A minimum creep life location (MCLL) on a blade for a turbine blade design is received. A temperature at the MCLL on the blade is monitored. When the temperature at the MCLL exceeds a predetermined threshold, a cooling air supply is adjusted to lower the temperature below the threshold during engine operation.

A minimum creep life location (MCLL) on a blade for a turbine blade design is received. A temperature at the MCLL on the blade is monitored. When the temperature at the MCLL exceeds a predetermined threshold, a cooling air supply is adjusted to lower the temperature below the threshold during engine operation.

Embodiments of a direct-drive fan system and a variable process control system are disclosed herein. The direct-drive fan system and the variable process control system efficiently manage the operation of fans in a cooling system such as a wet-cooling tower or air-cooled heat exchanger (ACHE), HVAC systems, mechanical towers or chiller systems.

A gas turbine engine has a cycle operability parameter β in a defined range to achieve improved overall performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined range of cycle operability parameter β may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.

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.

An aircraft has first and second fuel sources containing fuels with different characteristics, and one or more gas turbine engines powered by the fuels and each having a staged combustion system having pilot and main fuel injectors and being operable in pilot-only and pilot-and-main ranges of operation. The gas turbine engines each have a fuel delivery regulator arranged to control fuel delivery to the pilot and main fuel injectors. The method includes: obtaining a proposed mission description; obtaining nvPM impact parameters for the gas turbine engines, the impact parameters being associated with each operating condition of the proposed mission; calculating an optimised set of one or more fuel characteristics for each flight condition of the proposed flight defined in the flight description based on the nvPM impact parameters; and determining a fuel allocation based on the optimised set of one or more fuel characteristics.

An aircraft has first and second fuel sources containing fuels with different characteristics, and one or more gas turbine engines powered by the fuels and each having a staged combustion system having pilot and main fuel injectors and being operable in pilot-only and pilot-and-main ranges of operation. The gas turbine engines each have a fuel delivery regulator arranged to control fuel delivery to the pilot and main fuel injectors. The method includes: obtaining a proposed mission description; obtaining nvPM impact parameters for the gas turbine engines, the impact parameters being associated with each operating condition of the proposed mission; calculating an optimised set of one or more fuel characteristics for each flight condition of the proposed flight defined in the flight description based on the nvPM impact parameters; and determining a fuel allocation based on the optimised set of one or more fuel characteristics.

A aircraft gas turbine engine and operation method, the engine including: a staged combustion system having pilot and main fuel injectors, and operates in a pilot-only range wherein fuel delivers to pilot fuel injectors, and a pilot-and-main operation range wherein fuel is delivered to at least the main fuel injectors. The engine further includes a fuel delivery regulator to pilot and main fuel injectors, which receives fuel from a first and second source containing fuels each with different characteristics. The staged combustion system switches between pilot-only and pilot-and-main range operation when in steady cruise mode, the mode defining a boundary between first and second engine cruise operation range. The fuel delivery regulator delivers fuel to pilot fuel injectors during at least part of the first engine cruise operation with different fuel characteristics from fuel delivered to one or both pilot and main fuel injectors the second engine cruise operation range.

A first turboexpander generator defines a portion of a first conduit flow passage. The first turboexpander generator is configured to decrease a temperature or pressure of a process stream flowing through the first turboexpander generator by generating electrical power from the process stream. A second turboexpander generator defines a portion of a second conduit flow passage. The second turboexpander generator is configured to decrease a temperature or pressure of a process stream flowing through the second turboexpander generator by generating electrical power from the process stream. The first and second conduit flow passages are arranged to carry fluid flow in parallel to one another. The first and the second turboexpander generator are substantially identical in critical dimensions and performance.

A gas turbine includes: multiple fuel systems; a combustor that combusts fuels from the multiple fuel systems in compressed air to generate combustion gas; and a turbine that is driven by the combustion gas. A flow volume ratio calculation device which calculates the flow volume ratio of the fuels flowing in the multiple fuel systems includes calculators that receive values of a first parameter and a second parameter capable of expressing the combustion state in the combustor, and that calculate the flow volume ratio relative to the received values of the two parameters from a predetermined relationship between the two parameters and the flow volume ratio.