Patent Publication Number: US-2003223870-A1

Title: Method and apparatus for reducing turbine blade tip region temperatures

Description:
BACKGROUND OF THE INVENTION  
       [0001] This application relates generally to gas turbine engine rotor blades and, more particularly, to methods and apparatus for reducing rotor blade tip temperatures.  
       [0002] Gas turbine engine rotor blades typically include airfoils having leading and trailing edges, a pressure side, and a suction side. The pressure and suction sides connect at the airfoil leading and trailing edges, and span radially between the airfoil root and the tip, such that a cooling cavity is defined therein. To facilitate reducing combustion gas leakage between the airfoil tips and stationary stator components, the airfoils include a tip region that extends radially outward from the airfoil tip.  
       [0003] The airfoil tip regions include a first tip wall extending from the airfoil leading edge to the trailing edge, and a second tip wall also extending from the airfoil leading edge to connect with the first tip wall at the airfoil trailing edge. The tip region prevents damage to the airfoil if the rotor blade rubs against the stator components.  
       [0004] During operation, combustion gases impacting the rotating rotor blades transfer heat into the blade airfoils and tip regions. Over time, continued operation in higher temperatures may cause the airfoil tip regions to thermally fatigue. To facilitate reducing operating temperatures of the airfoil tip regions, at least some known rotor blades include slots within the tip walls to permit combustion gases at a lower temperature to flow through the tip regions.  
       [0005] To facilitate minimizing thermal fatigue to the rotor blade tips, at least some known rotor blades employ a plurality of tip openings which enable cooling air to discharge from the cooling cavity for cooling the rotor blade tips. After assembly in the rotor, the rotor blades are typically ground as an assembly to a specified diameter for the rotor. During the assembly grind, and during engine operation, the tip openings may become smeared shut, thus decreasing an amount of cooling air that may be discharged from the cooling cavity for cooling the rotor blade tips. Rotor operation with rotor blade tip openings that have been smeared shut may increase the operating temperature of the rotor blade tips, which may result in an increased amount of thermal creep, plastic deformation, and oxidation formation within the trailing edge tip region. Over time, continued operation with rotor blade tip openings that have been smeared shut may decrease the useful life of the rotor assembly.  
       BRIEF SUMMARY OF THE INVENTION  
       [0006] In one aspect of the invention, a method for fabricating a rotor blade for a gas turbine engine to facilitate reducing operating temperatures of a tip portion of the rotor blade is provided. The method comprises forming an airfoil including a first sidewall and a second sidewall connected at a leading edge and a trailing edge to define a cavity therein, wherein the first and second sidewalls extend radially between a rotor blade root and a rotor blade tip, and forming a first tip wall extending from the rotor blade tip plate along the first sidewall, such that at least a portion of the first tip wall is at least partially recessed with respect to the rotor blade first sidewall to define a tip shelf that extends from the airfoil trailing edge towards the airfoil leading edge. The method also comprises forming a second tip wall extending from the rotor blade tip plate along the second sidewall such that the second tip wall connects with the first tip wall at the rotor blade trailing edge.  
       [0007] In a further aspect, an airfoil for a gas turbine engine is provided. The airfoil includes a leading edge, a trailing edge, a tip plate, a first sidewall, a second sidewall, a first tip wall, and a second tip wall. The first tip wall extends in radial span between an airfoil root and the tip plate. The second sidewall is connected to the first sidewall at the leading edge and the trailing edge to define a cavity therein. The second sidewall extends in radial span between the airfoil root and the tip plate. The first tip wall extends radially outward from the tip plate along the first sidewall. The second tip wall extends radially outward from the tip plate along the second sidewall. The first tip wall is connected to the second tip wall at the leading edge. The first tip wall is at least partially recessed with respect to the rotor blade first sidewall to define a tip shelf that extends from the airfoil trailing edge towards the airfoil leading edge. The tip shelf includes at least one opening extending therethrough to the cavity.  
       [0008] In another aspect of the invention, a gas turbine engine including a plurality of rotor blades is provided. Each rotor blade includes an airfoil including a leading edge, a trailing edge, a first sidewall, a second sidewall, a first tip wall, and a second tip wall. The airfoil first and second sidewalls are connected axially at the leading and trailing edges to define a cavity within the airfoil. The first and second sidewalls extend radially from a blade root to the tip plate. The first tip wall extends radially outward from the tip plate along the first sidewall. The second tip wall extends radially outward from the tip plate along the second sidewall. The first tip wall is at least partially recessed with respect to the rotor blade first sidewall to define a tip shelf that extends from the airfoil trailing edge towards the airfoil leading edge. The tip shelf includes at least one opening extending therethrough to the airfoil cavity. 
     
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
     [0009]FIG. 1 is a schematic illustration of a gas turbine engine; and  
     [0010]FIG. 2 is a perspective view of a rotor blade that may be used with the gas turbine engine shown in FIG. 1; and  
     [0011]FIG. 3 is an enlarged partial perspective view of a portion of the rotor blade shown in FIG. 2.  
    
    
     DETAILED DESCRIPTION OF THE INVENTION  
     [0012]FIG. 1 is a schematic illustration of a gas turbine engine  10  including a fan assembly  12 , 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  28  and an exhaust side  30 . In one embodiment, engine  10  is a CT7 engine commercially available from General Electric Aircraft Engines, Cincinnati, Ohio.  
     [0013] In operation, air flows through fan assembly  12  and compressed air is supplied to high pressure compressor  14 . 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 .  
     [0014]FIG. 2 is a perspective view of a rotor blade  40  that may be used with gas turbine engine  10  (shown in FIG. 1). FIG. 3 is an enlarged partial perspective view of a portion of rotor blade  40 . In one embodiment, a plurality of rotor blades  40  form a high pressure turbine rotor blade stage (not shown) of gas turbine engine  10 . Each rotor blade  40  includes a hollow airfoil  42  and an integral dovetail  43  used for mounting airfoil  42  to a rotor disk (not shown) in a known manner.  
     [0015] Airfoil  42  includes a first sidewall  44  and a second sidewall  46 . First sidewall  44  is convex and defines a suction side of airfoil  42 , and second sidewall  46  is concave and defines a pressure side of airfoil  42 . Sidewalls  44  and  46  are connected at a leading edge  48  and at an axially-spaced trailing edge  50  of airfoil  42  that is downstream from leading edge  48 .  
     [0016] First and second sidewalls  44  and  46 , respectively, extend longitudinally or radially outward to span from a blade root  52  positioned adjacent dovetail  43  to a tip plate  54  which defines a radially outer boundary of an internal cooling chamber (not shown). The cooling chamber is defined within airfoil  42  between sidewalls  44  and  46 . Internal cooling of airfoils  42  is known in the art. In one embodiment, the cooling chamber includes a serpentine passage cooled with compressor bleed air. In the exemplary embodiment, sidewall  46  includes a plurality of film cooling openings  60 , extending therethrough to facilitate additional cooling of the cooling chamber and airfoil trailing edge  50 .  
     [0017] Airfoil  42  also includes a plurality of trailing edge openings  62 . More specifically, openings  62  extend radially between tip plate  54  and blade root  52  for discharging cooling fluid from the cooling chamber to facilitate cooling airfoil trailing edge  50 .  
     [0018] A tip region  70  of airfoil  42  is sometimes known as a squealer tip, and includes a first tip wall  72  and a second tip wall  74  formed integrally with airfoil  42 . First tip wall  72  extends from airfoil leading edge  48  along airfoil first sidewall  44  to airfoil trailing edge  50 . More specifically, first tip wall  72  extends from tip plate  54  to an outer edge  75  for a height  76 . In the exemplary embodiment, first tip wall height  76  is substantially constant along first tip wall  72 . In alternative embodiments height  76  is not substantially constant along first tip wall  72 .  
     [0019] Second tip wall  74  extends from airfoil leading edge  48  along second sidewall  46  towards airfoil trailing edge  50 . More specifically, second tip wall  74  is connected to first tip wall  72  at airfoil leading edge  48  and extends aftward towards trailing edge  50 . Because second tip wall  74  is laterally spaced from first tip wall  72 , an open-top tip cavity  80  is defined with tip walls  72  and  74 , and tip plate  54 . Second tip wall  74  also extends radially outward from tip plate  54  to an outer edge  82  for a height  84 . In the exemplary embodiment, second tip wall height  84  is approximately equal first tip wall height  76 . Alternatively, second tip wall height  84  is not equal first tip wall height  76 .  
     [0020] Second tip wall  74  is recessed at least in part from airfoil second sidewall  46 . More specifically, second tip wall  74  is recessed from airfoil second sidewall  46  to couple with first tip wall  72  a distance  88  from airfoil trailing edge  50  such that a tip shelf  90  is defined. More specifically, tip shelf  90  extends aftward from a front edge  94 , that is distance  88  from airfoil trailing edge  50 , to trailing edge  50  adjacent first tip wall  72 . Tip shelf  90  is a distance  96  from first tip wall outer edge  75 . In the exemplary embodiment, distance  96  is approximately equal first tip shelf height  76  such that tip shelf  90  is substantially co-planar with tip plate  54 . In alternative embodiments, tip shelf  90  is not substantially co-planar with tip plate  54 .  
     [0021] In the exemplary embodiment, tip shelf  90  includes a plurality of tip openings  100  spaced axially along tip shelf  90 . Openings  100  extend through tip shelf  90  for discharging cooling fluid from the cooling cavity for cooling airfoil trailing edge  50 . More specifically, openings  100  are for discharging cooling fluid towards first tip wall  72  to facilitate reducing operating temperatures within the region  110  of first tip wall adjacent trailing edge  50 . In one embodiment, openings  100  extend obliquely, rather than normal, through tip shelf  90  to facilitate discharging cooling fluid at first tip wall region  110 .  
     [0022] During operation, squealer tip walls  72  and  74  are positioned in close proximity with a conventional stationary stator shroud (not shown), and define a tight clearance (not shown) therebetween that facilitates reducing combustion gas leakage therethrough. Tip walls  72  and  74  extend radially outward from airfoil  42 . Accordingly, if rubbing occurs between rotor blades  40  and the stator shroud, only tip walls  72  and  74  contact the shroud and airfoil  42  remains intact. Furthermore, if rubbing occurs, because tip shelf cooling openings  100  are a distance  96  from first tip wall outer edge  75 , tip shelf distance  96  facilitates reducing smearing of tip shelf  90  and inadvertent sealing of openings  100 .  
     [0023] Because combustion gases assume a parabolic profile flowing through a turbine flowpath at blade tip region leading edge  48 , combustion gases near turbine blade tip region  70  are at a lower temperature than gases near a blade pitch line (not shown) of turbine blades  40 . As combustion gases flow from blade tip region leading edge  48  towards blade trailing edge  50 , hotter gases near the pitch line migrate radially towards rotor blade tip region  70  and first tip wall region  110  due to blade rotation.  
     [0024] Tip shelf  90  provides a discontinuity in airfoil pressure side  46  which causes the hotter combustion gases to separate from airfoil second sidewall  46 , thus facilitating a decrease in heat transfer thereof. Additionally, tip shelf openings  100  enable cooling air to be discharged from the internal cooling chamber to provide additional film and convective cooling of trailing edge  50  and first tip wall region  110 . Furthermore, sidewall film cooling openings  60  discharge additional film and convective cooling air from the cooling chamber to facilitate additional cooling of the cooling chamber and airfoil trailing edge  50 . The combination of the film cooling and the convection cooling, known as a double row cooling scheme, facilitates reducing the operating temperature of trailing edge  50  and first tip wall region  110 .  
     [0025] The above-described rotor blade is cost-effective and highly reliable. The rotor blade includes a tip shelf extending from the airfoil trailing edge towards the airfoil trailing edge. The tip shelf disrupts combustion gases flowing past the airfoil to facilitate the formation of a cooling layer against the tip shelf. As a result, cooler operating temperatures within the rotor blade facilitate extending a useful life of the rotor blades in a cost-effective and reliable manner.  
     [0026] While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.