an opening over a fluid path towards the internal duct. The internal surface includes a recess which defines at least partially a chamber communicating with the opening and with the intake duct.

Abstract: Systems and methods for adjusting airflow distortion in a gas turbine engine using a valved airflow passage assembly are provided. A gas turbine engine can include a compressor section, a combustion section, and a turbine section in series flow and defining at least in part an engine airflow path. The compressor section can include a compressor. The gas turbine engine can further include a valved airflow passage assembly comprising a valve and a duct, the duct defining an inlet in airflow communication with the engine airflow path at a location downstream of the compressor and an outlet in airflow communication with the engine airflow path at a location upstream of the compressor, the duct comprising an airflow passage extending between the inlet and outlet. The valve can be operable with the airflow passage for controlling an airflow through the airflow passage to adjust airflow distortion.

Abstract: A control system for a gas turbine includes a controller. The controller includes a processor configured to receive a plurality of signals from sensors disposed in the gas turbine engine system, wherein the gas turbine system engine comprises a compressor section fluidly coupled to a gas turbine section. The processor is additionally configured to derive a vanadium content in a gas turbine engine fuel based on at least one of the plurality of signals. The processor is also configured to determine if a control curve should be adjusted based on the vanadium content in the gas turbine engine fuel, and if it is determined that the control curve should be adjusted, then deriving an adjustment to the control curve based on the vanadium content, and applying the adjustment to the control curve to derive an adjusted control curve.

Abstract: Various embodiments include an offline leak detection system for a turbomachine fuel system. In some embodiments, the leak detection system includes: a fluid supply system fluidly connected to at least one fuel line of the turbomachine, the fluid supply system for delivering a non-flammable fluid to the combustor; a control system operably connected to the fluid supply system, the control system controlling a flow of the non-flammable fluid through the at least one fuel line, and controlling a pressure of the non-flammable fluid in the at least one fuel line to a pressure substantially equal to an operational fuel pressure of the turbomachine; and an optical monitor for determining a presence of the non-flammable fluid on an exterior of the at least one fuel line, the presence of the non-flammable fluid on the exterior of the at least one fuel line indicating a leak.

Abstract: An engine bleed control system for a gas turbine engine of an aircraft is provided. The engine bleed control system includes a multi-tap bleed array including engine bleed taps coupled to a compressor source of a lower pressure compressor section before a highest pressure compressor section of the gas turbine engine. A highest stage of the engine bleed taps has a maximum bleed temperature below an auto-ignition point of a fuel-air mixture of the aircraft at idle engine power at a maximum aircraft altitude and a pressure suitable for pressurizing the aircraft at the maximum aircraft altitude. The engine bleed control system also includes a plurality of valves operable to extract bleed air from each of the engine bleed taps. A controller is operable to selectively open and close each of the valves based on a bleed air demand and control delivery of the bleed air to an aircraft use.

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 aeroderivative gas turbine including an air intake plenum; a compressor with a compressor air intake in fluid communication with the air intake plenum; a combustor; a high pressure turbine; a power turbine. A forced air-stream generator is arranged in fluid communication with the air intake plenum. A shutter arrangement is provided in a combustion-air flow path, arranged and controlled to close the combustion-air flow path for pressurizing said air intake plenum by means of the forced air-stream generator to a pressure sufficient to cause pressurized air to flow through the aeroderivative air turbine.

Abstract: A method for controlling takeoffs of mechanical energy and/or air on a turbine engine for the propulsion of an aircraft. It is based on a protocol for the exchange of a request/authorization between and by the energy manager and a system for controlling the turbine engine. This protocol is implemented as a result of a modification of the takeoff requirement. It is intended to check whether the surge margin is compatible with the modification of the takeoff requirement and, if the need arises, to apply temporary measures in order to prevent surge in the turbine engine, for example by providing at least a part of the energy requirements by a buffer. This permits the optimization of the operation of a turbine engine for the propulsion of an aircraft, while avoiding the risk of surge of the turbine engine.

Abstract: This disclosure describes at least embodiments of an aircraft monitoring system for an electric or hybrid airplane. The aircraft monitoring system can be constructed to enable the electric or hybrid aircraft to pass certification requirements relating to a safety risk analysis. The aircraft monitoring system can have different subsystems for monitoring and alerting of failures of components, such as a power source for powering an electric motor, of the electric or hybrid aircraft. The failures that pose a greater safety risk may be monitored and indicated by one or more subsystems without use of programmable components.

Abstract: A system for managing a temperature of at least one bearing in an air cycle machine includes an inlet for allowing cooling air to enter the air cycle machine, at least one bearing downstream from the inlet and in fluidic connection with the inlet, an outlet downstream from the at least one bearing and in fluidic connection with the at least one bearing, a temperature sensor at a point between the inlet and the outlet, and an electronic controller electrically connected to and receiving temperature data from the temperature sensor with the electronic controller configured to modify an operation of the air cycle machine depending on the temperature data.

Abstract: A compound engine assembly with an inlet duct having an inlet surrounded by an inlet lip including at least one conduit extending therethrough, a compressor, an engine core including at least one internal combustion engine, a turbine section having a turbine shaft in driving engagement with the engine shaft, and an exhaust conduit providing a fluid communication between the outlet of the turbine section and the conduit(s) of the inlet lip. An exhaust duct and ant exhaust conduit providing a fluid communication between the outlet of the turbine section and the exhaust duct may also be provided. The internal combustion engine(s) may be rotary engine(s). A method of driving a rotatable load of an aircraft is also discussed.

Abstract: A system and method for monitoring icing conditions that are suitable ice formation on an aircraft and propulsion system. The system includes instrumentation that instantaneously detects ambient humidity, ambient temperature and ambient pressure. The sensed information is transmitted to a controller that evaluates the information to determine whether certain pressure, temperature and humidity criteria are favorable for icing and, declaring icing conditions. The system also includes an aircraft engine-mounted ice mitigation system. When conditions for ice formation are favorable, the controller either informs the pilot that conditions for ice formation are favorable or automatically activates the ice mitigation system, or both. The pilot optionally may inactivate the ice mitigation system. When sensed conditions indicate that conditions for ice formation are not favorable, the controller determines whether the ice mitigation system is activated and inactivates the system if activated.

Abstract: A power generation system includes a processor and memory storing instructions that cause the processor to receive a first set of sensor data indicative of one or more ambient conditions with respect to the power generation system, determine whether one of the one or more ambient conditions is above a respective threshold, and send a signal to a valve to open when the one of the one or more ambient conditions is below the respective threshold, such that the valve is configured to fluidly couple a first fluid exiting a compressor to an inlet of the compressor.

Abstract: The invention includes a method for predicting the operational state of equipment with turbulent flow characterized by time series data relating to its operation. The invention further includes a system and method for predicting the onset of an impending oscillatory instability. Further, the invention includes a system and method for identifying an impending absorbing transition such as flame blowout in combustion systems. A variable representing the dynamics of operation is measured with the help of a sensor, to obtain time series data. A complex network is then derived from the measured time series data. Network properties are then calculated using the complex network to identify the state of stability relating to operation of the equipment. The stability information may include one of thermoacoustic instability, aero-elastic instability such as flutter, flow-induced vibration, magneto-hydrodynamic, aerodynamic, aeromechanical, aero-acoustic instability or onset of flame blowout of a combustor.

Abstract: An aircraft jet propulsion system is disclosed. The aircraft jet propulsion system may comprise a thermoelectric generator array (“TEG” array) coupled to a portion of the aircraft jet propulsion system, wherein the TEG array converts heat energy to electrical energy, and supplies power to the aircraft jet propulsion system, wherein the electrical energy is supplied to a power supply. The aircraft jet propulsion system may comprise an alternator that generates less energy than is required to power the aircraft jet propulsion system. The TEG array may supplement the energy generated by the alternator. The energy generated by the TEG array and the energy generated by the alternator may be sufficient to power the aircraft jet propulsion system and/or the electrical energy generated by the TEG array may be sufficient to power to aircraft jet propulsion system.

Abstract: A fuel injection system includes a first injection valve, a second injection valve, a communication pipe, and an ECU. The communication pipe makes a second passage of the second injection valve and a back pressure chamber of the first injection valve communicate with each other. The ECU controls an electric actuator of the second injection valve so as to switch first injection control and second injection control according to the required injection amount. In the first injection control, a second valve body is opened for a predetermined time or more so that back pressure lowers to the pressure of opening a first valve body. In the second injection control, the second valve body is opened for less than the predetermined time so that the back pressure does not lower to the pressure of opening the first valve body.

Abstract: A method for use in a turbine control system includes controlling fuel supply to a gas turbine engine at least in part using a fuel supply limit determined as a first function of a rotational speed of a shaft of the gas turbine engine. The method also includes obtaining a first value representative of a rotational speed of the shaft, and differentiating the first value within a processing unit. The processing unit determines an adjusted fuel supply limit as an adjusted function of the first value. The adjusted function is based on the first function and the differentiated first value. The method further includes controlling the fuel supply to the gas turbine engine at least in part using the adjusted fuel supply limit.

Abstract: An example method of fuel delivery to a turbomachine includes reaching a light-off speed of a turbomachine and delivering fuel to a combustor of the turbomachine at a first rate. The method also increases the rate of the delivery. The combustor achieves light-off at a fuel flow rate on or in-between the first rate and a second rate that is greater than the first rate. An example turbomachine fuel delivery assembly includes a fuel delivery component and a controller configured to adjust the fuel pump to increase a rate of fuel delivery from a fuel supply to a combustor of a turbomachine.

Abstract: A method of operating a gas turbine engine compressor. The method includes: determining an operating point of the compressor, and modulating mass flow of environmental control system input air to maintain the operating point of the gas turbine engine compressor within predetermined limits.

Abstract: An arrangement for connecting a sensor assembly and a cable assembly. The arrangement includes a signal wire. The arrangement includes a transversely-extending insulating barrier. The signal wire extends through the insulating barrier. The arrangement includes a shield wire that is configured to provide at least part of a grounding pathway. The shield wire extends through the insulating barrier at a location spaced from the signal wire. The arrangement includes at least one conductive layer that is located upon the insulating barrier. The at least one conductive layer is spaced away from the signal wire and is in electrical contact with the shield wire. At least a portion of the insulating barrier has a non-linear profile. At least a portion of the at least one conductive layer that is located upon the at least a portion of the insulating barrier has the non-linear profile.

Abstract: In one embodiment, a system is provided. The system includes a combustor configured to combust a fuel, and a three-way valve fluidly coupled the combustor and disposed upstream of the combustor. The system also includes a fuel circuit comprising a fuel supply, wherein the fuel circuit is disposed upstream of the three-way valve and is configured to provide the fuel to the three-way valve. The system additionally includes a fuel conduit section fluidly coupling the fuel circuit to the combustor. The system further includes an inert fluid supply configured to provide an inert fluid to the three-way valve and a compressor discharge (CPD) fluid source configured to provide a purge fluid to the three-way valve. The three-way valve is configured to purge the fuel from a first portion of the fuel conduit section by using the purge fluid.

Abstract: A fuel injector for a gas turbine engine may include a flow path for a fuel-air mixture extending longitudinally through the fuel injector, and a fuel gallery extending circumferentially around the flow path. The fuel gallery may be adapted to inject a liquid fuel into the flow path. The fuel injector may also include an annular casing positioned circumferentially around the fuel gallery to define an insulating chamber around the gallery. The fuel injector may also include an annular cover extending around the fuel injector to define a metering chamber. The fuel injector may further include one or more purge holes fluidly coupling the metering chamber to the insulating chamber, and one or more metering holes fluidly coupling the metering chamber to a volume exterior to the fuel injector.

Abstract: A method tracking positive displacement efficiency of a high pressure positive displacement pump in a hydraulic regulator system of a turbomachine delivering a flow rate based on an engine speed of the turbomachine, the flow rate delivered to an actuator actuating variable geometry vanes of the turbomachine and to a bypass valve feeding engines of the turbomachine, the method including: starting engines of the turbomachine at a low engine speed, the valve being closed; using a computer to order a movement of the actuator; progressively increasing the engine speed until the flow rate reaches a predetermined value sufficient for opening the valve; storing in the computer a position of the actuator and the engine speed corresponding to opening of the valve; repeating the preceding operations at successive instants during a lifetime of the engines; and replacing the high pressure positive displacement pump when the engine speed exceeds a predetermined value.

Abstract: Systems and methods for learning torque estimate errors and updating torque estimation models are presented. In one example, torque errors are learned during an engine shut-down, after a disconnect clutch coupled between an engine and an electric machine has been released. An updated torque estimation model is then used to control torque during subsequent engine operation to improve drive feel and vehicle performance.

Abstract: A bifurcation assembly for a gas turbine engine is disclosed and includes a housing defining an inner cavity. A first partition and a second partition extend across the inner cavity and define a buffer area therebetween. Openings through the first and second partitions define passageway for supply conduits. A control device governs the flow of a purge gas into the buffer area for blocking the spread of fire. The purge gas provides a barrier to the spread of fire and heat through the bifurcation assembly.

Abstract: A vent for use in a gaseous fuel supply circuit of a gas turbine is provided. The vent includes an inlet in flow communication with the gaseous fuel supply circuit, a first outlet in flow communication with the gaseous fuel supply circuit and configured to release gaseous fuel at atmospheric pressure, a first valve coupled between the inlet and the first outlet, wherein the first valve includes a second outlet configured to channel the gaseous fuel towards a combustion device. The system also includes a second valve coupled between the inlet and the second outlet, and a control device configured to selectively open and close the first and second valves based on a pressure of the gaseous fuel.

Abstract: A gas turbine combustor is provided with a combustor basket where combustion gas flows, the combustion gas being produced by combustion of fuel injected from a nozzle, and a first resonance device and a second resonance device mounted on an outer surface of the combustor basket. The second resonance device is disposed on a downstream side from the first resonance device in a flow of the combustion gas and damps combustion oscillation of a frequency higher than the first resonance device. The first and second resonance devices are acoustic liners each having a housing mounted to the outer surface of the combustor basket. A resonance space surrounded by the housing and the outer surface of the combustor basket communicates with an interior space of the combustor basket via a plurality of acoustic holes formed in the combustor basket.

Abstract: A device for detecting a fault in a low-pressure fuel pump of a turbojet. The pump is driven by an accessory gearbox including a gear for mechanically driving the accessories. The device measures the vibration frequencies of the accessory gearbox and detects, from among the frequencies, at least one vibration frequency of the low-pressure fuel pump. The device allows a fault in the low-pressure fuel pump to be detected as soon as it occurs.

Abstract: An aircraft turbine engine comprising at least one spool rotating at speed N1 and a monitoring system (1) comprising: a regulating module (REG) comprising at least one regulation measurement channel (A, B) suitable for obtaining a measurement (N1A, N1B) of the speed N1 and means for comparing the obtained speed measurement (N1A, N1B) with a thrust setpoint (N1cons) to provide a thrust status (EREG); and a module (ENG) for engaging a protection function of the UHT or ATTCS type of the turbine engine, the turbine engine also comprising a system for protection against overspeed in order to prevent the ejection of high-energy debris outside said turbine engine, the protection system comprising at least one overspeed measurement channel (AS, BS) suitable for obtaining an overspeed (N1AS, N1Bs) of the rotating spool of the turbine engine, the turbine engine being characterised in that the engagement module (ENG) comprises means for comparison of at least one overspeed obtained (N1AS, N1BS) with at least one ref

Abstract: A turbine rotor with at least one rotor disk is provided. The turbine rotor includes a heating arrangement for heating at least a part of the rotor disk. The heating arrangement has a conductor for conducting electric current. The conductor is positioned such that in presence of the electric current in the conductor, Eddy current is generated in the part of the rotor disk. The electric current is an alternating current. The Eddy currents so generated in the part of the rotor disk result in inductive heating of the part of the rotor disk.

Abstract: A shell air recirculation system for use in a gas turbine engine includes one or more outlet ports located at a bottom wall section of an engine casing wall and one or more inlet ports located at a top wall section of the engine casing wall. The system further includes a piping system that provides fluid communication between the outlet port(s) and the inlet port(s), a blower for extracting air from a combustor shell through the outlet port(s) and for conveying the extracted air to the inlet port(s), and a valve system for selectively allowing and preventing air from passing through the piping system. The system operates during less than full load operation of the engine to circulate air within the combustor shell but is not operational during full load operation of the engine.

Abstract: The invention relates to a method for operating a gas turbine power plant, including a gas turbine, a HRSG following the gas turbine, an exhaust gas blower, and a carbon dioxide separation plant which separates the carbon dioxide contained in the exhaust gases and discharges it to a carbon dioxide outlet, the gas turbine, HRSG, exhaust gas blower, and carbon dioxide separation plant being connected by means of exhaust gas lines. According to the method a trip of the gas turbine power plant includes the steps of: stopping the fuel supply, switching off the exhaust gas blower, and controlling the opening angle of a VIGV at a position bigger or equal to a position required to keep a pressure in the exhaust gas lines between the HRSG and the exhaust gas blower above a minimum required pressure. The invention relates, further relates to a gas turbine power plant configured to carry out such a method.

Abstract: An aeroderivative gas turbine including an air intake plenum; a compressor with a compressor air intake in fluid communication with the air intake plenum; a combustor; a high pressure turbine; a power turbine. A forced air-stream generator is arranged in fluid communication with the air intake plenum. A shutter arrangement is provided in a combustion-air flow path, arranged and controlled to close the combustion-air flow path for pressurizing said air intake plenum by means of the forced air-stream generator to a pressure sufficient to cause pressurized air to flow through the aeroderivative air turbine.

Abstract: A system and method of operating an engine anti-ice system includes supplying heat from one or more heat sources to one or more components on or within a gas turbine engine, sensing data representative of one or more parameters related to ice crystal accretion, and based on the data, at least selectively inhibiting at least selected ones of the one or more heat sources from supplying heat to at least selected ones of the one or more components on or within the gas turbine engine.

Abstract: A method and a device are disclosed for the safe operation of a gas turbine plant, whose operation is both controlled by at least one process controller, which at least triggers and/or influences operationally relevant processes of the gas turbine plant, and is also monitored by a separate protection unit that is operated independently of the process controller on the basis of at least one first limit value for a safety-relevant operating parameter, wherein the gas turbine plant is subjected to an emergency switch-off once the at least one first limit value is exceeded. In that in the case of a defined transient operating state of the gas turbine plant that can be detected by the protection unit, the at least one first limit value of the safety-relevant operating parameter is raised to a second limit value, wherein the gas turbine plant is protected by an emergency switch-off of the gas turbine plant once said second limit value is exceeded.

Abstract: In a testing overspeed protection system: a) on receiving an order to start a turbomachine, an electronic regulation system (ERS) of the turbomachine sends an order to a control circuit of a fuel cutoff member to close the fuel cutoff member or to keep it in the closed position; b) the closed state of the FCM is verified on the basis of information transmitted to the ERS and representative of the position of the FCM; c) if the result of the verification in b) is positive, the ERS sends an order to the FCM control circuit to authorize opening of the FCM and enable the starting procedure to continue; and d) if the result of the verification in b) is negative, the ERS issues fault information concerning the overspeed protection system.

Abstract: A system is provided for detecting and controlling flashback and flame holding in a combustor of a gas turbine. The system includes at least one flame indicator disposed in a combustor and at least one detector disposed downstream from the flame indicator. The flame indicator may be configured to produce light when exposed to a flame and the detector may be configured to detect the light produced by the flame indicator.

Abstract: A combustion chamber (10, 20) for a gas turbine (1) having a housing (12, 23), a combustion zone (21, 22, 47) surrounded by the housing (12, 23), and at least one burner arrangement (11, 25), which has at least one pre-mixing passage (28) opening into the combustion zone (21, 22) for preparation of a fuel/air mixture. To enable suppression of combustion chamber pressure fluctuations, at least one interference body (50) is provided against which a flow can take place and to which a Strouhal number can be assigned, the flow generating fluidically induced sound waves. The interference body (50) is in a passage (48, 59) opening into the combustion zone (21, 22, 47). A fluid (51, 62) flowing through the passage (48, 59) in the direction of the combustion zone generates a sound wave with a frequency f by flowing against the interference body (50), wherein the frequency f substantially corresponds to a dominant frequency of the combustion chamber.

Abstract: The application relates to a method and a device for suppressing ice formation on intake structures of a compressor, particularly the compressor of a gas turbine. The technical aim of the present invention is to provide a method and a device for suppressing the formation of ice on said structures, which avoid the disadvantages of known solutions, such as a reduction of the performance of the gas turbine, and have a simple and broad applicability. According to the present invention the mechanical vibratory energy of said structures during operation is converted into electrical energy by a piezoelectric element, firmly applied to said structure, and in a connected electrical circuit the generated electrical energy is then converted into thermal energy by an ohmic resistor and this thermal energy is conducted to at least a portion of the structure for suppressing ice formation. Excess energy may be transmitted by a transmitter to other circuits in adjacent structures.