A method of controlling a gas turbine engine including receiving an instantaneous thrust demand for current operation of the engine, determining the inlet flow rate and/or the pressure ratio within the compressor of the engine and determining whether the inlet flow rate and/or the pressure ratio match the working line for the compressor. The angle of one or more vane of the compressor is adjusted according to a closed control loop if the inlet flow rate and/or pressure ratio lie outside said desired range in order to adjust the inlet inflow rate and/or pressure ratio to meet the working line. The fuel flow to the engine combustor is adjusted concurrently in order to meet the thrust demand.

A device for controlling clearance between the tips of rotating blades of an aircraft turbomachine and a casing. An air circulation path allows air to be blown onto the casing in order to change its temperature. In addition, heating means heat the circulating air.

An example thermal management system may include a first heat exchanger including a bleed air inlet configured to receive input bleed air from a gas turbine engine and a bleed air outlet configured to output cooled bleed air. A turbine including a turbine inlet may be fluidically coupled to the bleed air outlet. The turbine may be configured to drive a shaft mechanically coupled to the turbine in response to expansion of the cooled bleed air through the turbine. A second heat exchanger may include an expanded bleed air inlet fluidically coupled to a turbine outlet of the turbine. The second heat exchanger may be configured to extract heat from at least one heat source using the expanded bleed air.

A method of controlling a gas turbine engine including receiving an instantaneous thrust demand for current operation of the engine, determining the inlet flow rate and/or the pressure ratio within the compressor of the engine and determining whether the inlet flow rate and/or the pressure ratio match the working line for the compressor. The angle of one or more vane of the compressor is adjusted according to a closed control loop if the inlet flow rate and/or pressure ratio lie outside said desired range in order to adjust the inlet inflow rate and/or pressure ratio to meet the working line. The fuel flow to the engine combustor is adjusted concurrently in order to meet the thrust demand.

A device for controlling clearance between the tips of rotating blades of an aircraft turbomachine and a casing. An air circulation path allows air to be blown onto the casing in order to change its temperature. In addition, heating means heat the circulating air.

An example thermal management system may include a first heat exchanger including a bleed air inlet configured to receive input bleed air from a gas turbine engine and a bleed air outlet configured to output cooled bleed air. A turbine including a turbine inlet may be fluidically coupled to the bleed air outlet. The turbine may be configured to drive a shaft mechanically coupled to the turbine in response to expansion of the cooled bleed air through the turbine. A second heat exchanger may include an expanded bleed air inlet fluidically coupled to a turbine outlet of the turbine. The second heat exchanger may be configured to extract heat from at least one heat source using the expanded bleed air.

A turbine engine system includes a fire chamber and a helical fan. The fire chamber includes a first cylindrical body having a first central axis; a first elongated hollow shaft extending co-axially along the first central axis; a plurality of wings fixedly secured to an exhaust plate and the first elongated hollow shaft; a plurality of combustion chambers disposed within a thickness of the first cylindrical body; a plurality of fuel injectors; and a plurality of spark plug injectors. The helical fan includes a cylindrical body having a second central axis; a second elongated hollow shaft extending co-axially along the central axis and in gaseous communication with the first elongated hollow shaft; and a plurality of blades extending from an inner wall of the cylindrical body to an outer surface of the elongated hollow shaft.

A cooling structure for a radial turbine is equipped with a partition wall arranged between a compressor and the radial turbine, and a through hole formed in the partition wall supplying a part of air compressed by the compressor to the rear surface of the radial turbine as cooling air. The through hole is inclined in a rotational direction of the radial turbine in the range of 40? to 80? with respect to a rotation axis of the radial turbine.

A gas turbine engine includes a gear system that transfers power input from a fan drive turbine to a fan through a gear system with a gear reduction ratio greater than 2.3. A flexible mount supports the gear system and accommodates movement of the gear system relative to an engine static structure. A lubrication system provides lubricant flow to a plurality of rotating components including bearing systems and the gear system and has a maximum capacity for removing thermal energy less than or equal to no more than 2% of power input into the gear system.

A disclosed gas turbine engine includes a fan including a plurality of fan blades rotatable about an axis, a bypass duct, a compressor section and a bypass ratio greater than 10. A combustor in fluid communication with the compressor section and a fan drive turbine in communication with the combustor. A gear system provides a speed reduction between the fan drive turbine and the fan and transfer power input from the fan drive turbine to the fan at an efficiency greater than 98%.