Turbine airfoil with near-wall leading edge multi-holes cooling

An air cooled airfoil used in a gas turbine engine includes a showerhead arrangement to cool the leading edge. A cooling supply cavity supplies cooling air to the airfoil, and is metered into an impingement cavity through at least one metering hole from the cooling supply cavity. A plurality of second impingement cavities are located between the first impingement cavity and the leading edge, and each second impingement cavity includes at least one second metering hole to meter cooling air into the second impingement cavity. Each second impingement cavity includes a plurality of film cooling holes to discharge cooling air to the leading edge surface of the airfoil. The second impingement cavities are located adjacent to each other, with one on the leading edge, another on the pressure side, and the third on the suction side.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates generally to fluid reaction surfaces and more specifically to turbine airfoils with film cooling.

A gas turbine engine includes a turbine section in which a hot gas flow passes through the rotor blades and stationary vanes or nozzles to extract energy from the flow. The efficiency of the gas turbine engine can be increased by providing a higher flow temperature into the turbine. However, the temperature is limited to the material capabilities of the parts exposed to the hot gas flow.

One method of allowing for higher turbine temperatures is to provide cooling of the first and second stages of the turbine. Complex internal cooling passages have been disclosed to provide high cooling capabilities while using lower flow volumes of cooling air. Since the cooling air used in the turbine airfoils is typically bleed off air from the compressor, using less bleed off air for cooling also increases the efficiency of the engine.

A Prior Art airfoil leading edge is cooled with backside impingement in conjunction with a showerhead film cooling and is shown inFIG. 1. The airfoil10includes a cooling air supply cavity22, a metering hole23connecting the supply cavity22to an impingement cavity24, and a plurality of impingement holes27to discharge cooling air to the leading edge surface of the airfoil10. a mid-chord three-pass forward flowing serpentine cooling circuit includes a first leg channel40, a second leg channel36, and a third leg channel34. Film cooling holes35deliver cooing air to the suction side and film cooling holes37deliver cooling air to the pressure side from the third leg channel34. Turbulators50are positioned along the walls of the channels and cavity to promote heat transfer to the cooling air. A trailing edge cooling slots42and cooling air exit holes44discharge cooling air from the first leg channel40to the trailing edge of the airfoil.

The showerhead film rows are fed cooling air from a common cooling supply cavity22and discharged at various gas side pressures. The pressure at each of the gas side locations can vary substantially as the hot gas flow accelerates around the nose of the leading edge. The minimum pressure ratio across the showerhead holes27is typically set by back-flow margin requirements, and the pressure ratio (and flow) across all of the other film cooling rows becomes substantially a function of the gas-side pressure. Backflow occurs when the pressure of the hot gas flow outside the leading edge is higher than the cooling air pressure inside the cooling supply cavity24, resulting in the hot gas flowing into the inside of the airfoil. As a result of this cooling design, the cooling flow distribution and pressure ratio across the showerhead film holes27for the pressure side and suction side film row is predetermined by the supply pressure.

U.S. Pat. No. 7,011,502 B2 issued to Lee et al on Mar. 14, 2006 entitled THERMAL SHIELD TURBINE AIRFOIL. Discloses an airfoil with a longitudinal first inlet channel (56 in this patent) connected by a row of impingement holes (48 in this patent) to a longitudinal channel (42 in this patent), being connected by a plurality of film cooling holes (50 in this patent) to the leading edge surface of the airfoil. In the Lee et al patent, the longitudinal channel is also connected to bridge channels on the pressure side and suction sides of the longitudinal channel. Only one multi-impingement channel is used in the Lee et al invention to supply film cooling holes as opposed to the three separate multi-impingement channels of the present invention.

U.S. Pat. No. 4,859,147 issued to Hall et al on Aug. 22, 1989 entitled COOLED GAS TURBINE BLADE discloses an airfoil with a showerhead arrangement having a cooling air supply cavity (24 in this patent), an impingement cavity (28 in this patent) connected to the cooling supply cavity by three slots (26 in this patent), and three film cooling holes (30 in this patent) connected to the impingement cavity.

U.S. Pat. No. 6,379,118 B2 issued to Lutum et al on Apr. 30, 2002 entitled COOLED BLADE FOR A GAS TURBINE discloses a similar showerhead arrangement to that above of the Hall et al patent. A cooling air supply cavity (50 in this patent) is connected to an impingement cavity (47 in this patent) through two impingement cooling holes (49 in this patent), and three film cooling holes (48 in this patent) are connected to the impingement cavity. Both of the Hall et al and Lutum et al patents lack the first impingement cavity and the second impingement cavity in series, and the multi-impingement cavities of the present invention.

It is an object of the present invention to alleviate the problem associated with the turbine airfoil leading edge showerhead pressure ration or blowing ratio of the prior art. Another object of the present invention is to improve the efficiency of a gas turbine engine by providing improved cooling for the leading edge region of the airfoil.

BRIEF SUMMARY OF THE INVENTION

The present invention is a turbine airfoil having a showerhead film cooling holes arrangement to cool the leading edge, a cooling supply channel to supply cooling air, an impingement cavity connected to the supply cavity through a metering hole, and a plurality of multi-impingement cavities positioned between the showerhead film cooling holes and the impingement cavity, with multi-metering holes used to regulate the cooling air flow.

Cooling air is supplied through the leading edge cooling supply cavity and impinges onto the backside of the airfoil leading edge inner surface. This firstly provides impingement cooling of the airfoil leading edge section. Multiple metering holes are then used for each individual impingement cavity to provide a desired pressure and flow rate to the intermediate coolant pressure cavity. Multiple impingement cavities can be used in the spanwise direction for tailoring the blade spanwise hot gas side pressure and heat load conditions. In addition, the multiple metering holes also provide backside impingement cooling to the airfoil leading edge region at much closer distance to the airfoil exterior hot surface. Internal cooling pressure for each individual impingement cavity can also be regulated by the multiple metering holes. With the use of the cooling design of the present invention, a reduction of the overall pressure ratio across the leading edge film cooling holes can be achieved. At a given cooling flow rate and lower pressure ration across the film holes, it yields a large number of leading edge film holes thus forming a multi-hole cooling mechanism for the airfoil leading edge.

The advantages of the present invention over the prior art showerhead arrangement are numerous and include the following. The airfoil leading edge showerhead cooling flow and pressure are regulated in the blade chordwise and spanwise directions. Maximize the usage of the cooling air for a given airfoil inlet gas temperature and pressure profile is achieved. Increase number of leading edge film cooling holes will increase the leading edge film coverage. This translates to a better film effectiveness and lower leading edge metal temperature. Increase number of leading edge film cooling holes also increases the overall leading edge internal convection cooling capability and consequently reduces the blade leading edge metal temperature. Leading edge impingement cooling at reduced distance will lower the airfoil leading edge metal temperature. Lower the pressure ratio or blowing ratio across the showerhead film holes will minimize the coolant penetration into the gas path, yielding a good buildup of coolant sub-boundary layer next to the airfoil surface. This translates into higher leading edge film effectiveness and a lower metal temperature.

DETAILED DESCRIPTION OF THE INVENTION

The airfoil of the present invention is shown inFIG. 2and includes the features described above with respect to the prior artFIG. 1airfoil. In addition,FIG. 2includes three multi-impingement cavities26, one on the nose of the leading edge of the airfoil, a second one on the pressure side, and a third on the suction side of the airfoil. Multi-metering holes25connect the impingement cavity24to the multi-impingement and diffusion cavities26. Each multi-impingement and diffusion cavity26includes a plurality of multi-showerhead holes27that open onto the respective surface of the airfoil to provide film cooling.

On the pressure side multi-impingement cavity26, the downstream-most film cooling hole27discharges onto the pressure side surface at a location about where the forward-most side wall of the cooling supply cavity22ends. On the suction side multi-impingement cavity26, the downstream-most film cooling hole27discharges onto the suction side surface at a location about where the closest point of the impingement cavity24is to the suction side wall of the airfoil as shown inFIG. 2.

The operation of the showerhead design of the present invention inFIG. 2is described below. Cooling air is supplied from a source such as a compressor to the leading edge cooling supply cavity22, passes though a plurality of the first metering holes23and into the impingement cavity24. Cooling air flow through the first metering holes provides impingement cooling for the airfoil leading edge section. Cooling air then flows through the multi-metering holes25and into the respective multi-impingement cavity26. This flow provides backside impingement cooling to the airfoil leading edge region at a much closer distance to the airfoil exterior hot surface. Cooling air is then discharged to the airfoil external surface through a plurality of film cooling holes27associated with the three multi-impingement cavities26. In order to ensure adequate cooling to each portion of the leading edge region, the various metering holes and film cooling holes can be sized to direct the proper amount of cooling air to that particular portion of the airfoil.

The leading edge cooling supply channel22extends from the platform region of the airfoil up to the tip region. The impingement cavity24can be a single cavity extending from the platform to the tip region, or it can be formed of a plurality of individual cavities aligned with the cooling supply channel22. Each individual impingement cavity24can be connected to the cooling supply channel22by one or more of the first metering holes23.

The multi-impingement and diffusion cavities26can also be one long cavity extending along the leading edge of the airfoil, or be formed from a plurality of separate multi-impingement cavities aligned along the leading edge, each connected to the impingement cavity through one or more multi-metering holes25.