Patent Application
20050265844
1. Technical Field
This invention applies to gas turbine rotor blades in general, and to cooled gas turbine rotor blades in particular.
2. Background Information
Turbine sections within an axial flow turbine engine include rotor assemblies that include a rotating disc and a number of rotor blades circumferentially disposed around the disk. Rotor blades include an airfoil portion for positioning within the gas path through the engine. Because the temperature within the gas path very often negatively affects the durability of the airfoil, it is known to cool an airfoil by passing cooling air through the airfoil. The cooled air helps decrease the temperature of the airfoil material and thereby increase its durability.
Prior art cooled rotor blades very often utilize internal passage configurations that include a first radial passage extending contiguous with the leading edge, a second radial passage, and a rib disposed between and separating the passages. A plurality of crossover apertures is disposed within the rib, typically oriented perpendicular to the airfoil wall along the leading edge. A pressure difference across the rib causes a portion of the cooling air traveling within the second radial passage to pass through the crossover apertures and impinge on the leading edge wall. Cooling air passing through the crossover apertures typically travels in a direction perpendicular to the direction of the cooling airflow within the second radial passage. Hence, in the known prior art configurations cooling air is driven through the crossover apertures predominantly by static pressure, without little or no dynamic pressure contribution. Impingement cooling is efficient and desirable, but is provided in the prior art at the cost of a substantial static pressure drop across the rib.
The external gas path pressure is highest at the leading edge region during operation of the blade. In many turbine applications, airfoils are typically backflow margin limited at the leading edge of the airfoil. “Backflow margin” refers to the ratio of internal pressure to external pressure. To ensure an undesirable flow of hot gases from the gaspath does not flow into an airfoil, it is known to maintain a particular predetermined backflow margin that accounts for expected internal and external pressure variations. Hence, it is desirable to minimize pressure drops within the airfoil to the extent possible.
In addition to impingement cooling, it is also known to use trips strips within a cavity passage to enhance heat transfer between the cooling air and the airfoil. The trip strips enhance heat transfer by inducing the flow to become turbulent. Heat transfer in a boundary layer that is characterized by turbulent flow is typically greater than it is with one characterized by laminar flow. In addition to inducing turbulent flow, trip strips also provide additional surface area through which heat transfer may take place.
It is known to implement trip strips in a passage adjacent the crossover apertures (i.e., second radial passage). In the prior art of which we are aware, there is no specific positional relationship between the trip strips and crossover apertures. In fact, very often the trip strips are positioned where they impede cooling airflow through the crossover apertures.
What is needed, therefore, is an airfoil having an internal passage configuration that promotes desirable cooling of the airfoil and thereby increases the durability of the blade.
According to the present invention, a rotor blade is provided that includes a root, a hollow airfoil, and a conduit disposed within the root. The hollow airfoil has a cavity defined by a suction side wall, a pressure side wall, a leading edge, a trailing edge, a base, and a tip. An internal passage configuration is disposed within the cavity. The configuration includes a first radial passage, a second radial passage, a rib disposed between and separating the first radial passage and second radial passage, a plurality of crossover apertures disposed within the rib, and a plurality of trip strips disposed within the second radial passage. The trip strips are attached to an interior surface of one or both of the pressure side wall and the suction side wall. The trip strips are disposed within the second radial passage at an angle α that is skewed relative to a cooling airflow direction within the second radial passage, and positioned such that each of the plurality of trip strips converges toward the rib. The rib end of at least a portion of the plurality of trip strips is located between a pair of adjacent crossover apertures. The conduit is operable to permit airflow through the root and into the first passage.
One of the advantages of the present rotor blade and method is that airflow pressure losses within the airfoil are decreased relative to prior art airfoils having impingement cooling of which we are aware.
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.
Referring to
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The internal passage configuration includes a first conduit 42, a second conduit 44, and a third conduit 46 extending through the root 20 into the airfoil 22. Fewer or more conduits may be used alternatively. The first conduit 42 is in fluid communication with a first radial passage 48. A second radial passage 50 is disposed forward of the first radial passage 48, contiguous with the leading edge 32, and is connected to the first radial passage 48 by a plurality of crossover apertures 52. The crossover apertures 52 are disposed in a rib 53 that extends between and separates the first radial passage 48 and the second radial passage 50. The second radial passage 50 is connected to the exterior of the airfoil 22 by a plurality of cooling apertures 54 disposed along the leading edge 32. In some embodiments, the second radial passage 50 comprises one or more cavities. In other embodiments, the second radial passage 50 may be in direct fluid communication with the first conduit 42. At the outer radial end of the first radial passage 48 (i.e., the end of the first radial passage 48 opposite the first conduit 42), the first radial passage 48 is connected to an axially extending passage 56 that extends to the trailing edge 34 of the airfoil 22, adjacent the tip 30 of the airfoil 22.
The first radial passage 48 includes a plurality of trip strips 58 attached to the interior surface of one or both of the pressure side wall 36 and the suction side wall 38. The trip strips 58 are disposed within the passage 48 at an angle α that is skewed relative to the cooling airflow direction 60 within passage 48; i.e., at an angle between perpendicular and parallel to the airflow direction 60. Preferably, the trip strips 58 are oriented at angle of approximately 45° to the airflow direction 60. The orientation of each trip strip 58 within the passage 48 is such that the trip strip 58 converges toward the rib 53 containing the crossover apertures 52, when viewed in the airflow direction 60. Each of the trip strips 58 has an end 62 disposed adjacent the rib 53 (i.e., a “rib end”). At least a portion of the trip strips 58 have a rib end 62 radially located between a pair of crossover apertures 52, preferably approximately midway between the pair of crossover apertures 52. In a preferred embodiment, a majority of the trip strips 58 have a rib end 62 located radially between a pair of crossover apertures 52.
Referring to
An advantage of the above-described trip strip positioning is that the trip strips 58 provide two functions. First, the trip strips 58 perform a heat transfer function by causing desirable boundary layer conditions within the cooling airflow passing within the passage 48, and by providing additional surface area. Second, the trip strips 58 and their orientation relative to the crossover apertures 52 enable them to function as turning vanes, directing a portion of the cooling airflow toward the crossover apertures 52. As a result, the cooling air passing through the crossover apertures 52 is turning less than the 90° typical in the prior art. Indeed, in the preferred embodiment the 45° oriented trip strips 58 enable the cooling airflow to enter the crossover apertures 52 at an angle of approximately 45°. As a result, the pressure force driving the cooling airflow through the crossover apertures 52 includes a static pressure component and a dynamic pressure component, and the pressure drop across the rib is less than it would be in the aforesaid prior art configurations. The decreased pressure drop allows for a desirable higher backflow margin across the leading edge 32 of the airfoil 22.
Referring to
The third conduit 46 is in fluid communication with one or more passages 68 disposed between the serpentine passage 64 and the trailing edge 34 of the airfoil 22.
In the operation of the invention, the rotor blade airfoil 22 is disposed within the core gas path of the turbine engine. The airfoil 22 is subject to high temperature core gas passing by the airfoil 22. Cooling air, that is substantially lower in temperature than the core gas, is fed into the airfoil 22 through the conduits 42,44,46 disposed in the root 20.
Cooling air traveling through the first conduit 42 passes directly into the first radial passage 48, and subsequently into the axially extending passage 56 adjacent the tip 30 of the airfoil 22. A portion of the cooling air traveling within the first radial passage 48 encounters the trip strips 58 disposed within the passage 48. The trip strips 58 converging toward the rib 53 direct the portion of cooling airflow toward the rib 53. The position of the trip strips 58 relative to the crossover apertures 52 are such that the portion of cooling airflow directed toward the rib 53 is also directed toward the crossover apertures 52. The portion of cooling airflow travels through the crossover apertures 52 and into the second radial passage 50. The cooling air subsequently exits the second radial passage 50 via the cooling apertures 52 disposed in the leading edge 32 and the radial end of the second radial passage 48.
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.
