Gas turbine with cooled rotor blades

Abstract: In a gas turbine having a rotor (36) with blades (1) having root portions (3) and airfoil portions (2) with leading edge portions (7), center portions (39) and trailing edge portions (6), and also passageways extending therethrough, and means for supplying cooling air (29, 30) to the airfoil passageways, first radial passageways (11) are formed in said leading edge portions (7) in communication with a plurality of first holes (43) in said leading edge portion (7), second, radial holes (8) are formed in said trailing edge portion (6); and third, radial holes (9, 10) are formed in said center portion (39), and further, second, radial passageways (17) are formed in said root portion (3), for directing a first portion (18) of said cooling air (30) to said first passageway (11); and a plenum (16) is formed in said root portion (3), for distributing a second portion (19) of said cooling air (30) among said second radial holes (8) and said third radial holes (9, 10).

Description: The current invention relates to gas turbines. More specifically, the current invention relates to an arrangement for cooling the rotating blades of a gas turbine. In the turbine section of a gas turbine, the rotor is comprised of a series of disks to which blades are affixed. Hot gas from the combustion section flows over the blades, thereby imparting rotating power to the rotor shaft. In order to provide maximum power output from the gas turbine, it is desirable to operate with gas temperatures as high as possible. However, operation at high gas temperatures requires cooling the blades. This is so because the strength of the material from which the blades are formed decreases as its temperature increases. Typically, blade cooling is accomplished by flowing air, bled from the compressor section, through the blades. Although this cooling air eventually enters the hot gas flowing through the turbine section, little useful work is obtained from the cooling air, since it was not subject to heat up in the combustion section. Thus, to achieve high efficiency, it is crucial that the use of cooling air be kept to a minimum. In the past, the cooling of turbine blades by flowing cooling air through the blades was typically achieved using either of two blade cooling configurations. In the first configuration, a number of radial cooling holes are formed in the blade. These cooling holes span the length of the blade, beginning at the base of the blade root and terminating at the tip of the blade airfoil. Cooling air supplied to the base of the blade root flows through the holes, thus cooling the blade, and discharges into the hot gas flowing over the blade at its tip. Performance of a cooling air scheme can be characterized by two parameters - efficiency and effectiveness. Cooling efficiency reflects the amount of cooling air required to absorb a given amount of heat. High cooling efficiency is achieved by maximizing the quantity of heat each pound of cooling air absorbs. By contrast, cooling effectiveness reflects the total amount of heat absorbed by the cooling air, without the regard to the quantity of the cooling air utilized. The radial hole cooling configuration discussed above is very efficient because the small diameter of the radial holes, together with a high pressure drop across the holes, results in high cooling air velocity through the holes. This high velocity results in high heat transfer coefficients. Thus, each pound of cooling air absorbs a relatively large quantity of heat. Unfortunately, the cooling effectiveness of this configuration is low because the surface area of the radial holes is small. As a result, the radial hole configuration is incapable of providing the optimum cooling in the leading edge portion of the blade, where the gas temperatures and the heat transfer coefficients associated with the hot gas flowing over the blade are highest. Typically, in the second configuration, one or more large serpentine circuits are formed in the blade. Cooling air, supplied to the base of the blade root, enters the circuits and flows radially outward until it reaches the blade tip, whereupon it reverses direction and flows radially inward until it reaches the base of the airfoil, whereupon it changes direction again and flows radially outward, eventually exiting the blade through holes in the trailing edge or tip portions of the airfoil. As a result of the large surface area of the circuit and the large amount of cooling air flowing through the blade, the cooling effectiveness of this configuration is high. Moreover, heat transfer in the leading edge portion of the airfoil is often enhanced by forming one or more radially extending rows of approximately axially oriented holes through the leading edge of the airfoil. These holes connect with one of the serpentine circuits, allowing a portion of the cooling air entering the circuit to exit the blade at its leading edge. One arrangement of such leading edge holes used in the past, referred to as the "shower head" arrangement, involved arranging the holes into groups of three or more holes at each radial location. The middle hole directs the cooling air to the very center of the leading edge and the adjacent holes direct the cooling air to the convex and concave sides of the leading edge, respectively. It has been observed that the discharge of cooling air at the leading edge tends to disrupt the boundary layer in the hot gas flowing over the blade, resulting in an increase in the heat transfer coefficient associated with the hot gas flowing over the blade surface. To minimize this disturbance to the boundary layer, the holes in the leading edge are sometimes inclined with respect to the radial direction. It should be noted, however, that in the serpentine circuit configuration, all of the cooling air enters and flows through the circuits, so that the flow area of the circuits is large, resulting in low velocity flow and low heat transfer coefficients. Although axially oriented ribs have sometimes been incorporated into the serpentine circuits to increase turbulence, and hence the heat transfer coefficient, the cooling efficiency of the serpentine circuit configuration remains relatively low. As a consequence, excessive quantities of cooling air must be utilized to the detriment of the overall gas turbine efficiency. GB-A-2 112 468 shows a blade having a radial passageway formed in the leading edge portion communicating with a plurality of holes in the leading edge and further a serpentine cooling circuit. Thus, it is the principal object of the present invention to devise a scheme which allows the use of the efficient radial hole cooling configuration in most portions of the blade, but which provides a cooling effectiveness comparable to that of the serpentine circuit configuration in the critical leading edge portion of the blade without the large amount of cooling air usage associated with the serpentine configuration. With this object in view the present invention resides in a gas turbine having a rotor, with rotor discs, a plurality of blades affixed to the periphery of said discs, each of said blades having an airfoil portion and a root portion, each of said airfoil portions having a leading edge portion, a center portion and a trailing edge portion, and having passageways extending therethrough, and means for supplying cooling air to said airfoil passageways, characterized in that a first radial passageway is formed in said leading edge portion in communication with a plurality of first holes radially distributed in said leading edge portion, a plurality of second, radial holes are formed in said trailing edge portion; and a plurality of third, radial holes are formed in said center portion and that further a second radial passageway is formed in said root portion, said second radial passageway directing a first portion of said cooling air to said first passageway; and a plenum is formed in said root portion, said plenum distributing a second portion of said cooling air among said second radial holes and said third radial holes. Cooling air is supplied to each blade root and divided into two portions. The first portion flows through a radial passageway in a leading edge portion of the blade airfoil, thereby cooling the leading edge portion. The second portion of cooling air supplied to the blade root flows into a plenum formed in the blade root. The plenum distributes the air to small radial holes extending through the center and trailing edge portions of the blade. The cooling air flows through the radial holes and exits at the tip of the blade. The invention will become more readily apparent from the following description of a preferred embodiment thereof shown, by way of example only, in the accompanying drawings, wherein: Figure 1 is an isometric view, partially cut away, of a gas turbine. Figure 2 shows a portion of the turbine section in the vicinity of the row 1 rotating blades. Figure 3 is a cross-section of the airfoil portion of the blade taken through line III-III of Figure 2. Figure 4 is cross-section of the airfoil portion of the blade taken through line IV-IV of Figure 3. Figure 5 is a cross section of the root portion of the blade, taken through line V-V of Figure 4. There is shown in Figure 1 a gas turbine. The major components of the gas turbine are the inlet section 32, through which air enters the gas turbine; a compressor section 33 in which the entering air is compressed; a combustion section 34, in which the compressed air from the compressor section is heated by burning fuel in combustors 38, thereby producing a hot compressed gas 24; a turbine section 35 in which the hot compressed air from the combustion section is expanded, thereby producing rotating shaft power; and an exhaust section 37, through which the expanded gas is expelled to atmosphere. A centrally disposed rotor 36 extends through the gas turbine. The turbine section 35 of the gas turbine is comprised of alternating rows of stationary vanes and rotating blades. As shown in Figure 2, each rotating blade 1 is affixed to a disk 27. The disk 27 forms a portion of the rotor 36 which extends through the turbine section 35. Each blade has an airfoil portion 2 and a root portion 3. The blades are retained in the disk by sliding each root portion 3 into mating groove 52 in the periphery of the disk 27. As shown in Figure 2, a duct 55 directs hot gas 24 from the combustion section 34, which may be at a temperature in excess of 1100°C (2000°F), over the airfoil portion 2 of each blade, resulting in the vigorous transfer of heat into the blade. Cooling air 29, drawn from the compressor section 33, enters the rotor 36 through holes 31 in an outer shell 28 of the rotor structure. Radial passageways 26 in the disk 27 direct the cooling air into the disk groove 52. The cooling air 30 flows along the groove 52 and enters the blade root 3 at its base 53. As shown in Figure 3, the airfoil portion of the blade has a leading edge 13 and a trailing edge 40. In addition, the body of the airfoil portion can be seen as comprising a leading edge portion 7, which is approximately the upstream one fifth of the airfoil portion, a center portion 39 and a trailing edge portion 6, which is approximately the downstream one third of the airfoil portion. As shown in Figures 4 and 5, the blade root is essentially hollow. A radial rib 44 divides the interior portion of the root into a radial passageway 17 and a plenum 16. At the base 53 of the blade root, the cooling air 30 is divided by rib 44 into two portions 18, 19. Portion 18 enters the passageway 17 through a hole 15 in an orifice plate 14 affixed to the base 53 of the blade root. From hole 15 the cooling air 18 flows radially outward through passageway 17 in the blade root. Passageway 17 directs the cooling air to a radial passageway 11 in the airfoil. A number of holes 43 are arranged in a radially extending row along the leading edge 13 of the airfoil. The holes 43 connect the radial passageway 17 to the hot compressed gas 24 flowing through the turbine section and thereby allow a portion 23 of the cooling air 18 to flow through and cool the leading edge of the airfoil. As previously discussed, the holes 43 are inclined at an acute angle 46 to the radial direction 56 to minimize the harmful disturbance caused by the introduction of the cooling air 23 into the boundary layer of hot gas flowing over the airfoil. It should also be noted that by inclining the holes, their length, and hence their surface area, is increased, thereby increasing heat transfer to the cooling air 23. In the preferred embodiment, the angle 46 is approximately 30°. As previously discussed, the holes in the leading edge of the blade are preferentially arranged in the "shower head" arrangement shown in Figure 3. In this arrangement, there are three radially extending rows of holes - a center row formed by holes 43, a concave side row formed by holes 41 and a convex side row formed by holes 42. The holes in each row are aligned in the circumferential direction so that there are three holes 41, 42, 43, one from each of the radially extending rows, at each radial position 54 along the leading edge 13. Hole 43 is oriented toward the very center of the leading edge, whereas holes 41 and 42 are inclined toward the concave 4 and convex 5 sides of the airfoil, respectively. Of course, more than three holes could be used at each radial position in a similar arrangement. Typically, the heat transfer from the hot gas 24 into the airfoil is greater in the outboard portion 48 of the airfoil than in the inboard portion 49. This occurs because the temperature profile of the hot gas from the combustion section is often skewed so that the temperature of the gas is higher in the outboard portion. Also, the greater relative speed between the airfoil and the hot gas at the outboard portion results in higher heat transfer coefficients. Hence, in the preferred embodiment, although the radially extending rows of cooling holes 41, 42, 43 extend through both the inboard 49 and outboard 48 portions, the radial spacing 50 of the cooling holes 41, 42, 43 is less in the outboard portion 48 than in the inboard portion 49, so that the radial distribution of cooling air matches that of the temperature distribution along the leading edge. The portion of the cooling air which does not exit the blade through holes 41, 42, 43 flows through radial passageway 11 providing additional cooling to the leading edge portion 7 of the airfoil. A number of axially oriented ribs 12 are disposed along the passageway to increase the heat transfer coefficient at the surface of the passageway. The radial passageway 11 terminates at the tip 25 of the airfoil, the tip 25 being the most radially outboard portion of the airfoil. A hole 21 in the outboard end 45 of the passageway allows a portion 47 of the cooling air to flow out of the blade tip 25 to insure that dust particles entrained in the cooling air do not pile up in the passageway and eventually block the holes 41, 42, 43. As can be seen in Figure 4, the cross sectional flow area 22 of radial passageway 11 continuously decreases as it extends in the radially outward direction. This insures that the velocity of the cooling air is maintained as the quantity of cooling air is reduced due to the flow through holes 41, 42, 43. In the preferred embodiment, the flow area of passageway 11 at any cross-section along the leading edge 13 is inversely proportional to the number of holes 41, 42, 43 inboard of the cross-section - that is, the reduction in the cross-sectional area 22 depends on the number of holes 41, 42, 43 passed as the passageway extends radially outward, so that the rate of reduction in cross-sectional area is greatest in the outboard portion 48 of the airfoil where the radial spacing of holes 41, 42, 43 is the smallest. Thus, the velocity of the cooling air, and hence a high heat transfer coefficient, is maintained as the cooling air flows through passageway 11. For example, in the preferred embodiment, in a blade having an airfoil width-- that is, the distance from the leading edge to the trailing edge -- of approximately 9 cm (3.5 in), the cross-sectional flow area 22 at the entrance to passageway 11 is approximately 1.03 cm² (0.16 in²), whereas the cross-sectional flow area at outboard end 45 of the passageway is approximately 0.26 cm² (0.04 in²). Of course, other size passageways could also be utilized depending on the size and desired cooling characteristics of the blade. An orifice plate 14 is affixed to the portion of the base 53 of the blade root in the vicinity of the radial passageway 17. By adjusting the size of the hole 15 in the orifice plate, the quantity of cooling air supplied to the radial passageway can be adjusted. It can be appreciated that, according to the invention, highly effective cooling of the leading edge portion of the airfoil is achieved as a result of the combined effect of (1) the relatively large surface area of the radial passageway 11, (2) the large quantity of holes 41, 42, 43 connecting the passageway to the surface of the leading edge (inclined at an angle to increase surface area and minimize disturbance of the boundary layer, and spaced to provide cooling where it is most needed), (3) the high velocity of the cooling air throughout the passageway as a result of its tapered shape and (4) the turbulence enhancing ribs. As shown in Figures 3 and 4, according to the invention, the center portion 39 and the trailing edge portion 6 of the airfoil are cooled by the second portion 19 of the cooling air supplied to the base of the blade root. Groove 52 in disk 27 directs cooling air 19 along the base 53 of the blade root 3 to opening 51. From opening 51 cooling air 19 enters plenum 16 formed in the blade root. Radial holes 8, 9, 10 extend from the plenum 16 to the tip 25 of the airfoil. Although the invention could be practiced by dispensing with the plenum and extending the radial holes from the base of the blade root to the tip of the airfoil, or by reducing the size of the plenum so that it connected with only the radial holes 9, 10 in the center portion, in the preferred embodiment the plenum serves to distribute the cooling air evenly among the radial holes 8, 9, 10 in both the center and trailing edge portions of the airfoil. Cooling air 19 flows through the radial holes 8, 9, 10, after which the cooling air 20 discharges at the tip 25 into the hot gas 24 flowing over the airfoil. As previously discussed, the diameter of the radial holes 8, 9, 10 is relatively small so that the velocity of the cooling air through holes is high. This results in high heat transfer coefficients and efficient use of cooling air. As shown in Figure 3, a single row of radial holes 8 is formed in the trailing edge portion 6 of the airfoil. The row extends parallel to the surfaces 4, 5 of the airfoil. In the center portion 39, where the airfoil is thicker, two rows of holes 9, 10 are formed. Holes 10 are disposed close to the convex surface 4 of the airfoil and holes 9 are disposed close to the concave surface 5. As in the trailing edge portion, the rows of holes 9, 10 in the center portion extend parallel to the airfoil surfaces. As shown in Figure 3, the diameter of the holes 8 in the trailing edge portion are larger than the diameter of holes 9, 10 in the center portion, since only a single row of holes is utilized in the trailing edge portion. Moreover, according to the invention, the diameter of cooling air holes and their density could be varied throughout the center and trailing edge portions of the airfoil in response to variations in the temperature of the hot gas or heat transfer coefficients over the surfaces of the airfoil. For example, in the preferred embodiment, in a blade having an airfoil width of approximately 9 cm (3.5 in), the diameter of holes 8, 9, 10 is approximately in the 0.12-0.20 cm (0.05-0.08 in) range, thereby ensuring high velocity cooling air flow through the holes. By contrast, the cross-sectional area of passageway 11 is approximately 30-80 times greater than that of holes 8, 9, 10. Of course, holes of other size diameters could also be utilized depending on the size and desired cooling characteristics of the blade. According to the invention, a serpentine cooling circuit supplying large quantities of cooling air to the entire airfoil, as taught by prior art, is not employed. Instead, adequate cooling is achieved throughout the airfoil using a minimum quantity of cooling air by supplying a large flow of cooling air to only the leading edge portion of the airfoil, where such flow is required, and by making efficient use of such flow by maximizing the surface area and heat transfer coefficient associated with the cooling air in the leading edge portion. In the center and trailing edge portions, the use of cooling air is minimized by utilizing a large quantity of small radial holes, thereby achieving high heat transfer coefficients and efficient use of cooling air.

Claims: A gas turbine having a rotor (36), with rotor discs (27), a plurality of blades (1) affixed to the periphery of said discs (27), each of said blades (1) having an airfoil portion (2) and a root portion (3), each of said airfoil portions (2) having a leading edge portion (7), a center portion (39) and a trailing edge portion (6), and having passageways extending therethrough, and means for supplying cooling air (29, 30) to said airfoil passageways, a first radial passageway (11) is formed in said leading edge portion (7) in communication with a plurality of first holes (43) radially distributed in said leading edge portion (7), characterized in that a plurality of second, radial holes (8) are formed in said trailing edge portion (6); and a plurality of third, radial holes (9, 10) are formed in said center portion (39) and that further a second radial passageway (17) is formed in said root portion (3), said second radial passageway (17) directing a first portion (18) of said cooling air (30) to said first passageway (11); and a plenum (16) is formed in said root portion (3), said plenum (16) distributing a second portion (19) of said cooling air (30) among said second radial holes (8) and said third radial holes (9, 10). A gas turbine according to claim 2, characterized in that said first radial passageway (11) has a cross-sectional area (22) which decreases as said first radial passageway extends radially outward. A gas turbine according to claim 2, characterized in that said first radial passageway (11) has an end (45), with a fourth radial hole (21) extending from the end (45) of said first radial passageway (11) through the tip portion (25) of said blade. A gas turbine according to claim 3, characterized in that each of said second holes is inclined at an acute angle to the radial direction. A gas turbine according to any of claims 1 to 4, characterized in that an orifice (18) is formed at said second end of each of said second radial passageways (17). A gas turbine according to any of claims 1 to 5, characterized in that a plurality of axially oriented ribs (12) are formed in each of said first radial passageways (11). A gas turbine according to any of claims 1 to 6, characterized in that the cross-sectional area of each of said first radial passageways (11) is 30-80 times greater than the cross-sectional area of each of said first holes (43). A gas turbine according to claim 7, characterized in that the diameter of each of said first holes is in the 0.12-0.20 cm (0.05-0.08 in.) range.

Inventors: NORTH WILLIAM EDWARD, PISZ FRANK ANDREW, NORTH, WILLIAM EDWARD, PISZ, FRANK ANDREW
Applicants: WESTINGHOUSE ELECTRIC CORP, WESTINGHOUSE ELECTRIC CORPORATION
Application year: 1991
Main classification: F01D5
IPC codes: F01D5, F02C7