Patent Publication Number: US-8118547-B1

Title: Turbine inter-stage gap cooling arrangement

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 interstage gap between a blade outer air seal and an endwall of an adjacent stator 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. If proper cooling or design measures are not undertaken to prevent this hot gas ingress, exposure to the hot gas can result in severe damage to the front edges of the vane endwall as well as the turbine components located upstream of the endwall.  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. 
       FIG. 2  shows a prior art turbine with a first stage rotor blade located upstream from a row of second stage stator vanes. An interstage gap is formed between a blade ring for the rotor blade and a blade ring for the stator vane. This arrangement in  FIG. 2  includes a rotor blade  27  with a tip that forms a seal with a blade outer air seal (or BOAS)  24 , the BOAS  24  is supported by hooks on an isolation ring  22  on a forward side and a blade ring  21  on an isolation ring  25  on the aft side. A first blade ring  21  supports both isolation rings  22  and  25  and includes a cooling air passage that delivers cooling air to an impingement plate  23  that includes impingement holes  28  that discharge jets of impingement cooling air onto a top surface of the BOAS. 
     An adjacent stator vane assembly includes a second blade ring  26  that supports a guide vane  11  with an outer endwall  12 . an interstage gap  29  is formed between the isolation ring  25  and the vane outer diameter endwall  12  in which the hot gas ingress can occur due to the pressure differential described above. 
     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. 
     As a result of the design of  FIG. 2 , hot gas flows in and out along the inter-stage gaps and an over-temperature occurs at the blade outer air seal edges and the blade isolation ring corresponding to the hot gas injection location. This over-temperature issue is more pronounced when an insufficient amount of inter-stage gap purge air for the axial gap is available when a strong bow wave is induced by the low solidity vane airfoil creates a high circumferential pressure variation which acts to push the mainstream hot gas into the inter-stage gap  29 . 
     BRIEF SUMMARY OF THE INVENTION 
     It is an object of the present invention to provide for a turbine 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. 
     These objectives and more can be achieved by the turbine inter-stage gap cooling apparatus and method of the present invention. A row of cooling air holes are located on the BOAS upstream from the vane leading edge diameter that discharges cooling air into the airfoil leading edge section. The forced injection of the cooling air flow with the use of the blade outer air seal spent cooling air into the transition space between the vane leading edge airfoil and the vane outer diameter endwall will prevent the hot gas flow from ingesting into the interstage gap. 
    
    
     
       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 a cross section side view of an inter-stage seal arrangement for a prior art turbine rotor blade and adjacent stator vane design with an interstage gap. 
         FIG. 3  shows a cross section side view of an inter-stage seal arrangement of the present invention for the turbine rotor blade and adjacent stator vane with an inter-stage gap. 
         FIG. 4  shows a detailed close-up view of the BOAS cooling air holes for the gap of  FIG. 3 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The present invention is a turbine interstage gap cooling apparatus and method for an industrial gas turbine engine that can also be used in an aero engine for the same purpose.  FIG. 3  shows a stage of rotor blades adjacent to an upstream from a stage of guide vanes. The rotor blade  27  includes a tip that forms a seal with the BOAS  24  as in the prior art  FIG. 2 . The same parts in  FIG. 3  are labeled as the same reference numbers as in the prior art  FIG. 2  arrangement. The blade outer air seal (BOAS)  24  in the  FIG. 3  invention includes a row of cooling air holes  31  as seen in  FIG. 4  that connect the inner side of the BOAS to the aft side of the BOAS  24  such that spent impingement cooling air from the inner surface of the BOAS  24  will be discharged in the direction of the arrow shown in  FIG. 4 . The BOAS  24  includes an outward extending ledge  36  on the aft side that extends beyond the plane of the aft side that is flush with the isolation ring  25  as is the case in the prior art  FIG. 2  BOAS. The cooling air holes  31  are located above the ledge  36  and are directed to discharge the cooling air toward the transition between the concave shaped outer diameter endwall  12  and the leading edge of the airfoil  11 . The cooling air holes  31  extend along the aft side of the BOAS. A TBC is shown applied to the inner surface of the BOAS. A tangent line  32  is tangent to the concave shaped endwall surface as seen in  FIG. 4 . An arrow  33  represents the direction of the hot gas flow through the vane. The angle of the cooling air holes  31  and therefore the angle of injection of the cooling air  34  is half the difference between the two angles of the tangent  32  and the hot gas flow  33 . 
     The injection of the spent cooling air from the blade outer air seal trailing edge cooling through the row of metering holes  31  and into the vane leading edge nose region will eliminate the hot gas ingestion into the gap  29  that is present in the prior art inter-stage seal gap design. The spent cooling air form the blade outer air seal is discharged into the vane leading edge in-between the angle formed by the streamline of the hot gas flow and a tangent to the endwall corner diameter of the vane. This precise position of the spent cooling air discharge cooling holes  31  will provide proper cooling for the vane bow wave region in addition to prevent ingress of the hot gas into the gap  29 .