The invention relates to an architecture of a propulsion system of a multi-engine helicopter comprising turboshaft engines connected to a power transmission gearbox, characterised in that it comprises: at least one hybrid turboshaft engine (20) capable of operating in at least one standby mode during a stable cruise flight of the helicopter; at least two systems (30; 40) for controlling each hybrid turboshaft engine (20), each system (30; 40) comprising an electric machine (31; 41) connected to the hybrid turboshaft engine (20) and suitable for rotating the gas generator thereof, and at least one source (33; 43) of electrical power for said electric machine (31; 41), each reactivation system (30; 40) being configured such that it can drive said turboshaft engine (20) in at least one operating mode among a plurality of predetermined modes.

A gas turbine engine for an aircraft includes an engine core including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and including a plurality of fan blades extending from a hub. A turbine to fan tip temperature change ratio of a low pressure turbine temperature change to a fan tip temperature rise is in the range from 1.46 to 2.0.

A gas turbine engine for an aircraft includes an engine core including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and including a plurality of fan blades extending from a hub. A turbine to fan tip temperature change ratio of a low pressure turbine temperature change to a fan tip temperature rise is in the range from 1.46 to 2.0.

An aspect includes a system for pre-start motoring control for multiple engines of an aircraft. The system includes a first engine starting system of a first engine and a controller. The controller is operable to control a motoring time of the first engine starting system relative to one or more other engine starting systems of one or more other engines of the aircraft by adjusting the motoring time of the first engine starting system within a tolerance of the motoring time of the one or more other engine starting systems in a pre-start motoring sequence.

A gas turbine engine for an aircraft includes an engine core including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and including a plurality of fan blades extending from a hub. A low pressure turbine temperature change is defined as: $\frac{\mathrm{the}\mathrm{first}\mathrm{turbine}\mathrm{entrance}\mathrm{temperature}}{\mathrm{the}\mathrm{first}\mathrm{turbine}\mathrm{exit}\mathrm{temperature}}.$
A fan tip temperature rise is defined as: $ the ⁢ ⁢ fan ⁢ ⁢ tip ⁢ ⁢ rotor ⁢ ⁢ exit ⁢ ⁢ temperature the ⁢ ⁢ Engine assembly with a dedicated voltage bus 11124311 · 2021-09-21 · Pratt & Whitney Canada Corp. · Jean Thomassin, Eric Taillon There is provided an engine assembly for an aircraft, comprising a heat engine in driving engagement with an engine shaft having a first end coupled to a mechanical gearbox and a second end opposite the first end, an electric generator coupled to the second end to provide a generator output voltage, the electric generator separate from the mechanical gearbox, a power electronics module connected to the electric generator and configured to regulate the generator output voltage to provide a regulated output voltage that meets an electrical power demand of at least one aircraft accessory, and a voltage bus connected to the power electronics module and configured to supply the regulated output voltage to the at least one aircraft accessory. Motoring synchronization for multiple engines 11840968 · 2023-12-12 · RTX Corporation · Matthew D. Teicholz, James W. Dunn, Christopher P. Kmetz, Louis A. Celiberti An aspect includes a method for motoring control for multiple engines of an aircraft is provided. A controller can determine a motoring time of a first engine starting system to cool a first engine. The controller can compare the motoring time of the first engine starting system with a motoring time of one or more other engine starting systems of one or more other engines of the aircraft. The motoring time of the first engine starting system can be controlled relative to a tolerance of the motoring time of the one or more other engine starting systems by adjusting the motoring time of the first engine starting system relative to the one or more other engine starting systems in a motoring sequence based on comparing the motoring time of the first engine starting system with the motoring time of the one or more other engine starting systems. Gas turbine engine with a double wall core casing 11118470 · 2021-09-14 · ROLLS-ROYCE plc · Chathura K Kannangara, Jillian C Gaskell, Stewart T Thornton, Timothy Philp A gas turbine engine includes an engine core including: a compressor system including first, lower pressure, compressor, and a second, higher pressure, compressor; and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, wherein the first flange connection is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor; a nacelle surrounding the engine core and defining a bypass duct between the engine core and the nacelle; wherein an axial midpoint of the radially outer edge is defined as the fan OGV tip centrepoint. TOROIDAL LIFT FORCE ENGINE 20210270230 · 2021-09-02 · Orville J. Birkestrand A toroidal lift force engine is provided. Illustratively, the toroidal lift force engine operates in an enclosed environment without heat and/or expelling particles of any kind, utilizing asymmetric pressure distribution on lift turbine blades solely to generate thrust with the normal component of this lift force, while using the tangential component of this lift force to drive accessories, provide control to the fluid velocity, and/or provide motivation of the fluid's flow. The toroidal lift force engine may be utilized to generate thrust, heat and/or electricity for powering vehicles, homes, etc. COMBINED CYCLE PLANT AND METHOD FOR OPERATING SAME 20210095572 · 2021-04-01 · Daiki FUJIMURA In a combined cycle plant and a method for operating the same, the combined cycle plant is provided with a gas turbine, a waste heat recovery boiler, and a steam turbine, and is also provided with a low-pressure gland steam line for supplying steam to a low-pressure gland portion of a low-pressure turbine, and a first heat exchanging unit which performs heat exchange between gland steam flowing through the low-pressure gland steam line, and fuel gas to be supplied to a combustor. $