Patent Application
20170058678
1. Field
The present invention relates to gas turbine engines, and more specifically to a turbine blade with multiple internal cooling air circuits.
2. Description of the Related Art
In an industrial gas turbine engine, hot compressed gas is produced. The hot gas flow is passed through a turbine and expands to produce mechanical work used to drive an electric generator for power production. The turbine generally includes multiple stages of stator vanes and rotor blades to convert the energy from the hot gas flow into mechanical energy that drives the rotor shaft of the engine. Turbine inlet temperature is limited to the material properties and cooling capabilities of the turbine parts. This is especially important for first stages of turbine vanes and blades since these airfoils are exposed to the hottest gas flow in the system.
A combustion system receives air from a compressor and raises it to a high energy level by mixing in fuel and burning the mixture, after which products of the combustor are expanded through the turbine.
Since the turbine blades are exposed to the hot gas flow discharged from combustors within the combustion system, cooling methods are used to obtain a useful design life cycle for the turbine blade. Blade cooling is accomplished by extracting a portion of the cooler compressed air from the compressor and directing it to the turbine section, thereby bypassing the combustors. After introduction into the turbine section, this cooling air flows through passages or channels formed in the airfoil portions of the blades. The blade tip and the trailing edge of the blade are the most challenging locations in cooling.
The turbine second row blade is typically larger than the first row blade and has more surface area to cool. The second row blade is exposed to a lower gas temperature than the first row blade and therefore needs to allow for the use of less amounts of cooling air for better turbine efficiency.
In order to allow for higher temperatures, turbine blade designers have proposed several complex internal blade cooling circuits to maximize the blade cooling through the use of convection cooling, impingement cooling and film cooling of the blades. Conventionally, the focus of cooling improvement has been with the first row blade for more impact to turbine efficiency.
While this design provides good cooling to the majority of the airfoil, the blade tip section is much hotter than the other portions of the airfoil.
In one aspect of the present invention, a turbine rotor blade comprises: a leading edge and a trailing edge joined by a pressure side and a suction side, a tip end, and a root end; at least two integrated cooling circuits formed within the blade to provide cooling for the blade comprising; a leading edge circuit comprising a first cavity located along the leading edge of the blade and a second cavity positioned aft of the first cavity in an axial direction, wherein the second cavity opens forward into the first cavity; a trailing edge circuit comprising at least a third cavity located in a mid-chord area of the blade aft of the second cavity, wherein the trailing edge circuit flows aft with at least two substantially 180-degree turns at the tip end and the root end of the blade providing at least a penultimate cavity and a last cavity, wherein the last cavity is located along a trailing edge of the blade; and a tip axial cooling channel comprising a first opening and a second opening, wherein the first opening connects to the first cavity and the second opening connects to the penultimate cavity, wherein the tip axial cooling channel connects the leading edge circuit to the trailing edge circuit, integrating the at least two cooling circuits; and at least one crossover hole connecting the penultimate cavity to the last cavity substantially near the tip end of the blade.
In another aspect of the present invention, a method for increasing cooling to a trailing edge tip corner of a turbine blade, comprises: providing a tip axial cooling channel comprising a first opening and a second opening, connecting the first opening of the tip axial cooling channel to an end of a first cavity of a leading edge circuit of at least two integrated cooling circuits formed within the turbine blade, wherein the leading edge circuit comprises a first cavity located along a leading edge of the blade and a second cavity positioned aft of the first cavity in an axial direction, wherein the second cavity opens forward into the first cavity; connecting the second opening of the tip axial cooling channel to a trailing edge circuit, wherein the trailing edge circuit comprises at least a third cavity located in a mid-chord area of the blade aft of the second cavity, wherein the trailing edge circuit flows aft with at least two substantially 180-degree turns at the tip end and the root end of the blade providing at least a penultimate cavity and a last cavity, wherein the last cavity is located along a trailing edge of the blade, wherein the at least two integrated cooling circuits further comprise at least one crossover hole connecting the penultimate cavity to the last cavity substantially near the tip end of the blade; sending cooling air through the second cavity of the leading edge circuit and the third cavity of the trailing edge circuit, wherein the cooling air flowing through the leading edge circuit then flows through the tip axial cooling channel and into the penultimate cavity of the trailing edge circuit, merging with the cooling air entering into the penultimate cavity in the trailing edge circuit, wherein a portion of the cooling air flows through the at least one crossover hole into the trailing edge tip corner, and a portion of the cooling air flows through the rest of the penultimate cavity into and up through the last cavity through the rest of the trailing edge circuit into the trailing edge tip corner and/or out through the trailing edge of the turbine blade.
These and other features, aspects and advantages of the present invention will become better understood with reference to the following drawings, description and claims.
The invention is shown in more detail by help of figures. The figures show preferred configurations and do not limit the scope of the invention.
Broadly, an embodiment of the present invention provides a turbine rotor blade that includes at least two integrated cooling circuits that are formed within the blade that include a leading edge circuit having a first cavity and a second cavity and a trailing edge circuit that includes at least a third cavity located aft of the second cavity. The trailing edge circuit flows aft with at least two substantially 180-degree turns at the tip end and the root end of the blade providing at least a penultimate cavity and a last cavity. The last cavity is located along a trailing edge of the blade. A tip axial cooling channel connects to the first cavity of the leading edge circuit and the penultimate cavity of the trailing edge circuit. At least one crossover hole connects the penultimate cavity to the last cavity substantially near the tip end of the blade.
A blade of a gas turbine receives high temperature gases from a combustion system in order to produce mechanical work of a shaft rotation. Due to the high temperature gases, a cooling system may be provided to reduce the temperature levels throughout the blade.
A gas turbine cooling system may perform two basic functions. The first function may be to provide direct cooling of components exposed to gas path temperature that is higher than material temperature limits. The second function may be that of turbine environmental control. Air at correct pressure and temperature may be provided at various critical points to ensure that design environment is maintained throughout the turbine.
In certain embodiments, air for cooling the rotor and rotating blades may be extracted from the axial compressor discharge at a combustor shell. The compressor discharge air may pass through an air-to-air cooler and may be filtered for rotor cooling. Direct cooling may occur at the turbine spindle blade root end along one or more stages. The turbine stationary vanes may be cooled by both internal bypassing and external bleeding lines.
An effective step that can be taken to increase the power output and improve the efficiency of a gas turbine engine may be to increase the temperature at which heat is added to the system, that is, to raise the turbine inlet temperature of the combustion gases directed to the turbine. Increases in efficient turbines have lead to an increase in the temperature that must be withstood by the turbine blades and rotor. The result is that to use the highest desirable temperatures, some form of forced cooling may be desirable. This cooling may be in the form of air bled from the compressor at various stages, and ducted to critical elements in the turbine. Although emphasis is placed on cooling the initial stages of vanes and blades, air may be also directed to other vanes, blade rings and discs.
Being furthest from the inlet of cooling air, a trailing edge corner of the blade tends to be the one of the hottest portions of the blade after cooling. Better cooling to a tip section of a blade without using additional cooling air may be desirable. Embodiments of the present invention provide a blade that may allow for the reduction in temperature of the tip section of the blade without the use of additional cooling air. The cooling air that may be provided can stay within the turbine providing cooling for as long as possible.
The blades may be set in rows. A first row blade may receive the highest temperatures. A second row blade may receive reduced temperatures from the first row blade and may have a larger surface area. Due to the lower temperatures and larger surface area, efficiency may be based on a reduced amount of air being used for the second row blade.
As is illustrated in
The trailing edge circuit 14 may include a serpentine style path that may include multiple pass cooling channels, also referred to as cavities. In certain embodiments, there is a 3-pass serpentine cooling circuit. In certain embodiments, there is a 5-pass serpentine cooling circuit. In certain embodiments, there is a 7-pass serpentine cooling circuit. The trailing edge circuit 14 may include a third cavity 24. The entrance to the trailing edge circuit 14 may pass through the third cavity 24. The trailing edge circuit 14 may also include at least a penultimate cavity 30 and a last cavity 32. The trailing edge circuit 14 may include a third to last cavity 28. Below is described a 5-pass serpentine cooling circuit with the third to last cavity represented by the fifth cavity 28, the penultimate cavity represented by a sixth cavity 30 and the last cavity represented by a seventh cavity 32; however, in differently numbered cooling circuit the third to last cavity 28, the penultimate cavity 30, and the last cavity 32 may be different numbered cavities based on the total amount of passes within the serpentine cooling circuit.
In certain embodiments, the trailing edge circuit 14 may include, besides the third cavity 24, a fourth cavity 26, the fifth cavity 28, the sixth cavity 30, and the seventh cavity 32. The third cavity 24 may be in the radial direction towards a tip end 64 of the blade 10. The fourth cavity 26 may be in the radial direction towards a root end 62 of the blade 10. The fifth cavity 28 may be in the radial direction towards the tip end 64 of the blade 10. The sixth cavity 30 may be in the radial direction towards the root end 62 of the blade 10. The seventh cavity 32 may start in the radial direction towards the tip end 64 of the blade 10. The multiple pass cooling channels help move flow of air 50 from the leading edge 16 to the trailing edge 18 in order to help reduce the blade temperature throughout the blade 10.
The multiple cooling channels or cavities of the trailing edge circuit 14 are connected through substantially 180-degree turns along the tip end 64 and the root end 62 of the blade airfoil 10 that change the direction of cooling air flow 50 through the multiple cavities as the air flow 50 moves aft. The leading edge circuit 12 may include the first cavity 20 located along the leading edge 16 of the blade 10. The second cavity 22 is positioned aft of the first cavity 20 and may open forward into the first cavity 20. The trailing edge circuit 14 may include the third cavity 24 located in an approximately mid-chord area 46 of the blade 10 aft of the second cavity 22. The trailing edge circuit 14 may then flow aft with at least two substantially 180-degree turns at the tip end 64 and the root end 62 of the blade 10 providing the penultimate cavity 30 and the last cavity 32. The last cavity 32 may be located along the trailing edge 18 of the blade 10.
A tip axial cooling channel 34 may include a first opening 36 and a second opening 38. The tip axial cooling channel 34 may connect the leading edge circuit 12 to the trailing edge circuit 14, integrating the at least two cooling circuits. In certain embodiments, the trailing edge circuit 14 may be positioned below the tip axial cooling channel 34. The cooling air 50 flowing through the first cavity of the leading edge circuit 12 may then flow through the first opening of the tip axial cooling channel 34 and into the tip axial cooling channel 34. Cooling air 50 may also flow through the trailing edge circuit 14 starting by entering the third cavity 24 and moving aft through the trailing edge circuit 14. The air flow 50 may then pass through the tip axial cooling channel 34 second opening 38 into the penultimate cavity 30, or the sixth cavity in a 5-pass serpentine cooling circuit. The cooling air flow 50 from the tip axial cooling channel 34 may then merge with the air flow 50 from the fifth cavity 28 in the trailing edge circuit 14. This merging of air flow 50 may occur near a trailing edge tip corner 42. The cooling air 50 merging near the trailing edge tip corner 42 may decrease the temperature in that particular area without the need for additional cooling air.
The seventh cavity 32 of the trailing edge circuit 14 may open axially aft ward towards and through the trailing edge 18 of the blade 10. In certain embodiments, a plurality of trailing edge pins 44 and/or trailing edge exit holes 46 may be aligned along the trailing edge 18 allowing for the cooling air flow 50 to exit aft ward along the trailing edge 18 of the blade 10 and out of the blade 10. The trailing edge pins 44 and/or trailing edge exit holes 46 may promote heat transfer.
Further, at least one crossover hole 40 may connect the sixth cavity 30 and the seventh cavity 32. The at least one crossover hole 40 may be positioned substantially near the tip end 64 of the blade 10. The at least one crossover hole 40 may bring more cooling air 50 to the trailing edge tip corner 42. In certain embodiments, the at least one crossover hole 40 may have an elliptical shape.
The temperature of the blade 10 increases near the end of the trailing edge circuit 14 and along the tip end 64 along the leading edge 16 of the blade 10. Allowing additional cooling air 50 to enter the trailing edge tip corner 42 area through the tip axial cooling channel 34 and through the plurality of crossover holes 40, the temperature of the trailing edge tip corner 42 may decrease. A decrease in temperature may increase the life of the component and may provide increase efficiency.
The second opening 38 of the tip axial cooling channel 34 may connect with the penultimate cavity 30 in the trailing edge circuit 14. In certain embodiments, the tip axial cooling channel 34 may connect through the sixth cavity 30 of the trailing edge circuit 14. The cooling air 50 may then flow through the plurality of crossover holes 40 into the seventh cavity 32 along the trailing edge 18.
In certain embodiments, the trailing edge 18 may include zigzag pins and exit slots as detailed in
An inboard squealer tip 48 may be used in certain embodiments. The inboard squealer locates the squealer directly on the top of a tip cap of the tip end of the blade 10 to receive more effective conduction cooling.
In certain embodiments, refresher air 58 may be added to the airfoil 10. The refresher air 58 may be added through the seventh cavity 32. The refresher air 58 is not required, however the system may benefit by the use of the refresher air 58.
In certain embodiments, turbulators may be added to the body of the airfoil 10. Conventionally, full continuous turbulators may be used along the cavities. In certain embodiments, broken and offset turbulators 52 may cover the trailing edge 18. The pattern of broken and offset turbulators 52 may increase the flow of cooling air 50 radially upward along the cavity in order to provide additional cooling towards the tip end 64 of the blade 10. Upstream of the trailing edge 18, broken and staggered turbulators 54 may be used to cover the various channels.
In certain embodiments, the first cavity may include helical mini-grooves 56 along the leading edge 16. The helical mini-grooves 56 may increase the flow radially upward towards the tip axial cooling channel 34.
A method for increasing cooling to a trailing edge tip corner 42 may include providing the tip axial cooling channel 34 and connecting the first opening 36 of the tip axial cooling channel 34 to an end of the first cavity 20 of the leading edge circuit 12 of at least two integrated cooling circuits formed within the blade 10. The leading edge circuit 12 may include the first cavity 20 located along the leading edge 16 of the blade 10 and the second cavity 22 positioned aft of the first cavity 20 in an axial direction. The second cavity 22 may open forward into the first cavity 20. The second opening 38 of the tip axial cooling channel 34 may be connected to the trailing edge circuit 14 through the sixth cavity 30 of the trailing edge circuit 14.
The trailing edge circuit 14 may include at least a third cavity 24 located in a mid-chord area of the blade 10 aft of the second cavity 22. The trailing edge circuit 14 may flow aft with at least two substantially 180-degree turns at the tip end 64 and the root end 62 of the blade 10 providing at least the penultimate cavity 30 and the last cavity 32. The last cavity 32 may be located along the trailing edge 18 of the blade 10. The at least two integrated cooling circuits may further include at least one crossover hole 40 connecting the penultimate cavity 30 to the last cavity 32 substantially near the tip end 64 of the blade 10.
Cooling air 50 may be sent through the second cavity 22 of the leading edge circuit 12 and the third cavity 24 of the trailing edge circuit 14. The cooling air 50 flowing through the leading edge circuit 12 then may flow through the tip axial cooling channel 34 and into the penultimate cavity 30 of the trailing edge circuit 14, merging with the cooling air 50 moving from the third to last cavity 28 in the trailing edge circuit 14. A portion of the cooling air 50 may flow through the at least one crossover hole 40 into the trailing edge tip corner 42, and a portion of the cooling air 50 may flow through the rest of the penultimate cavity 30 into the last cavity 32 through the rest of the trailing edge circuit 14.
The cooling flow split between the leading edge circuit 12 and the trailing edge circuit 14 may be adjusted to achieve more uniform metal temperatures within the blade 10. The adjustment may be in the form of varying the thickness of the multiple channels, adjusting the length of the multiple channels, or the like. There may also be regenerative cooling for the platform 66 through the cooling circuits by routing some of the cooling air from the serpentine cooling circuit to the platform 66 cooling and then returning to the serpentine cooling circuit.
In certain embodiments, the trailing edge circuit 14 may have a smaller radial length than the prior art in order to position the tip axial cooling channel 34 above the trailing edge circuit 14.
While specific embodiments have been described in detail, those with ordinary skill in the art will appreciate that various modifications and alternative to those details could be developed in light of the overall teachings of the disclosure. Accordingly, the particular arrangements disclosed are meant to be illustrative only and not limiting as to the scope of the invention, which is to be given the full breadth of the appended claims, and any and all equivalents thereof.
Development of this invention was supported in part by the United States Department of Energy, Contract No. DE-FC26-05NT42644. Accordingly, the United States Government may have certain rights in this invention.
