The gas turbine includes a compressor to compress air introduced thereinto, a combustor to mix the compressed air with fuel for combustion, a main turbine having a plurality of turbine blades rotated by an energy produced by combustion gas in the combustor, a heat recovery boiler to produce steam by heat exchange with the combustion gas, and a fluid accelerator supplied with a first fluid compressed in the compressor to compress the first fluid and supply the compressed first fluid to the combustor, where the fluid accelerator includes a first inlet through which the first fluid is introduced, a second inlet through which a second fluid having a higher pressure than the first fluid is introduced, and an outlet through which the first and second fluids are mixed and discharged.

The premixed pilot nozzle includes axially elongated tubes defined within a plenum between an outer shroud and a first shroud disposed radially inward of the outer shroud. The tubes extend between tube inlets defined through a forward face and tube outlets defined through an aft face. A second shroud is disposed radially inward of the first shroud, thereby defining a fuel plenum between the first shroud and the second shroud, and the fuel plenum is in communication with a gaseous fuel supply. A fuel injection port, which is positioned between the tube inlet and the tube outlet of each tube, is in fluid communication with the fuel plenum. An air supply configured to fluidly communicate with the tube inlet of each tube. The second shroud defines a second plenum therein, the second plenum being coupled to a source of a non-combustible fluid.

A gas turbine engine that includes an inlet volume flow control appliance and methods of operating the same are provided. The method includes operating the gas turbine engine with the inlet volume flow control appliance supplying a compressor inlet volume flow that is below a maximum compressor inlet volume flow. A mass flow of a liquid agent is added to a compressor gas mass flow while the gas turbine engine is operated with a compressor inlet volume flow below a maximum compressor inlet volume flow. The mass flow of a liquid agent may be controlled as a function of the pitch of variable inlet guide vanes. The method further comprises adjusting the volume flow control appliance to increase the compressor inlet volume flow and increasing the mass flow of liquid agent added to the compressor gas mass flow while the inlet volume flow control appliance increases the compressor inlet volume.

Aircraft engines include a turbine engine comprising a compressor section, a burner section, and a turbine section arranged along a shaft, with a core flow path through the turbine engine such that exhaust from the burner section passes through the turbine section, a condensing assembly arranged downstream of the turbine section of the turbine engine along the core flow path, and an exhaust compressor arranged downstream of the condensing assembly along the core flow path. The condensing assembly is configured to reduce a mass flow of the exhaust compressor by condensing water vapor from the core flow and removing liquid water from the core flow.

A gas turbine engine includes a compressor section having a high pressure compressor and a core casing surrounding the compressor section and defining an inlet. The gas turbine engine also includes a cooling system for cooling air in or to the compressor section. The cooling system includes a fluid tank for storing a volume of cooling fluid and a fluid line assembly in fluid communication with the fluid tank. The fluid line assembly includes an outlet positioned upstream of the high pressure compressor and downstream of the inlet defined by the core casing for injecting cooling fluid into an airflow upstream of the high pressure compressor.

Propulsion systems for aircraft include a fan and a low pressure turbine operably coupled to a first shaft, a low pressure compressor and an intermediate pressure turbine operably coupled to a second shaft, and a high pressure compressor and a high pressure turbine operably coupled to a third shaft. A burner is arranged between the high pressure compressor and the high pressure turbine, with a main flow path defined through the propulsion system. A hydrogen fuel system is configured to supply hydrogen fuel to the burner. A condenser is arranged along the main flow path and configured to extract water from exhaust from the burner. An evaporator is arranged along the main flow path and configured to receive a portion of the water to generate steam which is injected into the main flow path upstream from the evaporator.

A method is provided during which an aircraft powerplant is provided. The aircraft powerplant includes a combustor and a water recovery system. The water recovery system includes a condenser and a reservoir. Fuel is combusted within the combustor to provide combustion products. Water is extracted from the combustion products using the condenser. The water recovery system is operated in one of a plurality of modes based on likelihood of contrail formation. The modes include a first mode and a second mode, where the water is collected within the reservoir during the first mode, and where the water passes through the water recovery system during the second mode.

A propulsion system for an aircraft includes a hydrogen fuel system supplying hydrogen fuel to the combustor through a fuel flow path. A condenser extracts water from an exhaust gas flow and includes a plurality of spiral passages disposed within a collector. The spiraling passages generate a transverse pressure gradient to direct water out of the exhaust gas flow toward the collector.

A device for temporarily increasing power in order to increase the power from at least one first turbine engine and from at least one second turbine engine, the device including a tank of coolant liquid, a first injection circuit connected to the tank and leading to at least one injection nozzle configured to be installed upstream from the first turbine engine, a second injection circuit connected to the tank and leading to at least one injection nozzle configured to be installed upstream from the second turbine engine, each of the first and second injection circuits including at least one first valve and at least one second valve arranged upstream from said at least one first valve, and a bridge pipe connecting together the first injection circuit and the second injection circuit upstream from their respective first valves and downstream from their respective second valves.

The present disclosure relates to an exhaust gas cooling device and method, and more particularly, to a device and method for installing an exhaust gas cooling device on the upper end of a duct of a heat recovery steam generator to cheaply cool the exhaust gas without occupying an additional dedicated area. An object of the present disclosure is to reduce the costs using a cheap cooling device in the cooling path for cooling the exhaust gas. In one aspect, the exhaust gas cooling device includes an exhaust gas cooling unit located on the upper end of a duct of a heat recovery steam generator connected with a gas turbine and for cooling the exhaust gas discharged from the gas turbine; and a control unit for controlling the exhaust gas cooling unit to lower the increase rate of the energy of the exhaust gas flowed into the heat recovery steam generator through the duct.