The present invention relates generally to gas turbine engine components, and more particularly to internally cooled airfoils used in gas turbine engine components.
Temperatures within gas turbines may exceed 3000° F. (1650° C.), and cooling of turbine blades is very important in terms of blade longevity. The gas turbine engine operates by utilizing a compressor portion to compress atmospheric air to 10-25 times atmospheric pressure and adiabatically heating the air to between about 800°-1250° F. (427° C.-677° C.) in the process. This heated and compressed air is directed into a combustor, where it is mixed with fuel. The fuel is ignited, and the combustion process heats the gases to very high temperatures, in excess of 3000° F. (1650° C.). These hot gases pass through the turbine, where airfoils fixed to rotating turbine disks extract energy to drive the fan and compressor of the engine and the exhaust system, where the gases provide sufficient thrust to propel the aircraft. To improve the efficiency of operation of the aircraft engine, combustion temperatures have been raised. Of course, as the combustion temperature is raised, steps must be taken to prevent thermal degradation of the materials forming the flow path for these hot gases of combustion.
Aircraft gas turbine engines have a so-called High Pressure Turbine (HPT) to drive the compressor. The HPT is located aft of the combustor in the engine layout and experiences the highest temperature and pressure levels (nominally—3000° F. (1850° C.) and 300 psia, respectively) developed in the engine. The HPT also operates at very high rotational speeds (10,000 RPM for large high-bypass turbofans, 50,000 for small helicopter engines). There may be more than one stage of rotating airfoils in the HPT. In order to meet life requirements at these levels of temperature and pressure, HPT components are air-cooled, typically from bleed air taken from the compressor, and are constructed from high-temperature alloys.
Without cooling, turbine blades would rapidly deteriorate. Improved cooling for turbine blades is very desirable, and much effort has been devoted by those skilled in the blade cooling arts to devise improved geometries for the internal cavities within turbine blades, in order to enhance cooling. Since the combustion gases are hot, the turbine vanes and blades are typically cooled with a portion of compressor air bled from the compressor for this purpose. Diverting any portion of the compressor air from use in the combustor necessarily decreases the overall efficiency of the engine. Accordingly, it is desired to cool the vanes and blades with as little compressor bleed air as possible.
Turbine rotor blades with internal cooling circuits are typically manufactured using an investment casting process commonly referred to as the lost wax process. This process comprises enveloping a ceramic core defining the internal cooling circuit in wax shaped to the desired configuration of the turbine blade. The wax assembly is then repeatedly dipped into a liquid ceramic solution such that a hard ceramic shell is formed thereon. Next, the wax is removed from the shell by heating so that the remaining mold consists of the internal ceramic core, the external ceramic shell and the space therebetween, previously filled with wax. The empty space is then filled with molten metal. After the metal cools and solidifies, the external shell is broken and removed, exposing the metal that has taken the shape of the void created by the removal of the wax. The internal ceramic core is dissolved via a leaching process. The resulting metal component has the desired shape of the turbine blade with the internal cooling circuit and cooling orifices.
In casting turbine blades with serpentine cooling circuits, the internal ceramic core is formed as a serpentine element having a number of long, thin branches. This presents the challenge of making the core sturdy enough to survive the pouring of the metal while maintaining the stringent requirements for positioning the core. Currently, the trail edge slots are cast utilizing substantially oval core insert projections that provide a slot size sufficiently large, typically greater than about 0.013 inches to provide strength to the core and provide sufficient cooling along the trail edge of the turbine component.
Accordingly, there is a need for an airfoil component in which cooling fluid flow through the trail edge slots is decreased, while the core during fabrication is sufficiently robust to withstand casting of the turbine component.
A first aspect of the present invention includes an airfoil component having a body having a leading edge and a trailing edge. The component includes an internal cooling passageway and an elongated opening in communication with the internal cooling passageway. The opening is configured with a geometry that provides structural stability during casting and has a cross-section that sufficiently restricts airflow through the opening to provide efficient component operation.
Another aspect of the present invention includes a gas turbine engine component casting insert having a ceramic insert body. The insert further includes core insert projections extending from the body having outer edge projections connected by a web portion. The outer edge projections have a thickness along the web portion that is greater than the thickness of the web portion. The core insert projections and web portion have sufficient structural stability to permit casting around the insert.
Still another aspect of the present invention includes a method for casting a gas turbine engine component. The method includes providing a core insert having core insert projections with outer edge projections connected by a web portion. The outer edge projections have a thickness along the web portion that is greater than the thickness of the web portion. A gas turbine engine component is cast over the core insert. The core insert is then removed to provide a gas turbine engine having cooling passages and elongated openings in communication with the cooling passages. The opening formed from removal of the core insert have a geometry that sufficiently restricts airflow through the opening to provide efficient component operation.
An advantage of an embodiment of the present invention is that the amount of bleed air from the compressor may be reduced and gas turbine engine operation may be more efficient.
Another advantage of an embodiment of the present invention is that the reduced cooling flow of cooling fluid from the trailing edge reduces or eliminates the need for other fluid flow restrictions, such as root plates.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings which illustrate, by way of example, the principles of the invention.
Wherever possible, the same reference numbers are used throughout the drawings to refer to the same or like parts.
Illustrated in
As shown in
The trailing edge openings 211 receive a flow of cooling fluid 204 wherein the cooling fluid 204 flows through the trailing edge openings 211 and is discharged from the airfoil 103. Cooling air discharge apertures or trailing edge openings 211 are preferably designed to provide impingement cooling of the trailing edge 107. The present invention utilizes a configuration of trailing edge openings 211 that provides efficient cooling, without the need for a root plate or other cooling fluid 204 restriction, allowing for efficient gas turbine engine operation.
Although an exemplary gas turbine blade 100 is illustrated in
An embodiment of the present invention utilizes a ceramic core 501 that is formed utilizing cores insert projections 503 having a geometry corresponding to the pinched geometry trailing edge openings 211. The pinched trail edge openings 211 are cast utilizing ceramic core 501 insert projections 503 that provide a slot geometry having a pinched geometry to provide strength to the ceramic core 501 and provide sufficient cooling along the trailing edge opening 211 of the turbine component.
In another embodiment of the invention, the trailing edge opening 211 may include a plurality of second minimum dimensions 803 between first end 805 and second end 807, for example, wherein the second maximum dimension 804 is located at a location near the center of first dimension 801 a substantially T-shaped opening 211. Likewise, the second maximum dimension 804 may extend in two directions past second minimum dimension 803. The present invention is not limited to the above configurations of the first dimension 801, the second minimum dimension 803 and second maximum dimension 804 and may include a plurality of each or both of the second minimum dimension 803 and second maximum dimension 804. The present invention utilizes the cross-sectional geometry formed to provide a reduced amount of cooling fluid 204 flow, while providing a sufficiently strong ceramic core 501 insert that allows casting of the blade 100. The cooling fluid 204 is therefore used more efficiently and less cooling fluid 204 is bled from the compressor for increasing the overall efficiency of operation of the gas turbine engine.
While the invention has been described with reference to a preferred embodiment, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this invention, but that the invention will include all embodiments falling within the scope of the appended claims.