The injector device comprises an elongated body with a leading edge and a trailing edge, gas nozzles and oil nozzles, an oil supply duct housed within the elongated body and connected to the oil nozzles, a gas supply duct housed within the elongated body and connected to the gas nozzles. The oil supply duct is connected to the gas supply duct only between one or more oil nozzles and one gas nozzles, and the gas supply duct is connected to the elongated body only via bridges.

The invention concerns a method for a part load CO reduction operation and a low-CO emissions operation of a gas turbine with sequential combustion. The gas turbine essentially includes at least one compressor, a first combustor which is connected downstream to the compressor. The hot gases of the first combustor are admitted at least to an intermediate turbine or directly or indirectly to a second combustor. The hot gases of the second combustor are admitted to a further turbine or directly or indirectly to an energy recovery. At least one combustor runs under a caloric combustion path having a can-architecture, and wherein the air ratio () of the combustion at least of the second combustor is kept below a maximum air ratio (.sub.max).

An aircraft propulsion system and computing system are provided. The propulsion system includes a low pressure (LP) spool and a core engine having a high pressure (HP) spool. A frame is positioned in serial flow arrangement between an HP turbine and an LP turbine. The frame includes an inter-turbine burner including a strut forming an outlet opening into a core flowpath of the propulsion system. A first fuel system is configured to flow a liquid fuel to a combustion section for generating first combustion gases. A second fuel system is configured to flow a gaseous fuel to the core flowpath via the inter-turbine burner for generating second combustion gases. The propulsion system forms a rated power output ratio of the core engine and the inter-turbine burner with the LP spool between 1.5 and 5.7.

An engine having a compressor for generating a flow of pressurized oxidizer, a fuel mixing system in fluid communication with the compressor for mixing fuel with the pressurized oxidizer creating a fuel-oxidizer mixture, a combustion chamber adapted to receive the fuel-oxidizer mixture, at least one ignition system connected to the combustion chamber for igniting the fuel-oxidizer mixture inside of the combustion chamber, an exhaust port in fluid communication with the combustion chamber for receiving exhaust generated by combustion of the fuel-oxidizer mixture, and a turbine having a rotating shaft and a plurality of turbine blades connected downstream of the combustion chamber for receiving the exhaust whereby the fluid force of the exhaust through the exhaust port causes the turbine blades to rotate the shaft.

A hybrid aircraft propulsion system. The system comprises a gas turbine engine having a turbine system comprising first and second turbine sections. The gas turbine engine further comprises a first combustor provided upstream in core flow from the turbine system, and a second combustor provided between the first and second turbine sections. The hybrid aircraft propulsion system further comprises an energy storage system and an electric propulsion system electrically coupled to the energy storage system.

A jet engine is provided having a front compression stage for pressurizing intake air, and a central superheating section for further increasing the temperature and pressure of the pressurized intake air from the front compression stage. The central superheating section is at least partially electrically powered by an internal energy generating power source. A rear exhaust nozzle stage recovers energy from a discharge of the superheating section and creates thrust.

A gas turbine engine may include a high pressure compressor coupled to a high pressure turbine by a high pressure shaft, a core combustor located downstream of the high pressure compressor and upstream of the high pressure turbine, and a low pressure compressor provided upstream of the high pressure compressor. The low pressure compressor may be configured to direct core airflow to the high pressure compressor and first bypass airflow which bypasses the high pressure compressor, core combustor and high pressure turbine through a first bypass duct. The engine may further include a mixer downstream of the high pressure turbine and low pressure compressor, the mixer being configured to mix the core and first bypass airflows. The engine also may include a re-heat combustor configured to combust fuel with both core airflow and first bypass airflow. A low pressure turbine may be provided downstream of the re-heat combustor and coupled to the low pressure compressor (14) by a low pressure shaft, the low pressure and high pressure shafts being independently rotatable. A shaft power transfer arrangement may be provided, which is configured to selectively transfer power between the low pressure and high pressure shafts.

Systems and methods of augmenting the thrust of the prime power engine(s) of an aircraft from a tank of compressed gas are described herein.

A gas turbine system has a source of ammonia and a source of an oxygen-containing gas, a first combustion chamber connected to receive ammonia, a hydrogen-rich gas stream and oxygen-containing gas, a turbine connected to receive an exhaust gas stream from the first combustion chamber; and a second combustion chamber connected to receive an exhaust gas from the turbine, ammonia and a hydrogen-rich gas stream.

In an embodiment, a method includes flowing an exhaust gas from a turbine of a gas turbine system to an exhaust gas compressor of the gas turbine system via an exhaust recirculation path; evaluating moist flow parameters of the exhaust gas within an inlet section of the exhaust gas compressor using a controller comprising non-transitory media programmed with instructions and one or more processors configured to execute the instructions; and modulating cooling of the exhaust gas within the exhaust recirculation path, heating of the exhaust gas within the inlet section of the exhaust gas compressor, or both, based on the evaluation.