ROTOR BLADE FORM FOR THE FIRST STAGE OF A COMBUSTION TURBINE

Abstract: A rotor blade for a land-based combustion turbine includes an airfoil portion, a root portion, and a platform. The airfoil portion comprises a unique and specific geometric and dimensional range shape structured to minimize aerody­ namic losses and maximize turbine efficiency.

Description: IMPROVED FIRST-STAGE ROTOR BLADE FOR A COMBUSTION TURBINE The present invention relates generally to combustion turbine rotor blades and, more particularly, to a cooled, first-stage rotor blade for a land-based cottibtis-- tion turbine. Rotor blades and stator vanes are key elements in the power generation cycle of a combustion turbine. The airfoil shape, which is unique for each row of blades or vanes, defines the shape of the airflow passage through that row and thereby governs the velocity of the expanding gases and distributes the work among the turbine stages. Although turbine airfoils typically share the same general shape, that is, each typically has a leading edge, a trailing edge, a concave surface, and a convey surface, the airfoil shape common to a particular row of blades or vanes differs from the airfoil shape for every other row within that turbine. Likewise, no two turbines of different designs share airfoils of the same shape The structural differences in airfoil shape, which Say appear minute to the untrained observer, result in signi- ficant variations in aerodynamic characteristics, stress patterns, operating temperature, and natural frequency of the airfoil. These variations, in turn, determine the operating life of the airfoil within the boundary condi- tions (turbine inlet temperature, compressor pressure ratio, and engine speed), which are generally determined prior to airfoil development. Hence, in the environment of a particular set of overall turbine design parameters, a particular combination of structural improvements to the airfoil shape may reduce aerodynamic losses so as to improve turbine efficiency and performance (increase power output) without need for increasing the turbine inlet temperature. Shape improvements in the airfoil structure may also increase the cooling efficiency of the airfoil and thereby improve the operating life and reliability of the airfoil. The invention in its broad form resides in an airfoil blade for a cooled first stage rotor of a landbased combustion turbine capable of driving an electric generator to generate electric power in the range of one hundred megawatts, said blade comprising: airfoil portion having a leading edge and a trailing edge with concave and convex surfaces therebetween providing a smooth airfoil surface, a root portion for affixing said blade to a turbine rotor, a platform portion joining said airfoil portion to said root portion, said airfoil portion having: a gradually decreasing cross-sectional area with increasing radius; a maximum cross-sectional thickness which between said convex and concave sides decreases with increasing radius from approximately 1.0 inch to approximately 0.78 inch; a cross-sectional width which decreases with increasing radius from approximately 3.853 inches to approximately 3.240 inches; a chord which decreases with increasing radius from approximately 4.051 to 3.965; said leading edge having a radius which decreases with increasing radius from approximately 0.141 inch to approximately 0.102 inch. In a preferred embodiment described herein, a combustion turbine rotor blade is described having improved turbine performance through greater aerodynamic efficiency and includes a root portion, a platform, and an airfoil portion. The airfoil portion comprises a gradually decreasing cross-sectional area with increasing radius, a maximum thickness which decreases with increas ing radius, a width which decreases with increasing razz us, a leading edge radius which decreases with increasing radius-and a trailing edge radius which increases slightly with increasing radius. The blade further comprise$ a plurality of cooling channels extending radially from the root portion through a tip of the airfoil portion. A more detailed understanding of the invention can be had from the following description of a preferred embodiment given by way of example and to be studied in conjunction with the accompanying drawing wherein: Figure 1 shows a plot of velocity distribution curves for a first stage turbine blade made in accordance with a preferred embodiment of the invention; Fig. 2 shows the profile of a typical cross section of the airfoil portion of the first stage blade structured according to a. preferred embodiment af the invention except that coolant channels are not shown; Fig. 3 shows trailing edge of the first stanza blade in an elevational view of one end of the blade; ; Fig. 4 shows a side elevational view of the blade; Fig. 5 illustrates the blade coolant channel locations in another side elevational view of the blade; Fig. 6 illustrates the twist of the blade tit a plan view from the top of the blade; -and Fig. 7 shows the blade tip in an enlarged plan view of the top of the blade airfoil portion. Development of the turbine section for a new commercial, power generation combustion turbine may require several years of painstaking effort to complete. In fact, the development of a single blade or vane structure for a turbine section (which might include four eadh of blades and vanes) may itself require several years of effort. The development effort comprises three basc divisions of labor: airfoil profile, airfoil cooling, and mechanical aspects. The development process begins with the specification of the airfoil profile, which is actually the definition of the-airflow channel between two airfoils of a given row. The developer is provided certain boundary conditions (turbine design parameters which must be satisfied), such as turbine inlet temperature, compressor pressure ratio, and engine speed. In the case of the present invention, the goal of the turbine design process was to achieve an improvement in turbine performance (increased power to fuel ratio) while accepting some decrease in turbine inlet temperature (2005QF to 19850E). Given that the turbine section will comprise four stages, that is, four pairs of vane-blade row combinations, the profile developer must establish the radial equilibrium pattern and the work distribution among the four stages. This is called the blade path calculation. The radial equilibrium pattern establishes the swirling flow path, as it varies axially and radially, of the expanding gases flowing through the gradually increasing area of the annular turbine section. The work distribution among the stages establishes the amount of work done by the expanding gases at each stage. The two are interrelated, as the radial equilibrium pattern is, in a sense, determined -by the angle of deflection and exit velocity of gases at each stage, this also governs the work done at each stage. The profile developer has complete freedom to vary the choices of blade path parameters in an effort to maximize power output within the boundary conditions. The complex mathematical calculations required to accomplish this task are solved with the aid of a computer, which adjusts the solutions according to loss models designed to mirror the effects of losses in actual turbine operation. The final result of the blade path calculation is a set of radially varying velocity vectors for both airfoil rows of each of the four stages. The velocity vectors establish the desired magnitude and direction of gas flow entering and leaving each row of airfoils. Once the blade path calculation is complete aed acceptable, the profile developer uses the velocity vectors to derive an airfoil profile. This portion of the blade development process is performed for each of frc- two to five plane "slices" perpendicular to a radial line for each of the eight rows of airfoils. As is described in greater detail below, these plane slices are later combined to present a three dimensional picture of each airfoil. Like that of varying parameters within the blade path calculation, this portion of the process requires engineering judgment derived from a thorough understanding and creative application of the underlying physical principles and extensive "hands-on" experience in the design -process. Given the change in velocity vectors from the upstream side to the downstream side of a row, the deve loper determines the approximate profile of the desired airfoil. Certain descriptive dimensions defining this hypothetical profile are entered in a computer program which generates a velocity distribution for that airfoil. A velocity distribution for the airfoil of the present invention is shown in Figure 1. It comprises a plot 10 of the gas velocity (or the ratio of gas velocity V to gas velocity VTE at the trailing edge) as a function of airfoil width. The airfoil width is, by definition, the maximum dimension of the airfoil along a line parallel to the rotor axis. The velocity distribution plot includes two curves: the velocity along the convex surface 12 and the velocity along the concave surface 14. The velocity distribution generated from the trial profile is next analyzed in light of principles gleaned from experience to determine its acceptability, Certain features of the curve, for example the decrease in velocity from the point of relative maximum velocity 16 to the exit velocity 18, may reveal flow tendencies which may give rise to problems. Drawing on creativity, experience, and skill, the developer may make alterations in the hypothetical profile to alleviate perceived shortcomings in the velocity distribution. In the example presented above, a relative maximum velocity 16 which exceeds the exit velocity 18 by more than approximately 30% may indicate a tendency of the flow to separate from the airfoil surface and generate additional losses. Where this is deemed a substantial enough problem, the profile developer may alter the profile to reduce the change in velocity. When an acceptable profile has been generated for each plane slice of the airfoil, the slices are combined and interpolated by a computer to produce a single three-dimensional airfoil. Conical slices of the threedimensional airfoil corresponding to stream surfaces are subjected to velocity distribution analysis to ensure that the airfoil conforms to the previously determined radial patterns. When the airfoil profile is determined to be complete, it is passed to the cooling developer for the next stage of the development process. The second stage of the development process involves the airfoil cooling. The basic procedure begins with the velocity distribution curves (Fig. 1) for the airfoil. A computer is used to convert the velocity distributions generated in the course of the profile development process into heat transfer curves for the outside wall for both the concave and the convex surfaces of the airfoil. The heat transfer curve is a plot of the heat transfer coefficient for the wall as a function of distance along the concave and convex surfaces of the airfoil. The outside wall heat transfer curve is-used to produce a corresponding inside wall heat transfer curve. The goal of the blade cooling developer is to achieve inner airfoil cooling which effects the previously calculated inside wall heat transfer curve. The actual development process, like that of deriving the trial airfoil profile, requires the exercise of creativity and skill and engineering judgment gleaned from actual development experience. In the case of the first stage blade, the cooling development process is simplified by the use of a single cooling technique, rather than the several techniques available in the case of vanes, to achieve the required level of airfoil cooling. The cooling technique for blades comprises a plurality of channels extending radially through the interior of the blade from a cooling air supply channel in the root to the blade tip. The cooling developer must arrange the channels within the airfoil to achieve the cooling level indicated by the inside wall heat transfer curve. Should it prove impossible, due to some characteristic of the airfoil profile, to cool adequately the airfoil, it may be necessary to require the profile developer to restructure portions of the airfoil profile so as to accommodate the structure needed for adequate cooling. The final division of responsibility respecting development of the first-stage blade structure and its shape is the mechanical aspect of the process. After receiving the airfoil profile and cooling arrangement, the mechanical developer specifies a platform for the blade and performs several mechanical analyses of the airfoil. The two primary checks, which are performed by computer, are a chordwise stress analysis and a natural frequency check. Should either of these analyses reveal problems, especially the more complex frequency check, it may be necessary to require the cooling developer or profile developer to make structural changes. Certain mechanical criteria must be met so as to guarantee the structural integrity of the airfoil. It is not unusual for an airfoil to be returned several times for rework by the profile or cooling developer. This interactive process is complex and time-consuming. When the airfoil successfully passes the mechanical computer analyses, the mechanical developer prepares a root to adjoin the airfoil and platform of the blade. Thereafter, a rotating disk test is performed to determine the actual resonant frequency of the blade. Again, if the test reveals problems, further structural refinement may be necessary. Variations in the structural shape of the airfoil with increasing radius determine the performance characteristics of the blade. Because the linear velocity of the airfoil increases with increasing radius, the flow angles of the expanding gases vary as well, giving rise to an aerodynamic requirement that the shape of the airfoil change with increasing radius, i.e. twist. Stress -rotational and gas bending forces -- imposes further constraints on the changing shape of the airfoil with radius. The first stage blade of the present invention represents a balanced improvement over previous first stage blades for the intended range of turbine design parameters. The changing shape of the airfoil, as set forth below, minimizes aerodynamic losses while maximizing structural integrity. By reducing losses and improving work distribution among the stages (in particular, reducing the exhaust gas exit velocity), the first-stage blade of the present invention, in cooperation with the other blades and vanes of the turbine (assignee's Model No. W501D5), effects an increase in output power rating from 95 megawatts to 100 megawatts while allowing a slight reduction in turbine inlet termperature (20050F to 1985 oF). In Figures 2-7 there are shown various views of a first stage turbine blade 20 made in accordance with the principles of the invention. Thus, the blade 20 comprises an airfoil portion 21, leading edge 22, a trailing edge 24, a concave surface 26, and a convex surface 28, with gas flow approximating that depicted at 30 (Figure 2). The airfoil portion 21 extends radially outwardly from a platform section 23 which in turn is based on a root extended portion 25. A blade root portion 27 is serrated for support on the rotor disc (not shown). As a preferred example, the radial height or span of the airfoil portion 21 of the blade 20, from platform to tip, is approximately 6.52 inches. The crosssectional area of the blade gradually decreases with increasing radius. Thus, the maximum thickness 32 (Figure 2) of the blade (TMAX) decreases from approximately 1,8 inch near the platform to 0.78 inch near the blade tip. The blade thickness is one characteristic which governs the structural integrity of the airfoil. Generally, the hub portion of the blade needs to be heavier and the blade tip needs to be thinner to provide the blade strength needed for increasing loading with decreasing radius. The width BD (Figure 2) of the blade 20, which is a measure of the extension of the profile along a line parallel to the rotor axis (CL) 36, decreases from approximately 3.853 inches near the platform to approximately 3.240 inches near the tip. The blade chord BA (Figure 2) also, decreases somewhat with radius and twists with increasing radius as observed in the top plan view of Figure 6 to cause the decrease in blade with increasing radius. The surface profile of the blade airfoil portion is defined by a series of X-Y coordinates for the twentytwo points shown in Figure 2 at each of the reference planes shown in Figure 3 in accordance with the following table: AIRFOIL CO-ORDINATE POINTS-BSC Distance SECTION From R1 R2 R3 R4 R5 R6 R7 R8 R9 R10 R11 AA-AA Y-Y -1.823 -1.458 -1.094 -.729 -.365 .000 +.406 +.812 +1.218 +1.624 +2.030 X-X + .306 - .381 - .660 -.798 -.843 -.801 -.641 -.342 + .110 + .731 +1.538 AB-AB Y-Y -1.752 -1.400 -1.048 -.696 -.344 +.008 +.401 +.794 +1.187 +1.580 +1.973 X-X + .192 - .449 - .708 -.825 -.847 -.783 -.601 -.292 + .153 + .746 +1.498 AC-AC Y-Y -1.680 -1.340 -1.001 -.662 -.323 +.016 +.396 +.776 +1.156 +1.535 +1.915 X-X + .075 - .518 - .755 -.852 -.851 -.764 -.561 -.242 + .197 + .762 +1.458 AD-AD Y-Y -1.607 -1.281 - .955 -.628 -.302 +.024 +.391 +.758 +1.124 +1.491 +1.858 X-X - .043 - .586 - .803 -.878 -.854 -.745 -.520 -. 192 + .240 + .777 +1.417 AE-AE Y-Y -1.552 -1.235 - .919 -.603 -.286 +.030 +.387 +.744 +1.100 +1.457 +1.814 X-X - .133 - .639 - .840 -.899 -.858 -.731 -.490 -.153 + .274 + .789 +1.386 AF-AF Y-Y -1.535 -1.221 - .908 -.595 -.281 +.032 +.386 +.739 +1.093 +1.446 +1.800 X-X - .160 - .655 - .851 -.905 -.858 -.726 -.480 -.141 + .284 + .793 +1.377 Distance SECTION From R12 R13 R14 R15 R16 R17 R18 R19 R20 R21 R22 AA-AA Y-Y +2.030 +1.624 +1.218 +.812 +.406 .000 -.365 -.729 -1.094 -1.458 -1.823 X-X +1.782 +1.258 +.842 +.536 +.328 +.203 +.154 +.161 + .225 + .365 + .620 AB-AB Y-Y +1.973 +1.580 +1.187 +.794 +.401 +.008 -.344 -.696 -1.048 -1.400 -1.752 X-X +1.737 +1.243 + .841 +.534 +.317 +.178 +.112 +.099 + .039 + .253 + .476 AC-AC Y-Y +1.915 +1.535 +1.156 +.776 +.396 +.016 -.323 -.662 -1.001 -1.340 -1.680 X-X +1.692 +1.228 + .840 +.533 +.306 +.152 +.069 +.035 + .051 + .139 + .329 AD-AD Y-Y +1.858 +1.491 +1.124 +.758 +.391 +.024 -.302 -.628 - .955 -1.281 -1.607 X-X +1.647 +1.214 + .839 +.531 +.294 +.126 +.026 -.028 - .037 + .026 + .183 AE-AE Y-Y +1.814 +1.457 +1.100 +.744 +.387 +.030 -.286 -.603 - .919 -1.235 -1.552 X-X +1.612 +1.202 + .838 +.529 +.286 +.107 -.007 -.077 - .104 - .061 + .071 AF-AF Y-Y +1.800 +1.446 +1.093 +.739 +.386 +.032 -.281 -.595 - .908 -1.221 -1.535 X-X +1.601 +1.199 + .838 +.529 +.284 +.101 -.017 -.092 - .125 - .088 + .036 DIMENSIONS FOR TYPICAL BLADE SECTIONS BA BD SECTION RL RT T-THK Z-DIH CHORD BB WIDTH 6 7 (REF) (REF) AA-AA .141 .074 1.000 .155 4.051 1.628 3.853 AB-AB .134 .074 .956 .155 3.956 1.589 3.725 AC-AC .126 .074 .910 .155 3.862 1.550 3.595 AD-AD .118 .075 .864 .156 3.772 1.511 3.465 AE-AE .111 .075 .828 .156 3.707 1.481 3.366 AF-AF .109 .075 .816 .157 3.687 1.472 3.335 The radius of curvature of the leading edge (RL) 38 is a measure of the sharpness of the leading edge transition from the concave 26 to the convex surface 28. The radius decreases (indicating a gradual increase in sharpness) from approximately 0.141 inch near the platform to 0.102 inch near the tip. The trailing edge radius (RT) 40 increases very slightly from approximately 0.074 inch to approximately 0.075 inch near the tip. The thickness of the trailing edge of the blade represents a compromise between competing aerodynamic, cooling, and mechanical considerations. The ideal trailing edge from an aerodynamic point of view would narrow to a point, but cooling considerations require a cooling channel (of a minimum 0.046 inch diameter) along the trailing edge. To provide structural integrity to a trailing edge supporting a cooling channel, as an example, the present blade is provided with a trailing edge thickness (measured 0.10 inch'from the trailing edge tip) which increases slightly from approximately 0.155 inch near the platform to approximately 0.157 inch near the tip. Figure 7 shows a plan view of the blade tip where a wall 31 surrounds a tip cavity 33 wherein the span ends of blade cooling channels are disposed. The channels are generally linear along the blade span and substantially circular in cross-section, are spaced as shown, and have the following approximate diameters and locations at the reference planes indicated: : HOLE LOCATIONS Distance SECTION From CA CB CD CE CF CH CJ CK CL CM AF-AF Y-Y -1.418 -1.239 -1.012 -.709 -.415 -.144 +.108 +.346 +.572 +.735 X-X - .137 - .348 - .542 -.633 -.600 -.530 -.348 -.179 +.026 +.194 AA-AA Y-Y -1.693 --- --- --- --- --- --- --- --- -- X-X + .351 --- --- --- --- --- --- --- --- -- Distance SECTION From CN CP CQ CR CS CT AF-AF Y-Y +.947 +1.116 +1.317 +1.486 +1.617 +1.725 X-X +.425 + .606 + .849 +1.066 +1.251 +1.395 AA-AA Y-Y --- --- --- --- --- -- X-X --- --- --- --- --- --- Thus, the surface profile, cooling, and mechanical characteristics of the present airfoil combine to define a unique blade structure, which operates in cooperation with other turbine blades and vanes to produce greater turbine efficiency within the predetermined set of turbine boundary conditions. The resultant increase in power-to-fuel ratio is achieved without resort to an increase in turbine inlet temperature.

Claims: CLAIMS: 1. An airfoil blade for a cooled first stage rotor of a land-based combustion turbine capable of driving an electric generator to generate electric power in the range of one hundred megawatts, said blade comprising: airfoil portion having a leading edge and a trailing edge with concave and convex surfaces therebetween providing a smooth airfoil surface, a root portion for affixing said blade to a turbine rotor, a platform portion joining said airfoil portion to said root portion, said airfoil portion having: : a gradually decreasing cross-sectional area with increasing radius; a maximum cross-sectional thickness which between said convex and concave sides decreases with increasing radius from approximately 1.0 inch to approximately 0.78 inch; a cross-sectional width which decreases with increasing radius from approximately 3.853 inches to approximately 3.240 inches; a chord which decreases with increasing radius from approximately 4.051 to 3.965; said leading edge having a radius which decreases with increasing radius from approximately 0.141 inch to approximately 0.102 inch. 2. A blade according to claim 1 wherein said airfoil portion has a tip and sixteen cooling channels of substantially circular cross-section extending generally radially from said root portion through said airfoil portion to said tip. 3. A blade according to claim 1, wherein said trailing edge radius increases with increasing blade radius from approximately 0.074 inches to 0.075 inches. 4. A blade according to claim 2, wherein said coolant channels are generally linear and the location of said coolant channels are defined by X-Y coordinates in preselected cross reference planes as defined by the holes location table in the specification. 5. A blade according to claim I wherein the airfoil surface along said convex and concave sides is defined by a series of X-Y coordinates in each of a plurality of preselected cross reference planes spaced along the span of said airfoil portion as defined by the airfoil coordinate points table in the specification. 6. A blade according to claim 4 wherein the airfoil surface along said convex and concave sides is defined by a series of X-Y coordinates in each of a plurality of preselected cross reference planes spaced along the span of said airfoil portion as defined by the airfoil coordinate points table in the specification.

Inventors: BROWN, DAVID LEE, FENTRESS, WILLIAM KILBY, TAYLOR, MELVIN DALLAS
Applicants: WESTINGHOUSE ELECTRIC CORPORATION
Application year: 1983
Main classification: F01D5
IPC codes: F01D5