TURBINE COOLING SYSTEM AND METHOD

Abstract

A turbine blade of a gas turbine engine is disclosed. The blade may comprise a first cooling passage and a second cooling passage, where the second cooling passage may be located adjacent a leading edge of the blade. The blade may further comprise at least one slot formed in a partition disposed between the first and second cooling passages, and at least one turbulator disposed within the second passage and downstream the at least one slot.

Description

TECHNICAL FIELD

The present disclosure relates generally to gas turbine engine cooling, and more particularly to the cooling of turbine blades in a gas turbine engine (GTE).

BACKGROUND

GTEs produce power by extracting energy from a flow of hot gas produced by combustion of fuel in a stream of compressed air. In general, turbine engines have an upstream air compressor coupled to a downstream turbine with a combustion chamber (“combustor”) in between. Energy is released when a mixture of compressed air and fuel is burned in the combustor. In a typical turbine engine, one or more fuel injectors direct a liquid or gaseous hydrocarbon fuel into the combustor for combustion. The resulting hot gases are directed over blades of the turbine to spin the turbine and produce mechanical power.

High performance GTEs include cooling passages and cooling fluid to improve reliability and cycle life of individual components within the GTE. For example, in cooling the turbine section, cooling passages are provided within the turbine blades to direct a cooling fluid therethrough. Conventionally, a portion of the compressed air is bled from the air compressor to cool components such as the turbine blades. The amount of air bled from the air compressor, however, is usually limited so that a sufficient amount of compressed air is available for engine combustion to perform useful work.

U.S. Pat. No. 5,603,606 to Glezer et al. (the '606 patent) describes a system for cooling airfoils, such as turbine blades and nozzles, in a GTE. According to the system described in the '606 patent, cooling fluid from the compressor section of the GTE is directed through a cooling path of a turbine blade. The turbine blade includes a plurality of holes or slots, which act as a means for swirling a portion of the cooling fluid through the turbine blade. The cooling fluid flows out of the plurality of holes or slots and into an empty gallery, where the cooling fluid swirls radially along the gallery.

SUMMARY

In one aspect, a turbine blade of a gas turbine engine is disclosed. The blade may comprise a first cooling passage and a second cooling passage, where the second cooling passage may be located adjacent a leading edge of the blade. The blade may further comprise at least one slot formed in a partition disposed between the first and second cooling passages, and at least one turbulator disposed within the second passage and downstream of the at least one slot.

In another aspect, a system for cooling a turbine blade of a gas turbine engine is disclosed. The system may include a first cooling path disposed in an interior of the turbine blade and configured to receive a flow of cooling fluid, and a second cooling path disposed in the interior of the turbine blade and configured to receive the flow of cooling fluid. A plurality of passages may form each of the first and second cooling paths, and a plurality of slots may be formed in a partition disposed between two of the plurality of passages forming the first cooling path. The plurality of slots may be configured to guide the flow of cooling fluid through the first cooling path, and at least one turbulator may be disposed downstream of the flow of cooling fluid flowing through the plurality of slots.

In yet another aspect, a method of cooling a turbine blade of a gas turbine engine is disclosed. The method may include flowing a cooling fluid through a plurality of passages and through at least one slot formed in a partition disposed between two of the plurality of passages. The method may further include adding turbulence to at least part of the flow of cooling fluid downstream of the at least one slot, the turbulence being adjacent a leading edge of the turbine blade.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a sectional view of a portion of a turbine section of a gas turbine engine;

FIG. 2 is an enlarged sectional view of a turbine blade taken along lines 2-2 of FIG. 1;

FIG. 3 is an enlarged view of a portion of the turbine blade of FIG. 2 along line 3;

FIG. 4 is an enlarged sectional view of the turbine blade of FIG. 2 taken alone line 4-4.

FIG. 5 is an enlarged view of a portion of a turbine blade showing another embodiment of a turbulator; and

FIG. 6 is an enlarged view of a portion of a turbine blade showing yet another embodiment of a turbulator.

DETAILED DESCRIPTION

FIG. 1 illustrates a sectional view of a portion of a GTE, specifically a turbine section 10 of the GTE. The turbine section 10 includes a first stage turbine assembly 12 disposed partially within a first stage shroud assembly 20.

During operation, a cooling fluid, designated by the arrows 14, flows from the compressor section (not shown) to the turbine section 10. Furthermore, each of the combustion chambers (not shown) are radially disposed in a spaced apart relationship with respect to each other, and have a space through which the cooling fluid 14 flows to the turbine section 10. The turbine section 10 further includes a fluid flow channel 16 through which the cooling fluid 14 flows.

The turbine section 10 shown in FIG. 2 includes the first stage turbine assembly 12, which includes a rotor assembly 18 radially aligned with the shroud assembly 20. The rotor assembly 18 may be of a conventional design including a plurality of turbine blades 22. Each of the turbine blades 22 may be made from a variety of materials. The rotor assembly 18 further includes a disc 24 having a plurality of circumferentially arranged root retention slots 30. The plurality of turbine blades 22 are replaceably mounted within the disc 24. Each of the plurality of blades 22 may include a first end 26 having a root section 28 extending therefore which engages with one of the corresponding root retention slots 30. The first end 26 may be spaced away from a bottom of the root retention slot 32 in the rotor assembly 18 to form a cooling fluid inlet opening 34 configured to receive cooling fluid 14. Each turbine blade 22 may further include a platform section 36 disposed radially outward from a periphery of the disc 24 and the root section 28. Additionally, an airfoil 38 may extend radially outward from the platform section 36. Each of the plurality of turbine blades 22 may include a second end 40, or tip, positioned opposite the first end 26 and adjacent the airfoil 38.

As is more clearly shown in FIGS. 2, 3, and 4, each of the plurality of turbine blades 22 includes a leading edge 42, and a trailing edge 44 positioned opposite the leading edge 42. Interposed the leading edge 42 and the trailing edge 44 is a pressure, or concave, side 46 and a suction, or convex, side 48 of the turbine blade 22. Each of the plurality of turbine blades 22 may have a generally hollow configuration forming a peripheral wall 50, which, in some embodiments, may have a uniform thickness.

As shown in FIGS. 2, 3, and 4, an arrangement 52 for internally cooling each of the turbine blades 22 is provided. The arrangement 52 for internal cooling may include a pair of cooling paths 64 and 76 (FIG. 4), each described in more detail below, positioned within the peripheral wall 50, and separated from one another; however, any number of cooling paths could be used.

A first cooling path 64 positioned within the peripheral wall 50 is interposed the leading edge 42 and the trailing edge 44 of each of the blades 22. The first cooling path 64 includes a first passage 54 extending between the first end 26 and the second end 40 of the turbine blade 22. Further included in the first cooling path 64 is a second passage 56, which extends between the first end 26 and the second end 40 of the turbine blade 22. The second passage 56 is interposed the leading edge 42 and the first passage 54 by a first partition 60, which is connected to the peripheral wall 50 at both the pressure side 46 and the suction side 48 of the turbine blade 22.

As shown in FIGS. 2, 3 and 4, at least one slot 58 is positioned in the first partition 60 in fluid communication with the first passage 54 and the second passage 56. The slot, which may be a plurality of slots 58 as shown in the figures, is positioned adjacent the peripheral wall 50 near the pressure side 46 of each of the turbine blades 22.

A turbulator or turbulator element 62 configured to produce a turbulent fluid flow is disposed on the peripheral wall 50 in the second passage 56. In some embodiments, the turbulator element 62 may be formed integrally with the peripheral wall 50. As shown in FIGS. 2 and 3, the turbulator element 62 is disposed in a flow path of the cooling fluid 14 downstream of the slot 58 and upstream of the leading edge 42, such that the turbulator element 62 is aligned with the flow of cooling fluid 14 exiting the slots 58. In one instance, as shown in FIGS. 2, 3, and 4, a single turbulator element in the form of a radial trip-strip may be disposed in the second passage 56. If provided as a trip-strip, the turbulator element 62 may have any cross-section, for example a rectangular cross-section, as shown in FIGS. 2 and 3. As shown in FIG. 4, the turbulator element 62 may extend radially within the second passage 56 directly downstream of each of a plurality slots 58. Although FIG. 4 shows the turbulator element 62 downstream of each one of a plurality of slots 58, the turbulator element 62 may have any length so as to be downstream of any number of slots 58 provided in the first partition 60.

In some embodiments, a plurality of turbulator elements 62 may be provided in the second passage 56. For example, a plurality of radially disposed trip strips, like the trip strip shown in FIGS. 2, 3, and 4, could be disposed on the peripheral wall 50 in the second passage 56 downstream of the slots 58. In other embodiments, the turbulator element 62 may be a plurality of broken ribs arranged on the peripheral wall 50 at different angles within the second passage 54. In yet other embodiments, as shown in FIGS. 5 and 6 and as described in more detail below, the turbulator 62 may take the form of one or more concave cavities, or dimples 62a, in the peripheral wall 50 and/or one or more convex protrusions 62b formed on the peripheral wall 50 of the second chamber 56. The one or more concave dimples 62a or the one or more convex protrusions 62b may otherwise be located and arranged as discussed above with respect to the turbulator elements 62, including arranged as a plurality of concave dimples 62a or convex protrusions 62b. The plurality of concave dimples 62a may be formed at various locations in the peripheral wall 50 in the second passage 56, and the plurality of convex protrusions 62b may be formed at various locations on the peripheral wall 50 in the second passage 56.

As shown in FIG. 3, the turbulator element 62 has a turbulator element height D1, representing the distance that the turbulator element 62 extends from the peripheral wall 50 into the second passage 56. As further shown in FIG. 3, the slot 58 has a slot height D2, representing the open space through which the cooling fluid 14 can flow from the first passage 54 to the second passage 56, as described in more detail below. Furthermore, the turbulator element 62 is spaced from the slot inlet opening by a distance D3, which may be measured from where the cooling fluid 14 first enters the slot 58 to a front face of the turbulator element 62.

In some embodiments, a turbulator element 62 is provided having a turbulator element height D1 depending on the slot height D2. For example, a specified ratio of slot height D2 to turbulator element height D1 may be used to determine the turbulator element 62 to provide in the second passage 56. In one instance, a ratio of the slot height D2 to the turbulator element height D1 may be about 2 to 1, such that the slot height D2 is approximately two times the turbulator element height D1. For example, a turbine blade 22 may be provided having a slot height D2 of about 0.6100 mm (0.024 inches), in which case a corresponding turbulator element 62 may be provided having a turbulator element height D1 of about 0.305 mm (0.012 inches).

The distance D3 between the turbulator element 62 and the slot inlet opening may also depend on both the turbulator element height D1 and the slot height D2. For example, as described above, a larger turbulator element 62 (i.e. a turbulator element 62 having a greater turbulator element height D1) may be provided for a larger slot (i.e. a slot 58 having a greater slot height D2). Additionally or alternatively, the distance D3 may increase proportionately to an increase in slot height D2. For example, a ratio of the distance D3 to the slot height D2 may be less than about 4 to 1, or about 3.5 to 1. Thus, the distance D3 from the slot inlet opening to the turbulator or turbulator element 62 may be less than about four times the slot height D2, or about three and one half times the slot height D2. In one instance, for a turbine blade having a slot 58 with a slot height D2 of about 0.6100 mm (0.024 inches), a corresponding distance D3 may be about 2.16 mm (0.085 inches).

As shown in FIGS. 2 and 3, a cross-sectional view of the second passage 56 has a pre-established cross-sectional configuration. A generally arcuate portion 96 adjacent the leading edge 42, a generally straight portion 98 following along the first partition 60, and the intersection therebetween form an angle 100. In some embodiments, the angle 100 may be acute, for example between 45 and 60 degrees. In other embodiments, however, the angle may be less than 45 degrees, for example 30 degrees or less, or the angle may be greater than 60 degrees, for example 75 degrees or more. As further shown in FIGS. 2 and 3, a plurality of apertures 102, only one of which is shown, may each have a predetermined cross-sectional area, and may each communicate between the second passage 56 and the suction side 48 of the turbine blade 22. In some embodiments, the cross-sectional area of the apertures 102 may be about half the size of a cross-sectional area of the second passage 56. The plurality of apertures 102 provide outlets on the suction side 48 at an inclined angle generally directed from the leading edge 42 toward the trailing edge 44. The plurality of apertures 102 may be arranged as showerhead apertures on the suction side 48 for showerhead film cooling of the turbine blade 22.

As mentioned above with respect to FIG. 1, a cooling fluid inlet opening 34 provided adjacent the first end 26 of the turbine blade 22 is configured to receive cooling fluid 14 from first flow channel 16. The cooling fluid inlet opening 34 is connected to the first and second cooling paths 64 and 76, respectively. As shown in FIG. 4, the first cooling path 64 includes a first cooling path inlet opening 66 originating at the first end 26 of the turbine blade 22. The first passage 54 of the first cooling path 64 extending outwardly substantially an entire length of the turbine blade 22 toward the second end 40 of the turbine blade 22. The second passage 56 includes the turbulator element 62 and is in communication, for example fluid communication, with a first horizontal passage 68. The first horizontal passage 68 may be at least partially interposed the second end 40 of the turbine blade 22 and the first passage 54 by a second partition 70, which is connected to the peripheral wall 50 at both the pressure side 46 and the suction side 48. The second passage 56 has an end 72 adjacent the first end 26 of the turbine blade 22. The first horizontal passage 68 communicates with a first cooling path outlet opening 74 disposed in the trailing edge 44 of the turbine blade 22.

FIG. 4 also shows a plurality of the slots 58 extending radially between the end 72 of the second passage 56 and an end 192 of the first passage 54, where the end 192 of the first passage 54 is positioned opposite the first end 26 of the turbine blade 22. In some embodiments, an angled passage 194 may extend between the first passage 54 and the second passage 56. The angled passage 194 enters the end 72 of the second passage 56 at an angle, which, in some embodiments, may be about 30 to 60 degrees.

FIG. 4 further illustrates a second cooling path 76 positioned with the peripheral wall 50 and interposed the first cooling path 64 and the trailing edge 44 of each turbine blade 22. The second cooling path 76 is separated from the first cooling path 64 by a first wall member 80. The second cooling path 76 includes a second cooling path inlet opening 78 originating at the first end 26 of the turbine blade 22 and has a radial third passage 82 extending outwardly substantially the entire length of the turbine blade 22 toward the second end 40 of the turbine blade 22. The second cooling path inlet opening 78 and the third passage 82 are interposed the first cooling path 64 and the trailing edge 44. Further included is a second horizontal passage 84 positioned inwardly of the first horizontal passage 68 of the first cooling path 64 and is in communication with the third passage 82 and a fourth passage 86. The fourth passage 86 extends radially inwardly from the second horizontal passage 84 to a third horizontal passage 88. The third horizontal passage 88 communicates with a generally radial second cooling path outlet opening 90 disposed at the trailing edge 44 of the turbine blade 22. The third passage 82 is separated from the fourth passage 86 by a second wall member 92 which is connected to the peripheral wall 50 at the pressure side 46 and the suction side 48. The fourth passage 86 is separated from the second cooling path outlet opening 90 by a third wall member 94 that is also connected to the peripheral wall 50 at the pressure side 46 and the suction side 48.

The aforementioned description is of the first stage turbine assembly 12; however, it should be known that the construction could be typical of the remainder of the turbine stages within the turbine section 10 where cooling may be employed.

INDUSTRIAL APPLICABILITY

The described system may be applicable to turbine blades of a GTE. Additionally, although the system has been described with respect to turbine blades in the first stage turbine assembly, the system may be applied to any turbine blade in any stage of the turbine section of a GTE. Furthermore, although the above-mentioned cooling system has been described for cooling a turbine blade, the system may be adapted to any airfoil, for example a first stage nozzle and shroud assembly. Additionally, the cooling system may be applied to any localized structure subject to heat that involves air flows and hot fluid temperatures. Moreover, the described cooling system may be applied in a variety of industries, for example, heat exchange, energy, aerospace, or defense.

The following operation will be directed to the first stage turbine assembly 12; however, the cooling operation of other airfoils and stages (turbine blades or nozzles) could be similar.

A portion of the compressed fluid from the compressor section of the GTE is bled from the compressor section and forms the cooling fluid 14 used to cool the first stage turbine blades 22. The compressed fluid exits the compressor section, flows through a combustor discharge plenum and enters into a portion of the fluid flow channel 16 as cooling fluid 14. The flow of cooling fluid 14 is used to cool and prevent ingestion of hot gases into the internal components of the GTE. For example, the air bled from the compressor section flows into a compressor discharge plenum, through spaces between a plurality of combustion chambers, and into the fluid flow channel 16 (FIG. 1). After passing through the fluid flow channel 16 in the turbine blade 22, the cooling fluid enters the cooling fluid inlet opening 34 between the first end 26 of the turbine blade 22 and the bottom 32 of the root retention slot 30 in the disc 24. The cooling fluid inlet opening 34 is fluidly connected to the first and second cooling paths 64 and 76, respectively, in the interior of the turbine blade 22.

As shown in FIG. 4, a portion of the cooling fluid 14, after having passed through the cooling fluid inlet opening 34 (FIG. 1), enters the first cooling path 64. The cooling fluid 14 enters the first cooling path inlet opening 66 from the cooling fluid inlet opening 34, and travels radially along the first passage 54, absorbing heat from the peripheral wall 50 and the first wall member 80. An amount of the cooling fluid 14, for example a majority of the cooling fluid 14, exits the first passage 54 through the plurality of slots 58.

The cooling fluid 14 flows through the slots 58 and over the turbulator element 62 creating a turbulent swirling flow that travels at least partially along the arcuate portion 96 of the second passage 56 in a radial direction, as shown in FIG. 3, absorbing heat from the leading edge 42 of the peripheral wall 50. The cooling fluid 14 may flow directly from the slots 58 to the turbulator element 62, such that the cooling fluid 14 flows in an unimpeded path from the slots 58 to the turbulator element 62. Thus, the turbulator element 62 may be said to be directly downstream of the slots 58. The cooling fluid 14 flows over the turbulator element 62 such that laminar flow is turned into turbulent flow through at least the second passage 56. As mentioned above, the turbulator element 62 may be disposed on the peripheral wall within the second passage 56 such that all of the cooling fluid 14 flowing through the slots 58 flows over the turbulator element 62. In some embodiments, however, only a portion of the cooling fluid 14 flowing through the slots 58 from the first passage 54 into the second passage 56 may flow over the turbulator element 62. As noted above, a plurality of turbulator elements 62 may be disposed in the second passage 56 such that the cooling fluid 14 flows over the plurality of turbulator elements 62.

A portion of the cooling fluid 14 may also exit the plurality of apertures 102, cooling a skin of the peripheral wall 50 in contact with combustion gases on the suction side 48 prior to mixing with the combustion gases. The remainder of the cooling fluid 14 in the first cooling path 64 exits the first cooling path outlet opening 74 in the trailing edge 44 to also mix with the combustion gases.

A second portion of the cooling fluid 14, after having passed through the cooling fluid inlet opening 34 (FIG. 1), enters the second cooling path 76. For example, cooling fluid 14 enters the second cooling path inlet opening 78 from the cooling fluid inlet opening 34, and travels radially along the third passage 82 absorbing heat from the peripheral wall 50, the first wall member 80, and the second wall member 92 before entering the second horizontal passage 84, where more heat is absorbed from the peripheral wall 50. As the cooling fluid 14 enters the fourth passage 86, additional heat is absorbed from the peripheral wall 50, the first wall member 80, and the second wall member 92 before entering the third horizontal passage 88 and exiting the second cooling path outlet opening 90 along the trailing edge 44 to be mixed with the combustion gases.

Current swirl-cooling technology may be prone to manufacturing challenges due to a small slot height D2 (FIG. 3) provided for optimum operation. For example, having a small slot height D2 may shorten turbine blade core die life due to accelerated core die wear at the portion of the die forming the slot inlet. The small slot height D2 may also cause high pressure losses that penalize the back-flow margin for blade tip film-cooling. Increasing the slot height D2 alone would lead to a decrease in the velocity of cooling fluid flowing through the slots 58 and a decrease in pressure, which may in turn lead to less effective cooling of the turbine blade 22.

The above-described system provides more efficient use of the cooling air bled from the compressor section of a GTE in order to facilitate increased component life and efficiency of the GTE. The swirling and turbulence of the cooling fluid 14 contributes to the efficiency of the cooling fluid 14 flow as the cooling fluid 14 passes through the turbine blade 22. Increasing the slot height D2 and providing a turbulator element 62 downstream of the slot 58 may decrease pressure losses, which may in turn improve cooling effectiveness downstream of the slot 58. Furthermore, efficiency is facilitated within the internal portion of the turbine blade 22 along the leading edge 42, downstream of the turbulator element 62.

Specifically, the swirling action caused by the arrangement of slots 58 in the first partition 60 and the turbulator element 62 in the second passage 56, the arcuate configuration of the arcuate portion 96 of the second passage 56, along with the flow of cooling fluid 14 through the angled passage 194, cause the cooling fluid 14 to generate a vortex flow in the second passage 56. The vortex flow caused by the slots 58 and turbulator element 62 leads to high local turbulence (vortices) along the arcuate portion 96 adjacent the leading edge 42 of the turbine blade 22. Turbulent flow, as opposed to laminar flow, has more energy for cooling components such as turbine blades 22. The portion of the cooling fluid 14 entering the angled passage 194 between the first passage 54 and the second passage 56 adds to the vortex flow by directing the cooling fluid 14 generally radially outward from the second passage 56 into the first horizontal passage 68. The cooling fluid 14 takes on a screw-type flow from the end 72 of the second passage 56 toward the first horizontal passage 68, which adds to the cooling efficiency in order to better cool the turbine blade 22.

The above-described system enables use of slot having a greater height, and thus easier and less costly to manufacture, without suffering from undesirable reductions in cooling performance, by use of a turbulator element. The turbulator element compensates for a reduced swirl cooling fluid velocity through the slot due to a larger slot height, and also increases the back-flow margin for blade tip film-cooling. As mentioned above, the size, specifically the height, of the turbulator element provided may depend on the slot height. Using a predetermined desired ratio of slot height to turbulator element height, for example a ratio of 2 to 1, may allow for improved cooling performance, while avoiding the manufacturing challenges associated with smaller slots. For example, where a turbulator element is provided as a single radial trip-strip, improvement in cooling may be achieved. Additionally, the above-described system may reduce the tendency of core die wear due to the increased slot height, thereby maintaining turbine blade manufacturing tolerances and increasing turbine blade core die life. Due to a decreased metal temperature of the turbine blade, the blade life may be extended, in some instances by three times. Improved cooling and extended blade life may thus be achieved with minimal core changes to the internal flow passages of the turbine blade by incorporation of a turbulator element downstream of the slots between first and second flow passages.

As noted above, in some embodiments the turbulator or turbulator element may be provided as at least one convex protrusion, or a plurality of convex protrusions, disposed on the peripheral wall in the second chamber. In some embodiments, the protrusions may be formed integrally with the peripheral wall. In yet other embodiments, as also noted above, the turbulator element may be provided as at least one concave cavity, or dimple, or a plurality of concave dimples, formed in the peripheral wall of the second chamber. A plurality of protrusions and/or dimples may be provided as the turbulator or turbulator element where, for example, a radially disposed trip strip may cause excessive pressure loss in the second chamber. Protrusions and/or dimples may provide a desired turbulent flow without causing too great of a reduction in pressure.

It will be apparent to those skilled in the art that various modifications and variations can be made to the disclosed turbine cooling system. Other embodiments will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed system and method. It is intended that the specification and examples be considered as exemplary only, with a true scope being indicated by the following claims and their equivalents.

Claims

1. A turbine blade of a gas turbine engine, the blade comprising:

a first cooling passage;

a second cooling passage, the second cooling passage being located adjacent a leading edge of the blade;

at least one slot formed in a partition disposed between the first and second cooling passages; and

at least one turbulator disposed within the second passage and downstream of the at least one slot.

2. The turbine blade of claim 1, wherein the at least one turbulator comprises a single strip radially disposed within the second cooling passage.

3. The turbine blade of claim 1, wherein the at least one turbulator includes a plurality of turbulator elements disposed within the second cooling passage.

4. The turbine blade of claim 1, wherein the at least one turbulator is disposed directly downstream of the at least one slot.

5. The turbine blade of claim 1, wherein the at least one slot comprises a plurality of slots, and wherein the at least one turbulator is disposed downstream of each of the plurality of slots.

6. The turbine blade of claim 1, wherein the at least one turbulator is connected to a peripheral wall of the turbine blade within the second cooling passage.

7. The turbine blade of claim 1, wherein a height of the at least one slot is greater than a height of the at least one turbulator.

8. The turbine blade of claim 1, wherein a height of the at least one slot is about two times greater than a height of the at least one turbulator.

9. The turbine blade of claim 1, wherein a distance from an inlet opening of the at least one slot to the at least one turbulator is less than about four times a height of the at least one slot.

10. The turbine blade of claim 1, wherein the at least one turbulator comprises a concave dimple formed in a peripheral wall of the turbine blade.

11. A system for cooling a turbine blade of a gas turbine engine, the system comprising:

a first cooling path disposed in an interior of the turbine blade and configured to receive a flow of cooling fluid;

a second cooling path disposed in the interior of the turbine blade and configured to receive the flow of cooling fluid;

a plurality of passages forming each of the first and second cooling paths;

a plurality of slots formed in a partition disposed between two of the plurality of passages forming the first cooling path, wherein the plurality of slots are configured to guide the flow of cooling fluid through the first cooling path; and

at least one turbulator disposed downstream of the flow of cooling fluid flowing through the plurality of slots.

12. The system of claim 11, wherein the at least one turbulator comprises a single strip radially disposed within one of the plurality of passages forming the first cooling path.

13. The system of claim 12, wherein the at least one turbulator is disposed directly downstream of the flow of cooling fluid flowing through the plurality of slots.

14. The system of claim 13, wherein the at least one turbulator is disposed downstream of each of the plurality of slots.

15. The system of claim 11, wherein the at least one turbulator is connected to a peripheral wall of the turbine blade within one of the plurality of passages forming the first cooling path and located adjacent a leading edge of the turbine blade.

16. The system of claim 11, wherein a height of each of the plurality of slots is greater than a height of the at least one turbulator.

17. The system of claim 11, wherein a distance from an inlet opening of each of the plurality of slots to the at least one turbulator is less than about four times a height of each of the plurality of slots.

18. A method of cooling a turbine blade of a gas turbine engine, the method comprising:

flowing a cooling fluid through a plurality of passages and through at least one slot formed in a partition disposed between two of the plurality of passages; and

adding turbulence to at least part of the flow of cooling fluid downstream of the at least one slot, the turbulence being adjacent a leading edge of the turbine blade.

19. The method of claim 18, wherein turbulence is added to the flow of cooling fluid directly downstream of the at least one slot.

20. The method of claim 18, further comprising adding turbulence to the entire flow of cooling fluid flowing through the at least one slot.