7 results about "Overall pressure ratio" patented technology

A gas turbine engine has a turbomachine comprising compressor, combustion, and turbine sections. The gas turbine engine defines a maximum exhaust gas temperature, a maximum drive turbine shaft torque, and a corrected specific power. The gas turbine engine includes a nozzle vane and an engine casing defining a fluid supply plenum. The engine casing is configured to route cooling fluid from the fluid supply plenum to the nozzle vane. The routing of the cooling fluid facilitates cooling of the nozzle vane and increases an overall pressure ratio of the gas turbine engine. Integration of the cooling fluid routing within the engine casing reduces parasitic bleed air losses, enables tighter turbine stage spacing, and improves the thermal efficiency of the engine, thereby enhancing performance at elevated pressure ratios.

A turbofan engine is provided. The turbofan engine includes a low pressure spool; a fan mechanically coupled with the low voltage spool; a high-voltage wire shaft; the accessory gear box is mechanically connected with the high-voltage wire shaft; a hydraulic pump mechanically coupled with the accessory gearbox; and one or more heat exchangers connected to the accessory gearbox, the one or more heat exchangers having a heat exchange capacity defined by a power of half of a product determined by multiplying an average heat exchanger effectiveness of the one or more heat exchangers by a thermal conductivity coefficient, the thermal conductivity is related to an accessory gearbox thermal load, a hydraulic pump power of the hydraulic pump, a fan diameter of the fan, an engine length of the turbofan engine, and a total pressure ratio of the turbofan engine.

An ignition and combustion assist system and method comprising a plasma igniter and electronic driver unit for use with gas turbine engines operating under low air densities, reduced voltage conditions and overall pressure ratios of 3:1 to 7:1. The plasma igniter has an inner chamber housing a centrally positioned and electrically isolated electrode attached to an electrical lead, driver unit, and AC or DC power supply. The electrode features a corner positioned near an outlet end of the igniter, where a plasma arc ignites a fuel-air mixture creating a flame extending into a primary burn region of a combustor of the gas turbine. The driver unit is in two embodiments and configured with low-cost microsecond voltage wave time periods or energy-efficient nano-second pulses. The method uses the plasma igniter and the electronic driver units described herein separately with other components or together.

A gas turbine engine includes, among other things, a propulsor section including a propulsor, a core engine, a gear arrangement that drives the propulsor. A compressor section includes a first compressor and a second compressor. A turbine section includes a first turbine and a second turbine. An overall pressure ratio is provided by the combination of a pressure ratio across the first compressor and a pressure ratio across the second compressor, and greater than 40. The pressure ratio across the high pressure compressor is no more than 15, and the pressure ratio across the low pressure compressor is at least 4.

This application proposes a method for correcting the interstage bleed air performance parameters of a heavy-duty gas turbine compressor component, comprising: obtaining the overall pressure ratio of the gas turbine; obtaining the thermodynamic performance parameters of the gas turbine compressor; based on the overall pressure ratio and thermodynamic performance parameters, obtaining the first isentropic compression power and the first actual power consumed by the compressor outlet air; based on the thermodynamic performance parameters, obtaining the second isentropic compression power and the second actual power consumed by the air bleed out from the compressor interstage; based on the numerical relationship between the first isentropic compression power, the second isentropic compression power and the first actual power consumed, the second actual power consumed, obtaining the temperature rise efficiency considering interstage bleed air; and based on the numerical relationship between the first isentropic compression power, the second isentropic compression power and the compressor torque value, obtaining the torque efficiency considering interstage bleed air. The technical solution of this application can reduce the deviation in bleed air performance calculations.

A gas turbine engine includes, among other things, a propulsor section including a propulsor, a core engine, a gear arrangement that drives the propulsor. A compressor section includes a low pressure compressor and a high pressure compressor. A turbine section includes a high pressure turbine and a low pressure turbine. An overall pressure ratio is provided by the combination of a pressure ratio across the low pressure compressor and a pressure ratio across the high pressure compressor, and greater than 40. The pressure ratio across the high pressure compressor is between 7 and 15, and the pressure ratio across the low pressure compressor is between 4 and 8.

A gas turbine engine includes, among other things, a propulsor section including a propulsor, a core engine, a gear arrangement that drives the propulsor. A compressor section includes a first compressor and a second compressor. A turbine section includes a first turbine and a second turbine. An overall pressure ratio is provided by the combination of a pressure ratio across the first compressor and a pressure ratio across the second compressor, and greater than 40. The pressure ratio across the high pressure compressor is no more than 15, and the pressure ratio across the low pressure compressor is at least 4.