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 multi-stage hydraulic rotating machine (MSHRM) maintains near-optimal efficiency over widely varying conditions of service (COS) when controlling a fluid having a gas volume fraction (GVF) greater than 50% and large changes in volumetric flow rate (VFR) between stages. The MSHRM includes separately controlled stages having at least two different designs with different VFR ranges. Stage impellor differences can include impellor diameter, blade pitch, blade width, blade number, inlet diameter, and outlet diameter. Diffusers can differ in similar ways between stages. VFR ranges can be progressively higher or lower in successive stages. The stages can share a common VFR range within which incompressible liquids can be controlled. The MSHRM can function as a pump or turbine, and can be applicable to energy storage and recovery in “green” energy systems.

A multi-stage hydraulic rotating machine (MSHRM) maintains near-optimal efficiency over widely varying conditions of service (COS) when controlling a fluid having a gas volume fraction (GVF) greater than 50% and large changes in volumetric flow rate (VFR) between stages. The MSHRM includes separately controlled stages having at least two different designs with different VFR ranges. Stage impellor differences can include impellor diameter, blade pitch, blade width, blade number, inlet diameter, and outlet diameter. Diffusers can differ in similar ways between stages. VFR ranges can be progressively higher or lower in successive stages. The stages can share a common VFR range within which incompressible liquids can be controlled. The MSHRM can function as a pump or turbine, and can be applicable to energy storage and recovery in “green” energy systems.

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 method of operating a gas turbine engine in a multi-engine aircraft, the gas turbine engine having an engine shaft mounted for rotation in a bearing and a gearbox connected to the engine shaft for torque transmission therebetween, includes axially preloading the bearing using an axially biasing element disposed between the gas turbine engine and the gearbox. The axially biasing element reacts against the gearbox to exert an axial preload force on the bearing and the engine shaft of the gas turbine engine.

In the present invention, a coal gasification combined power generation facility comprises: a feed water supply line (72); a condensate pump (39) and an intermediate-pressure feed water pump (40); a turbine bypass line (32) that bypasses a steam turbine and supplies steam to the condenser (73); and a spray water line (76) that supplies the feed water to the turbine bypass line (32). The coal gasification combined power generation facility has a normal operation mode and a bypass operation mode, and in the bypass operation mode, a control device (80) supplies the feed water to the turbine bypass line (32) and performs a first opening degree control to control the opening degree of the supply adjustment valve so that the amount of feed water supplied to the exhaust heat recovery boiler becomes less than that in the normal operation mode.

In the present invention, a coal gasification combined power generation facility comprises: a feed water supply line (72); a condensate pump (39) and an intermediate-pressure feed water pump (40); a turbine bypass line (32) that bypasses a steam turbine and supplies steam to the condenser (73); and a spray water line (76) that supplies the feed water to the turbine bypass line (32). The coal gasification combined power generation facility has a normal operation mode and a bypass operation mode, and in the bypass operation mode, a control device (80) supplies the feed water to the turbine bypass line (32) and performs a first opening degree control to control the opening degree of the supply adjustment valve so that the amount of feed water supplied to the exhaust heat recovery boiler becomes less than that in the normal operation mode.

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: $\frac{\mathrm{the}\mathrm{fan}\mathrm{tip}\mathrm{rotor}\mathrm{exit}\mathrm{temperature}}{\mathrm{the}<p></p>\; <div\; class="d-none\; d-xl-block\; offcanvas\; offcanvas-end\; h-100"\; data-bs-backdrop="false"\; data-bs-scroll="true"\; data-turbo-permanent=""\; id="chat">\; <turbo-frame\; loading="lazy"\; class="vstack\; h-100"\; id="sidebar\_chat"\; src="/chats"></turbo-frame>\; </div>\; <div\; class="toast-container\; position-fixed\; top-0\; end-0\; p-3"\; data-toast-container-target="container"></div>}$