Patent Application
20160319681
1. Technical Field
The present invention relates generally to gas turbine engine turbine blade cooling and, more specifically, cooled turbine blades and slots for mounting the blades.
2. Background Information
Turbine blades in gas turbine engine turbines and, particularly, high pressure turbine blades are often cooled by a portion of pressurized air from a compressor of the engine. Each turbine stage includes a row of turbine rotor blades extending radially outwardly from a supporting rotor disk with the radially outer tips of the blades being mounted inside a surrounding turbine shroud. Typically, turbine rotor blades of at least the first turbine stage are cooled by the bled portion of the pressurized air from the compressor. The blades include roots slid into and secured by axial slots in a turbine disk.
The blades are typically cooled using a portion of high pressure compressor discharge air bled (also known as compressor discharge pressure or CDP air) from the last stage of the compressor. The air is suitably channeled through internal cooling channels inside the hollow blades and discharged through the blades in various rows of film cooling holes from the leading edge and aft therefrom, and also typically including a row of trailing edge outlet holes or slots on the airfoil pressure side.
Blade cooling air is gathered and transferred from static portions of the engine to the rotating disk supporting the blades. The cooling air passes through the slot and into the blade root from where it is distributed through a cooling circuit having cooling passages in an airfoil of the blade.
The typical turbofan aircraft engine initially operates at a low power, idle mode and then undergoes an increase in power for takeoff and climb operation. Upon reaching cruise at the desired altitude of flight, the engine is operated at lower or intermediate power setting. The engine is also operated at lower power as the aircraft descends from altitude and lands on the runway, following which thrust reverse operation is typically employed with the engine again operated at high power. In the various transient modes of operation of the engine where the power increases or decreases, the turbine blades heat up and cool down respectively.
A slot bottom of the disk is exposed to blade cooling air during engine operation. The cooling air increases the thermal response of the slot bottom creating a large thermal gradient between the slot bottom and bore of the disk. This gradient creates large thermal stresses in both the acceleration and deceleration of the engine. These large thermal stresses reduces the low cycle fatigue life of the disk.
Accordingly, it is desired to provide a gas turbine engine having turbine blade cooling with a design which reduces a thermal gradient in a bottom of a root mounting slot. It is further desired to reduce large thermal stresses in the bottom of the root mounting slot caused by the thermal gradient. It is also desired to increase the low cycle fatigue life of the disk by reducing these thermal stresses.
A gas turbine engine turbine blade assembly includes a hollow airfoil integrally joined to a blade root, a dovetail slot heat shield attached to a bottom surface of the root, and a shield outlet from the dovetail slot heat shield open to at least one inlet aperture extending radially through a radially inner root end of the root. The heat shield may be bonded to the bottom surface.
The heat shield may include a body with a heat shield bottom and sides or legs extending upwardly or radially outwardly from the heat shield bottom. The heat shield may have a slanted open forward or upstream end and free ends of the legs may be longer than the heat shield bottom.
An axially extending straight flange may be located along a free end of each of the legs and the flanges may be bonded to the bottom surface. The heat shield may have a slanted open forward or upstream end of the heat shield and the flanges and the free ends of the legs may be longer than the heat shield bottom. The body may be rounded. The heat shield bottom and/or the legs may be rounded.
A gas turbine engine turbine disk assembly may include a disk including a web extending radially outwardly from a hub to a rim; a plurality of dovetail slots in the rim; a complimentary plurality of turbine blades removably retained in the plurality of dovetail slots; slot bottoms of the dovetail slots and the dovetail slots extending circumferentially between disk posts in the rim on the disk assembly, and each of the turbine blades including a hollow airfoil integrally joined to a blade root, a dovetail slot heat shield attached to a bottom surface of the root, and a shield outlet from the dovetail slot heat shield open to at least one inlet aperture extending radially through a radially inner root end of the root.
The gas turbine engine turbine disk assembly may include a clearance between the heat shield bottoms of the heat shields and respective ones of the slot bottoms. The heat shield bottoms may be radially spaced apart from respective ones of the slot bottoms and the heat shields may be bonded to the bottom surfaces.
Illustrated schematically in
Referring to
Referring to
The cooling air 11 flows into the dovetail slot 29, through the root end 35, and then radially outwardly through cooling channels 70 in the cooling air circuit 52 in the airfoil 16. The cooling air 11 is then discharged through rows of outlet holes in the pressure and suction sides of the blade airfoil in a conventional manner. Further referring to
Referring to
Referring to
Referring to
An open forward or upstream end 100 of the heat shield 40 is bevelled or slanted upstream indicated by a bevel 102 on the upstream end 100. The upstream end 100 is bevelled or slanted such that the flanges 96 and the free ends 98 of the legs 92 are longer than the heat shield bottom 90 of the heat shield 40. The bevelled or slanted upstream end 100 of the heat shield 40 helps direct the cooling air 11 into a hollow interior 89 of the body 88 of the heat shield 40. The cooling air 11 exits the hollow interior 89 through a shield outlet 93 between the flanges 96 and the free ends 98 of the legs 92 and through the plurality of inlet apertures 50. The cooling air 11 flows through the dovetail slot and through the inner root end 35 of the dovetail root 18 with minimal contact of the slot bottom 60 disposed along the rim 24 on the rotor disk 30.
Illustrated in
While there have been described herein what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention. Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims.
