EP0924384A2

EP0924384A2 - Airfoil with leading edge cooling

Info

Publication number: EP0924384A2
Application number: EP98310191A
Authority: EP
Inventors: George P. Liang; Thomas A. Auxier
Original assignee: United Technologies Corp
Current assignee: Raytheon Technologies Corp
Priority date: 1997-12-17
Filing date: 1998-12-11
Publication date: 1999-06-23
Also published as: JPH11247611A; EP0924384A3; KR19990063131A

Abstract

A hollow airfoil (16) is provided which includes a body, a trench (30), and a plurality of cooling apertures (36) disposed within the trench (30). The body extends chordwise between a leading edge (26) and a trailing edge (29), and spanwise between an outer radial surface and an inner radial surface, and includes an exterior wall surrounding a cavity (20). The trench (30) is disposed in the exterior wall along the leading edge (26), extending in a spanwise direction.

Description

This invention relates to cooled rotor blades and/or stator vanes for gas turbines in general, and to apparatus that cools the leading edge and establishes film cooling along the surface of the rotor blade or stator vane in particular.
In the turbine section of a gas turbine engine, core gas travels through a plurality of stator vane and rotor blade stages. Each stator vane or rotor blade has an airfoil with one or more internal cavities surrounded by an exterior wall. The suction and pressure sides of the exterior wall extend between the leading and trailing edges of the airfoil. Stator vane airfoils extend spanwise between inner and outer platforms and the rotor blade airfoils extend spanwise between a platform and a blade tip.
High temperature core gas (which includes air and combustion products) encountering the leading edge of an airfoil will diverge around the suction and pressure sides of the airfoil, or impinge on the leading edge. The point along the leading edge where the velocity of the core gas flow goes to zero (i.e., the impingement point) is referred to as the stagnation point. There is a stagnation point at every spanwise position along the leading edge of the airfoil, and collectively those points are referred to as the stagnation line. Air impinging on the leading edge of the airfoil is subsequently diverted around either side of the airfoil.
Cooling air, typically bled off of a compressor stage at a temperature lower and pressure higher than the core gas passing through the turbine section, is used to cool the airfoils. The cooler compressor air provides the medium for heat transfer and the difference in pressure provides the energy required to pass the cooling air through the stator or rotor stage.
In many cases, it is desirable to establish film cooling along the surface of the stator or rotor airfoil. A film of cooling air traveling along the surface of the airfoil transfers thermal energy away from the airfoil, increases the uniformity of the cooling, and insulates the airfoil from the passing hot core gas. A person of skill in the art will recognize, however, that film cooling is difficult to establish and maintain in the turbulent environment of a gas turbine. In most cases, film cooling air is bled out of cooling apertures extending through the external wall of the airfoil. The term "bled" reflects the small difference in pressure motivating the cooling air out of the internal cavity of the airfoil. One of the problems associated with using apertures to establish a cooling air film is the films sensitivity to pressure difference across the apertures. Too great a pressure difference across an aperture will cause the air to jet out into the passing core gas rather than aid in the formation of a film of cooling air. Too small a pressure difference will result in negligible cooling air flow through the aperture, or an in-flow of hot core gas. Both cases adversely affect film cooling effectiveness. Another problem associated with using apertures to establish film cooling is that cooling air is dispensed from discrete points along the span of the airfoil, rather than along a continuous line. The gaps between the apertures, and areas immediately downstream of those gaps, are exposed to less cooling air than are the apertures and the spaces immediately downstream of the apertures, and are therefore more susceptible to thermal degradation. Another problem associated with using apertures to establish film cooling is the stress concentrations that accompany the apertures. Film cooling effectiveness generally increases when the apertures are closely packed, and skewed at a shallow angle relative to the exterior surface of the airfoil. Skewed, closely packed apertures, however, create stress concentrations.
What is needed is an apparatus that provides adequate cooling along the leading edge of an airfoil, one that creates a uniform and durable cooling air film downstream of the leading edge on both sides of the airfoil, and one that creates minimal stress concentrations within the airfoil wall.
According to the present invention, a hollow airfoil is provided which includes a body, a trench, and a plurality of cooling apertures disposed within the trench. The body extends chordwise between a leading edge and a trailing edge,
and includes an exterior wall surrounding a cavity. The trench is disposed in the exterior wall along the leading edge, extending in a spanwise direction.
An advantage of the present invention is that uniform and durable film cooling downstream of the leading edge is provided on both sides of the airfoil. The cooling air bleeds out of the trench on both sides and creates continuous film cooling downstream of the leading edge. The trench minimizes cooling losses characteristic of cooling apertures, and thereby provides more cooling air for film development and maintenance.
Another advantage of the present invention is that stress is minimized along the leading edge and areas immediately downstream of the leading edge. One characteristic responsible for minimizing stress is the trench of cooling air extending continuously along the leading edge. The trench substantially eliminates discrete cooling points separted by uncooled areas, and thereby eliminates the thermally induced stress associated therewith. The trench also minimizes stress by distributing cooling air along the leading edge. The cooling air bleeds out of the trench on both sides and creates continuous film cooling downstream of the leading edge. The continuous film eliminates uncooled zones between and downstream of cooling apertures, and thereby eliminates the thermally induced stress associated therewith.
A preferred embodiment will now be described by way of example only and with reference to the accompanying drawings in which:
FIG.1 is a diagrammatic perspective view of a turbine rotor blade for a gas turbine engine.
FIG.2 is a partial sectional view of the airfoil portion of the rotor blade shown in FIG. 1, having a single trench. The partial sectional view of the airfoil shown in this drawing also represents the airfoil of a stator vane.
FIG.3 is the partial sectional view of an airfoil shown in FIG.2, having a plurality of trenches.
Referring to FIG. 1, a gas turbine engine turbine rotor blade 10 includes a root portion 12, a platform 14, an airfoil 16, and a blade tip 18. The airfoil 16 comprises one or more internal cavities 20 (see FIGS. 2 and 3) surrounded by a external wall 22, at least one of which is proximate the leading edge 24 of the airfoil 16. The suction side 26 and the pressure side 28 of the external wall 22 extend chordwise 27 between the leading edge 24 and the trailing edge 29 of the airfoil 16, and spanwise 31 between the platform 14 and the blade tip 18. The leading edge 24 has a smoothly curved contour which blends with the suction side 26 and the pressure side 28 of the airfoil 16.
Referring to FIGS. 2 and 3, a trench 30 having a base 32 and a pair of side walls 34 is disposed in the external wall 22 of an airfoil 16 along the leading edge 24. FIG. 3 shows an embodiment having a plurality of trenches 30. Each trench 30 extends substantially the entire span 31 (see FIG. 1) of the airfoil 16 leading edge 24. A plurality of cooling apertures 36 are disposed in the trench 30, extending between an internal cavity 20 and the trench 30. The shape of the cooling apertures 36 and their position within the trench 30 will vary depending upon the application. In most cases, however, the cooling apertures 36 are uniformly distributed in the base 32 of the trench 30 throughout the span 31. In a preferred embodiment, the cooling apertures 36 include a diffusion portion 38.
In the operation of the invention, cooling air typically bled off of a compressor stage (not shown) is routed into the airfoil 16 of the rotor blade 10 (or stator vane) by means well known in the art. Cooling air disposed within the internal cavity 20 proximate the leading edge 24 of the airfoil 16 is at a lower temperature and higher pressure than the core gas flowing past the external wall 22 of the airfoil 16. The pressure difference across the airfoil external wall 22 forces the internal cooling air to enter the cooling apertures 36 and subsequently pass into the trench(es) 30 located in the external wall 22 along the leading edge 24. The cooling air exiting the cooling apertures 36 diffuses into the cooling air already in the trench 30 and distributes within the trench 30. In the preferred embodiment where the cooling apertures 36 include diffusion portions 38, the diffusion portions 38 increase cooling air diffusion and distribution and therefore uniformity within the trench 30.
One of the advantages of distributing cooling air within the trench 30 is that the pressure difference problems characteristic of conventional cooling apertures are minimized. For example, the difference in pressure across a cooling aperture 36 is a function of the local internal cavity 22 pressure and the local core gas pressure adjacent the aperture 36. Both of these pressures vary as a function of time. If the core gas pressure is high and the internal cavity pressure is low adjacent a particular cooling aperture in a conventional scheme, undesirable hot core gas in-flow can occur. The present invention minimizes the opportunity for the undesirable in-flow because the cooling air from all apertures 36 collectively distributes within the trench 30, thereby decreasing the opportunity for any low pressure zones to occur. Likewise, the distribution of cooling air within the trench 30 also avoids cooling air pressure spikes which, in a conventional scheme, would jet the cooling air into the core gas rather than add it to the film of cooling air downstream.
Cooling air subsequently exits the trench 30 in a uniform manner along both spanwise sides of the trench 30. The exiting flow forms a film of cooling air on both sides of the trench 30 that extends downstream. In the case of multiple trenches 30, the cooling air exiting a trench 30 positioned downstream of the stagnation point 40 of the airfoil 16 may exit predominantly on the downstream side of the trench 30. In that case, film cooling emanating from a upstream trench 30 predominantly cools the external wall 22 of the airfoil 16 between the two adjacent trenches 30.
From the above it will be seen that there has been described an airfoil having improved cooling along the leading edge, with leading edge cooling apparatus that establishes uniform and durable film cooling downstream of the leading edge on both sides of the airfoil and that creates minimal stress concentrations within the airfoil wall.
Although this invention has been shown and described with respect to the detailed embodiments thereof, it will be understood by those skilled in the art that various changes in form and detail thereof may be made without departing from the scope of the invention. For example, FIGS. 2 and 3 show a partial sectional view of an airfoil. The airfoil 16 may be that of a stator vane or a rotor blade.

Claims

An airfoil (16), comprising:

a body, extending chordwise between a leading edge (24) and a trailing edge (29), said body having an exterior wall surrounding a cavity (20);

a trench (30), disposed in said exterior wall (22) along said leading edge (26), extending in a spanwise direction; and

a plurality of cooling apertures (36), disposed within said trench (30) and extending through said exterior wall (22).
An airfoil according to claim 1, wherein said trench (30) comprises:

a first side wall (34);

a second side wall (34);

a base (32), extending between said first and second side walls (34);
wherein said cooling apertures (36) are disposed in said base (32) and extend through said exterior wall (22).
An airfoil according to claim 1 or 2, wherein said airfoil (16) comprises a plurality of said trenches (30).
An airfoil according to any preceding claim wherein said cooling apertures (36) include diffusion portions (38).
An airfoil according to any preceding claim, wherein said airfoil (16) is part of a stator vane.
An airfoil according to any of claims 1 to 5, wherein said airfoil (16) is part of a rotor blade.