Abstract:
Film cooling of turbine component surfaces, such as high-lift airfoil surfaces, is achieved by a flow of cooling air through film cooling holes of generally constant cross-sectional area, which extend from the surfaces to a radial cooling passage within the airfoil. The film cooling holes are angularly offset from that portion of the radial cooling passage immediately upstream of the film cooling holes by an acute angle which effects a radial reversal of the flow of cooling air into the film cooling holes from the radial cooling passage to reduce the momentum of airflow through the film cooling holes to reduce separation of cooling air film from the surface.

Description:
U.S. GOVERNMENT RIGHTS 
       [0001]    This invention was made with U.S. government support under contract F33615-03-D-2354-0017 awarded by the U.S. Navy. The U.S. government has certain rights in this invention. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    1. Technical Field 
         [0003]    This invention relates generally to gas turbine engines and particularly to film cooled turbine components such as airfoil rotor blades and stator vanes employed in such engines. 
         [0004]    2. Background Information 
         [0005]    Gas turbine engines, such as those which power aircraft and industrial equipment, employ a compressor to compress air which is drawn into the engine, and a combustor, which receives air from the compressor, injects fuel into the compressed air and ignites the fuel air mixture, converting the fuel air mixture into a high temperature working fluid which is exhausted to the engine&#39;s turbine. The working fluid flows through a turbine which captures the energy associated with the working fluid. The working fluid temperature is typically higher than the melting point of the materials used to manufacture turbine components such as rotor blades and stator vanes. Therefore, it has become the practice to cool such components internally and externally with cooling air, typically from the engine&#39;s compressor. In the case of the aforementioned turbine rotor blades and stator vanes, there are three well-known methods for cooling such components. One method is the convective cooling of such components by flowing compressor discharge air through radial passages within the interiors of those components. A second method of cooling such components involves the impingement of air against the interior surfaces of those components. A third method critical to achieve adequate cooling of turbine rotor blades and stator vanes, involves the establishment of a film of air on the surfaces of those components exposed to the engine&#39;s working fluid. The film of air flowing over the surfaces of the rotor blades and vanes acts as a thermal barrier, shielding the surfaces of those components from the destructive thermal effects of the working fluid flowing thereover. Typically, air flowing through the internal radial convective cooling passages in the turbine blades and vanes is channeled to the blade and vane surfaces through film cooling holes which open onto the blade and vane surfaces and connect to the internal radial convective cooling passages which feed compressor air to the film cooling holes for establishing the cooling air film at the blade and vane surfaces. With such film cooling, it is imperative that the cooling air film remain attached to the blade and vane surfaces to achieve the requisite protection of those surfaces from the deleterious thermal effects of the working fluid passing over the component surfaces. It will be appreciated that separation of the cooling air film from the blade and vane surfaces, also known as lift-off or blow-off of the cooling film due to the contact of the working fluid therewith, will severely compromise the cooling air film&#39;s ability to shield the blade and vane surfaces from the destructive effects thereon of working fluid heat. One known technique for reducing cooling film lift off or blow off from the blade and vane surfaces involves the use of diffuser cooling hole openings at the surfaces. Such diffuser openings are usually conical in shape for reducing the velocity and dynamic pressure of the cooling air as it exits the film cooling holes at the blade and vane surfaces. However, intricate and therefore somewhat sophisticated and costly manufacturing techniques such as electro-discharge machining, are required to form the diffuser openings of the film cooling holes and therefore contributes significantly to the cost of manufacture of the turbine blades and vanes. Moreover, known diffuser hole opening configurations, while being suitable for rotating turbine blades, may not be suitable for stationary turbine stator vanes because such diffuser hole configurations may require the influence of centrifugal force on air flowing through the diffuser hole opening for the diffuser hole opening to function properly in establishing a fluid film at the surface over which the working fluid flows. 
         [0006]    As aircraft and industrial equipment powered by gas turbine engines become larger and more complex, the thrust output requirements of gas turbine engines is continually increasing. One of the most well-known techniques for increasing the thrust output of a gas turbine engine to meet the ever-increasing thrust output demands thereof is increasing the mass flow of working fluid passing through the engine&#39;s turbine and increasing the temperature of the working fluid exhausted to the turbine from the engine&#39;s combustor. Increasing the mass flow of the working fluid passing through the turbine requires larger and heavier turbine component parts and may therefore increase the weight of the engine beyond that required to meet the design requirements and weight constraints of associated aircraft or other equipment which the engine powers. 
         [0007]    Modern gas turbine engine turbine blades and vanes sometimes employ high-lift airfoil shapes to maximize the ability of the turbine to capture energy from the working fluid, thereby reducing the need for increasing the flow of working fluid through the turbine for increased turbine performance and the attendant larger and heavier turbine components required thereby. However, working fluid flow over such high lift airfoil surfaces results in high ratios of working fluid static pressure to total pressure (typically greater than 0.9) across substantially the entire airfoil surface. Such high static to total pressure ratios cross the airfoil surfaces of the turbine blades and vanes results in an increased risk of cooling air film blow-off from the airfoil surfaces. Accordingly, there is a continuing effort to achieve high lift turbine airfoils for gas turbine engines wherein the risk of cooling air film blow off from the airfoil surfaces is minimized without the aforementioned complex and costly shaped film cooling hole openings onto the airfoil surfaces which may not be suitable for use with stationary stator vanes. 
       SUMMARY OF THE DISCLOSURE 
       [0008]    In accordance with the present invention, turbine components such as stator vanes and rotor blades are provided with simplified film cooling hole configurations wherein shaped cooling hole openings at the airfoil surfaces are not required, the cooling holes comprising cores of substantially constant cross-sectional area from the cooling hole openings on the surface to the inner end of the cooling holes which receive cooling air from radial cooling passages within the interior of the components. The elimination of the necessity for shaped cooling hole openings is achieved by a radial reversal of cooling air flow from a radial cooling passage within the interior of the component by a cooling air flow diverter at the intersection of the film cooling hole with the radial cooling air passage within the interior of the turbine component. The flow diverter comprises the intersection of the film cooling hole wall with the wall of the radial cooling air supply passage wherein the film cooling hole intersects that portion of the radial cooling passage which supplies the film cooling hole with cooling air, medially upstream of the film cooling hole at an acute angle whereby cooling air entering the film cooling hole from the radial cooling passage is angularly displaced greater than 90 degrees from the flow of cooling air through the radial cooling passage thereby decreasing the momentum of the air flow through the film cooling hole and reducing any tendency of the cooling air discharged from the film cooling hole at the outer component surface to separate therefrom due to blow-off. Since the necessity for shaped film cooling hole openings is eliminated, the turbine component cooling film of the present invention is achieved simply and economically. Moreover, since diffuser openings are not required with the film cooling hole configuration of the present invention, centrifugal force is not required for the film cooling holes to function properly, the film cooled turbine component of the present invention is well-suited for stationary turbine stator vanes as well as rotor blades. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]      FIG. 1  is a schematic view of a turbofan gas turbine engine of the type employing the present invention; 
           [0010]      FIG. 2  is an isometric view of a gas turbine engine turbine rotor blade of the type employing the present invention; and 
           [0011]      FIG. 3  is a sectional view of the rotor blade of  FIG. 2  taken in the direction of line  3 - 3  of  FIG. 2 . 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0012]    Referring to  FIG. 1 , a turbofan gas turbine engine  5  has a longitudinal axis  7  about which the rotors  8  within stator  9  rotate. Stator  9  circumscribes the rotors. A fan  10  disposed at the engine inlet draws air into the engine. A low pressure compressor  15  located immediately downstream of fan  10  compresses air exhausted from fan  10  and a high pressure compressor  20  located immediately downstream of low pressure compressor  15 , further compresses air received therefrom and exhausts such air to combustors  25  disposed immediately downstream of high pressure compressor  20 . Combustors  25  receive fuel through fuel injectors  30  and ignite the fuel/air mixture. The burning fuel-air mixture (working medium fluid) flows axially to a high pressure turbine  35  which extracts energy from the working medium fluid and in so doing, rotates hollow shaft  37 , thereby driving the rotor of high pressure compressor  20 . The working medium fluid exiting the high pressure turbine  35  then enters low pressure turbine  40 , which extracts further energy from the working medium fluid. The low pressure turbine  40  provides power to drive the fan  10  and low pressure compressor  15  through low pressure shaft  42 , which is disposed interiorly of the hollow shaft  37 , coaxial thereto. Working medium fluid exiting the low pressure turbine  40  provides axial thrust for powering an associated aircraft (not shown) or a free turbine (also not shown). 
         [0013]    Bearings  43 ,  45 ,  50  and  53  radially support the concentric high pressure and low pressure turbine shafts from separate frame structures  52 ,  54 ,  55  and  56  respectively, attached to engine case  57 , which defines the outer boundary of the engine&#39;s stator  9  which circumscribes rotors  8 . However, the present invention is also well suited for mid-turbine frame engine architectures wherein the upstream bearings for the low and high pressure turbines are mounted on a common frame structure disposed longitudinally (axially) between the high and low pressure turbines. 
         [0014]    Referring to  FIG. 2 , a high pressure turbine blade  60  comprises an airfoil shaped surface  65  having a concave pressure portion  66  and a convex suction portion  67  extending radially outwardly from a platform  70  which defines the radially innermost boundary of the working fluid flow path through high pressure turbine  35 . A dovetail shaped root portion  75  is provided at the radially innermost end of blade  60  and is accommodated within a mating slot provided in the radially outer portion of a disk shaped blade retainer (not shown) mounted on the turbine shaft. A number of film cooling holes  80  open onto the airfoil surface  65  to provide a film of cooling air which flows over the airfoil surface  65  thereby providing a thermal boundary layer to protect the surface  65  from the deleterious thermal effects associated with working fluid which passes over blade  60  as the working fluid flows through the turbine. 
         [0015]    Turbine blade  60  is a high lift airfoil which, as described hereinabove, maximizes the energy captured from the working fluid by the turbine blade. As further set forth hereinabove, working fluid flowing over the airfoil surface of such high-lift turbine blades exhibits a ratio of static pressure to total pressure in proximity to airfoil surface  65  greater than approximately 0.9 across a substantial portion of the airfoil surface of blade  60 . As further set forth hereinabove, such levels of static pressure to total pressure ratio can, with prior art film cooling hole configurations, cause the cooling air film flowing over the airfoil surface to separate or blow off the airfoil surface thereby severely jeopardizing the film&#39;s ability to protect the airfoil surface from the extreme destructive thermal effects of working fluid heat. 
         [0016]    Referring to  FIG. 3 , the interior of turbine blade  60  is provided with a radial cooling air passage  85  which accommodates a radially outward flow of cooling air provided by the engine&#39;s compressor in the direction of arrow  90 . Cooling air flowing through passage  85  convectively cools the body of the blade. Each of film cooling holes  80  extends between airfoil surface  65  and radial cooling passage  85 , intersecting radial that portion of cooling passage  85  immediately upstream of film cooling hole  80  at an acute angle a of approximately 25° whereby cooling air flowing through radial cooling passage  85  is angularly displaced greater than 90 degrees as it enters film cooling holes  80 . This angular displacement of cooling air as it enters film cooling holes  80  from radial cooling passage  85 , lowers the momentum of cooling air flowing through film cooling hole  80  prior to reaching the airfoil surface and therefore enhances the cooling air film&#39;s ability to remain attached to airfoil surface  65  to maximize the thermal protection of airfoil surface  65  afforded by the film from the deleterious thermal effects of the working fluid passing over the blade surface. 
         [0017]    Still referring to  FIG. 3 , the sidewalls of radial passage  85  and film cooling holes  80  intersect to define a flow diverter  95  which effects the turning of the cooling air as it enters film cooling holes  80  from radial passage  85 . 
         [0018]    As shown in  FIG. 3 , film cooling holes  80  arc of a generally constant cross-sectional area throughout the length thereof since the momentum reducing turning of the cooling air as it enters film cooling holes  80  from radial passage  85  reduces the momentum of the cooling air flowing through holes  80  and thus renders shaped outer diffuser openings of the film cooling holes unnecessary to maintain the cooling air film attached to airfoil surface  65 . Accordingly, since shaped diffuser film cooling hole openings are unnecessary with the momentum reducing turning of the cooling air as it enters the film cooling holes from the radial cooling passage, the film cooling hole configuration of the present invention is equally well-suited for stationary stator vane and outer air seals since centrifugal force is not necessary to keep any shaped cooling hole openings filled with cooling air to prevent blow off of the cooling air film from airfoil surface  65 . Furthermore, since shaped diffuser cooling hole openings are not employed with the present invention, the film cooling holes may be conveniently and economically produced in the turbine blades and vanes by well-known and economical drilling methods, thereby rendering the expensive and intricate electro-discharge machining of shaped film cooling hole openings unnecessary. 
         [0019]    While the present invention has been described within the context of a high lift gas turbine engine turbine rotor blade, it will be appreciated that the invention herein is equally well-suited for conventional (not high lift) turbine blades wherein the risk of cooling air film blow off from the blade&#39;s airfoil surfaces is not as high. Also, while the invention herein has been described in connection with a turbine blade, as set forth hereinabove, this invention is equally well-suited for stationary turbine stator vanes and outer air seals. Furthermore, while a specific number of film cooling holes opening onto a specific portion of the rotor blade&#39;s airfoil surface have been illustrated, the invention herein may be employed with any required number of film cooling holes opening onto any portion of the blade&#39;s airfoil surface (pressure or suction) as required to achieve required film cooling of the surface. 
         [0020]    Accordingly, it will be understood that various modifications to the preferred embodiment described herein may be made without departing from the present invention, and it is intended by the appended claims to cover such modifications as fall within the true spirit and scope of the invention.