Embodiments of the disclosure provide a method for controlling steam pressure within a turbine component. The method includes calculating a predicted stress on a rotor of the turbine component based on a predicted steam flow with the inlet valve in a minimum load position, a rotor surface temperature, and an inlet steam temperature, and determining whether the predicted stress exceeds a threshold. If the predicted stress exceeds the threshold, the inlet valve adjusts to a warming position. When steam in the discharge passage reaches a target pressure, the exhaust valve partially closes while maintaining the warming position of the inlet valve. If a safety parameter of the turbine component violates a boundary, the exhaust valve partially opens while maintaining the warming position of the inlet valve. When the predicted stress does not exceed the threshold, the inlet valve opens to at least the minimum load position.

A solar power generating system for generating electricity and providing heat includes; at least one generator for generating the electricity; a heating element for heating a heat transfer fluid; a turbocharger having at least one turbocharger turbine and at least one turbocharger compressor, wherein the at least one turbocharger compressor is adapted to receive and pressurize the heat transfer fluid, and the at least one turbocharger turbine is coupled to the at least one turbocharger compressor, wherein the at least one turbocharger compressor receiving and expanding a heated compressed heat transfer fluid coming from the heating element to drive the at least one turbocharger compressor and; a control unit configured to control the solar power generating system by comparing thermophysical properties obtained from more than one sensors placed in the solar power generating system with predetermined data in the control unit.

An aircraft turbomachine includes a lubricating circuit notably including a reservoir of lubricating fluid, a pump drawing lubricating fluid from the reservoir to inject it toward several components of the turbomachine, and an intake pipe connecting the reservoir to an intake orifice of the pump, the turbomachine further including a pneumatic starting circuit including a starting tube through which there circulates a flow of compressed air intended to supply a pneumatic starter of the turbomachine, wherein it includes a pneumatic line extending from the starting tube as far as the reservoir so as to supply the reservoir with compressed air so as to cause the lubricating fluid to circulate through the intake pipe towards the pump.

An aircraft turbomachine includes a lubricating circuit notably including a reservoir of lubricating fluid, a pump drawing lubricating fluid from the reservoir to inject it toward several components of the turbomachine, and an intake pipe connecting the reservoir to an intake orifice of the pump, the turbomachine further including a pneumatic starting circuit including a starting tube through which there circulates a flow of compressed air intended to supply a pneumatic starter of the turbomachine, wherein it includes a pneumatic line extending from the starting tube as far as the reservoir so as to supply the reservoir with compressed air so as to cause the lubricating fluid to circulate through the intake pipe towards the pump.

This steam turbine is provided with: a steam turbine rotor which extends in an axial direction; a pair of bearings which support the steam turbine rotor in such a way as to be capable of rotating about the axial direction; a steam turbine casing which encloses the steam turbine rotor between the pair of bearings; a casing support unit which supports the steam turbine casing from below; and a first heating unit: which is provided on the casing support unit and which is capable of heating the casing support unit.

A gas turbine engine includes a compressor section, the compressor section including a variable inlet guide vane which is movable between distinct angles to control the airflow approaching the compressor section. A control is programmed to position the vane at startup of the engine to direct airflow across the compressor section. The engine includes a fan for delivering bypass air into a bypass duct positioned outwardly of a core engine including the compressor section. The position of the vane is configured to direct airflow across the compressor section while an aircraft associated with the gas turbine engine is in the air, and to increase a windmilling speed of the compressor section and the turbine rotors. A method and variable inlet vane are also disclosed.

A gas turbine engine includes a compressor section, the compressor section including a variable inlet guide vane which is movable between distinct angles to control the airflow approaching the compressor section. A control is programmed to position the vane at startup of the engine to direct airflow across the compressor section. The engine includes a fan for delivering bypass air into a bypass duct positioned outwardly of a core engine including the compressor section. The position of the vane is configured to direct airflow across the compressor section while an aircraft associated with the gas turbine engine is in the air, and to increase a windmilling speed of the compressor section and the turbine rotors. A method and variable inlet vane are also disclosed.

A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.

A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.

An air turbine start system includes an air supply duct, an air turbine starter, a starter air valve, a stepper motor, and a controller. The air turbine starter is coupled to the air supply duct to selectively receive a flow of pressurized air therefrom. The starter air valve is mounted on the air supply duct and is movable between a closed position and a plurality of open positions. The stepper motor is coupled to the starter air valve and is configured, in response to valve position commands, to move the starter air valve between the closed position and one or more of the plurality of open positions. The controller is coupled to the stepper motor and is configured to supply the valve position commands to the stepper motor and determine a position of the starter air valve based on the valve position commands supplied to the stepper motor.