A gas turbine engine includes a first spool, a second spool, a primary combustor, and a diffuser. The first spool includes a first compressor rotationally driven by a first turbine via a first shaft. The second spool includes a second compressor driven by a second turbine via a second shaft. The first compressor, the diffuser, and the primary combustor are arranged in series to provide a compressed airflow discharged from the first compressor to the primary combustor via the diffuser, which includes walls that diverge towards the primary combustor. The second compressor is fluidly coupled to the diffuser to receive at least a portion of the compressed airflow from the diffuser. The second turbine is fluidly coupled to the diffuser to discharge an expanded airflow to the diffuser.

A hybrid multi-spool gas turbine engine, has: a LP spool and a HP spool rotatable about a central axis, the LP spool having an LP compressor and an LP turbine engaged to the LP compressor via an LP shaft, the LP shaft engaged to a rotatable load at a first end thereof, the HP spool having an HP turbine and an HP compressor engaged to the HP turbine via a HP shaft; an accessory gearbox (AGB) engaged to both of the LP shaft and the HP shaft and located proximate a second end thereof, the AGB having at least one accessory output drivingly engageable to at least one accessory and at least one input; and at least one electric motor drivingly engaged to the at least one input of the AGB, the at least one electric motor drivingly engaged to the rotatable load via the AGB and the LP shaft.

There is described a method and system for in-flight start of an engine. The method comprises rotating a propeller; generating electrical power at an electric generator embedded inside a propeller hub from rotation of the propeller; transmitting the electrical power from the electric generator to an engine starter mounted on a core of the engine via an electric power link; and driving the engine with the engine starter to a sufficient speed while providing fuel to a combustor to light the engine to achieve self-sustaining operation of the engine.

A gas turbine engine has a first spool having a low pressure compressor section in fluid communication with an air inlet, the low pressure compressor section including a first plurality of variable guide vanes therein, and a low pressure turbine section drivingly engaged to the low pressure compressor section. A second spool has a high pressure compressor section in fluid communication with the low pressure compressor section to receive pressurized air therefrom, the high pressure compressor section including a second plurality of variable guide vanes at an entry thereof, and a high pressure turbine section drivingly engaged to the high pressure compressor section, the high pressure turbine section disposed upstream of the low pressure turbine section and in fluid communication therewith. An output drive shaft drivingly engages the low pressure turbine section and is adapted to drivingly engage a rotatable load of the gas turbine engine.

A reverse flow gas turbine engine has a low pressure (LP) spool and a high pressure (HP) spool arranged sequentially in an axial direction. The LP spool comprises an LP compressor disposed forward of an LP turbine and drivingly connected thereto via an LP compressor gear train. The HP spool comprises an HP compressor in flow communication with the LP compressor, and an HP turbine disposed forward of the HP compressor and drivingly connected thereto via an HP shaft.

A turbine engine includes, in a serial flow relationship, an axially oriented compressor, a combustion section, a turbine section, and an exhaust section. An air flowpath extends from an inlet duct to the exhaust section such that the compressor, combustion section, turbine, and the exhaust section are in fluid communication. The inlet duct is positioned upstream of the compressor and defines an inlet portion of the air flowpath. The inlet duct is generally radially oriented. A variable inlet guide vane extends at least partially through the inlet duct.

An apparatus and method for a reverse flow combustor, the reverse flow combustor including a straight portion, a dilution portion and a curved portion. The reverse flow combustor receives a flow of fuel that is ignited and mixed with cooling air to form a flow of combustion gases. The flow of combustion gases travels through the reverse flow combustor to a turbine section of an engine.

A shrouded impeller and a method of additively manufacturing the same are provided. In one example aspect, the shrouded impeller includes a hub and a shroud spaced from the hub. The shrouded impeller also includes a plurality of vanes extending between and connecting the hub and the shroud. The vanes are spaced circumferentially apart from one another. Flow passages are defined between adjacent vanes. In some implementations, the shrouded impeller is additively manufactured. During printing, one or more support structures are formed within and fill a portion of one or more of the flow passages to support the unsupported walls of the shrouded impeller, e.g., the shroud. Further, the support structures are removable from the shrouded impeller, e.g., after the shrouded impeller has been printed.

A multi-spool gas turbine engine comprises a low pressure (LP) spool and a high pressure (HP) spool independently rotatable about a central axis extending through an accessory gear box (AGB). The LP spool has an LP compressor, which is axially positioned between the HP compressor of the HP spool and the AGB. A tower shaft drivingly connects the HP spool to the AGB.

A gas turbine engine has an engine case housing a low pressure compressor drivingly connected to a low pressure turbine by a low pressure compressor shaft extending along an engine axis. The low pressure turbine is disposed forward of the low pressure compressor. A low pressure turbine shaft is drivingly connected to the low pressure turbine and extends forwardly of the low pressure turbine. A reduction gear box (RGB) is drivingly connected to the low pressure turbine shaft. The RGB is offset from the engine axis to free an access to low pressure compressor shaft connection. The offset positioning of the RGB allows to provide an access port in an axially forwardly facing surface of the engine case to access the low pressure compressor shaft and more particularly a connection thereof to the low pressure turbine.