Abstract:
The present invention relates to a method for preventing backflow and forming a cooling layer in an airfoil by creating separation regions at a cooling slot inlet and flowing cooling fluid through the cooling slot.

Description:
FIELD OF THE INVENTION 
       [0001]    The invention relates to a method for preventing backflow and forming a cooling layer in an airfoil. More specifically, the invention relates to method for preventing backflow and forming a cooling layer where separation regions are formed at a cooling slot inlet. 
       BACKGROUND OF THE INVENTION 
       [0002]    Gas turbine engines extract energy from a stream of hot combustion gases that flow through a flow path defined by the turbine. A typical turbine engine includes at least one stage of turbine blades and one stage of vanes spaced from the turbine blades. Each turbine stage comprises a plurality of turbine blades or airfoils spaced circumferentially around, and extending radially outward from, a rotatable hub or disk so that a portion of each turbine blade extends into the flow path and comes in contact with the flow of the combustion gases through the flow path. In practice, turbine engines comprise multiple stages of vanes and blades. 
         [0003]    During engine operation it is necessary to cool turbine blades and vanes to improve their ability to endure extended exposure to the hot combustion gases. Frequently, blade cooling is achieved by creating a cooling film along the blade. In order to develop the desired cooling film, the turbine blades include one or more rows of spanwisely distributed cooling air supply holes, referred to as film holes and these holes are located along the surface of the blade. The film holes penetrate the walls of the airfoil to establish fluid flow communication between cooling fluid passing through the interior of the blade and the externally located hot combustion gases. Additionally, the blade includes a plurality of cooling slots spaced along the trailing edge of the blade. The slots are located within the blade and have outlet openings spaced along the trailing blade edge. During engine operation, cooling fluid or air is typically supplied to the blade by a compressor upstream of the airfoil compressor. The cooling air passes through the interior of the blade, including the slots, and exits the blade through the film holes and outlet openings. The cooling air flows from the holes and the cooling slots as a series of discrete jets. The air discharged from the slots and holes is intended to form the cooling film along the blade surface. 
         [0004]    A conventional airfoil in  FIG. 2  provides an example of a turbine blade  70  of the prior art. As shown in  FIG. 2 , the blade  70  includes leading edge  71 , trailing edge  72  and a plurality of parallel cooling slots  75  at the blade trailing edge. In prior art blade  70 , each of the cooling slots has an associated axially extending slot reference line  80 . Each slot has an inlet  62  and an outlet  63 . The outlet is located at the trailing edge of the blade. The inlet and outlet are located at substantially the same radial position along the radially extending blade length. For simplicity, in  FIG. 2  reference lines  80  are provided for fewer than all of the slots however, the reference lines apply to all of the cooling slots  75 . Each of the cooling slots is parallel to its respective reference line  80 . 
         [0005]    Film cooling provides an effective means for controlling the temperature of airfoil surfaces, however in practice, cooling films are difficult to effectively produce. One shortcoming associated with the conventional parallel cooling slot orientation is that the blade is susceptible to the backflow of combustion gases through the cooling slots. Backflow occurs when the static pressure of the cooling air does not exceed the static pressure of the combustion gases flowing through the flow path. When backflow occurs, the combustion gases flow through the cooling holes and into the cooling slots 
         [0006]    In order to overcome the susceptibility to backflow in conventional blades, the high cooling air is discharged from the slots and holes at a high pressure to prevent backflow. The relatively high pressure cooling air can cause the cooling air to be discharged from the cooling slots with a velocity that prevents the cooling air from effectively adhering to the surface and edges of the airfoil. As a result, the desired cooling film does not form on the blade. Instead the cooling air is directly flowed into and entrained with the combustion gases. As a result, a portion of the blade airfoil surface immediately downstream of each cooling hole or cooling slot is exposed to the combustion gases and is not protected by a cooling film. Additionally, each of the cooling air jets may locally intersect and bifurcate the stream of combustion gases into a pair of minute, oppositely swirling vortices. The combustion gases enter the exposed portion of the airfoil and can cause irreparable damage to the airfoil. The intense heat of backflow gases can quickly and irreparably damage an airfoil. 
         [0007]    What is therefore needed is an airfoil with cooling slots arranged in a manner that promotes effective formation of a cooling film along the airfoil surface. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0008]    A method for preventing backflow and forming a cooling layer in an airfoil, said airfoil comprising a leading edge, a trailing edge, a blade tip at a first blade end and a blade root at a second blade end, the tip and root being separated by a radial distance, a cooling passage extending between the leading and trailing edges, and at least one cooling slot having an inlet end in fluid receiving communication with the cooling passage and an outlet end proximate the trailing edge, and wherein for the at least one slot the inlet and outlet are located at different radial locations within the airfoil, the method comprising flowing a cooling fluid in a first direction through said cooling passage toward a cooling slot, flowing the cooling medium in a second direction through said cooling passage toward the cooling slot, forming a separation region proximate the cooling slot inlet and flowing the cooling medium through said cooling slot and out the slot to form a layer at the trailing edge of the airfoil. 
         [0009]    Thus, by the described invention improved the cooling of an airfoil is achieved. This improvement is accomplished by metering airflow through a plurality of angled cooling slots. Also, instead of drilling cooling slots into an airfoil, one may cast cooling slots into an airfoil and thus decrease manufacturing costs and increase the beneficial variability of cooling slots at their creation. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0010]    While the specification concludes with claims particularly pointing out and distinctly claiming the invention, it is believed that the embodiments set forth herein will be better understood from the following description in conjunction with the accompanying figures, in which like reference numerals identify like elements and in which: 
           [0011]      FIG. 1  shows a schematic representation of a gas turbine; 
           [0012]      FIG. 2  is a sectional view of a prior art turbine blade comprising a conventional cooling slot configuration; 
           [0013]      FIG. 3  is a sectional view of a turbine blade comprising a cooling slot arrangement according to an embodiment of the present invention; 
           [0014]      FIG. 4  is a sectional view of a turbine blade comprising an alternate embodiment of the invention; and 
           [0015]      FIG. 5  is an enlarged detailed view of the portion of  FIG. 4  within the circle identified as  5 . 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0016]      FIG. 1  is a schematic representation of an exemplary gas turbine engine  10 . Engine  10  includes a fan assembly  12 , a core engine  13 , a high-pressure compressor  14 , and a combustor  16 . Engine  10  also includes a high-pressure turbine  18 , a low-pressure turbine  20 , and a booster  22 . Fan assembly  12  includes an array of fan blades  24  extending radially outward from a rotor disc  26 . Engine  10  has an intake side  27  through which air flows into and an exhaust side  29  through which air flows out of the engine. In one embodiment, the gas turbine engine is a GE90-115B that is available from General Electric Company, Cincinnati, Ohio. Fan assembly  12  and turbine  20  are coupled by shaft  31 . Compressor  14  and turbine  18  are coupled by shaft  33 . 
         [0017]    During operation, air flows axially through fan assembly  12  in a direction that is substantially parallel to central axis  34  extending through engine  10 . Compressed air is supplied primarily to combustor  16  by high-pressure compressor  14 . Most of the highly compressed air is delivered to combustor  16 . Airflow (not shown in  FIG. 1 ) from combustor  16  drives turbines  18  and  20 , and turbine  20  drives fan assembly  12  through shaft  31 . High-pressure turbine  18  includes an array of blades  60 . 
         [0018]    Blade or airfoil  60  is shown in greater detail in  FIG. 3 . Additionally the airfoil may be a vane. Airfoil  60  comprises leading edge  74 , and a trailing edge  76  opposite the leading edge. The blade also comprises radially opposed blade tip  81  and root  79 . The tip and root are separated by a radially extending distance. The blade is coupled with the rotor (not shown) at the root. Air flowing through the gas turbine engine along the flow path flows across the blade  60  in an axial direction from the leading edge  74  to the trailing edge  76 . Compressed cooling air flows into the blade through openings at the leading edge  74  of the airfoil and also through inlet passages  77 . The cooling air that flows through passages  77  flows radially outward toward blade tip  81 . As the inlet passages extend toward tip  81 , they combine into a single cooling passage  91 . The cooling passage extends in a serpentine manner through the interior of the blade. As shown in  FIG. 3 , blade  60  includes two inlets but it should be understood that blade  60  may include any suitable number of inlet passages  77 . Arrows in  FIG. 3  generally represent the flow direction of cooling air within blade  60 . 
         [0019]    A plurality of spaced apart vanes  92  are located in cooling passage  91  between inlet passages  77  and tip  81 . The vanes are oriented in a parallel array, with each vane being substantially parallel to the other vanes in the array. Each vane has a first end  94  and a second end  95 . For each vane the first end  94  of each discrete vane is located closer to root  79  than second end  95  of the same vane. For each discrete vane each second vane end  95  is located closer to tip  81  than first vane end  94  for the same vane. The vanes are fixed to the wall that defines the portion of cooling passage  91  at the trailing blade edge. The vanes are oriented at an angle relative to generally axially extending axis  99 . Each vane is oriented relative to axis  99  at an angle that is less than ninety degrees. By orienting the vanes in this manner, with the first and second ends for each vane at different radial locations, cooling air is more effectively directed into cooling slots  45 . 
         [0020]    As shown in  FIG. 3 , blade  60  includes a plurality of cooling slots  45 . The cooling slots are oriented in a generally parallel array. For purposes of disclosing a preferred embodiment of the invention blade  60  comprises seven slots however it should be understood that any suitable number of slots  45  may be provided in the blade. Each slot has an inlet  96  and an outlet  97 . The outlets  97  are located at the trailing edge  76  of blade  60 . The slots are formed in the blade proximate the trailing edge. The inlet is in flow communication with the cooling slot  91  and cooling air in the cooling passage  91  enters the cooling slot through inlet  96 . The slots  45  of blade  60  are of substantially constant radial dimension and the radial dimension may be a diameter for example. For each cooling slot, the outlet  97  is located closer to the root  79  than the slot inlet  96  for the same cooling slot. For each discrete slot, the slot inlet  96  is located nearer the blade tip  81  than the slot outlet  97  for the same cooling slot. As a result of positioning the inlet and outlet for each cooling slot at a different radial locations along the blade, the airfoil of the present invention more effectively produces a cooling film along the blade. More specifically, airfoil  60  more effectively forms a cooling film along the trailing edge  76  of the blade. 
         [0021]      FIG. 4  discloses an alternate embodiment blade  61  that comprises slots  48 , similar to slots  45 . Slots  48  include inlet  106  and outlet  107 . Like slots  45 , the inlet and outlet for each slot is located at a different radial location along the blade with each inlet  106  located closer to tip  81  than outlet  107 . The outlet  107  is located closer to root  79  than inlet  106 . The radial dimensions for inlets  106  and  107  are not the same. As shown in  FIG. 4 , the inlet has a smaller radial dimension than the outlet. The radial dimension may be a diameter for example with the diameter of inlet  106  being smaller than the diameter of outlet  107 . Blade  61  includes passages  77 ,  91  leading edge  74 , trailing edge  76 , tip  81 , root  79  and vanes as described in blade  60 . 
         [0022]    Note that unless specifically indicated to the contrary, as the description proceeds the description relating to slot  45  shall also apply to slot  48 . For simplicity, the description shall refer to slot  45 . As is shown in  FIGS. 3 and 4 , substantially all of the cooling slots  45 ,  48  may be oriented in a parallel array, at substantially the same angle Alpha (α) as shown in detail in  FIG. 5 . The angle alpha, identified at  110  is measured between reference line  35  and the central axis of slot  45 . The central axis is identified as  120 . The reference line  35  is substantially horizontal. In an alternate embodiment, fewer than substantially all of the slots may be arranged in parallel. For example, fifty percent of the slots may be arranged in parallel at the same angle  110 . Angle  110  of cooling slot  45  is shown in which the angle is less than 90° and greater than 0°. 
         [0023]    In practice, the flow of air through the cooling slot  45  of the present embodiment invention is distinguishable from the flow of air through conventional slots where the slot inlet and outlet are located at the same radial positions along the length of the blade. Cooling slots  45  minimize the mass flow of air through the slots  45  thus providing a controlled flow through the blade that is discharged from the slot outlet  97  at a velocity that is greatly reduced relative to prior art cooling slots. Such metered or controlled airflow creates a partial restriction of cooling air passing through the cooling slots  45 . It should be understood that such restriction does not diminish the quality of the cooling layer formed on blade  60 . Rather, the controlled, metered flow serves to enhance the formation of cooling film layer  30  and also to prevent both the escape of cooling air into the flow path of combustion gases and the formation of a backflow condition. By decreasing the cooling air mass flow through cooling slot  45  the velocity of the cooling air exiting the slots is reduced, thereby providing a cooler, slower moving boundary layer. As a result, upon exiting the slot the cooling air remains close to the surface and edges of turbine blade  60 , ensuring that a suitable cooling layer is formed. 
         [0024]      FIG. 5  provides a more detailed view of cooling airflow entering, traveling through and exiting cooling slot  45 . Although the flow of cooling air entering, flowing through and exiting is only shown relative to one slot  45 , the flow represents the flow for all slots  45  and  48 . Cooling air flows to slot  45  through passage  91 , from a first flow position  126  toward cooling slot inlet  96 . Oppositely, cooling air flows through passage  91 , in from second flow position  127  toward cooling slot inlet  96 . First flow position cooling air enters the blade through openings at the blade leading edge  74  and passes through upstream portion of passage  91  toward the slots. As cooling air flows to the cooling slot from flow position  126  it may substantially move unobstructed into cooling slot  45 . As cooling air enters from second flow position  127  the flow may be obstructed by one or more separation regions  136  created at or proximate cooling slot inlet  96 . A separation region  136  occurs in a region adjacent cooling slot inlet  96 . When cooling air from the flow position  127  approaches slot  45 , cooling air from flow position  127  abruptly meets the flow  126 , and thus creates one or more areas in which the air swirls or separates from its original flow stream, producing separation region  136 . 
         [0025]    In addition to the angled orientation of cooling slot  45 , the separation region  136  can aid in metering the flow of cooling air through cooling slot  45  since it can at least partially block the flow of air from flow position  127  from moving into cooling slot  45 . This prevents the formation of backflow as well as controlling the flow of cooling air into the slot. Cooling film layer  130  is formed by the cooling air exiting from cooling slot outlet  45 . Cooling film layer  130  is formed on the leading edge  76  of blade  60  and serves to help cool the surface of turbine blade  60  and protect the blade against the harmful effects associated with hot combustion gases. 
         [0026]    Cooling slot  45  is oriented at an angle  110  that may range from about 1 degree (1°) to about 88 degrees (88°). In another embodiment the angle  110  may range from about 10 degrees (10°) to about 75 degrees (75°). In still another embodiment the angle may range from about 20 degrees (20°) to about 60 degrees (60°)(30°) to about 50 degrees (50°). 
         [0027]    The pressure ratio for each turbine blade  60  at the inlet  96  of each cooling slot  45  ranges from a pressure ratio of about 1.05 to about 2.0. The term “pressure ratio” means the ratio of the internal blade pressure to the external flow path pressure. It is desired to produce a pressure ratio greater than 1.0 since a pressure ratio lower than that would produce a backflow condition. Also, the movement of air within the airfoil through the cooling passage, slots and vanes is desired to have a Mach number ranging from about 0.03 Mach number to about 1.0 Mach number. The Mach number is defined as a ratio of the speed of an object or flow relative to the speed of sound in the medium through which it is traveling. In the present invention the Mach number falls into the desired range. 
         [0028]    Additional benefits associated with the blade of the present invention include the fact that more cooling slots  45  can be used in engines having smaller turbine blades. By the term “smaller turbine blades” it is meant herein a turbine blade in an aircraft engine application in which the engine core flow rate is less than 13.61 kg/s at take-off power level. An exemplary engine having smaller turbine blades of the type discussed is a CT7 or T700 available from General Electric Company, Cincinnati, Ohio. 
         [0029]    The blade of the present invention allows cooling slots  45  to be cast rather than drilled. The use of cast slots instead of drilled holes presents a significant cost savings in manufacturing, use of resources and material usage. In one embodiment, at least a portion of cooling slots  45  may be cast along trailing edge  76  of turbine blade  60 . 
         [0030]    Cooling slots  45  of the invention also allow for beneficial variability. The term “beneficial variability” means that one or more cooling slots  45  may have a varying diameter along its length and/or because of casting may have much larger diameters in comparison to drilled cooling slots  75 . One example of beneficial variability is the use of larger holes, i.e., the exits of the cooling slots along the trailing edge of the turbine blades  70  (see  FIG. 4 ). By having larger exit holes than those provided by drilling, e.g., laser drilling, greater cooling film coverage is achieved about the surface of turbine blade  60 . Also, since outlets  107  can be made to be larger, than current slot technology, fewer cooling slots  45  may be used than in blades where constant radial dimension/diameter slots are used. 
         [0031]    This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to make and use the invention. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.