A core engine article includes a combustor liner defining a combustion chamber therein and a turbine nozzle. The combustor liner includes a plurality of injector ports, and the plurality of injector ports have a shape that tapers to a corner on a forward side of the injector ports. The turbine nozzle includes a plurality of airfoils. The combustor liner and turbine nozzle are integral with one another. A method of making a core engine article is also disclosed.

A method for starting and operating a plant for the liquefaction of a gaseous product comprising the steps of electrically connecting a variable frequency drive to a motor of a first machine string; increasing the speed of the motor of the first machine string up until a first predefined threshold; electrically disconnecting the variable frequency drive from the motor of the first machine string; electrically connecting the variable frequency drive to a motor of a second machine string; the first predefined threshold is function of said frequency of the power supply grid. The variable frequency drive can be switched during operation of the plant among the strings according to process requirements.

An exoskeletal gas turbine engine having a rotatable outer shaft and an inner stationary case enclosed in a casing. The engine comprises a compressor section at an inlet end, a combustor section, and a turbine section at an outlet end. Rotating compressor blades and turbine blades are attached to, and extend radially inward from, an inner surface of the outer shaft. Stationary vanes are attached, and extend radially outward from, an outer surface of the inner stationary case. The outer shaft rotates around a front bearing and a rear bearing. An inlet compressor blade arrangement is attached to the outer race of the front bearing. An outlet turbine blade arrangement is attached to the outer race of the rear bearing. The inner race of the front and rear bearings attach to the inner stationary case.

A micro-combustion device generating electrical power raises global performance of the system, is compact, and reduces losses by utilizing an induced helical path. The device includes: injection ducts inserting a combustion agent, a fuel and/or a mixture thereof wherein the injection of the combustion agent takes place tangentially to the internal cylindrical wall, inducing a helical combustion path, the internal cylindrical walls of the chamber having a deposition of catalytic material to accelerate the combustion reaction; a turbo compressor group, including a compressor, feeding under pressure the combustion chamber through the injection ducts, and a turbine, receiving the flue gases from the discharge duct, compressor and turbine being keyed on the same axis, whereon a generator of electrical power, in turn, is keyed; and a fuel cell, fed by the flue gases through the turbine and by an oxidizing agent, implementing an electrochemical process generating additional electrical power.

A gas turbine engine according to an example of the present disclosure includes a fan situated at an inlet of a bypass passage. The fan has a fan diameter, Dfan. A low pressure turbine section is configured to drive the fan and a first compressor section. The low pressure turbine section has a greater number of stages than the first compressor section. The low pressure turbine section has a maximum rotor diameter, Dturb. A ratio of the maximum rotor diameter Dturb divided by the fan diameter Dfan is less than 0.6.

A gas turbine engine according to an example of the present disclosure includes a fan situated at an inlet of a bypass passage. The fan has a fan diameter, Dfan. A low pressure turbine section is configured to drive the fan and a first compressor section. The low pressure turbine section has a greater number of stages than the first compressor section. The low pressure turbine section has a maximum rotor diameter, Dturb. A ratio of the maximum rotor diameter Dturb divided by the fan diameter Dfan is less than 0.6.

A gas turbine engine includes an engine core includes at least one compressor, a combustor downstream of the compressor, and at least one turbine downstream of the combustor. A primary flowpath fluidly connects each of the compressor, the combustor, and the turbine. At least one particle extraction duct has an extraction duct inlet connected to the primary flowpath fore of the compressor and an extraction duct outlet connected to a bypass flowpath.

A gas turbine engine includes an engine core includes at least one compressor, a combustor downstream of the compressor, and at least one turbine downstream of the combustor. A primary flowpath fluidly connects each of the compressor, the combustor, and the turbine. At least one particle extraction duct has an extraction duct inlet connected to the primary flowpath fore of the compressor and an extraction duct outlet connected to a bypass flowpath.

A gas turbine engine includes a fan and a compressor section that is fluidly connected to the fan. The compressor includes a high pressure compressor and a low pressure compressor. A combustor is fluidly connected to the compressor section. A turbine section is fluidly connected to the combustor. The turbine section includes a high pressure turbine coupled to the high pressure compressor via a first shaft. A low pressure turbine is coupled to the low pressure compressor via a second shaft. A geared architecture interconnects between the second shaft and the fan. The gas turbine engine is a high bypass geared aircraft engine having a bypass ratio of greater than six (6). The low pressure turbine has a pressure ratio that is greater than 5, and the geared architecture includes a gear reduction ratio of greater than 2.5:1.

A gas turbine engine includes a fan and a compressor section that is fluidly connected to the fan. The compressor includes a high pressure compressor and a low pressure compressor. A combustor is fluidly connected to the compressor section. A turbine section is fluidly connected to the combustor. The turbine section includes a high pressure turbine coupled to the high pressure compressor via a first shaft. A low pressure turbine is coupled to the low pressure compressor via a second shaft. A geared architecture interconnects between the second shaft and the fan. The gas turbine engine is a high bypass geared aircraft engine having a bypass ratio of greater than six (6). The low pressure turbine has a pressure ratio that is greater than 5, and the geared architecture includes a gear reduction ratio of greater than 2.5:1.