A method for controlling a gas turbine assembly includes: a compressor in which compression of the outside air occurs for producing a flow of compressed air; a sequential combustor including a first combustor, in which combustion of a mixture of fuel and compressed air arriving from the compressor occurs for producing a flow of hot gasses, and a second combustor which is located downstream of the first combustor and in which combustion of a mixture of fuel and hot gasses arriving from the first combustor occurs; an intermediate turbine in which a partial expansion of the hot gasses arriving from the first combustor occurs; and a second combustor in which combustion of a mixture of fuel and hot gasses arriving from the intermediate turbine occurs; the method further includes, on a start-up transient operating phase of the gas turbine assembly, the step of controlling the fuel mass flow-rate supplied to the first and/or the second combustor on the basis of the flame temperature inside the first combustor.

A cooling system includes: a high pressure bleed line configured to bleed high pressure compressed air from a first bleed position of a compressor and to send the air to a first hot part; a low pressure bleed line configured to bleed low pressure compressed air from a second bleed position of the compressor and to send the air to a second hot part; an orifice provided in the low pressure bleed line; a connecting line configured to connect the high pressure bleed line and the low pressure bleed line; a first valve provided in the connecting line; a bypass line configured to connect the connecting line and the low pressure bleed line; and a second valve provided in the bypass line.

A system for operating a gas turbine engine to mitigate the risk of ice formation within the engine, the system including a controller arranged to control at least one operational parameter of the engine such that the engine operates in a safe zone; and, a processor configured to function as a determining module to make a comparison between values and determine whether the engine is operating within a safe zone based on at least a core pressure parameter relating to the pressure within the engine and a core temperature parameter relating to the temperature within the engine, wherein the safe zone is defined by the product (multiplied) of the core pressure parameter and core temperature parameter being above a safe threshold.

One example aspect of the present disclosure is directed to a method for measuring engine performance. The method includes receiving first parameters related to engine performance prior to an engine wash event. The method includes receiving second parameters related to engine performance after the engine wash event. The method includes determining an engine performance prior to the engine wash event based on the first parameters. The method includes determining an engine performance after the engine wash event based on the second parameters. The method includes determining an effectiveness of the engine wash event based on the engine performance prior to the engine wash event and the engine performance after the engine wash event.

In one embodiment, a plant includes a combustor to burn fuel with oxygen from an inlet guide vane (IGV) to generate a gas for a gas turbine (GT), and a heat recovery steam generator to use an exhaust gas from GT to generate steam for a steam turbine (ST). An apparatus controls an IGV opening degree to a first degree and a GT output value to a value larger than a first value between GT start and ST start. The first value is an output value at which exhaust gas temperature can be kept at a first temperature that depends on ST metal temperature, when the IGV opening degree is the first degree. The apparatus increases the IGV opening degree from the first degree based on steam temperature or the GT output value, while the GT output value is controlled to the value larger than the first value.

The present application discloses a method of determining one or more fuel characteristics of an aviation fuel used for powering a gas turbine engine of an aircraft. The method comprises: determining one or more performance parameters of the gas turbine engine during a first time period of operation of the gas turbine engine; and determining one or more fuel characteristics of the fuel based on the one or more performance parameters. A method of operating an aircraft, a fuel characteristic determination system, and an aircraft are also disclosed.

A fuel conditioning and control system provides dynamic control and steady state operations of a gas turbine provided fueled by supercritical liquefied petroleum gas (LPG). The fuel conditioning and control system comprises a storage for LPG fuel; a fuel delivery sub-system connecting the storage to turbomachinery; and a control system. The gas turbine includes a gas turbine core control that provides at least one operational data of the gas turbine to the control system. The fuel delivery sub-system includes at least one sensor for sensing at least one property of the LPG fuel in the fuel delivery sub-system, where the at least one sensor providing data on the at least one property of the LPG fuel to the control system. The control system analyzes the data on the at least one property of the LPG fuel and at least one operational data of the gas turbine for dynamic control of LPG fuel to the gas turbine under dynamic and steady state conditions.

Cantilevered airfoils and methods of forming the same are disclosed herein. An example airfoil disclosed herein includes an airfoil including an airfoil body including a first face and a second face, a first recessed portion formed in the first face and internal temperature-regulating features and a first insert disposed within the first recessed portion, the first insert causing the airfoil body to assume a first predefined curvature profile at a first temperature, the first insert causing the airfoil body to assume a second predefined curvature profile at a second temperature.

A method of maintaining at least one gas turbine engine includes monitoring a compressor of the gas turbine engine. The compressor includes a compressor case at least partially defining a flow path, a plurality of stages and a vane actuator system configured to move at least one of the stages. The vane actuator system includes a vane mover having one or more slots formed therein and configured to actuate the at least one stage. The vane mover may be replaced after the gas turbine engine has experienced engine degradation.

A motor driving device for driving a motor is provided. The motor driving device includes: a driver outputting a drive signal for driving the motor to the motor; and a controller controlling a duty ratio of the drive signal based on three control parameters which are an initial duty ratio, a duty ratio increasing speed, and a target duty ratio. The controller has a normal control mode in which each of the three control parameters is a predetermined value, and a cryogenic control mode in which at least one of the three control parameters is a smaller value than in the normal control mode.