Patent Publication Number: US-6991430-B2

Title: Turbine blade with recessed squealer tip and shelf

Description:
GOVERNMENT INTERESTS 
   The U.S. Government may have certain rights in this invention in accordance with Contract No. N00019-96-C-0080 awarded by the Dept. of the Navy. 

   BACKGROUND OF THE INVENTION 
   1. Field of the Invention 
   The invention relates generally to gas turbine engine turbine blade squealer tip cooling and, more specifically, to turbine blade squealer tips cooled using cooling holes through a tip shelf. 
   2. Description of Related Art 
   Gas turbine engine turbine blades extract energy from hot combustion gas for powering the compressor and providing output power. Since the turbine blades are directly exposed to the hot combustion gas, they are typically provided with internal cooling circuits which channel a coolant, such as compressor bleed air, through the airfoil of the blade and through various film cooling holes around the surface thereof. One type of airfoil extends from a root at a blade platform, which defines the radially inner flowpath for the combustion gas, to a radially outer tip cap, and includes opposite pressure and suction sides extending axially from leading to trailing edges of the airfoil. The cooling circuit extends inside the airfoil between the pressure and suction sides and is bounded at its top by the airfoil tip cap. A squealer tip blade has a squealer tip wall extending radially outwardly from the top of the tip cap and completely around the perimeter of the airfoil on the tip cap to define a radially outwardly open tip cavity. 
   The squealer tip is a short radial extension of the airfoil wall and is spaced radially closely adjacent to an outer turbine shroud to provide a relatively small clearance gap therebetween for gas flowpath sealing purposes. Differential thermal expansion between the blade and the shroud, centrifugal loading, and radial accelerations cause the squealer tips to rub against the turbine shroud and abrade. Since the squealer tips extend radially above the tip cap, the tip cap itself and the remainder of the airfoil is protected from damage, which maintains integrity of the turbine blade and the cooling circuit therein. 
   However, since the squealer tips are solid metal projections of the airfoil, they are directly heated by the combustion gas which flows thereover. They are cooled by heat conduction with the heat then being removed by convection into the tip cap and cooling air injected into the cavity by passages through the tip. The cooling air from within the airfoil cooling circuit is used to convect heat away from tip and to inject into cavity. The squealer tip typically operates at temperatures above that of the remainder of the airfoil and can be a life limiting element of the airfoil in a hot turbine environment. 
   Since the pressure side of an airfoil typically experiences the highest heat load from the combustion gas, a row of conventional film cooling holes is typically provided in the pressure side of the airfoil outer wall immediately below the tip cap for providing a cooling film which flows upwardly over the pressure side of the squealer tip. U.S. Pat. No. 6,164,914 discloses a turbine blade including a hollow airfoil having a squealer tip wall extending outboard from a tip cap enclosing the airfoil. Film cooling holes extend through the junction of the tip cap below the pressure-side portion of the tip rib for discharging the coolant in a layer of film cooling air for flow along the exposed pressure side of the squealer tip wall. It is difficult to entrain the cooling air flow in a boundary layer along the exposed pressure side of the squealer tip wall. Often the film cooling holes will direct the cooling air to impinge on the pressure side of the squealer tip wall and a large portion will bounce off and not be entrained in the boundary layer. 
   However, cooling of the squealer wall is limited in effectiveness, and thermal gradients and stress therefrom are created which also affect blade life. The exposed squealer wall runs hotter than the airfoil sidewalls with the tip cap therebetween running cooler. Tip cooling must therefore be balanced against undesirable thermal gradients. 
   SUMMARY OF THE INVENTION 
   A turbine blade includes an airfoil having an airfoil outer wall extending longitudinally outwardly from a root, pressure side and suction sides extending laterally from a leading edge to a trailing edge of the airfoil, and a squealer tip at a radially outer end of the airfoil. The squealer tip includes a radially outer tip cap attached to the airfoil outer wall, a continuous squealer tip wall extending radially outwardly from and continuously around the tip cap forming a radially outwardly open tip cavity, and a recessed tip wall portion recessed inboard from the pressure side of the airfoil outer wall forming a tip shelf therebetween. An internal cooling circuit extends longitudinally outwardly from the root to the tip cap and a plurality of film cooling shelf holes are disposed through the tip shelf to the internal cooling circuit and spaced away from a junction between the recessed tip wall portion and the tip shelf. 
   In an exemplary of the turbine blade, the film cooling shelf holes have shelf hole centerlines passing through pierce points in the shelf angled at compound angles with respect to vertical lines passing through the pierce points. The compound angles have orthogonal first and second component angles. The first component angles lie in first planes defined by the vertical lines and first coordinate lines that are normal to the vertical lines and extend between the vertical lines and the recessed tip wall portion. The second component angles lie in second planes defined by the vertical lines and second coordinate lines that are normal to the vertical lines and normal to the first coordinate lines. At least a majority of the shelf hole centerlines are angled in outboard directions away from and outboard of the squealer tip wall. Their shelf hole centerlines are angled at the second component angles in downstream lateral directions with respect to vertical lines wherein the downstream lateral directions are normal to corresponding ones of the outboard directions and the vertical lines. 
   In a more particular embodiment of the turbine blade the first coordinate lines lie along transverse lines which are substantially shortest distances between the vertical lines are shortest distances between the vertical lines and the recessed tip wall portion. The shelf hole centerlines are spaced away from a fillet at the junction. The film cooling shelf holes extend into the fillet no more than 50 percent of a fillet width of the fillet as measured along the tip shelf. The majority of first component angles are in a range between 2 degrees and 16 degrees. A first plurality of the film cooling shelf holes have shelf hole centerlines with the positive first component angles in a range between 0.5 degrees and 5 degrees. The turbine blade is made with a nickel-base superalloy having a free sulfur content less than about 1 part per million by weight. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The foregoing aspects and other features of the invention are explained in the following description, taken in connection with the accompanying drawings where: 
       FIG. 1  is an isometric view illustration of an exemplary gas turbine engine turbine blade having a squealer blade tip with a tip shelf and film cooling shelf holes disposed through the tip shelf and spaced away from a tip wall. 
       FIG. 2  is a partial cut-away illustration of the gas turbine engine turbine blade in  FIG. 1 . 
       FIG. 3  is an enlarged isometric view illustration of the squealer blade tip and tip shelf illustrated in  FIG. 1 . 
       FIG. 4  is a cross-sectional view illustration through  4 — 4  in  FIG. 1  and through tip shelf illustrated in  FIG. 1 . 
       FIG. 5  is an enlarged cut-away isometric view illustration of a portion of the squealer blade tip and tip shelf illustrated in  FIG. 1 . 
       FIG. 6  is a cross-sectional view illustration through  6 — 6  in  FIG. 5 . 
       FIG. 7  is a cross-sectional view illustration through  7 — 7  in  FIG. 5 . 
       FIG. 8  is a table of angles of tip film cooling holes illustrated in  FIG. 1 . 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   Illustrated in  FIGS. 1 and 2  is an exemplary gas turbine engine turbine rotor blade  10  configured for use as a first stage high pressure turbine blade. The blade  10  includes a dovetail  12  having suitable tangs  13  for mounting the blade in corresponding dovetail slots in the perimeter of a rotor disk (not shown). The blade  10  further includes an airfoil  16  joined to the dovetail  12  at an airfoil base  19  at an integral platform  20  and a squealer tip  38  at a radially outer end  23  of the airfoil. The squealer tip  38  includes an airfoil shaped squealer tip cap  22 . The airfoil  16  further includes a continuous outer wall  15  with laterally opposite pressure and suction sides  24  and  26 , respectively, extending longitudinally between a leading edge  28  and an opposite trailing edge  30  and radially from the airfoil base  19  to the tip cap  22 . The airfoil is designed to withstand the deteriorating effects of a hot flowpath gas  32 . 
   The airfoil  16  further includes an internal cooling channel or circuit  34  which extends from the tip cap  22  to the root and through the dovetail  12  for circulating or channeling a suitable coolant  36 , such as air which may be bled from a conventional compressor (not shown) for cooling the blade  10 . The internal cooling channel or circuit  34  is radially outwardly bound by the tip cap  22 . The exemplary embodiment of the blade  10  is formed as a one-piece casting of the dovetail  12 , airfoil  16 , and platform  20  of a suitable high temperature metal such as nickel-base superalloys in a single crystal configuration which enjoys suitable strength at high temperature operation. A particular embodiment of the blade  10  is made of a more particular nickel-base superalloy having a free sulfur content less than about 1 part per million by weight (ppmw) which is disclosed in greater detail in U.S. Pat. No. 6,333,121. This low sulfur nickel-base superalloy (also referred to as N5) material helps reduce oxidation of the squealer tip  38 . 
   The squealer tip  38  includes a continuous squealer tip wall  39  extending radially outwardly from and entirely around the airfoil shaped tip cap  22  along the pressure and suction sides  24  and  26 , respectively, of the airfoil  16 . The squealer tip wall  39  and tip cap  22  may be integrally formed or cast with the airfoil or be brazed or welded or otherwise attached to the airfoil. The squealer tip wall  39  extends around the tip cap  22  between laterally spaced apart leading and trailing edges  28  and  30  of the airfoil  16  to define a radially outwardly open tip cavity  40 . 
   Further referring to  FIGS. 3–5 , a recessed tip wall portion  45  is recessed inboard from the pressure side  24  of the airfoil outer wall  15  forming a tip shelf  47  between the recessed tip wall portion  45  and the pressure side  24  of the airfoil outer wall  15 . Thus the internal cooling circuit  34  is bounded in part by the recessed tip wall portion. A plurality of film cooling shelf holes  52  are disposed through the tip shelf  47  to the internal cooling circuit  34 . The shelf holes  52  are spaced away from a junction  57  between the recessed tip wall portion  45  and the tip shelf  47 . The shelf hole centerlines  73  are spaced away from a fillet  59  having a fillet radius R at the junction  57 . The film cooling shelf holes  52  may extend into the fillet no more than 50 percent of a fillet width W of the fillet as measured along the tip shelf  47  from the end of the fillet to the recessed tip wall portion  45 . The location of the film cooling shelf holes  52  away from the recessed tip wall portion  45  reduces or avoids crack initiation. The exemplary embodiment of the turbine blade is designed to have between 18 and 23 shelf holes  52  each having a hole diameter DH in a range of about 14–18 mils (0.014–0.018 inches). 
   Further referring to  FIGS. 6–7 , the film cooling shelf holes  52  have shelf hole centerlines  73  passing through pierce points  200  in the shelf  47  and angled at compound angles C with respect to the vertical lines  79  passing through the pierce points  200 . The compound angles C have orthogonal first and second component angles A and B. The first component angles A lie in first planes E defined by the vertical lines  79  and first coordinate lines  81 , normal to the vertical lines  79 , between the vertical lines  79  and the recessed tip wall portion  45 . The second component angles B lie in second planes F defined by the vertical lines  79  and second coordinate lines  83  that are normal to the vertical lines  79  and normal to first coordinate lines  81 . In the exemplary embodiment of the blade  10  illustrated herein the first coordinate lines  81  lie along transverse lines which are substantially shortest distances  204  between the vertical lines  79  and the recessed tip wall portion  45 . 
   A majority of the film cooling shelf holes  52  have shelf hole centerlines  73  have positive first component angles A and which point in generally outboard directions  61  away from and outboard of the squealer tip wall  39 . Thus, the majority of the shelf hole centerlines  73  are angled in outboard directions  61  away from and outboard of the squealer tip wall  39 . The shelf hole centerlines  73  are angled at the second angles B in downstream lateral directions  63  with respect to vertical lines  79  and the downstream lateral directions  63  are normal to corresponding ones of the outboard directions  61 . 
   Referring to exemplary Table 1 illustrated in  FIG. 8 , the exemplary embodiment of the blade  10  has 23 tip wall film cooling holes H 1 –H 23  of which H 4 –H 19  are the film cooling shelf holes  52 . The tip wall film cooling holes H 1 –H 23  are used to film cool the pressure side  24  of the airfoil outer wall  15  including the recessed tip wall portion  45 . The shelf hole centerlines  73  of the film cooling shelf holes  52  have positive first component angles A in a range between O degrees and 16 degrees. The tip wall film cooling holes H 3 –H 6 , H 9 –H 12 , and H 17 –H 18  illustrate a majority of the film cooling shelf holes  52  having shelf hole centerlines  73  with the positive first component angles A between 2 degrees and 16 degrees. The tip wall film cooling holes H 6 –H 8 , H 11 –H 14 , and H 16 –H 17  illustrate a plurality of the film cooling shelf holes  52  having shelf hole centerlines  73  with the positive first component angles A between 0.5 degrees and 5 degrees. The pressure side  24  of the airfoil outer wall  15  including the recessed tip wall portion  45  is angled away from the shelf hole centerlines  73  in inboard directions at tip angles D as illustrated in  FIGS. 1 and 6  and Table 1. The positive first component angles A of the shelf hole centerlines  73  of the film cooling shelf holes  52  direct the cooling air to be entrained in the boundary layer and not impinge on the pressure side of the squealer tip wall so as to cause a large portion of the cooling air to bounce off the wall and not be entrained in the boundary layer. 
   Referring to  FIGS. 1 and 6 , an external surface  17  of the outer wall  15  of airfoil  16  is film cooled by flowing cooling air through leading edge shower head cooling holes  72  and downstream angled film cooling airfoil holes  74  along the outer wall  15 . The tip wall film cooling holes H 1 –H 2  and H 20 – 23  are radially outwardly angled shaped cooling holes  76  disposed through the pressure side  24  of the airfoil  16  immediately below the tip cap  22  for flowing cooling air radially outwardly along an outboard side  60  of squealer tip wall  39 . The squealer tip wall  39  includes a flat top  62  for maintaining a relatively small radial gap G between the tip wall and a turbine shroud  44  for reducing leakage of the flowpath gas  32  therebetween during operation. During portions of the engine&#39;s operation, the squealer tip wall  39  will rub against the shroud  44  protecting the remainder of the airfoil  16  and tip cap  22  from damage. This will cause an acceptable and planned amount of cracking in the tip wall  39  which is periodically replaced during overhauls. A plurality of chordally spaced apart tip cap supply holes  46  extend radially through the tip cap  22  in flow communication with the cooling circuit  34  inside the airfoil  16  for channeling respective portions of the coolant  36  therefrom and into the tip cavity  40  for cooling the tip, the cavity, and inboard side  66  of the tip wall  39  by convection. The tip cap supply holes  46  are located near the tip wall  39  along the suction side  26  of the continuous outer wall  15  to help purge the cavity  40  of hot gases and cool the tip wall  39 . 
   Illustrated in  FIG. 6  is a thermal barrier coating (TBC)  48  which coats the entire inner surface bounding the tip cavity  40  along inboard side  66  of the squealer tip wall  39 , a radially outwardly facing surface  41  of the tip cap  22  within the squealer tip wall  39 , the flat top  62 , the outboard side  60 , and the external surface  17  of the airfoil  16  along both the pressure and suction sides  24  and  26 , respectively, from the root  18  to the squealer tip  38 . The TBC coatings may be of any well known and conventional composition, such as yttria stabilized zirconia, which is a thermally insulating ceramic material. The thermal barrier coating (TBC) inside and outside the tip cavity  40  and on the tip cap  22  and the flat top  62  of the squealer tip wall  39  insulates the squealer tip  38  from hot gas ingestion and spikes in temperature during engine transients. 
   The present invention has been described in an illustrative manner. It is to be understood that the terminology which has been used is intended to be in the nature of words of description rather than of limitation. While there have been described herein, what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is, therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention. 
   Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims.