The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. The gases flow downstream through a high pressure (HPT) turbine that extracts energy for powering the compressor.
And, a low pressure turbine (LPT) follows the HPT for extracting additional energy from the combustion gases for powering an upstream fan in a turbofan aircraft engine application, or for powering an external drive shaft for turboprop applications.
Engine efficiency is proportional to the maximum temperature of the combustion gases. However, combustion gas temperature is limited by the material strength of the various gas turbine engine components which are heated by the combustion gases.
The various flowpath components that adjoin the combustion gases during operation are typically cooled by bleeding a portion of the pressurized air from the compressor. Each component has a specifically configured and dedicated cooling circuit for locally maximizing cooling efficiency with a limited amount of cooling air.
Any cooling air diverted from the combustion process correspondingly reduces engine efficiency and is balanced against the desired life expectancy for the various engine components.
Since the combustion gases are born in the annular combustor immediately downstream of corresponding fuel injectors spaced circumferentially apart from each other, the combustion gases have a corresponding circumferentially sinusoidal temperature pattern.
And, since the annular flowpath for the combustion gases as they travel axially through the engine has radially outer and inner boundaries, the combustion gases also experience a radial profile which is initially parabolic with a maximum temperature near the radial midspan of the flowpath and lower temperatures near the outer and inner flowpath boundaries.
After entering the first stage turbine nozzle, the combustion gases are mixed in the various stages of turbine rotor blades downstream therefrom which changes both the circumferential and radial temperature distributions of the combustion gases due to the aerodynamic and centrifugal forces created thereby.
Each turbine nozzle stage includes a row of hollow vanes extending radially between outer and inner supporting bands. Each rotor stage includes a row of typically hollow turbine rotor blades extending radially outwardly from a supporting blade platform and dovetail mounted in a supporting rotor disk. And a stationary annular turbine shroud surrounds each row of turbine blades.
The nozzle vanes and turbine blades have corresponding airfoil configurations for guiding and extracting energy from the combustion gases. The nozzle bands, blade platforms, and turbine shrouds define the radially outer and inner flowpath boundaries for the combustion gases.
And each of these flowpath components typically includes a corresponding cooling circuit therefore. The inner root and outer tip of the turbine blades are typically more difficult to cool than the radial midspan portions of the airfoils. The nozzle inner and outer bands have different environments affecting the cooling configurations therefore. And, the turbine shrouds are suitably suspended above the blade tips and also have different cooling configurations.
Since the radially outer and inner turbine bands, blade platforms, and turbine shrouds define the outer and inner flowpath boundaries, the velocity of the combustion gases is relatively low therealong as compared to the midspan of the airfoils where gas velocity is at its maximum. Correspondingly, the heat flux from the combustion gases varies substantially in the radial direction and additionally varies as the circumferential pattern and radial profile of the combustion gases vary downstream from stage to stage.
Accordingly, the various cooling circuits for the various flowpath components typically include various forms of cooling holes for convection cooling, impingement cooling, and film cooling as dictated by the local conditions of the combustion gases and the temperature or heat generated thereby.
The prior art contains innumerable configurations of cooling circuits and cooling holes for gas turbine engine components all striving to maximize cooling efficiency while minimizing cooling air for effecting long life and durability of the engine components.
The specific configuration of the cooling circuits is designed for accommodating the local profile of the combustion gases to limit component temperature and stress within acceptable limits for durability and life.
Accordingly, it is desired to provide a turbine nozzle having improved cooling for the nozzle itself, as well as improving the temperature profile of the combustion gases discharged therefrom.
USPTO Backgrounds
USPTO Backgrounds