A flow rate per unit time of fuel fed to a gas turbine is calculated. A flow rate per unit time of air fed to the gas turbine is calculated. A turbine inlet temperature is calculated through use of a physical model formula expressing a relationship of input and output of thermal energy relating to a combustor of the gas turbine. A fuel distribution ratio for each of a plurality of fuel supply systems connected to the combustor is calculated based on the turbine inlet temperature.

A control system for a gas turbine includes a controller. The controller includes a processor configured to receive a plurality of signals comprising a temperature signal, a pressure signal, a speed signal, a mass flow signal, or a combination thereof, from sensors disposed in the gas turbine system. The processor is further configured to apply the plurality of signals as input to a heating value model. The processor is also configured to execute the heating value model to derive a heating value for a fuel combusted by the gas turbine system. The processor is additionally configured to control operations of the gas turbine system based on the heating value for the fuel.

A method of operating an aircraft that includes a propulsion system. The propulsion system includes a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine. The method includes: determining at least one fuel characteristic of the fuel arranged to be provided to the gas turbine engine; and controlling the propulsion system based on the at least one fuel characteristic.

A loading/unloading method for a gas turbine system is disclosed. The gas turbine system includes a combustion section featuring a primary combustion stage with a first plurality of fuel nozzles and a downstream, secondary combustion stage with a second plurality of fuel nozzles. For loading, the method progresses through each of a plurality of progressive combustion modes that sequentially activate a higher number of at least one of the first or second plurality of fuel nozzles; and for unloading, the method progresses through each of a plurality of progressive combustion modes that sequentially activate a lower number of at least one of the first or second plurality of fuel nozzles. During each combustion mode, regardless of whether loading or unloading, a primary combustion stage exit temperature of a combustion gas flow is controlled to be within a predefined target range corresponding to the respective combustion mode.

A system for a gas turbine engine includes an igniter event conductor. The igniter event conductor has a first portion adjacent an igniter of the engine and in a combustion chamber of the engine and a second portion apart from the igniter and apart from the combustion chamber. The conductor is adapted to conduct an aspect of an igniter event at the igniter from the first portion to the second portion. A sensor is coupled to the second portion of the conductor to sense the aspect of the igniter event.

A method of operating a control unit for controlling at least two different input fuel flows to a combustion device, e.g. a gas turbine includes the step of determining on the basis of at least one operating parameter whether the combustion device is in a predefined operating stage. In response hereto, generating a control signal configured for setting a ratio of at least two different input fuel flows to a predetermined value (psc1, psc3) for a predetermined time (dt) in case the combustion device is in the predefined operating stage.

A multivariable fuel control and estimator (MFCE) of a gas turbine engine for preventing combustor blowout is provided. The MFCE includes a first input port that receives controller requests and provide system usage commands, a second input port that receives measured disturbance values, a third input that receives system and component limits, a fourth input port that receives sensed parameters, a fuel system model of the fuel system of the gas turbine engine and an engine model of the engine system that includes the combustor of the gas turbine engine, a processor that generates a control signal for controlling the fuel valve and generates a control signal for controlling the actuator using the fuel system and engine model based on the controller requests, the measured disturbance values, the system and component limits, and the sensed parameters, and an output port that transmits the control signals to the fuel system.

A controller for a gas turbine is arranged to supply a load L. The gas turbine includes a fuel supply arranged to supply fuel at a fuel flow rate FF to a combustor, wherein the fuel supply includes a first fuel supply and a second fuel supply. The controller is arranged to determine one or more ratios R of one or more combustor operating parameters COP respectively at the load L to respective reference combustor operating parameters COPR at a reference load LR. The controller is further arranged to control a proportion P of the fuel flow rate FF supplied via the first fuel supply based, at least in part, on the determined one or more ratios R. A gas turbine with the controller and a method controls the gas turbine.

The present invention relates to a method for operating a lean premix burner of an aircraft gas turbine, where fuel and primary supporting air are supplied by means of a supporting burner (pilot burner) arranged centrically to the burner axis, where secondary air surrounding the supporting burner is supplied, and where fuel and air are supplied by means of a main burner, characterized in that the primary supporting air is supplied in an amount of 5 vol % to 10 vol % of the total air quantity, that the secondary supporting air is supplied in an amount of 5 vol % to 20 vol % and that 35 vol % to 75 vol % of the total air quantity are supplied via the main burner in the partial load range and in the full load range.

In a gas turbine combustor 3 of the present invention, an air hole plate 20 includes a center air hole group 51 configured from a plurality of air holes 51A and 51B and a plurality of outer circumferential air hole groups 52 configured from a plurality of air holes 52A, 52B, and 52C and formed to surround the center air hole group 51. The gas turbine combustor 3 includes a hole part 601 and a temperature sensor 401 provided on the air hole plate 20 to be located in a region surrounded by two outer circumferential air hole groups 52 adjacent to each other and the center air hole group 51, a supply source 220 of coolant, a cooling pipeline 205 that connects the hole part 601 and the supply source 220, valves 67 and 68 provided in the cooling pipeline 205, and a control system 500 that drives the valves 67 and 68 on the basis of a measured value of the temperature sensor 401.