Patent Application
20100124508
This invention is directed generally to turbine airfoils, and more particularly to cooling systems in hollow turbine airfoils usable in turbine engines.
Typically, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures. In addition, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures.
Typically, turbine blades are formed from a root portion having a platform at one end and an elongated portion forming a blade that extends outwardly from the platform coupled to the root portion. The blade is ordinarily composed of a tip opposite the root section, a leading edge, and a trailing edge. The inner aspects of most turbine blades typically contain an intricate maze of cooling channels forming a cooling system. The cooling channels in a blade receive air from the compressor of the turbine engine and pass the air through the blade. Some of the cooling fluids are passed through the root and into the cavity between adjacent turbine blades to cool the platforms of the blades. The cooling fluids may be exhausted through gaps between adjacent blades and may create film cooling. The gaps are typically formed between side surfaces of the platforms that are generally parallel to each other and parallel to a longitudinal axis of the turbine blade. Oxidation and erosion of the side surfaces of the platforms often occurs and results in a greater flow of cooling fluids through the gap. The excessive fluid flow creates more turbulence in the film cooling layer and prevents adequate formation of the film cooling layer. Thus, a need exists for reducing the oxidation and erosion problems that typically occur on the side surfaces of platforms of turbine blades.
This invention relates to a turbine airfoil that is used in turbine engines and includes an internal cooling system with a portion of the cooling system positioned on side surfaces of a platform of the turbine airfoil. In particular, the turbine airfoil may include side surfaces proximate to the suction side and pressure side of the turbine airfoil that enhance cooling of the platform and promote the creation of film cooling boundary layers proximate to an upper surface of the platform. The side surfaces may be angled relative to the upper surface to increase the effectiveness of the interface between platforms of adjacent turbine airfoils regulating the flow of cooling fluids to reduce oxidation and erosion.
The turbine airfoil may be formed from a generally elongated, hollow airfoil having a leading edge, a trailing edge, a tip section at a first end, a root coupled to the airfoil at an end generally opposite the first end for supporting the airfoil and for coupling the airfoil to a disc, and a cooling system formed from at least one cavity in the elongated, hollow airfoil. The airfoil may also include a platform positioned at the intersection of the generally elongated, hollow airfoil and the root. The platform may include a leading edge, a trailing edge opposite the leading edge, a pressure side edge positioned proximate to a pressure side of the generally elongated, hollow airfoil and a suction side edge positioned proximate to a suction side of the generally elongated, hollow airfoil. The suction side edge may be positioned at an acute angle relative to a longitudinal axis of the platform.
The suction side edge may be formed from at least two surfaces including a first surface positioned at an obtuse angle relative to an upper surface of the platform and a second surface positioned at an obtuse angle relative to a bottom surface of the platform and intersecting the first surface such that the first and second surfaces are positioned in different planes. The first surface may include one or more film cooling slots. The film cooling slot may include a diffusion portion positioned adjacent to the upper surface of the platform. The diffusion portion may include a first backside wall angled between about five degrees and about twenty five degrees from the first surface, and in particular, about ten degrees from the first surface. The diffusion portion may also include a first angled sidewall angled between about five degrees and about twenty five degrees from a first sidewall of the film cooling slot. The diffusion portion may also include a second angled sidewall angled between about five degrees and about twenty five degrees from a second sidewall of the film cooling slot that is positioned generally opposite to the first sidewall of the film cooling slot. In one embodiment, the first and second angled sidewalls may be angled at about ten degrees relative to the first and second sidewalls, respectively.
The turbine airfoil may also include one or more dampers positioned between the second surface adjacent to the bottom surface of the platform and a side surface of a platform of an adjacent turbine blade. The damper may have a suction side surface that is generally aligned with the second surface of the suction side edge and may have a pressure side surface that is generally aligned with a side surface of a pressure side edge of an adjacent turbine blade. One or more cooling slots may be positioned in the suction side and may extend generally parallel to the suction side of the damper.
An advantage of this invention is that the angled suction side and pressure side edges limit the flow of cooling fluids through the gap between platforms of adjacent turbine blades.
Another advantage of this invention is that the suction side edge or the pressure side edge, or both, may include film cooling slots for cooling the platforms.
Yet another advantage of this invention is that the film cooling slots alleviate oxidation and erosion problems associated with conventional turbine blade platform edges.
Another advantage of this invention is that the configuration produces a good film sub-layer with a highly effective local film layer.
These and other embodiments are described in more detail below.
The accompanying drawings, which are incorporated in and form a part of the specification, illustrate embodiments of the presently disclosed invention and, together with the description, disclose the principles of the invention.
As shown in
As shown in
The platform 16 may be positioned at the intersection of the generally elongated, hollow airfoil 24 and the root 26. In one embodiment, the platform 16 may extend generally orthogonally to the generally elongated, hollow airfoil 24. As shown in
In one embodiment, the suction side edge 44 may be positioned at an acute angle 47 relative to a longitudinal axis 46 of the platform 16 where the suction side edge 44 is formed from a single surface. For instance, the suction side edge 44 may be positioned at an angle between about 30 and about 45 degrees. In another embodiment, as shown in
The turbine airfoil 10 may also include one or more film cooling slots 60 for cooling the platform 16 and allowing cooling fluids to form a film cooling boundary layer proximate to the upper surface 22. In one embodiment, the film cooling slots 60 may be positioned on the first surface 48. The film cooling slots 60 may extend for all of or a portion of the suction or pressure side edges 44, 42, or both, and may extend from the bottom surface 52 to the upper surface 22. The film cooling slots 60 may also include a diffusion portion 62 positioned adjacent to the upper surface 22 of the platform 16. The diffusion portion 62 may include side walls 64 at angles relative to sidewalls 66 forming the film cooling slots 60 to decrease the velocity of the cooling fluids flowing therethrough to reduce disruption of the layer of film cooling fluids proximate to the upper surface 22 of the platform 16 and the upper surface 56 of the adjacent platform 58. The diffusion portion 62 have an ever increasing cross-sectional area moving in a direction from the bottom surface 52 to the upper surface 22. As shown in
The turbine airfoil 10 may also include one or more dampers 74, as shown in
The damper 74 may also include one or more cooling slots 86, as shown in
During use, the damper 74 may substantially block the flow of cooling fluids through the gap 76 between adjacent platforms 16, 58. The cooling fluids may flow through the cooling slots 86 and the gap 76 proximate to the second surface 50. The cooling fluids then impinge on the pressure side edge 54 to provide backside impingement cooling for the platform 58. The cooling fluids may then flow proximate to the first surface 48 through the gap 76 and the film cooling slots 60. As the cooling fluids flow through the diffusion portions 62 of the film cooling slots 60, the velocity of the cooling fluids is reduced because of the increasing cross-sectional areas of the diffusion portions 62 of the film cooling slots 60 moving toward the upper surface 22. The diffusion portions 62 also enable the cooling fluids to be exhausted at a shallow angle relative to the upper surface 22 of the platform 16. The cooling fluids may then flow out of the gap 76 and form a layer of film cooling fluids proximate to the upper surface 22 of the platform 16 and the upper surface 56 of the platform 58. This configuration produces a good film sub-layer with a high local film effectiveness level and minimizes the local heat transfer coefficient augmentation due to film blowing effect.
The foregoing is provided for purposes of illustrating, explaining, and describing embodiments of this invention. Modifications and adaptations to these embodiments will be apparent to those skilled in the art and may be made without departing from the scope or spirit of this invention.
