Patent Publication Number: US-2013236325-A1

Title: Blade tip profile

Description:
BACKGROUND 
     A gas turbine engine is a rotating device that uses the action of a fluid to produce work. In a gas turbine engine, a pressurized, high temperature gas is the driving force. One of the reasons that gas turbine engines are widely used to power aircraft is they are light and compact and have a high power-to-weight ratio. 
     Gas turbine engines typically include several stages including a fan, a compressor, a combustor, and a turbine. Some of these stages utilize rotating airfoils with shaped blades arranged in series. The blades convert energy from combustion gases produced by the combustor into mechanical work used to turn a rotor. The blades positioned forward of the combustor are turned by the rotor to compress air entering the combustor. 
     Blades usually have a clearance gap between the blade tip and the stationary casing or the shroud surface adjacent the blade tip. This clearance gap is typically called a tip gap. This clearance gap is necessary to allow for rotation of the blade and to allow for mechanical and thermal growth of the blade. Due to the pressure difference between the pressure side and the suction side of the blade, hot gas leaks across this gap from the pressure side to the suction side. This phenomenon is known as tip leakage flow. 
     The tip leakage flow results in a reduction in the blade force, (e.g., the work done) and therefore, the overall efficiency of the gas turbine engine. In fact, the losses that are attributed to leakage flow can be very substantial. Leakage flow also increases the thermal loading on the blade tip, leading to high local temperatures adjacent the tip, and thus, is considered one of the primary sources of blade failure. Typically, leakage flow is reduced by using a recessed or squealer tip blade or by modifying the blade tip to incorporate film cooling. 
     Blades, including turbine blades in particular, can utilize a pocket recess which comprises a recess cavity that extends radially through the length of the blade. The pocket recess creates an opening at the tip of the blade. The pocket recess is used for efficiency purposes to reduce the weight of the blade and to reduce blade creep. However, due to excessive blade tip leakage, the efficiency of the pocket recess is limited. 
     SUMMARY 
     An airfoil includes, a blade having a leading edge, a trailing edge, a pressure side, a suction side, a tip, and a scoop. The scoop extends along the tip of the blade. The scoop comprises a difference in a radial height of the blade from a pressure side to a suction side of the blade. The radial height of the blade at the pressure side is less than the radial height of the blade at the suction side. 
     In another aspect, a blade includes an outer radial surface disposed along a suction side of a tip of the blade, an inner radial surface disposed along a pressure side of the tip of the blade, and a pocket recess. The inner radial surface has a different radial height than the outer radial surface. The pocket recess is open at the tip of the blade and is disposed between the inner radial surface and the outer radial surface. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a perspective view of a first embodiment of an airfoil including a blade tip scoop and a pocket recess. 
         FIGS. 2A and 2B  are perspective views of the airfoil of  FIG. 1 . 
         FIGS. 3A and 3B  are perspective views of a second embodiment of an airfoil including a blade tip scoop and a pocket recess. 
         FIGS. 4A and 4B  are perspective views of a third embodiment of an airfoil including a blade tip scoop and a pocket recess. 
         FIGS. 5A and 5B  are perspective views of a fourth embodiment of an airfoil including a blade tip scoop and a pocket recess. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  shows a first embodiment of an airfoil  8 A for a gas turbine engine including a blade  10 A, a blade tip  12 A, and a pocket recess  14 A disposed along a camber line  16 A of blade  10 A. Blade  10 A includes a leading edge  18 A, a trailing edge  20 A, a pressure surface  22 A, and a suction surface  24 A. Blade  10 A additionally includes a scoop  26 A disposed along blade tip  12 A. Due to scoop  26 A, blade tip  12 A includes an outer radial surface  28 A and an inner radial surface  30 A. During operation, when mounted on a rotor (not shown), airfoil  8 A would rotate in the direction D illustrated to work air A flowing past airfoil  8 A, as shown in  FIG. 2A . During such operation, tip leakage T L  would occur. Scoop  26 A would generate a rotational vortex V ( FIGS. 2A and 2B ) that would rotate in an opposing direction to the direction D of rotation of airfoil  8 A and in an opposing direction to the direction of tip leakage T L . This counter rotating vortex V reduces tip leakage T L , and thereby improves the efficiency of the airfoil  8 A and gas turbine engine. 
     Airfoil  8 A is of conventional design and includes a blade  10 A extending generally radially outward from a platform section (not numbered) and a root section (not numbered) to blade tip  12 A. When installed, blade tip  12 A is disposed adjacent gas turbine engine stator case (not shown). Pocket recess  14 A extends into blade  10 A at blade tip  12 A. In the embodiment shown in  FIGS. 1 ,  2 A, and  2 B, pocket recess  14 A is symmetric with respect to camber line  16 A of blade  10 A. Thus, pocket recess  14 A straddles and is bifurcated by camber line  16 A, is disposed adjacent leading edge  18 A in a thicker region of blade  10 A, and is substantially equidistant from pressure surface  22 A and suction surface  24 A of blade  10 A. 
     Blade  10 A extends from leading edge  18 A along concave pressure surface  22 A and along convex suction surface  24 A to trailing edge  20 A. For reference purposes, camber line  16 A extends along blade tip  12 A from leading edge  18 A to trailing edge  20 A. Pocket recess  14 A is separated from exterior of blade  10 A and pressure surface  22 A by first wall  32 A. Similarly, pocket recess  14 A is separated from exterior of blade  10 A and suction surface  24 A by second wall  34 A. Because pocket recess  14 A is symmetric with respect to camber line  16 A, first wall  32 A has substantially a same thickness as second wall  34 A along a corresponding extent of pocket recess  14 A. 
     Scoop  26 A extends along blade tip  12 A from leading edge  18 A to trailing edge  20 A. Thus, in the embodiment shown both outer radial surface  28 A and inner radial surface  30 A of blade tip  12 A extend from leading edge  18 A to trailing edge  20 A. Scoop  26 A comprises a cutout along a pressure side of blade  10 A. Thus, scoop  26 A forms step feature that extends along blade tip  12 A. In the embodiment shown, the step or change in radial height H in blade  10 A between outer radial surface  28 A and inner radial surface  30 A is generally aligned on camber line  16 A. Outer radial surface  28 A extends along a suction side of blade tip  12 A and inner radial surface  30 A extends along a pressure side of blade tip  12 A. 
     Because the construction and operation of gas turbine engines is known in the art, gas turbine engine will not be discussed in great detail. While blade  10 A is shown as a separate component removable from a rotor (not shown) in other embodiments airfoil can be integrated with the rotor. Although described with reference to a turbine airfoil, in other embodiments blade can be utilized in the compressor or other stage of the gas turbine engine. 
     Consider an incompressible inviscid turbine tip clearance flow field; the flow through the tip gap is obtained from Bernoulli equation is as shown in Equation 1: 
     
       
         
           
             
               
                 
                   
                     W 
                     L 
                   
                   = 
                   
                     
                       2 
                        
                       
                         
                           
                             P 
                             p 
                           
                           - 
                           
                             P 
                             s 
                           
                         
                         ρ 
                       
                     
                   
                 
               
               
                 
                   ( 
                   1 
                   ) 
                 
               
             
           
         
       
     
     Where Pp=Pressure side pressure, Ps=Suction side pressure, ρ=. Gas density, and W L =Leakage flow. 
     As shown in  FIGS. 2A and 2B , leakage flow W L  is reduced if a tip vortex V is generated at the blade tip in the opposite direction to tip leakage flow T L  and in the opposite direction to blade  10 A rotation. The rolling up of the counter rotation of tip vortex V is due to scoop  26 A with step from pressure to suction side of blade tip  12 A. Additionally, vortex creation is enhanced by pocket recess  14 A. Results regarding vortex creation can be verified and analyzed utilizing commercially available computational fluid dynamics software. 
     As previously discussed, creation of counter rotating vortex V at blade tip  12 A leads to a reduction in blade tip leakage T L , which results in improved gas turbine engine efficiency and fuel savings. Operating temperatures within the turbine engine can be increased as a result of decreased temperature localization, which results in reduced emissions. Additionally, reduced blade tip leakage T L  results in extended hot section durability due to a reduction in the thermal loading on the blade tip  12 A. In instances of disc-blade coupling mistuning, a turbine wheel with tip modification as described would result in tuning out the interfered frequency and also aids in maintaining a balanced rotor. 
       FIGS. 2A and 2B  provide additional perspective views of blade  10 A with scoop  26 A. As discussed previously, the radial height. H of blade  10 A differs between outer radial surface  28 A and inner radial surface  30 A. Scoop  26 A has a fillet radius R along radial height H between outer radial surface  28 A and inner radial surface  30 A. In the embodiments shown in  FIGS. 2A and 2B , scoop  26 A additionally includes an aft length L 1  and a forward length L 2  along the step transition between outer radial surface  28 A and inner radial surface  30 A. Outer radial surface  28 A has a thickness T 1  along camber line  16 A that is substantially the same as a corresponding thickness T 2  of inner radial surface  30 A (T 1 ≈T 2 ). 
     Scoop  26 A comprises a curved step between outer radial surface  28 A and inner radial surface  30 A and extends the entire length of blade tip  12 A from leading edge  18 A to trailing edge  20 A. Radial height H between outer radial surface  28 A and inner radial surface  30 A will vary from embodiment to embodiment. In one embodiment, the radial height H is greater than or equal to the tip clearance (the distance between the stator casing and the outermost radial extent of blade tip  12 A) and is less than or equal to about 15% of the blade span. Radius R is greater than or equal to height H. Aft length L 1  and a forward length L 2  along step feature between outer radial surface  28 A and inner radial surface  30 A will vary from embodiment to embodiment and vary as a function of pocket axial length, pocket configuration, and step lateral thickness (expressed in term of T 1  and T 2 ). 
       FIGS. 3A and 3B  show a second embodiment of an airfoil  8 B for a gas turbine engine including a blade  10 B, a blade tip  12 B, a pocket recess  14 B, and a camber line  16 B of blade  10 B. Blade  10 B includes a leading edge  18 B, a trailing edge  20 B, a pressure surface  22 B, and a suction surface  24 B. Blade  10 B additionally includes a scoop  26 B disposed along blade tip  12 B. Due to scoop  26 B, blade tip  12 B includes an outer radial surface  28 B and an inner radial surface  30 B. During operation when mounted on a rotor (not shown) airfoil  8 B would rotate in the direction D illustrated. During such operation, tip leakage T L  would occur. Scoop  26 B generates a rotational vortex V that would rotate in an opposing direction to the direction D of rotation of airfoil  8 B and in an opposing direction to the direction of tip leakage T L . This counter rotating vortex V would reduce tip leakage T L , and thereby, improve the efficiency of the airfoil  8 B and gas turbine engine. 
     Pocket recess  14 B extends into blade  10 B from blade tip  12 B. In the embodiment shown in  FIGS. 3A and 3B , pocket recess  14 B is asymmetric with respect to camber line  16 B of blade  10 B. Thus, a first wall  32 B of the blade  10 B has a thickness that differs from a corresponding thickness of second wall  34 B at a substantially similar location with respect to camber line  16 B. Pocket recess  14 B is biased toward the pressure side of camber line  16 B leaving first wall  32 B a thinner thickness than second wall  34 B along a corresponding extent of pocket recess  14 B. 
     For reference purposes, camber line  16 B extends along blade tip  12 B from leading edge  18 B to trailing edge  20 B. Blade  10 B extends from leading edge  18 B along concave pressure surface  22 B and along convex suction surface  24 B to trailing edge  20 B. Scoop  26 B is biased away from camber line  16 B toward pressure side. Scoop  26 B extends along blade tip  12 B from leading edge  18 B and trailing edge  20 B. Thus, in the embodiment shown inner radial surface  30 B and outer radial surface  28 B of blade tip  12 B extend from leading edge  18 B to trailing edge  20 B. Scoop  26 A comprises a cutout or step feature that extends along blade tip  12 B. 
     Due to the changes in the configuration of scoop  26 B and pocket recess  14 B compared to the embodiments of  FIGS. 2A and 2B , lateral thickness T 1  of outer radial surface  28 B along camber line  16 B greater than corresponding lateral thickness T 2  of inner radial surface  30 A. 
       FIGS. 4A and 4B  show a third embodiment of an airfoil  8 C for a gas turbine engine including a blade  10 C, a blade tip  12 C, a pocket recess  14 C, and a camber line  16 C of blade  10 C. Blade  10 C includes a leading edge  18 C, a trailing edge  20 C, a pressure surface  22 C, and a suction surface  24 C. Blade  10 C additionally includes a scoop  26 C disposed along blade tip  12 C. Due to scoop  26 C, blade tip  12 C includes an outer radial surface  28 C and an inner radial surface  30 C. During operation when mounted on a rotor (not shown) airfoil  8 C would rotate in the direction D illustrated. During such operation, tip leakage T L  would occur. Scoop  26 C generates a rotational vortex V that would rotate in an opposing direction to the direction D of rotation of airfoil  8 C and in an opposing direction to the direction of tip leakage T L . This counter rotating vortex V would reduce tip leakage T L , and thereby, improve the efficiency of the airfoil  8 C and gas turbine engine. 
     Pocket recess  14 C extends into blade  10 C from blade tip  12 C. In the embodiment shown in  FIGS. 4A and 4B , pocket recess  14 C is asymmetric with respect to camber line  16 C of blade  10 C. Thus, a first wall  32 C of the blade  10 C has a thickness that differs from a corresponding thickness of second wall  34 C at a substantially similar location with respect to camber line  16 C. Pocket recess  14 C is biased toward the suction side of camber line  16 C leaving first wall  32 C a thicker thickness than second wall  34 C along a corresponding extent of pocket recess  14 C. 
     Blade  10 C extends from leading edge  18 C along concave pressure surface  22 C and along convex suction surface  24 C to trailing edge  20 C. For reference purposes, camber line  16 C extends along blade tip  12 C from leading edge  18 C to trailing edge  20 C. Step between outer radial surface  28 C and inner radial surface  30 C runs along and straddles camber line  16 C. Scoop  26 C is of a reduced size and is offset from camber line  16 C. Thus, scoop  26 C is asymmetric with respect to camber line  16 C. Therefore, outer radial surface  28 C has a lateral thickness T 1  along camber line  16 C that differs from a corresponding lateral thickness T 2  of inner radial surface  30 C (T 1 &lt;T 2 ) and scoop  26 C extends from leading edge  18 C to trailing edge  20 C. Due to the changes in position of pocket cavity  14 C compared to the embodiments of  FIGS. 1 ,  2 A,  2 B,  3 A, and  3 B aft length L 1  and forward length L 2  sections differ in extent. 
       FIGS. 5A and 5B  show a fourth embodiment of an airfoil  8 D for a gas turbine engine including a blade  10 D, a blade tip  12 D, a pocket recess  14 D, and a camber line  16 D of blade  10 D. Blade  10 D includes a leading edge  18 D, a trailing edge  20 D, a pressure surface  22 D, and a suction surface  24 D. Blade  10 D additionally includes a scoop  26 D disposed along blade tip  12 D. Due to scoop  26 D, blade tip  12 D includes an outer radial surface  28 D and an inner radial surface  30 D. During operation when mounted on a rotor (not shown) airfoil  8 D would rotate in the direction D illustrated. During such operation, tip leakage T L  would occur. Scoop  26 D generates a rotational vortex V that would rotate in an opposing direction to the direction D of rotation of airfoil  8 D and in an opposing direction to the direction of tip leakage T L . This counter rotating vortex V would reduce tip leakage T L , and thereby, improve the efficiency of the airfoil  8 D and gas turbine engine. 
     Pocket recess  14 D extends into blade  10 D from blade tip  12 D. In the embodiment shown in  FIGS. 5A and 5B , pocket recess  14 D is asymmetric with respect to camber line  16 D of blade  10 D. Pocket recess  14 D is disposed such that pocket recess  14 D is angled with respect to camber line  16 D. Because pocket recess  14 D is disposed asymmetrically with respect to camber line  16 D, first wall  32 D of the blade  10 D has increasing thickness along the axial length of pocket recess  14 D from forward to aft and second wall  34 D with a decreasing thickness along the axial length of pocket recess  14 D. First wall  32 D has a lesser thickness adjacent leading edge  18 D than aft near a trailing termination edge of pocket recess  14 D. Similarly, a corresponding thickness of second wall  34 D is greater near the leading edge  18 D and decreases in thickness with travel aft along pocket recess  14 D. Thus, second wall  34 D has decreasing thickness along the length of pocket recess  14 D from forward to aft. 
     Scoop  26 D is disposed asymmetrically with respect to camber line  16 D such that step feature does not straddle or follow camber line  16 D. Rather, outer radial surface  28 D has a thickness T 1  along camber line  16 D that differs from a corresponding thickness T 2  of inner radial surface  30 D. During aft length L 1  of scoop  26 D outer radial surface  28 D thickness T 1  is less than corresponding thickness T 2  of inner radial surface  30 D. During forward length L 2  of scoop  26 D, outer radial surface  28 D thickness T 1  is greater than corresponding thickness T 2  of inner radial surface  30 D. 
     An asymmetric pocket cavity such as pocket cavities  14 B,  14 C, and  14 D alters the mass/stiffness of the blade  10 B,  10 C, and  10 D, thereby shifting or tuning away the natural frequency of the pocket cavity and blade from the frequency of acoustic pressure oscillation or the frequency of the aero-excitation source. 
     More particularly, blade can be tuned at blade anti-nodes as further discussed in United States Patent Application Publications 2010/0278632A and  2010 / 0278633 A, which are incorporated herein by reference. Tuning is performed by modifying the stiffness/mass (i.e. wall thickness) at one or more blade anti-nodes. Increasing the mass at the blade anti-node decreases natural frequency, and decreasing mass at blade anti-node increases natural frequency. Wall thickness as a result of pocket recess geometry can be modified until the natural frequency of the blade resonant mode shapes that have interferences are moved out of the expected acoustic pressure oscillation frequency or out of the aero-excitation source frequency. Wall thickness as a result of the pocket recess geometry can be further modified to further increase a substantially resonance-free running range. If further tuning is desired, the pocket recess geometry can be modified on one or more additional blade anti-nodes until the blade has no natural frequencies that excite at the expected acoustic pressure oscillation frequency or aero-excitation frequency. The natural frequency of the blade resonant mode shapes can be modeled using a finite element method. 
     While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.