A method for regulating the temperature of the exhaust gases of a turbomachine, the method including regulation of the injection of fuel into a combustion chamber of the turbomachine so that the turbomachine generates a target thrust; regulation of the injection of mechanical power by an electric motor onto a shaft driven in rotation by a turbine, the electric motor being activated when a clearance between a casing and the blades of the turbine exceeds a threshold value.

An engine control system may be configured to perform a method of controlling thermal bias in a turbomachine. An exemplary method may include determining a thermal bias-value for the turbomachine, and performing a cooling treatment based at least in part on the thermal bias-value. The thermal bias-value may include a difference between an upward temperature-value corresponding to a first one or more temperature measurements of an upward portion of the turbomachine and a downward temperature-value corresponding to a second one or more temperature measurements of a downward portion of the turbomachine. The cooling treatment may include at least one of: circulating air through at least a portion of the turbomachine, and rotating a shaft of the turbomachine with a motoring system.

A gas turbine engine has a compressor section and a turbine section. A secondary cooling air includes a first fluid connection to tap cooling air and pass the cooling air through a plurality of tubes, and a second fluid connection for returning air from the tubes back to at least one of the compressor and turbine for cooling. A sensor senses a condition of the cooling air downstream of the tubes and a control compares the sensed condition of the cooling air to an expected condition, and to identify a potential concern in the cooling air system should the sensed condition differ from the expected condition by more than a predetermined amount.

The invention relates to an aircraft propulsion system, comprising a main transmission unit (12) and at least two turbojet engines connected to the main transmission unit (12), respectively a first turbojet engine (14a) and a second turbojet engine (14b), each turbojet engine comprising a free turbine (24a, 24b), characterized in that the first turbojet engine (14a) comprises a heat exchanger (30) configured to recover some of the thermal energy from the exhaust gas at the outlet of the free turbine, and in that the propulsion system comprises at least one computer (28a, 28b) configured to control the two turbojet engines and to limit the acceleration and the deceleration of the first turbojet engine (14a) when neither of the turbojet engines is broken down, in order to limit the reactor power transients at the heat exchanger (30).

A first fuel flow rate command value indicating a command value of a fuel input amount is calculated so that an output of a gas turbine matches a target output. An upper limit value of the first fuel flow rate command value is calculated based on a deviation obtained by subtracting, from an estimated value of a turbine inlet temperature of the gas turbine, a second limit value relating to the estimated value set such that the estimated value does not exceed a first limit value of the turbine inlet temperature.

An operational support device sets an execution time of overfiring serving as an operation of a power generation facility at an output higher than a rated output. The device includes a life index value acquisition unit that acquires a life index value at a start time, the life index value being an index indicating a life of the power generation facility and changing in value in one direction with the output of the power generation facility; an output pattern setting unit that sets an output pattern per unit time of the power generation facility from the start time to a stop time based on the life index value such that the life index value reaches a predetermined value; and an overfiring setting unit that sets, based on the output pattern, a time in a period from the start to the stop time at which the overfiring is to be performed.

A method is provided of controlling a cooled cooling air system for an aeronautical gas turbine engine. The method includes: receiving data indicative of an ambient condition of the aeronautical gas turbine engine, data indicative of a deterioration parameter of the aeronautical gas turbine engine, data indicative of an operating condition of the aeronautical gas turbine engine, or a combination thereof; and modifying a cooling capacity of the cooled cooling air system in response to the received data indicative of the ambient condition of the aeronautical gas turbine engine, data indicative of the deterioration parameter of the aeronautical gas turbine engine, data indicative of an operating condition of the aeronautical gas turbine engine, or the combination thereof.

A method for starting and stopping a gas turbine in a combined cycle power plant, wherein the gas turbine includes a compressor having adjustable guide vanes and the gas turbine power can also be controlled by opening the guide vanes. When the gas turbine is started, it is driven up to a base load or up to an emission-compliant load point, and the guide vanes are opened before the base load or the emission-compliant load point is reached.

A full authority digital engine controller (FADEC) based system is also disclosed. The system includes a processor, and a tangible, non-transitory memory configured to communicate with the processor, the tangible, non-transitory memory having instructions stored thereon that, in response to execution by the processor, cause the FADEC to perform operations. The operations may include measuring a first temperature at a first sensor disposed at a first known location of an engine, measuring a second temperature at a second sensor disposed at a second known location of the engine, and estimating at least one of a stress or a strain of a part or component in the engine based on the first temperature and the second temperature. The system may control fuel flow and/or other engine effectors during a thrust transient to limit the estimated stress or the estimated strain of the component from exceeding a predetermined threshold.

Herein provided are methods and systems for detecting a high temperature condition of a gas turbine engine. A fuel flow to a combustor of the engine and a compressor outlet pressure of the engine are obtained. A ratio of the fuel flow to the compressor outlet pressure is determined. The ratio is compared to a threshold and a high temperature condition of the engine is detected when the ratio exceeds the threshold.