Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means – Changing state mass within or fluid flow through working...
Reexamination Certificate
2000-01-11
2001-04-03
Verdier, Christopher (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
With heating, cooling or thermal insulation means
Changing state mass within or fluid flow through working...
C415S115000, C029S889721
Reexamination Certificate
active
06210112
ABSTRACT:
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
Auxier Thomas A.
Downs James P.
Soechting Friedrich O.
Tabbita Martin G.
Getz Richard D.
United Technologies Corporation
Verdier Christopher
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