Patent Publication Number: US-8118554-B1

Title: Turbine vane with endwall cooling

Description:
GOVERNMENT LICENSE RIGHTS 
     None. 
     CROSS-REFERENCE TO RELATED APPLICATIONS 
     None. 
     BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates generally to a gas turbine engine, and more specifically to a turbine vane. 
     2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98 
     A gas turbine engine, such as an industrial gas turbine (IGT) engine, includes a turbine with multiple rows or stages or stator vanes that guide a high temperature gas flow through adjacent rotors of rotor blades to produce mechanical power and drive a bypass fan, in the case of an aero engine, or an electric generator, in the case of an IGT. In both cases, the turbine is also used to drive the compressor. 
     It is well known that the efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine parts, such as the first stage guide vanes and rotor blades. Also, the turbine inlet temperature is limited to an amount of cooling that can be produced on a turbine vane or blade. Improved cooling capability will also allow for the turbine airfoils to be exposed to higher temperatures. Improved cooling will also allow for longer part life which results in longer engine run times or longer periods between engine breakdowns. 
     Another problem with the turbines is hot flow ingestion into a section of the turbine that is sensitive to the high temperatures such as the rim cavities or interstage gaps. Bow wave driven hot gas flow ingestion is created when the hot gas core flow enters a vane row where a leading edge of the vane induces a local blockage and thus creates a circumferential pressure variation at an intersection of the airfoil leading edge location of the vane. The leading edge of a turbine vane generates upstream pressure variations which can lead to hot gas ingress into the front gap. The leading edge of the turbine stator vanes generates an upstream pressure that is higher than the pressure inside the cavity.  FIG. 1  shows a prior art turbine vane with a bow wave effect located upstream of the turbine vanes. The high pressure upstream of the vane leading edge is greater than the pressure inside the cavity formed by the gap. As a result of the pressure differential, the hot gas will flow radially inward into the cavity. The ingested hot gas flows through the gap circumferentially inside the cavity towards the lower pressure zones. The hot gas then flows out at locations where the cavity pressure is higher than the local hot gas pressure. 
     In general, the size of the bow wave is a strong function of the vane leading edge diameter and distance of the vane leading edge to the endwall edge. The pressure variation in the tangential direction with the gap is sinusoidal. The amount of hot gas flow penetrating the axial gap increases linearly with the increasing axial gap width. It is therefore necessary to reduce the axial gap width to a minimum allowable by tolerance limits in order to reduce the hot gas ingress. 
     BRIEF SUMMARY OF THE INVENTION 
     It is an object of the present invention to provide for a turbine vane with an interstage gap in which the hot gas ingress into the gap is eliminated. 
     It is another object of the present invention to eliminate the ingress of hot gas flow caused by a differential pressure between the hot gas pressure and the cavity pressure from the bow-wave effect. 
     It is another object of the present invention to provide for a turbine stator vane with an endwall having a lower metal temperature in a bow wave region than in the prior art endwall. 
     It is another object of the present invention to provide for a turbine stator vane with an endwall having a much higher convective cooling than the backside impingement cooling method of the prior art. 
     These objectives and more can be achieved by the turbine vane with a directed cooling system in the airfoil leading edge section. The over-temperature caused by the bow wave ingress hot gas flow issued described above can be reduced or eliminated by the use of a directed cooling system into the airfoil leading edge section design of the vane. A backside impingement cooling in conjunction with multiple hole film cooling is used along a forward section of the airfoil leading edge root section. The multiple rows of film cooling holes is formed around the airfoil leading edge peripheral that will inject the film cooling air to form a film sub-layer for a baffle against the hot gas ingestion region from the downward draft of the hot core gas stream. Due to the cooling being inline with the endwall external heat load, the impingement onto the backside of the hot wall is then discharged as film cooling air will yield a very efficient method of cooling the hot wall surface. 
     The backside impingement and multiple hole film cooling circuit is formed around the airfoil leading edge root section at the endwall junction region by means of machining circumferential slots into the endwall. Impingement and film cooling holes are then machined into the inner and outer walls prior to welding a cap onto the edge of the cooling slot. The present embodiment retains an original design intent load path for the airfoil. The circumferential slots form multiple compartments that divide the endwall into multiple cooling zones. The multiple compartments of the endwall will minimize a pressure gradient effect for the cooling flow mal-distribution. Micro pin fins are also used on the backside of the impingement cavity to enhance the convection cooling effect. 
    
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
         FIG. 1  shows a cross section side view of a prior art turbine stator vane with the hot gas flow pattern and hot gas ingress flow into the outer diameter endwall and inner diameter endwall of the vane. 
         FIG. 2  shows cross section side view of a section of the vane endwall with the cooling circuit of the present invention. 
         FIG. 3  shows an isometric view of a close up of the leading edge endwall of the vane in  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The present invention is a turbine stator vane for an industrial gas turbine engine. However, the present invention is also usable in an aero engine stator vane as well  FIG. 2  shows a cross section side view of a stator vane leading edge endwall with the cooling circuit of the present invention. The stator vane includes a leading edge  11  that extends from an outer endwall (not shown) to an inner endwall  12 . The inner endwall  12  extends beyond the leading edge and curves downward to form the flow path for a hot gas flow that is passing through the turbine. This region is referred to as the endwall edge  13 . 
     The endwall edge  13  includes an outer surface and an inner surface that forms a cooling air supply cavity  21 . The endwall edge  13  includes a compartment slot or channel  14  that is machined from the endwall edge  13 . A number of compartment divider ribs  15  separate a number of compartments  14  from each other. The inner surface includes a number of impingement holes  22  that connect the cooling air supply cavity  21  to the compartment slots  14 . Each compartment slot  14  can have one row of impingement holes  22  or several rows of impingement holes  22 . The outer endwall surface includes a number of film cooling holes  16  that connect the compartment slots  14  to the outer surface of the endwall edge and discharge film cooling air. A cover plate  24  is secured to a bottom of the rail of the endwall edge to enclose the compartment slots  14 . 
       FIG. 3  shows an isometric view of the endwall edge  13  with an arrangement of the film cooling holes  16 . The airfoil leading edge  11  is shown intersecting with the endwall edge  13 , and the arrangement of film cooling holes  16  are shown in front of the leading edge and on the endwall edge that are all connected to the cooling air supply cavity  21 . The arrows pointing toward the film cooling holes and away from the leading edge  11  represent the hot gas ingress flow. The arrows exiting from the film cooling holes  16  represent the film cooling air discharged. Micro pin fins can be used on the inner walls of the compartment slots  14  to enhance the heat transfer effect from the hot metal to the cooling air. 
     Pressurized cooling air supplied to the cooling air supply cavity  21  will flow through the impingement cooling holes  22  and into the compartment slots  14  to provide impingement cooling for the backside wall of the endwall edge  13  outer surface that is exposed to the hot gas flow. The spent impingement cooling air within the compartment slots  14  then flows out through the film cooling holes  16  to form a layer of film cooling air on the outer endwall surface. The film cooling air will push up the bow wave hot gas to reduce or eliminate the prior art effects described above. 
     Major design features and advantages of the endwall cooling circuit of the present invention over the prior art film blowing design are described below. The backside impingement in conjunction with multiple hole film cooling provides improved cooling along the bow wave region of the endwall. Film cooling holes on the bow wave section provides convective and film cooling for the airfoil endwall as well as to baffle the down draft hot gas core air for the vane leading edge. The ejected film cooling air will then migrate down to the airfoil endwall and provide purge air for the end gap of the endwall. The backside impingement cooling onto the backside of the hot wall with built in circumferential micro fins will generate a much higher convective cooling than the prior art backside impingement cooling method. Current cooling system increases the uniformity of the endwall region cooling and insulates the airfoil endwall structure from the passing hot core gas and thus establishes a durable airfoil cooling for the entire endwall and lowers the temperature for the airfoil endwall region. The multiple cooling hole at the bow wave region of the vane leading edge injects cooling air in line with the mainstream flow. This minimizes cooling losses or degradation of the film layer and therefore provides a more effective film cooling for the film layer formation. The multiple film cooling holes at the bow wave region of the endwall provides local film cooling all around the vane leading edge endwall location and thus greatly reduces the local metal temperature and improves the airfoil LCF (low cycle fatigue) capability. Micro fins used in the backside of the hot wall will enhance the bow wave region convective cooling and thus reduce the endwall section metal temperature that will then increase the airfoil ability.