Turbine blade with recessed squealer tip and shelf

A turbine blade squealer tip has a continuous squealer tip wall extending radially outwardly from and continuously around a tip cap. A recessed tip wall portion of the tip wall is recessed inboard from a pressure side of an airfoil outer wall of an airfoil of the blade forming a tip shelf therebetween. A plurality of film cooling shelf holes are disposed through the tip shelf to an internal cooling circuit of the blade and are spaced away from a junction between the recessed tip wall portion and the tip shelf. The exemplary embodiment of the airfoil includes shelf hole centerlines of the holes passing through pierce points in the shelf. At least a majority of the shelf hole centerlines are angled in outboard directions away from and outboard of the squealer tip wall. A majority of centerlines are angled away from vertical lines passing through the pierce points at first component angles in a range between 2 degrees and 16 degrees.

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.

DETAILED DESCRIPTION OF THE INVENTION

Illustrated inFIGS. 1 and 2is an exemplary gas turbine engine turbine rotor blade10configured for use as a first stage high pressure turbine blade. The blade10includes a dovetail12having suitable tangs13for mounting the blade in corresponding dovetail slots in the perimeter of a rotor disk (not shown). The blade10further includes an airfoil16joined to the dovetail12at an airfoil base19at an integral platform20and a squealer tip38at a radially outer end23of the airfoil. The squealer tip38includes an airfoil shaped squealer tip cap22. The airfoil16further includes a continuous outer wall15with laterally opposite pressure and suction sides24and26, respectively, extending longitudinally between a leading edge28and an opposite trailing edge30and radially from the airfoil base19to the tip cap22. The airfoil is designed to withstand the deteriorating effects of a hot flowpath gas32.

The airfoil16further includes an internal cooling channel or circuit34which extends from the tip cap22to the root and through the dovetail12for circulating or channeling a suitable coolant36, such as air which may be bled from a conventional compressor (not shown) for cooling the blade10. The internal cooling channel or circuit34is radially outwardly bound by the tip cap22. The exemplary embodiment of the blade10is formed as a one-piece casting of the dovetail12, airfoil16, and platform20of 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 blade10is 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 tip38.

The squealer tip38includes a continuous squealer tip wall39extending radially outwardly from and entirely around the airfoil shaped tip cap22along the pressure and suction sides24and26, respectively, of the airfoil16. The squealer tip wall39and tip cap22may be integrally formed or cast with the airfoil or be brazed or welded or otherwise attached to the airfoil. The squealer tip wall39extends around the tip cap22between laterally spaced apart leading and trailing edges28and30of the airfoil16to define a radially outwardly open tip cavity40.

Further referring toFIGS. 3–5, a recessed tip wall portion45is recessed inboard from the pressure side24of the airfoil outer wall15forming a tip shelf47between the recessed tip wall portion45and the pressure side24of the airfoil outer wall15. Thus the internal cooling circuit34is bounded in part by the recessed tip wall portion. A plurality of film cooling shelf holes52are disposed through the tip shelf47to the internal cooling circuit34. The shelf holes52are spaced away from a junction57between the recessed tip wall portion45and the tip shelf47. The shelf hole centerlines73are spaced away from a fillet59having a fillet radius R at the junction57. The film cooling shelf holes52may extend into the fillet no more than 50 percent of a fillet width W of the fillet as measured along the tip shelf47from the end of the fillet to the recessed tip wall portion45. The location of the film cooling shelf holes52away from the recessed tip wall portion45reduces or avoids crack initiation. The exemplary embodiment of the turbine blade is designed to have between 18 and 23 shelf holes52each having a hole diameter DH in a range of about 14–18 mils (0.014–0.018 inches).

Further referring toFIGS. 6–7, the film cooling shelf holes52have shelf hole centerlines73passing through pierce points200in the shelf47and angled at compound angles C with respect to the vertical lines79passing through the pierce points200. 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 lines79and first coordinate lines81, normal to the vertical lines79, between the vertical lines79and the recessed tip wall portion45. The second component angles B lie in second planes F defined by the vertical lines79and second coordinate lines83that are normal to the vertical lines79and normal to first coordinate lines81. In the exemplary embodiment of the blade10illustrated herein the first coordinate lines81lie along transverse lines which are substantially shortest distances204between the vertical lines79and the recessed tip wall portion45.

A majority of the film cooling shelf holes52have shelf hole centerlines73have positive first component angles A and which point in generally outboard directions61away from and outboard of the squealer tip wall39. Thus, the majority of the shelf hole centerlines73are angled in outboard directions61away from and outboard of the squealer tip wall39. The shelf hole centerlines73are angled at the second angles B in downstream lateral directions63with respect to vertical lines79and the downstream lateral directions63are normal to corresponding ones of the outboard directions61.

Referring to exemplary Table 1 illustrated inFIG. 8, the exemplary embodiment of the blade10has 23 tip wall film cooling holes H1–H23of which H4–H19are the film cooling shelf holes52. The tip wall film cooling holes H1–H23are used to film cool the pressure side24of the airfoil outer wall15including the recessed tip wall portion45. The shelf hole centerlines73of the film cooling shelf holes52have positive first component angles A in a range between O degrees and 16 degrees. The tip wall film cooling holes H3–H6, H9–H12, and H17–H18illustrate a majority of the film cooling shelf holes52having shelf hole centerlines73with the positive first component angles A between 2 degrees and 16 degrees. The tip wall film cooling holes H6–H8, H11–H14, and H16–H17illustrate a plurality of the film cooling shelf holes52having shelf hole centerlines73with the positive first component angles A between 0.5 degrees and 5 degrees. The pressure side24of the airfoil outer wall15including the recessed tip wall portion45is angled away from the shelf hole centerlines73in inboard directions at tip angles D as illustrated inFIGS. 1 and 6and Table 1. The positive first component angles A of the shelf hole centerlines73of the film cooling shelf holes52direct 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 toFIGS. 1 and 6, an external surface17of the outer wall15of airfoil16is film cooled by flowing cooling air through leading edge shower head cooling holes72and downstream angled film cooling airfoil holes74along the outer wall15. The tip wall film cooling holes H1–H2and H20–23are radially outwardly angled shaped cooling holes76disposed through the pressure side24of the airfoil16immediately below the tip cap22for flowing cooling air radially outwardly along an outboard side60of squealer tip wall39. The squealer tip wall39includes a flat top62for maintaining a relatively small radial gap G between the tip wall and a turbine shroud44for reducing leakage of the flowpath gas32therebetween during operation. During portions of the engine's operation, the squealer tip wall39will rub against the shroud44protecting the remainder of the airfoil16and tip cap22from damage. This will cause an acceptable and planned amount of cracking in the tip wall39which is periodically replaced during overhauls. A plurality of chordally spaced apart tip cap supply holes46extend radially through the tip cap22in flow communication with the cooling circuit34inside the airfoil16for channeling respective portions of the coolant36therefrom and into the tip cavity40for cooling the tip, the cavity, and inboard side66of the tip wall39by convection. The tip cap supply holes46are located near the tip wall39along the suction side26of the continuous outer wall15to help purge the cavity40of hot gases and cool the tip wall39.

Illustrated inFIG. 6is a thermal barrier coating (TBC)48which coats the entire inner surface bounding the tip cavity40along inboard side66of the squealer tip wall39, a radially outwardly facing surface41of the tip cap22within the squealer tip wall39, the flat top62, the outboard side60, and the external surface17of the airfoil16along both the pressure and suction sides24and26, respectively, from the root18to the squealer tip38. 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 cavity40and on the tip cap22and the flat top62of the squealer tip wall39insulates the squealer tip38from 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.