Patent Application
20140099190
The present disclosure generally relates to gas turbine engine turbine housings and more particularly to a gas turbine engine turbine housing with cooling holes.
Gas turbine engines include-compressor, combustor, and turbine sections. Portions of a gas turbine engine are subject to high temperatures. In particular, the first stages of the turbine section are subject to such high temperatures that these stages are often cooled by directing relatively cool air through internal cooling passages.
U.S. Pat. No. 4,820,123, to K. Hall, describes a dirt removal means for air cooled blades of a gas turbine engine. The dirt removal means uses louvers stamped out of sheet metal that overlie the inlets of the blades' internal cooling passages. The louvers deflect dirt entrained in cooling air through a high velocity air stream and allow a cleaner portion of the cooling air to flow through the cooling passages of the blades.
The present disclosure is directed toward overcoming one or more of the problems discovered by the inventors.
A turbine housing includes a support housing with an annular shape. The support housing includes an outer surface and an inner surface. The inner surface is located radially inward from the outer surface. A plurality of cooling holes extends through the support housing from the outer surface to the inner surface. The cooling holes include a stadium shape. The stadium shape includes a rectangle with a top and bottom of a first length and ends capped with semicircles with a radius of a second length. The cooling holes are disposed at an angle from thirty to sixty degrees. The angle is between a line projected to the outer surface from an axis of the turbine housing and a line extending through a center of each cooling hole parallel to the first length.
A method for modifying a turbine housing with an annularly shaped first stage turbine support housing is also provided. The first stage turbine support housing includes a plurality of round cooling holes circumferentially overlapping with at least one opening to an airfoil cooling passage when the first-stage turbine nozzle support housing is assembled in a gas turbine engine. The method includes plugging the round cooling holes. The method also includes determining locations for new cooling holes. The new cooling holes cannot circumferentially overlap with the openings to airfoil cooling passages when the turbine housing is installed in a gas turbine engine. The method further includes machining new holes through the first stage turbine support housing.
The systems and methods disclosed herein include cooling holes of a gas turbine engine turbine housing. In embodiments the cooling holes are elongated or stadium shaped and configured to provide cooling air to the gas turbine engine turbine nozzles. The cooling holes can be angled so as to not circumferentially overlap with the openings far the cooling air passages of the turbine nozzles when the turbine housing is installed in the gas turbine engine. This configuration can prevent large particles from entering the turbine nozzle cooling air passages. Large particles entering the turbine nozzle cooling air passages may clog or block the cooling air passages, inhibiting the cooling of the turbine nozzles.
Air 10 enters an inlet 15 as a “working fluid” and is compressed by the compressor 200. Fuel 35 is added to the compressed air in the combustor 300 and then ignited to produce a high energy combustion gas. Energy is extracted from the combusted fuel/air mixture via the turbine 400 and is typically made usable via a power output coupling 5. The power output coupling 5 is shown as being on the forward side of the gas turbine engine 100, but in other configurations it may be provided at the aft end of gas turbine engine 100. Exhaust 90 may exit the system or be further processed (e.g., to reduce harmful emissions or to recover heat from the exhaust 90).
The compressor 200 includes a compressor rotor assembly 210 and compressor stationary vanes (“stators”) 250. The compressor rotor assembly 210 mechanically couples to shaft 120. As illustrated, the compressor rotor assembly 210 is an axial flow rotor assembly. The compressor rotor assembly 210 includes one of more compressor disk assemblies 220. Each compressor disk assembly 220 includes a compressor rotor disk that is circumferentially populated with compressor rotor blades. Stators 250 axially precede each of the compressor disk assemblies 220.
The turbine 400 includes a turbine rotor assembly 410, turbine nozzles 450, and a turbine housing 430. The turbine rotor assembly 410 mechanically couples to shaft 120. As illustrated, the turbine rotor assembly 410 is an axial flow rotor assembly. The turbine rotor assembly 410 includes one or more turbine disk assemblies 420. Each turbine disk assembly 420 includes a turbine rotor disk that is circumferentially populated with turbine rotor blades. Turbine nozzles 450 axially precede each of the turbine rotor assemblies 420. The turbine nozzles 450 have circumferentially distributed turbine nozzle vanes. The turbine nozzle vanes helically reorient the combustion gas that is delivered to the turbine rotor blades where the energy in the combustion gas is converted to mechanical energy and rotates the shaft 120.
The turbine nozzles 450 may be coupled to turbine housing 430. The turbine nozzle 450 closest to the combustor 300 may be considered the first stage turbine nozzle 451. The turbine nozzles 450 may be cooled by routing cooling air from the compressor 200. The cooling air may be routed through cooling holes 431 (shown in
The various components of the compressor 200 are housed in a compressor case 201 that may be generally cylindrical. The various components of the combustor 300 and the turbine 400 are housed, respectively, in a combustor case 301 and a turbine case 401.
As can be seen in
Turbine housing 430 is configured to include cooling holes 431 distributed circumferentially about first stage housing 432. Cooling holes 431 radially pass through first stage housing 432 from outer surface 433 to inner surface 434.
Forward flange 436 may protrude radially outward from an axially forward end of turbine housing 430. In one embodiment forward flange 436 protrudes from first stage housing 432. Forward flange 436 may include mounting holes 437. Mounting holes 437 are circumferentially located about flange 436. Each mounting hole 437 axis may be parallel to the axis of turbine housing 430. A coupler 445, such as a bolt, may be inserted into each mounting hole 437 to secure the turbine housing 430 to the gas turbine engine as shown in
Referring now to
First stage turbine nozzle (“nozzle”) 451 is located radially inward from first stage housing 432. Nozzle 451 includes an outer wall 452, an inner wall 453, and one or more airfoils 454. Inner wall 453 is adjacent to a diaphragm 414 and is located radially inward from outer wall 452. Outer wall 452 and inner wall 453 are connected by one or more airfoils 454. Outer wall 452 is located radial adjacent to first stage housing 432. First stage housing 432 and outer wall 452 may be configured to define cavity 449 there between. Cooling holes 431 are configured to be in flow communication with cavity 449. In one embodiment, cavity 449 is defined by first stage housing 432 and the outer wall 452 of multiple nozzles 451 and circumferentially extends completely around first stage housing 432.
As can be seen in
Cooling holes 431 are circumferentially offset or clocked relative to airfoils 454 and are configured to not overlap with a radially outward projection of the openings into passages 455. Each of the cooling holes 431 may supply cooling air to multiple, for example, two, of the airfoils 454. Referring to
Cooling holes 431 are elongated holes. Cooling holes 431 may be of any elongated shape. The length of cooling holes 431 in an elongated direction may be more than three-quarters the axial length of the first stage housing 432. In one embodiment, cooling holes 431 are slots with a stadium shape, a rectangle with a top and bottom of a first length and ends capped with semicircles with a radius of a second length. In the depicted embodiment of
Each cooling hole 431 may be disposed at an angle 98. Angle 98 may be defined as the angle between line 96, a line projected to outer surface 433 from the axis of turbine housing 430, and line 97, a line extending through the center of cooling holes 431 in the elongated direction at outer surface 433. In one embodiment angle 98 is from thirty to sixty degrees so as to not overlap with a radially outward projection of passages 455. In the configuration shown, one cooling hole 431 provides cooling air to two airfoils 454.
The embodiments discussed above describe the first stage turbine nozzle support housing 432 being configured with cooling holes 431. However, any of support housings 440 may be configured with cooling holes 431 for providing a cooling path to the various stage nozzles of a gas turbine engine. Support housings 440 may also include an outer surface and an inner surface.
Referring again to
One or more of the above components (or their subcomponents) may be made from stainless steel and/or durable, high temperature materials known as “superalloys”. A superalloy, or high-performance alloy, is an alloy that exhibits excellent mechanical strength and creep resistance at high temperatures, good surface stability, and corrosion and oxidation resistance. Superalloys may include materials such as HASTELLOY, INCONEL, WASPALOY, RENE alloys, HAYNES alloys, INCOLOY, MP98T, TMS alloys, and CMSX single crystal alloys.
Gas turbine engines may be suited for any number of industrial applications such as various aspects of the oil and gas industry (including transmission, gathering, storage, withdrawal, and lifting of oil and natural gas), the power generation industry, cogeneration, aerospace, and other transportation industries.
Operating efficiency of a gas turbine engine generally increases with a higher combustion temperature. Thus, there is a trend in gas turbine engines to increase the temperatures. Gas reaching first stage turbine nozzles 451 from combustor outlet 330 may be 1000 degrees Fahrenheit or more.
To operate at such high temperatures, the first stage turbine nozzles 451 have internal cooling passages. A portion of the compressed air from the compressor 200 of the gas turbine engine is diverted from entering the combustor 300 and is routed to the internal cooling passages. The cooling air lowers the temperature of the first stage turbine nozzles 451 so as to deter corrosion, deformation, or melting. The internal cooling passages often include many small holes.
Particles can become entrained in the cooling air. The particles may be ingested by the gas turbine engine from its environment or self-generated within the gas turbine engine. The particles can accumulate in the internal cooling passages and interfere with cooling of turbine components such as the first stage turbine nozzles 451. The small holes in the internal cooling passages can clog and block the flow of cooling air in some areas. Accumulated particles can also cover surfaces in the internal cooling passages and form an insulating layer that reduces cooling effectiveness.
As shown in
Cooling holes 431 may deter deterioration of the cooling of the first stage turbine nozzles 451 due to particle contamination. The configuration of cooling holes 431 relative to passages 455 of the first stage turbine nozzles 451 may create a torturous path for cooling air and any entrained particles. The torturous path may avoid accumulation of particles in passages 455. Cooling holes 431 may be circumferentially offset or clocked relative to the internal cooling passages 455. Without a direct path from cooling holes 431 into passages 455 the particles may be broken into smaller pieces that may pass through passages 455 without accumulating. Other particles may accumulate within cavity 449 rather than in passages 455. Still other particles may be deflected away from passages 455.
Some gas turbine engines have used a screen to shield the first stage turbine nozzles 451 from particles. However, the screens themselves are subject to clogging that can block the flow of cooling air to the nozzle and interfere with cooling. Furthermore, the screens are prone to deterioration that can contribute to the particles reaching the first stage turbine nozzles 451.
Other gas turbine engines, such as the one depicted in
Plugging the existing round cooling holes is followed by determining locations for new cooling holes at step 520. The locations are selected such that the new cooling holes do not circumferentially overlap passages 455 when turbine housing 430 is installed into a gas turbine engine. The new cooling holes may be cooling holes 431. Each cooling hole 431 may be centered between the trailing edge 459 of the leading airfoil 454a and the leading edge 458 of the trailing airfoil 454b as illustrated in
At step 530, determining the location for new cooling holes is followed by machining the new cooling holes through first stage housing 432. Machining the new cooling holes may include machining a portion of the plug. Cooling holes 431 may have a stadium shape. The stadium shape may be easily machined and may reduce production costs.
The method of modifying turbine housing 430 may include removing an existing screen or particle deflector 470. In some embodiments the particle deflector 470 may be reinstalled. In other embodiments, a new particle deflector 470 is installed. The method may also include removing turbine housing 430 from a gas turbine engine prior to plugging the existing round cooling holes and reinstalling turbine housing 430 into a gas turbine engine after machining the new cooling holes.
The preceding detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. The described embodiments are not limited to use in conjunction with a particular type of gas turbine engine. Hence, although the present disclosure, for convenience of explanation, depicts and describes cooling holes, it will be appreciated that the cooling holes in accordance with this disclosure can be implemented in various other configurations and used in other types of machines. Furthermore, there is no intention to be bound by any theory presented in the preceding background or detailed description. It is also understood that the illustrations may include exaggerated dimensions to better illustrate the referenced items shown, and are not consider limiting unless expressly stated as such.
