Patent Application
20160245095
The subject matter disclosed herein relates to turbine systems and, more particularly, to a turbine rotor blade with enhanced cooling and reduced tip leakage losses.
In a gas turbine engine, air pressurized in a compressor is used to combust a fuel in a combustor to generate a flow of hot combustion gases, whereupon such gases flow downstream through one or more turbines so that energy can be extracted therefrom. In accordance with such a turbine, generally, rows of circumferentially spaced turbine rotor blades extend radially outwardly from a supporting rotor disk. Each blade typically includes a dovetail that permits assembly and disassembly of the blade in a corresponding dovetail slot in the rotor disk, as well as an airfoil that extends radially outwardly from the dovetail and interacts with the flow of the working fluid through the engine.
The airfoil has a generally concave pressure side and generally convex suction side extending axially between corresponding leading and trailing edges and radially between a root and a tip. It will be understood that the blade tip is spaced closely to a radially outer turbine shroud for minimizing leakage therebetween of the combustion gases flowing downstream between the turbine blades. Improved efficiency of the engine is obtained by minimizing the tip clearance or gap such that leakage is prevented, but this strategy is limited somewhat by the different thermal and mechanical expansion and contraction rates between the rotor blades and the turbine shroud and the motivation to avoid an undesirable scenario of having the tip rub against the shroud during operation.
In addition, because turbine blades are bathed in hot combustion gases, effective cooling is required for ensuring a useful part life. Typically, the blade airfoils are hollow and disposed in flow communication with the compressor so that a portion of pressurized air bled therefrom is received for use in cooling the airfoils. Airfoil cooling is quite sophisticated and may be employed using various forms of internal cooling channels and features, as well as cooling holes through the outer walls of the airfoil for discharging the cooling air. Nevertheless, airfoil tips are particularly difficult to cool since they are located directly adjacent to the turbine shroud and are heated by the hot combustion gases that flow through the tip gap. Accordingly, a portion of the air channeled inside the airfoil of the blade is typically discharged through the tip for the cooling thereof.
Tip portions of blades often include a pocket that the cooling air is discharged to, but the cooling air is typically forced to be expelled radially outwardly over the top of the pocket walls, thereby not utilizing the high pressure cooling flow to contribute to produce work/torque.
According to one aspect of the invention, a turbine rotor blade includes a tip portion having a pressure tip wall and a suction tip wall, a tip leading edge and a tip trailing edge, wherein the pressure tip wall and the suction tip wall define a trailing edge tip thickness. Also included is a squealer cavity at least partially defined by the pressure tip wall and the suction tip wall, the squealer cavity including a trench extending fully to the tip trailing edge to form an open flow path out of the tip trailing edge. Further included is a suction side wall and a pressure side wall extending from a root portion of the turbine rotor blade to the tip portion, wherein the suction side wall and the pressure side wall define a trailing edge blade thickness that is less than the trailing edge tip thickness.
According to another aspect of the invention, a turbine section of a turbine system includes a plurality of turbine rotor blades forming a plurality of turbine stages, wherein each of the plurality of turbine rotor blades includes a leading edge a trailing edge, a suction side wall and a pressure side wall, wherein the suction side wall and the pressure side wall define a trailing edge blade thickness. Also included is a tip portion of at least one of the plurality of turbine rotor blades having a pressure tip wall and a suction tip wall, a tip leading edge and a tip trailing edge, wherein the pressure tip wall and the suction tip wall proximate the tip trailing edge define a trailing edge blade thickness, the trailing edge tip thickness being greater than the trailing edge blade thickness.
According to yet another aspect of the invention, a gas turbine engine includes a compressor section, a combustion section, and a turbine section. The turbine section includes a plurality of turbine rotor blades forming a plurality of turbine stages, wherein each of the plurality of turbine rotor blades includes a leading edge a trailing edge, a suction side wall and a pressure side wall, wherein the suction side wall and the pressure side wall define a trailing edge blade thickness. The turbine section also includes a tip portion of at least one of the plurality of turbine rotor blades having a pressure tip wall and a suction tip wall, a tip leading edge and a tip trailing edge, wherein the pressure tip wall and the suction tip wall proximate the tip trailing edge define a trailing edge tip thickness that is greater than the trailing edge blade thickness.
These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.
The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which:
The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings.
Referring to
In operation, air flows into the compressor 12 and is compressed into a high pressure gas. The high pressure gas is supplied to the combustor assembly 14 and mixed with fuel, for example natural gas, fuel oil, process gas and/or synthetic gas (syngas), in the combustor chamber 18. The fuel/air or combustible mixture ignites to form a high pressure, high temperature combustion gas stream, which is channeled to the turbine 24 and converted from thermal energy to mechanical, rotational energy.
Referring now to
The pressure side wall 48 and the suction side wall 50 are spaced apart in the circumferential direction over the entire radial span of the turbine rotor blade 40 to define at least one internal flow chamber or channel for channeling cooling air through the turbine rotor blade 40 for cooling thereof. Cooling air is typically bled from the compressor section 12 in any conventional manner. The inside of the turbine airfoil blade 40 may have any configuration including, for example, serpentine flow channels with various turbulators therein for enhancing cooling air effectiveness, with cooling air being discharged through at least one, but typically a plurality of outlet holes 56 located at the tip portion 46 of the turbine rotor blade 40 and, more particularly, proximate a squealer cavity 80 that will be described in detail below in conjunction with the tip portion 46.
The tip portion 46 includes a tip plate 60 disposed atop the radially outer ends of the pressure side wall 48 and the suction side wall 50, where the tip plate 60 bounds the internal cooling cavities. The tip plate 60 may be integral to the turbine rotor blade 40 or may be welded into place. A pressure tip wall 62 and a suction tip wall 64 may be formed on the tip plate 60. Generally, the pressure tip wall 62 extends radially outwardly from the tip plate 60 and extends axially from a tip leading edge 68 to a tip trailing edge 70. Generally, the pressure tip wall 62 forms an angle with the tip plate 60 that is approximately 90°, though this may vary. The path of pressure tip wall 62 is adjacent to or near the termination of the pressure side wall 48 (i.e., at or near the periphery of the tip plate 60 along the pressure side wall 48).
Similarly, the suction tip wall 64 generally extends radially outwardly from the tip plate 60 and extends axially from the tip leading edge 68 to the tip trailing edge 70. The path of the suction tip wall 64 is adjacent to or near the termination of the suction side wall 50 (i.e., at or near the periphery of the tip plate 60 along the suction side wall 50). The height and width of the pressure tip wall 62 and/or the suction tip wall 64 may be varied depending on best performance and the size of the overall turbine assembly. As shown, the pressure tip wall 62 and/or the suction tip wall 64 may be approximately rectangular in shape, although other shapes are also possible.
The pressure tip wall 62 and the suction tip wall 64 generally form what is referred to herein as the squealer cavity 80. The squealer cavity 80 may include any radially inward extending depression or cavity formed on or within the tip portion 46. Generally, the squealer cavity 80 has a similar shape or form as the turbine rotor blade 40, though other shapes are possible, and is typically bound by the pressure tip wall 62, the suction tip wall 64, and an inner radial floor, which herein has been described as the tip plate 60.
As best illustrated in
A local increase in thickness along the trailing edge is provided, including the tip trailing edge 70 and possibly the trailing edge 54 of the main body portion 42 of the blade. The increase in thickness is gradual and widens in a radially outward direction of the turbine rotor blade 40. The increase may be made in a linear manner or in a curve of higher order (
The widened region of the trailing edge a trench 84 that is part of the squealer cavity 80, the advantages of which will be described in detail below. Inclusion of the at least one winglet 82 provides additional benefits. One benefit associated with the outwardly flared region, particularly in embodiments associated with the tip suction wall 64, the tip region leakage is reduced, thereby improving efficiency of the turbine section 24. This is due to weakening of tip leakage vortices proximate the tip portion 46 of the turbine rotor blade 40, which tends to inhibit flow at this region. Another benefit associated with the at least one winglet 82 relates to further thickening of the tip trailing edge 70. This enhanced thickening of the tip trailing edge 70 further accommodates the trench 84 that is part of the squealer cavity 80.
The trench 84 comprises a depression, groove, notch, trench, or similar formation that is positioned at an aft end of the squealer cavity 80 and extends fully to the tip trailing edge 70 of the tip portion 46, thereby forming a flow path for a cooling flow that opens directly out of the tip trailing edge 70 into a main flow path of the turbine section 24. The trench 84 may comprise several different shapes, sizes, alignments, and configurations. For example, as illustrated in
The cross-sectional profile of the trench 84 may be approximately semi-elliptical in nature. Alternatively, though not depicted in the figures, the profile of the trench 84 may be rectangular, semi-circular, triangular, trapezoidal, “V” shaped, “U” shaped and other similar shapes, as well as other combinations of profiles and filet radii. The edge formed between the top of the pressure tip wall 62, the suction tip wall 64 and the radially aligned walls of the trench 84 may be sharp (i.e., a 90 degree corner) or, in some cases, more rounded in nature.
The depth of the trench 84 may be substantially constant as it extends toward the tip trailing edge 70. Note that as used herein, the depth of the trench 84 is meant to refer to the maximum radial height of the trench 84 at a given location on its path. Thus, in the case of a semi-elliptical profile, the depth of the trench 84 occurs at the inward apex of the elliptical shape. In other embodiments, the depth of the trench 84 may vary to become less or more deep relative to the upstream, originating location of the trench 84. Similarly, the width of the trench 84 may be constant or vary along an entire length of the trench 84.
Regardless of the precise configuration of the trench 84, the localized thickened tip trailing edge 70 and the at least one winglet 82 facilitates a widening of the trench 84 at the tip trailing edge 70. In particular, a space between outer portions of the pressure tip wall 62 and the suction tip wall 64 at the tip trailing edge 70 is defined and is referred to as a trailing edge tip thickness. Similarly, a space between outer portions of the pressure side wall 48 and the suction side wall 50 at the trailing edge 54 of the main body portion 42 is defined and is referred to as a trailing edge blade thickness. The trailing edge tip thickness is greater than the trailing edge blade thickness. In other words, a localized thickening of the trailing edge region of the overall turbine rotor blade 40. In one embodiment, the trailing edge tip thickness is about 1.1 times to about 3.0 times the thickness of the trailing edge blade thickness. In another embodiment, the trailing edge tip thickness is about 1.5 times to about 2.5 times the thickness of the trailing edge blade thickness. In yet another embodiment, the trailing edge tip thickness is about 1.95 times to about 2.05 times the thickness of the trailing edge blade thickness. The preceding examples are merely illustrative of the fact that the trailing edge tip thickness is greater than the trailing edge blade thickness. The local increase in the thickness at the tip of the blade adds extra local mass at the tip trailing edge portion. Addition of this mass on the tip will change the frequency of the blade in a favorable direction which assists in meeting aerodynamics requirements of the blade. The local increase in thickness targets a local mass addition at the trailing edge portion of the tip. Due to the location's high kinetic energy, it is very sensitive to changes in mass and stiffness, which will change the airfoil's mode shapes and frequencies. These changes in mode shapes and frequencies are used to the blade's advantage to avoid aeromechanic drivers and to meet design requirements.
As noted above, the squealer cavity 80 includes the plurality of outlet holes 56 for expulsion of cooling flow. The plurality of outlet holes 56 is also present within the trench 84 for the provision of cooling air to this region of the squealer cavity 80 to keep the surrounding surface area of the tip portion 46 cool by convecting away heat and insulating the part from the extreme temperatures of the working fluid. More particularly, the coolant may better cool the tip portion 46 proximate the tip trailing edge 70. As shown, the trench cooling apertures may be regularly spaced through the trench 84 and positioned on the floor of the trench 84, i.e., near the deepest portion of the trench 84.
Advantageously, the embodiments described above decrease the tip leakage flow and weaken the tip leakage vortex, thereby reducing losses that directly impact overall turbine system efficiency. Increasing the trailing edge thickness at the tip portion 46, in combination with the winglet 82, allows higher width of the trench 84 at the squealer cavity 80. A wider trench facilitates a wider space for the cooling flow to escape from the trench opening at the immediate tip trailing edge 70, in contrast to closed squealer cavities that require the cooling flow to escape from the radially outward portion of the squealer cavity 80. By increasing the trailing edge thickness only proximate the tip portion, an aerodynamic benefit is achieved by accommodating the wider trench. In particular, the trench 84 better utilizes the cooling flow to extract work from the cooling flow as the cooling flow imparts a circumferential force along the trench wall as it flows toward the trailing edge. Rather than wasting the cooling flow by simply expelling it from the squealer cavity, the cooling flow assists in the rotation of the blade.
While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.
