Patent Publication Number: US-8535004-B2

Title: Four-wall turbine airfoil with thermal strain control for reduced cycle fatigue

Description:
STATEMENT REGARDING FEDERALLY SPONSORED DEVELOPMENT 
     Development for this invention was supported in part by Contract No. DE-FC26-05NT42644, awarded by the United States Department of Energy. Accordingly, the United States Government may have certain rights in this invention. 
    
    
     FIELD OF THE INVENTION 
     This invention is related generally to turbine airfoils, and more particularly to hollow turbine airfoils such as blades and vanes with internal cooling channels for passing fluids such as air to cool the airfoils. 
     BACKGROUND OF THE INVENTION 
     Gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade and vane assembly for producing power. Combustors operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose the turbine vane and blade assemblies to these high temperatures. Turbine vanes and blades must be made of materials capable of withstanding such temperatures. Turbine vanes and blades often contain cooling systems for prolonging their life and reducing the likelihood of failure as a result of excessive temperatures. 
     A turbine blade is a rotating airfoil attached to a disk on the turbine rotor by a platform and blade shank. A turbine vane is a stationary airfoil that is radially oriented with respect to a rotation axis of the turbine rotor. The vanes direct the combustion gas flow optimally against the blades. One or each end of a vane airfoil is coupled to a platform, also known as an endwall. A radially outer vane platform is connected to a retention ring on the engine casing. An inner vane platform, if present, is supported by the vane. 
     Blades and vanes often contain cooling circuits forming a cooling system. The cooling circuits receive a cooling fluid such as air bled from the compressor of the turbine engine via a plenum and supply port in one or each platform. The cooling circuits often include multiple flow paths inside the airfoil designed to maintain all portions of the airfoil at a relatively uniform temperature. At least some of the air passing through these cooling circuits may be exhausted through film cooling holes in the leading edge, trailing edge, suction side, and pressure side of the airfoil. 
     Some turbine airfoils have a dual wall structure formed of inner and outer walls. This is called a 4-wall airfoil construction, since the pressure and suction sides of the airfoil each have two walls. The outer wall is exposed to hotter temperatures, so it is subject to greater thermal expansion, and stress develops at the connection between the inner and outer walls. 
     It is known that high cooling efficiency can be achieved by near-wall cooling in which cooling air flows in channels between the inner and outer walls of a 4-wall airfoil. However, differential thermal expansion between the hot outer walls and the cooler inner walls can cause Low Cycle Fatigue (LCF) limitations for reasons later described. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention is explained in the following description in view of the drawings that show: 
         FIG. 1  is a sectional view of prior art 4-wall turbine airfoil such as a vane or blade. 
         FIG. 2  is a sectional view of a turbine airfoil showing aspects of the invention. 
         FIG. 3  is a sectional view taken along line  3 - 3  of  FIG. 2 . 
         FIG. 4  is an outline of an airfoil in a cold state (solid lines) and under operational heating (dashed lines), also showing a camber of the airfoil in each state. 
         FIG. 5  is a sectional view as in  FIG. 2 , showing a relocated stress area. 
         FIG. 6  is a sectional view of a turbine airfoil showing additional embodiments of aspects of the invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The invention reduces and relocates stress on a 4-wall turbine airfoil by controlling the thermal expansion mismatch between the relatively hotter outer walls and the relatively cooler inner walls to reduce low cycle fatigue (LCF) in the airfoil. 
       FIG. 1  shows a known construction of a 4-wall airfoil  20 A. The purpose of a 4-wall airfoil is to provide near-wall cooling, in which the cooling air flows in channels  31 ,  33  adjacent to the outer walls  26 ,  32  of the airfoil. The cooling channels  31 ,  33  are formed between the double walls  26 ,  28  and  32 ,  34 . Near-wall cooling is advantageous because the cooling air is in close proximity of the hot outer surfaces of the airfoil, and the resulting heat transfer coefficients are high due to the high flow velocity achieved by restricting the flow through narrow channels. 
     The airfoil  20 A of  FIG. 1  has a leading edge  22 , a trailing edge  24 , a pressure side outer wall  26 , a pressure side inner wall  28 , pressure side ribs  30 , pressure side near-wall cooling channels  31 , a suction side outer wall  32 , a suction side inner wall  34 , suction side ribs  36 , suction side near-wall cooling channels  33 , a central forward plenum  37 , a central aft plenum  40 , a rib or septum  42  that separates the central plenums, a leading edge cooling channel  44 , and one or more trailing edge cooling channels  46 . Such designs experience low cycle fatigue especially in the circled area  47 . This is because the suction side outer wall  32  thermally expands more than the cooler suction side inner wall  34 . This differential expansion tends to increase the camber of the airfoil. However, the pressure side outer wall  26  also thermally expands more than the cooler pressure side inner wall  28 . This tends to decrease the airfoil camber, which opposes the forces created by the differential expansion of the suction side walls  32 ,  34 . As a result, the suction side outer wall  32  will tend to bow outward at its apex around area  47 , and thus tries to pull away from the connecting ribs  36 , creating cyclic stress in that area. 
     Many different 4-wall airfoil constructions have been evaluated in the past. One hurdle has been manufacturability. However, with advances in metal investment casting and ceramic core processing, this limitation can be overcome. Another problem has been differential thermal growth stress between the hot outer walls  26 ,  32  and cooler inner walls  28 ,  34 . Previous 4-wall airfoils as in  FIG. 1  often use relatively thinner outer walls  26 ,  32  rigidly attached to relatively thicker inner walls  28 ,  32  by ribs  30 ,  36  or pedestals. However, a thin outer wall  26 ,  32  loses the fight of differential thermal expansion against a thicker inner wall  28 ,  34 , thus creating the type of LCF described above. 
     Attempts have been made to solve this by either: 1) overcooling the outer wall, or 2) using better wall materials and fabrication technology such as advanced single-crystal casting. These solutions improve the airfoil life by changing the fabrication and additional cooling, but they do not address the design geometry. In contrast, the present invention reduces thermal stress via an airfoil sectional geometry combined with a particular cooling flow pattern, which together control macro deflections in the airfoil due to thermal expansion in a way not previous known in the art. 
       FIG. 2  shows an airfoil section including aspects of the invention. The pressure side inner wall  28 B may be at least as thick as the combined thickness of the pressure side outer wall  26  and the suction side inner wall  34 . This allows the pressure side inner wall  28 B to dominate the other two walls  26 ,  34 B in camber deformation, in cooperation with the suction side outer wall  32 . For example, the pressure side inner wall  28 B may be at least twice as thick as the pressure side outer wall  26 , and at least twice as thick as the suction side inner wall  34 B. As another example, the pressure side inner wall  28 B may be at least twice as thick as the pressure side outer wall  26 , and at least three times as thick as the suction side inner wall  34 B.  FIG. 2  is not necessarily drawn to scale, however, it is meant to illustrate an embodiment where the pressure side inner wall  28 B is at least 30% thicker than the combined thicknesses of the pressure side outer wall  26  and the suction side inner wall  34 B to assure its dominance in controlling the camber deflection as the airfoil heats up during operation in a gas turbine. 
     The near-wall channels are designated as forward pressure-side channels  31 F, aft pressure-side channels  31 A, forward suction-side channels  33 F, and aft suction-side channels  33 A. One or more forward passages  38  may transfer cooling air  50 H from the forward central plenum  37  to the leading edge cooling channel  44 . Film-cooling holes  39  may be provided anywhere on the exterior surface of the airfoil  20 B, including ones such as shown passing from the leading edge cooling channel  44  to provide film cooling flows  51  and coolant exhaust. One or more aft coolant passages  41  may communicate from the central aft plenum  40  through the trailing edge  24  as shown. 
       FIG. 3  illustrates a two-pass radial 4-wall cooling scheme according to aspects of the invention. A cooling fluid such as air in a relatively cool state  50 C enters the pressure side near-wall cooling channels  31 F,  31 A through one or more ports  55  in the platform  54 . The coolant travels up the channels  31 F,  31 A along the pressure side of the airfoil. The coolant turns around in the blade or vane end  56  opposite the inlet port  55 , then travels down the respective suction side channels  33 F,  33 A. Along the way, the cooling fluid gains heat and is illustrated as relatively warmer  50 W proximate the vane end  56  and heated cooling fluid  50 H as it passes from the suction side near-wall cooling channels  33 F,  33 A into the respective central plenums  37 ,  40  of the airfoil. The forward edge near-wall channels  33 F are dumped into the leading edge plenum  37 , and the trailing edge channels  33 A are dumped into the trailing edge plenum  40 . This forms a forward cooling circuit  31 F- 33 F- 37 - 44  and an aft cooling circuit  31 A- 33 A- 40 - 46 . The aft circuit is shown in  FIG. 3 . The fore and aft cooling circuits may be independent in some embodiments, with no communication between them, providing independent metering. The coolant  50 H in the central plenums  37 ,  40  respectively cools the leading edge  22  and trailing edge  24  via the leading and trailing edge cooling channels  44 ,  46  as shown in  FIG. 2 . The coolant  50 C,  50 W,  50 H heats as it flows within the airfoil  20 A from the pressure side  26  to the suction side  32 . 
     The difference in temperature of the cooling air is used to relieve thermal stress in the airfoil by creating an inverse temperature gradient across the pressure side inner wall  28 B. In prior art designs, this wall is normally hotter toward the pressure side outer wall  26  and colder toward the central cooling plenums  37 ,  40 . However, in the present flow paths the cooling air  50 C is coldest in the pressure side near-wall channels  31 F,  31 A, and is hotter  50 H in the central plenums  37 ,  40 . As a result, the pressure side inner wall  28 B is colder toward the pressure side outer wall  26  and hotter toward the central plenums  37 ,  40 , reversing the normal gradient (i.e. inverse gradient). The resulting differential thermal expansion across this wall causes its curvature to increase. A thermal gradient of only about 20° C. (for example 435 to 455° C.) is enough to control the strain state of the airfoil in one embodiment. 
       FIG. 3  represents either a rotating turbine blade or a stationary vane. Stationary vanes may have a platform  54  at each end of the airfoil not shown. Sometimes a separate cooling flow  50 C is supplied to each of these platforms. In this case, the forward cooling circuit  31 F,  33 F,  37  and the aft cooling circuit  31 A,  33 A,  40  may optionally start at respective inlet ports  55  in opposite platforms. In each circuit the coolant flow still starts on the pressure side of the airfoil, turns around in the end of the airfoil opposite the inlet port, passes to the suction side, then to the central plenums. 
       FIG. 4  shows a comparison of the original cold airfoil shape in solid outline and the deformed hot airfoil shape in dashed outline, with a respective original camber line  60  and deformed camber line  61 . The pressure side outer wall  26  increases its curvature in the hot state due to the temperature inversion in the pressure side inner wall previously described. This allows the suction side outer wall  32  to grow naturally thermally with less stress as it increases its curvature also. 
     The pressure side outer wall  26  also tends to grow and tries to reduce its concavity in the dual-wall geometry. However, the curling of the thicker pressure side inner wall  28 B dominates, increasing the concavity of the pressure side outer wall  26 . The pressure side outer wall  26  and the suction side inner wall  34  oppose curling  70  of the pressure side inner wall  28 B. These opposing walls  26 ,  34  are made thin enough not to negate the curling effect of the pressure side inner wall and to have some compliance. The pressure side inner wall  28 B may be at least as thick as the combined thicknesses of the pressure side outer wall  26  and the suction side inner wall  34 B as previously described. Stress states and predicted thermal growth geometries in various airfoil embodiments of the present invention can be calculated with commonly available design tools. 
     The net effect is that thermal strain is off-loaded from the suction side outer wall  32  onto the pressure side outer wall  26  and the suction side inner wall  34 . This is a net advantage for the following reasons:
         Due to the difference in moment arm, the thermal curling effect relieves more strain on suction side than it adds on the pressure side.   The pressure side inner wall  26  is cooler than the suction side outer wall  32  due to the lower temperature of the cooling air  50 C on that side, so it has better LCF properties.   The suction side outer wall  32  tends to grow away from the airfoil, while the pressure side outer wall  26  tends to grow into the airfoil. This causes tensile stress between the outer wall  32  and ribs  36  on the suction side and compressive stress on the pressure side. Compressive stress is favorable for life. Past problems observed in 4-wall designs were due to cracking on the suction side of the airfoil.       

     In  FIG. 5 , the suction side inner wall  34 B is stretched by both the thermal growth of the suction side outer wall  32  and the thermal curling  70  of the pressure side inner wall  28 B. As a result, this wall  34 B may experience the highest thermal strain, for example in area  72 . Therefore, it is important that this wall have relatively good compliance. This stress is mitigated by the following:
         The suction side inner wall  34 B is relatively cool; therefore it has excellent LCF properties.   The suction side inner wall  34 B may be thin to provide compliance.   For greater compliance features such as undulations may be added to this wall.       

       FIG. 6  illustrates an embodiment  20 C having a suction side inner wall  34 C with a generally sinusoidal undulation between each rib  36  as a compliance mechanism. This may allow the suction side inner wall  34 C to be thicker than otherwise necessary to get the same degree of compliance, and therefore being easier to cast. In view of the mitigation factors above, the illustrated stress area  72  is a more favorable location than stress area  47  of  FIG. 1 . 
       FIG. 6  also illustrates a pressure side outer wall  26 C that is formed separately from the ribs  30 C, and is attached thereto. For example, this wall may be formed by metal spraying onto the ends of the ribs with a fugitive material in the channel areas. The pressure side outer wall  26 C has ends bracketed by abutments  74 ,  76  at the leading and trailing edges of the airfoil. These abutments may converge slightly when the airfoil camber  61  increases. This causes the wall  26 C to bow toward the ribs  30 C, compressing the bonds between the wall  26 C and the ribs  30 C. This wall  26 C may be made of a metal with a lower elastic modulus than that of the ribs  30 C and the pressure side inner wall  28 B for increased compliance. 
     While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. For example, the invention has been described as a gas turbine engine airfoil including thermal strain state control arrangement effective to allow the suction side outer wall to increase its curl during operation of the gas turbine engine so that a region of peak strain in the airfoil during operation of the gas turbine engine is located remote from the suction side outer wall. The airfoil may have a thermal expansion control mechanism causing its camber to increase under differential thermal expansion of the airfoil during operational heating in order to improve its LCF life. Herein, camber means the degree of curvature of a line halfway between the pressure side and the suction side of an airfoil section. In one embodiment, the airfoil sectional geometry and an internal cooling flow pattern cause the airfoil camber to increase by controlling a temperature gradient on an internal wall structure of the airfoil. In the embodiments described above, it was the relatively thicker pressure side inner wall that curled and controlled thermal strain to off-load one of the outer walls, but in other embodiments it may be the suction side inner wall that is sized to control thermal strain and to off-load an outer wall. Other embodiments may utilize a temperature difference between the average metal temperature of the pressure side and suction side of the airfoil. This may be accomplished with a difference in the cooling air temperature between the pressure and suction sides of the airfoil. This could also be accomplished by using thermal barrier coatings having different insulating abilities on opposed sides of the airfoil. Alternatively, active heating of the backside of the strain-controlling wall may be used instead of the passive cooling scheme described above. Alternatively, bi-material may be used to achieve a desired thermal curl, for example by spraying a low or high coefficient of thermal expansion (CTE) alloy on only one side of the strain-controlling wall. 
     Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.