WO1996012874A1

WO1996012874A1 - Gas turbine blade with enhanced cooling

Info

Publication number: WO1996012874A1
Application number: PCT/US1995/012649
Authority: WO
Inventors: Paul H. Davis; William E. North
Original assignee: Westinghouse Electric Corporation
Priority date: 1994-10-24
Filing date: 1995-10-02
Publication date: 1996-05-02
Also published as: JPH09507550A; IL115665A0

Abstract

A gas turbine rotating blade having an airfoil (21) portion containing four radially extending cooling air passages (51-54). A first opening (58) in the blade root (20) allows cooling air to enter a first passage (51) which is located adjacent the leading edge of the blade. In the first passage (51) the cooling air flows radially outward and then exits at the blade tip (45). Fins (60) project into the first passage (51) and are oriented so as to direct the cooling air against the leading edge (25) of the blade as it flows outward. The second (52), third (53) and fourth (54) passages are arranged sequentially in a serpentine fashion. A second opening (59) in the blade root allows additional cooling air to enter the second passage (52), from which it flows to the third and fourth passages by making approximately 180 turns. Fins (61, 62, 63) project into the passages and are angled so as to inhibit the tendency of the cooling air to experience flow separation as it makes the turns.

Description

GAS TURBINE BLADE WITH ENHANCED COOLING

BACKGROUND OF THE INVENTION The present invention relates to a rotating blade in a gas turbine. More specifically, the present invention relates to a gas turbine rotating blade having a cooling air flow path with fins oriented to provide enhanced cooling effectiveness.

The turbine section of a gas turbine includes a rotor that 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 airfoils of the blades. In the past, the cooling of turbine blades by flowing cooling air through the blade airfoil 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.

The radial hole cooling configuration discussed above has certain advantages 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.

Typically, in the second configuration, a number of large serpentine passages, typically three, are formed in the blade. Cooling air, supplied to the base of the blade root, enters the first passage and flows radially outward until it reaches the blade tip, whereupon it reverses direction and flows radially inward through the second passage until it reaches the base of the airfoil, whereupon it changes direction again and flows radially outward through the third passage, eventually exiting the blade through holes in the trailing edge or tip portions of the airfoil.

Various methods have been tried to increase the effectiveness of the cooling air flowing through the serpentine passages. One such approach involves the use of fins extending from the walls that form the passages. The use of both fins that extend perpendicular to the direction of flow and fins that are angled to the direction of flow have been tried. However, the ability of such schemes to adequately cool the blade airfoils is impaired in gas turbines in which the airfoils have a large cross-sectional area since this reduces the velocity, and hence the heat transfer coefficient, of the cooling air flowing through the passages. The cooling ability of such schemes is also impaired when used in conjunction with higher pressure ratio compressors, since the cooling air bled from such compressors is at a relatively high temperature. One potential solution to this problem is to dramatically increase the cooling air supplied to the airfoil, thereby increasing the flow rate of the cooling air flowing through the passages. However, such a large increase in cooling air flow is undesirable. Although such 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.

It is therefore desirable to provide a cooling scheme that significantly increases the cooling effectiveness of the cooling air flowing through the airfoil of a rotating blade in a gas turbine.

SUMMARY OF THE INVENTION Accordingly, it is the general object of the current invention to provide a cooling scheme that significantly increases the cooling effectiveness of the cooling fluid flowing through the airfoil of a rotating blade in a gas turbine.

Briefly, this object, as well as other objects of the current invention, is accomplished in a rotating blade for a turbine comprising (i) leading and trailing edges, (ii) a root portion for attaching the blade to a rotor, and a tip portion remote from the root portion, the root portion having means for receiving a first flow of cooling fluid, (iii) a first passage disposed adjacent one of the edges, the first passage having means for directing the first flow of cooling fluid to flow in a direction away from the root portion toward the tip portion, and (iv) a plurality of first fins extending into the first passage, the first fins angled so as to extend toward the tip portion of the blade as they extend toward the one of the edges to which said first passage is adjacent. Preferably, the one of the edges to which the first passage is adjacent is the leading edge. In one embodiment of the invention, the root portion has means for receiving a second flow of cooling fluid. In this embodiment, the blade further comprises (i) second and third passages having means for directing the second flow of cooling fluid in first and second directions, respectively, the second and third passages being in sequential flow communication, whereby the second flow of cooling fluid flows from the second passage to the third passage, (ii) first turning means for turning the second flow of cooling fluid from the first direction to the second direction as it flows from the second passage to the third passage, and (iii) means for retarding flow separation in the second flow of cooling fluid as it is turned by the first turning means. The invention also encompasses a turbomachine comprising (i) a compressor for producing compressed fluid, (ii) a combustor for heating a first portion of the compressed fluid, thereby producing a hot compressed gas, and (iii) a turbine for expanding the hot compressed gas. The turbine includes a rotating blade having a plurality of successive cooling fluid passages arranged in a serpentine configuration, whereby each of the passages except a first one of the passages is preceded by an adjacent passage. One of the passages has means for receiving a second portion of the compressed fluid from the compressor and each of the passages has means for directing the second portion of the compressed fluid to flow in a direction opposite to that of the preceding passage. A plurality of fins projects into each of the passages. The fins are angled so that the direction in which the fins extend along their respective passages is reversed with respect to the preceding passage.

BRIEF DESCRIPTION OF THE DRAWINGS Figure 1 is a longitudinal cross-section, partially schematic, through a portion of the gas turbine having a row 2 turbine blade made in accordance with the current invention. Figure 2 is a longitudinal cross-section through the row 2 turbine blade shown in Figure 1.

Figure 3 is a transverse cross-section taken through line III-III shown in Figure 2. Figure 4 is a cross-section taken through line

IV-IV shown in Figure 2.

Figure 5 is detailed view of the portion of the blade shown in Figure 2 in the vicinity of the airfoil tip.

Figure 6 is detailed view of the portion of the blade shown in Figure 2 in the vicinity of the airfoil base.

DESCRIPTION OF THE PREFERRED EMBODIMENT

Referring to the drawings, there is shown in Figure 1 a longitudinal cross-section through a portion of a gas turbine. The major components of the gas turbine are a compressor section 1, a combustion section 2, and a turbine section 3. As can be seen, a rotor 4 is centrally disposed and extends through the three sections. The compressor section 1 is comprised of cylinders 7 and 8 that enclose alternating rows of stationary vanes 12 and rotating blades 13. The stationary vanes 12 are affixed to the cylinder 8 and the rotating blades 13 are affixed to discs attached to the rotor 4.

The combustion section 2 is comprised of an approximately cylindrical shell 9 that forms a chamber 14, together with the aft end of the cylinder 8 and a housing 22 that encircles a portion of the rotor 4. A plurality of combustors 15 and ducts 16 are contained within the chamber 14. The ducts 16 connect the combustors 15 to the turbine section 3. Fuel 35, which may be in liquid or gaseous form -- such as distillate oil or natural gas -- enters each combustor 15 through a fuel nozzle 34 and is burned therein so as to form a hot compressed gas 30.

The turbine section 3 is comprised of an outer cylinder 10 that encloses an inner cylinder 11. The inner cylinder 11 encloses rows of stationary vanes and rows of rotating blades. The stationary vanes are affixed to the inner cylinder 11 and the rotating blades are affixed to discs that form a portion of the turbine section of the rotor .

In operation, the compressor section 1 inducts ambient air and compresses it. The compressed air 5 from the compressor section 1 enters the chamber 14 and is then distributed to each of the combustors 15. In the combustors 15, the fuel 35 is mixed with the compressed air and burned, thereby forming the hot compressed gas 30. The hot compressed gas 30 flows through the ducts 16 and then through the rows of stationary vanes and rotating blades in the turbine section 3, wherein the gas expands and generates power that drives the rotor 4. The expanded gas 31 is then exhausted from the turbine 3. A portion 19 of the compressed air 5 from the compressor 1 is extracted from the chamber 14 by means of a pipe 39 connected to the shell 9. Consequently, the compressed air 19 bypasses the combustors 15 and forms cooling air for the rotor 4. If desired, the cooling air 19 may be cooled by an external cooler 36. From the cooler 36, the cooled cooling air 32 is then directed to the turbine section 3 by means of a pipe 41. The pipe 41 directs the cooling air 32 to openings 37 formed in the housing 22, thereby allowing it to enter a cooling air manifold 24 that encircles the rotor 4. The cooling air 32 exits the manifold 24 through passages 38 and then travels through a series of passages within the rotor 4 to the various rows of rotating blades. The current invention will be described in detail with reference to the cooling of the second row of rotating blades 18, one of which is shown in Figures 2-6.

As shown in Figures 2 and 3, each row two turbine blade 18 is comprised of an airfoil portion 21 and a root portion 20. The airfoil portion 21 has a base portion 46 adjacent the root 20 and a tip portion 45. Thus, the tip portion 45 of the airfoil 21 forms one end of the blade 18 and the root portion 20 forms the other end of the blade. As shown in Figure 3, the airfoil portion 21 of the blade 18 is formed by a generally concave shaped wall 42, which forms the pressure surface of the airfoil, and a generally convex wall 43, which forms the suction surface of the airfoil. At their upstream and downstream ends, the walls 42 and 43 form the leading and trailing edges 25 and 26, respectively, of the airfoil 21. The blade root 20 has a plurality of serrations (not shown) that engage with grooves formed in the rotor 4 so as to secure the blades 18 to the rotor.

The airfoil 21 is substantially hollow. Radially extending walls 55-57 extend between the walls 42 and 43 and separate the interior of the airfoil 21 into four radially extending cooling air passages 51-54. As shown in Figure 2, a first opening 58 in the root 20 allows a first portion 80 of the cooling air 32 to enter the first passage 51. The first passage 51 is disposed adjacent the leading edge 25 and extends from the root 20 to the airfoil tip 45. After flowing radially outward through the first passage 51, the cooling air 80 exits the blade via a radially extending hole 40 formed in the tip portion 45 of the airfoil 21.

A second opening 59 in the root 20 allows a second portion 81 of the cooling air 32 to enter the second passage 52, which also extends from the root 20 to the airfoil tip 45. The wall 55 that divides the first and second passages 51 and 52, respectively, extends from the root 20 to the airfoil tip 45. However, the walls 56 and 57 that divide the second, third and fourth passages 52-54, respectively, do not extend all the way from the root 20 to the tip 45. Instead, wall 56 stops short of the tip 45 so as to form a U-shaped connecting passage 64 that allows the cooling air 81 to flow from the second passage 52 to the third passage 53. Similarly, wall 57 stops short of the root so as to form a U-shaped connecting passage 65 that allows the cooling air 81 to flow from the third passage 53 to the fourth passage 54. Consequently, the second, third and fourth passages 52-54 are arranged in a serpentine fashion so that the cooling air 81 flows sequentially from passage 52 to passage 53 to and finally to passage 54, which is adjacent the trailing edge 26. The connecting passages 64 and 65 cause the cooling air 81 to turn approximately 180° before it enters the adjacent passage. From the trailing edge passage 54, the cooling air 81 is divided into a plurality of small streams 83 that exit the blade 18 through a plurality of axially extending passages 44 that are drilled in the trailing edge 26 of the airfoil 21. Upon exiting the blade 18, the streams of cooling air 83 mix with the hot gas 30 flowing through the turbine section 3.

According to the current invention, a plurality of fins 60-63 -- sometimes referred to as turbulating ribs -- project from the walls 42 and 43 into the passages 51- 54. As shown in Figure 2, the fins 60-63 are preferably distributed along the entire height of the passages 51-54 in the airfoil portion 21 of the blade 18. Moreover, as shown in Figure 3, the fins 60-63 preferably extend along substantially the entire axial length of the passages.51- 54. Figure 4 shows the fins 61 in the second passage 52 but is typical of the arrangement of the fins in each of the passages. As shown in Figure 4, the fins 61 project transversely into the second passage 52 from opposing walls and, preferably, have a height equal to approximately 10% of the width of the passage. The fins 61 are staggered so that the fins projecting from the wall 42 are disposed between the fins projecting from the wall 43. The fins 60- 63 serve to increase the turbulence in the cooling air 80 and 81 flowing through the passages 51-54, thereby increasing its cooling effectiveness.

According to an important aspect of the current invention, the fins 60-64 are angled with respect to the direction of flow of the cooling air 80 and 81 through the passages 51-54 -- which is essentially in the radial direction. Thus, as shown in Figure 5, the fins form an acute angle A with respect to the radial direction. In the preferred embodiment, the angle A with respect to the radially outward direction is in the range of approximately 45-60°, most preferably 45°. This is so whether the fins are angled radially outward as they extend upstream to the direction of the flow of hot compressed gas 30, as in the first, second and fourth passages 51, 52, and 54, or whether they are angled radially inward as they extend upstream, as in the third passage 53. In the first passage 51, the cooling air 80 flows radially outward from the airfoil base 46 to the tip 45. Therefore, according to another important aspect of the current invention, the fins 60 in the first passage 51 are angled so that they extend radially outward -- that is, toward the airfoil tip 45 -- as they extend in the upstream direction toward the leading edge 25, as shown in Figures 2 and 5. As a result, the cooling air 80 is guided so that it flows toward the leading edge 25 as it flows radially outward, as shown best by the arrows indicated by reference numeral 84 in Figure 5. Thus, the fins 60 not only increase the turbulence of the cooling air 80 but also serve to direct it against the leading edge 25, thereby increasing the effectiveness of the cooling of the leading edge. This is important since the hot gas 30 flowing through the turbine section 3 impinges directly on the leading edge 25 so that it is one of the portions of the airfoil 21 most susceptible to over heating.

As shown in Figures 2 and 5, in flowing from the second passage 52 to the third passage 53, the U-shaped connecting passage 64 causes the cooling air 80 to turn 180°, as previously discussed. Such an abrupt change in direction has a tendency to cause flow separation of the cooling air as it flows around the turn. Such flow separation is undesirable since it reduces the flow rate of cooling air through the passages.

Therefore, according to still another important aspect of the current invention, the tendency of the cooling air to experience flow separation is retarded by angling the fins 62 in the third passage 53 so that they extend radially inward -- that is, toward the airfoil base 46 -- as they extend in the upstream direction toward the wall 56 dividing the second and third passages. This causes the cooling air 80 to be guided so that it flows toward the dividing wall 56 as it completes its travel around the turn, as shown best by the arrows indicated by reference numeral 82 in Figure 5. Such guiding of the cooling air 80 toward, rather than away from, the dividing wall 56 -- and, hence, toward the direction of rotation of the cooling air as it makes the turn -- inhibits the tendency for flow separation.

According to the current invention, this scheme for orienting the fins is implemented in the fourth passage 54 as well. Consequently, as shown in Figures 2 and 6, the fins 63 in the fourth passage 54 are angled so that they extend radially outward -- that is, toward the airfoil tip 45 -- as they extend in the upstream direction toward the wall 57 dividing the third and fourth passages.

Thus, the fins 62 and 63 in both the third and fourth passages 53 and 54 are angled so that they extend toward the direction of cooling air flow though that passage as they extend toward the upstream passage -- that is, as viewed in Figure 2, the fins 62 are angled down (radially inward) as they extend to the left (toward upstream passage 52) and the fins 63 are angled up (radially outward) as they extend to the left (toward the upstream passage 53) . Consequently, the orientation of the fins 61-63 -- that is, the angle at which the fins extend as they extend along the length of the passage -- is reversed with each succeeding passage.

As can be seen, according to the current invention, the fins not only increase the turbulence of the cooling air 81 but also serve to increase its flow rate through the passages 52-54 by inhibiting flow separation. Although the present invention has been discussed with reference to the second row of turbine blades in a gas turbine, the invention is also applicable to other rows of blades, as well as to other types of turbomachines in which airfoil cooling effectiveness is important. Accordingly, the present invention may be embodied in other specific forms without departing from the spirit or essential attributes thereof and, accordingly, reference should be made to the appended claims, rather than to the foregoing specification, as indicating the scope of the invention.

Claims

CLAIMS:

1. A rotating blade for a turbine, comprising: a) leading and trailing edges; b) a root portion for attaching said blade to a rotor, and a tip portion remote from said root portion, said root portion having means for receiving a first flow of cooling fluid; c) a first passage disposed adjacent one of said edges, said first passage having means for directing said first flow of cooling fluid to flow in a direction away from said root portion and toward said tip portion; and d) a plurality of first fins extending into said first passage, said first fins angled so as to extend toward said tip portion of said blade as they extend toward said one of said edges to which said first passage is adjacent.

2. The turbine blade according to claim 1, wherein said first fins are angled so as to form an angle in the range of approximately 45-60° with respect to the radially outward direction.

3. The turbine blade according to claim 1, wherein said one of said edges to which said first passage is adjacent is said leading edge.

4. The turbine blade according to claim 1, wherein said root portion has means for receiving a second flow of cooling fluid, and further comprising: a) second and third passages having means for directing said second flow of cooling fluid in first and second directions, respectively, said second and third passages being in sequential flow communication, whereby said second flow of cooling fluid flows from said second passage to said third passage; b) first turning means for turning said second flow of cooling fluid from said first direction to said second direction as it flows from said second passage to said third passage; and c) means for retarding flow separation in said second flow of cooling fluid as it is turned by said first turning means.

5. The turbine blade according to claim 4, further comprising a wall disposed between said second and third passages, and wherein said means for retarding flow separation comprises a plurality of second fins extending into said third passage, said second fins being angled so as to extend toward said second direction as they extend within said second passage toward said wall.

6. The turbine blade according to claim 5, wherein said second fins are oriented so as to form an angle in the range of approximately 45-60° with respect to the radially outward direction.

7. The turbine blade according to claim 6, wherein said first and second directions differ by approximately 180°.

8. The turbine blade according to claim 7, wherein said first direction is toward said root portion and said second direction is toward said tip portion.

9. The turbine blade according to claim 4, wherein said first turning means comprises a fourth passage connecting said second and third passages.

10. The turbine blade according to claim 4, further comprising: a) a fourth passage having means for directing said second flow of cooling fluid in said first direction, said third and fourth passages being in sequential flow communication, whereby said second flow of cooling fluid flows from said third passage to said fourth passage; b) second turning means for turning said second flow of cooling fluid from said second direction to said first direction as it flows from said third passage to said fourth passage; and c) means for retarding flow separation in said second flow of cooling fluid as it is turned by said second turning means.

11. The turbine blade according to claim 4, further comprising second and third fins projecting into said second and third passages, respectively.

12. The turbine blade according to claim 11, wherein said second fins are angled toward said root portion as they extend toward said leading edge, and wherein said third fins are angled toward said tip portion as they extend toward said leading edge.

13. The turbine blade according to claim 11, wherein said second fins are angled toward said tip portion as they extend toward said leading edge, and wherein said third fins are angled toward said root portion as they extend toward said leading edge.

14. A rotating blade for a turbine, comprising: a) leading and trailing edges and first and second ends; b) first and second cooling fluid passages, said second passage being connected to said first passage so as to be in sequential flow communication therewith; c) a plurality of first fins projecting into said first passage, said first fins angled so as to extend toward said first end of said blade as they extend within said first passage toward said leading edge; and d) a plurality of second fins projecting into said second passage, said second fins angled so as to extend toward said second end of said blade as they extend within said second passage toward said leading edge.

15. The turbine blade according to claim 14, further comprising: a) a third cooling fluid passage connected to said second passage so as to be in sequential flow communication therewith; and b) a plurality of third fins projecting into said third passage, said third fins angled so as to extend toward said first end of said blade as they extend within said third passage toward said leading edge.

16. The turbine blade according to claim 14, wherein said first passage has means for directing cooling fluid from said second end of said blade toward said first end.

17. The turbine blade according to claim 14, wherein said second passage is disposed adjacent said trailing edge.

18. The turbine blade according to claim 17, further comprising a plurality of cooling holes extending from said second passage through said trailing edge.

19. A turbomachine comprising: a) a compressor for producing compressed fluid; b) a combustor for heating a first portion of said compressed fluid, thereby producing a hot compressed gas; and c) a turbine for expanding said hot compressed gas, said turbine including a rotating blade having:

(i) a plurality of successive cooling fluid passages arranged in a serpentine configuration, whereby each of said passages except a first one of said passages is preceded by another one of said passages, one of said passages having means for receiving a second portion of said compressed fluid from said compressor, each of said passages having means for directing said second portion of said compressed fluid to flow in a direction opposite to that of said preceding passage; and

(ii) a plurality of fins projecting into each of said passages, said fins in each of said passages except said first passage being angled so that the direction in which said fins extend in their respective passages is reversed from the direction in which said fins extend in said preceding passage.