Patent Publication Number: US-6210112-B1

Title: Apparatus for cooling an airfoil for a gas turbine engine

Description:
This application is a continuing application of U.S. patent application Ser. No. 08/992,322, having a filing date of Dec. 17, 1997, now U.S. Pat. No. 6,050,777. 
    
    
     BACKGROUND OF THE INVENTION 
     1. Technical Field 
     This invention relates to cooled rotor blades and/or stator vanes for gas turbines in general, and to apparatus and methods for cooling the leading edge and establishing film cooling along the surface of the rotor blade or stator vane in particular. 
     2. Background Information 
     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 external wall. The suction and pressure sides of the external 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. 
     The precise location of each stagnation point along the length of the leading edge is a function of the angle of incidence of the core gas relative to the chordline of the airfoil, for both rotor and stator airfoils. In addition to the angle of incidence, the stagnation point of a rotor airfoil is also a function of the rotational velocity of the airfoil and the velocity of the core gas. Given the curvature of the leading edge, the approaching core gas direction and velocity, and the rotational speed of the airfoil (if any), the location of the stagnation points along the leading edge can be readily determined by means well-known in the art. In actual practice, rotor speeds and core gas velocities vary depending upon engine operating conditions as a function of time and position along the span of the airfoil. As a result, the stagnation points (or collectively the stagnation line) along the leading edge of an airfoil will move relative to the leading edge. 
     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 external surface of the airfoil. Skewed, closely packed apertures, however, create stress concentrations. 
     Some prior art discloses the use of a porous transpiration strip disposed in a recess as a means to create a plenum in a forward portion of an airfoil. The transpiration strip has an arcuate outer profile that, when attached to the recess, provides the airfoil with an aerodynamic leading edge profile. Air entering the plenum through metering holes diffuses through the transpiration strip. A problem with this approach, particularly in those instances where the transpiration strip extends between the pressure and suction sides through the leading edge, is that pressure gradients along the leading can influence where cooling air exits the transpiration strip along the leading edge. The high pressure region that typically resides adjacent the stagnation line of an airfoil during operation, for example, will force cooling air to exit the transpiration strip in regions of lesser pressure. As a result, the leading edge region aligned with the stagnation line, which is typically subjected to some of the highest temperatures, may not be cooled as effectively as other regions of the transpiration strip. Another problem with transpiration cooling occurs when the strip becomes clogged with debris. The debris can inhibit or prevent cooling air from reaching portions of the strip, leaving those portions susceptible to undesirably high temperatures and consequent thermal degradation. 
     What is needed is an apparatus that provides adequate cooling along the leading edge of an airfoil, one that accommodates a variable position stagnation line, 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 in the airfoil wall. 
     DISCLOSURE OF THE INVENTION 
     It is, therefore, an object of the present invention to provide an airfoil having improved cooling along the leading edge. 
     It is another object of the present invention to provide an airfoil with leading edge cooling apparatus that accommodates a plurality of stagnation lines. 
     It is another object of the present invention to provide an airfoil with leading edge cooling apparatus that establishes uniform and durable film cooling downstream of the leading edge on both sides of the airfoil. 
     It is another object of the present invention to provide an airfoil with leading edge cooling apparatus 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 leading and trailing edges and spanwise between inner and outer radial surfaces, and includes an external wall surrounding an internal cavity. The trench is disposed in the external wall along the leading edge, extends in a spanwise direction, and is aligned with a stagnation line extending along the leading edge. 
     According to one aspect of the present invention, a method for cooling an airfoil is provided wherein a trench is provided disposed in the external wall of the airfoil. The trench is aligned with a stagnation line for the airfoil. 
     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. The trench of cooling air that extends continuously along the leading edge minimizes thermally induced stress by eliminating the discrete cooling points separated by uncooled areas characteristic of conventional cooling schemes. The uniform film of cooling air that exits from both sides of the trench also minimizes thermally induced stress by eliminating uncooled zones between and downstream of cooling apertures characteristic of conventional cooling schemes. 
     Another advantage of the present invention is that the leading edge cooling apparatus accommodates a plurality of stagnation lines. In the most preferable embodiment, the trench is preferably centered on the stagnation line which coincides with the largest heat load operating condition for a given application, and the width of the trench is preferably large enough such that the stagnation line will not travel outside of the side walls of the trench under all operating conditions. As a result, the present invention provides improved leading edge cooling and cooling air film formation relative to conventional cooling schemes. 
     These and other objects, features and advantages of the present invention will become apparent in light of the detailed description of the best mode embodiment thereof, as illustrated in the accompanying drawings. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     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, including core gas flow lines to illustrate the relative position of the trench and the stagnation point of the airfoil. The partial sectional view of the airfoil shown in this drawing also represents the airfoil of a stator vane. 
     FIG. 3 is a diagrammatic sectional view of a trench disposed in the leading edge of an airfoil. 
    
    
     BEST MODE FOR CARRYING OUT THE INVENTION 
     Referring to FIG. 1, a gas turbine engine turbine rotor blade  10  includes a root portion  12 , a platform  14 , an airfoil  16 , a trench  18  disposed in the airfoil  16 , and a blade tip  20 . The airfoil  16  comprises one or more internal cavities  22  (see FIG. 2) surrounded by an external wall  24 , at least one of which is proximate the leading edge  26  of the airfoil  16 . The suction side  28  and the pressure side  30  of the external wall  24  extend chordwise between the leading edge  26  and the trailing edge  32  of the airfoil  16 , and spanwise between the platform  14  and the blade tip  20 . The leading edge  26  has a smoothly curved contour which blends with the suction side  28  and pressure side  30  of the airfoil  16 . 
     Referring to FIG. 2, the trench  18  includes a base  34  and a pair of side walls  36  disposed in the external wall  24  along the leading edge  26 , preferably extending substantially the entire span of the airfoil  16 . A plurality of cooling apertures  38  provide passages between the trench  18  and the forward most internal cavity  22  for cooling air. The shape of the cooling apertures  38  and their position within the trench  18  will vary depending upon the application. FIG. 2 includes streamlines  40  representing core gas within the core gas path to illustrate the direction of core gas relative to the airfoil  16 . 
     As stated earlier, the stagnation point  42  (or in collective terms, the stagnation line) at any particular position along the span will move depending upon the engine operating condition at hand. The trench  18  is preferably centered on those stagnation points  42  which coincide with the largest heat load operating condition for a given application, and the width  44  of the trench  18  is preferably large enough such that the stagnation line  42  will not travel outside of the side walls  36  of the trench  18  under all operating conditions. If, however, it is not possible to provide a trench  18  wide enough to accommodate all possible stagnation line  42  positions, then the width  44  and the position of the trench  18  are chosen to accommodate the greatest number of stagnation lines  42  that coincide with the highest heat load operating conditions. The most appropriate trench width  44  and depth  46  for a given application can be determined by empirical study. Referring to FIG. 3 for example, empirical studies indicate that a trench  18  for a rotor airfoil  16  having a depth  46  substantially equal to one (1) cooling aperture  38  diameter (“D”) and a width  44  substantially equal to three (3) cooling aperture  38  diameters (“3D”), where the cooling aperture  38  is that which is disposed within the trench  18 , provides favorable leading edge  26  cooling and downstream cooling air film formation. 
     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  22  proximate the leading edge  26  of the airfoil  16  is at a lower temperature and higher pressure than the core gas flowing past the external wall  24  of the airfoil  16 . The pressure difference across the airfoil external wall  24  forces the internal cooling air to enter the cooling apertures  38  and subsequently pass into the trench  18  located in the external wall  24  along the leading edge  26 . The cooling air exiting the cooling apertures  38  diffuses into the air already in the trench  18  and distributes within the trench  18 . The cooling air subsequently exits the trench  18  in a substantially uniform manner over the side walls  36  of the trench  18 . The exiting flow forms a film of cooling air on both sides of the trench  18  that extends downstream. 
     One of the advantages of distributing cooling air within the trench  18  is that the pressure difference problems characteristic of conventional cooling apertures (not shown) are minimized. For example, the difference in pressure across a cooling aperture  38  is a function of the local internal cavity  22  pressure and the local core gas pressure adjacent the aperture  38 . 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 (not shown), 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  38  distributes and increases in uniformity within the trench  18 , thereby decreasing the opportunity for any low pressure zones to occur. Likewise, the distribution of cooling air within the trench  18  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. 
     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 spirit and the scope of the invention. For example, FIG. 2 shows a partial sectional view of an airfoil  16 . The airfoil  16  may be that of a stator vane or a rotor blade.