A turbine nozzle includes a plurality of vanes joined at opposite ends to outer and inner bands. The inner band has a forward hook which is segmented to reduce thermal mismatch. And, in additional embodiments the vane includes an impingement baffle having preferential cooling.
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1. A turbine nozzle for a gas turbine engine, comprising:
a plurality of vanes integrally joined at opposite ends to outer and inner bands; and
said inner band having a forward hook inboard of a leading edge thereof, said hook including an inner lip and a plurality of circumferentially spaced apart outer lips spaced radially from said inner lip to define a retention slot therein.
0. 48. A turbine nozzle for a gas turbine engine, said nozzle comprising:
a radially inner band;
a radially outer band; and
at least one vane extending between said radially inner and outer bands, said inner band comprising a leading edge and a forward hook radially inward of said leading edge, said forward hook comprising a radially inner lip and a plurality of circumferentially-spaced apart radially outer lips such that a retention slot is defined therebetween.
0. 30. A turbine nozzle for a gas turbine engine, said nozzle comprising:
a radially inner band;
a radially outer band; and
at least one vane extending between said inner and outer bands, said vane comprising a first sidewall and a second sidewall connected together at a leading edge and a trailing edge to define a cavity therebetween, said at least one vane formed integrally with said radially outer band such that a compound fillet extends between said radially outer band and said at least one vane, said compound fillet circumscribes said at least one vane, and comprises a first radius and a second radius that is larger than said first radius, said first radius extending between said second radius and at least one of said radially outer and inner bands.
0. 43. An impingement baffle for a turbine nozzle for use in a gas turbine engine, said impingement baffle comprising a first sidewall and a second sidewall connected together at a leading edge and a trailing edge, said impingement baffle further comprising a plurality of rows of impingement holes extending therethrough, said plurality of rows of impingement holes comprising at least a first row of first impingement holes extending along said baffle leading edge and a pattern of remaining impingement holes, at least some of said impingement holes have a first diameter and said remaining impingement holes have a second diameter, said first diameter is different than said second diameter, said impingement baffle leading edge has a radius sufficient for forming said first diameter first impingement holes substantially planar therein.
0. 34. A turbine nozzle for a gas turbine engine, said nozzle comprising at least one vane comprising a first sidewall and a second sidewall connected together at a leading edge and a trailing edge to define a cavity therebetween, the cavity sized to receive an impingement baffle therein, said impingement baffle comprising a plurality of impingement holes extending therethrough, said plurality of impingement holes comprising at least a row of first impingement holes for directing cooling airflow along an inner surface of said at least one vane leading edge, and a pattern of remaining impingement holes, at least some of said impingement holes have a first diameter, said remaining impingement holes have a second diameter, said first diameter is different than said second diameter, said impingement baffle leading edge has a radius sufficient for forming said first impingement holes substantially planar therein.
15. A turbine nozzle for a gas turbine engine, comprising:
a plurality of vanes integrally joined at opposite ends to outer and inner bands;
said inner band having a forward hook inboard of a leading edge thereof, said hook including an inner lip and a plurality of circumferentially spaced apart outer lips spaced radially from said inner lip to define a retention slot therein;
said vanes have generally concave, pressure sides and opposite, generally convex, suction sides extending between leading and trailing edges thereof and between said outer and inner bands along respective spans of said vanes; and
said vanes include cavities each having an impingement baffle, and said baffles have a plurality of spaced apart impingement holes extending therethrough and arranged in different patterns facing an inner surface of said cavities for preferentially channeling cooling air in impingement jets toward midspan of said vane.
0. 20. A turbine nozzle for a gas turbine engine, said nozzle comprising:
a radially inner band;
a radially outer band; and
at least one vane extending between said inner and outer bands, said vane comprising a first sidewall and a second sidewall connected together at a leading edge and a trailing edge to define a cavity therebetween; and
an impingement baffle positioned within the vane cavity, said impingement baffle comprising a plurality of first impingement baffle holes extending therethrough and a plurality of second impingement baffle holes extending therethrough, at least some of said first impingement baffle holes extending along said vane leading edge and having a first diameter, said second impingement baffle holes having a second diameter, said first diameter is larger than said second diameter, said impingement baffle leading edge having a radius sufficient for forming said first impingement baffle holes substantially planar therein, said at least one vane leading edge having a corresponding larger radius that substantially compliments said impingement baffle leading edge.
2. A nozzle according to
3. A nozzle according to
4. A nozzle according to
5. A nozzle according to
said vanes have a generally concave, pressure side and an opposite, generally convex, suction side extending between leading and trailing edges and between said outer and inner bands along respective spans of said vanes; and
said vanes include cavities having an impingement baffle, and said baffles have a plurality of spaced apart impingement holes extending therethrough and arranged in different patterns facing an inner surface of said cavities for preferentially channeling cooling air in impingement jets toward midspan of said vanes.
6. A nozzle according to
7. A nozzle according to
8. A nozzle according to
said impingement holes on said pressure and suction sides have equal size; and
said baffles include a row of larger impingement holes extending along leading edges thereof.
9. A nozzle according to
10. A nozzle according to
11. A nozzle according to
12. A nozzle according to
13. A nozzle according to
14. A nozzle according to
16. A nozzle according to
17. A nozzle according to
18. A nozzle according to
19. A nozzle according to
0. 21. A turbine nozzle in accordance with
0. 22. A turbine nozzle in accordance with
0. 23. A turbine nozzle in accordance with
0. 24. A turbine nozzle in accordance with
0. 25. A turbine nozzle in accordance with
0. 26. A turbine nozzle in accordance with
0. 27. A turbine nozzle in accordance with
0. 28. A turbine nozzle in accordance with
0. 29. A turbine nozzle in accordance with
0. 31. A turbine nozzle in accordance with
0. 32. A turbine nozzle in accordance with
0. 33. A turbine nozzle in accordance with
0. 35. A turbine nozzle in accordance with
0. 36. A turbine nozzle in accordance with
0. 37. A turbine nozzle in accordance with
0. 38. A turbine nozzle in accordance with
0. 39. A turbine nozzle in accordance with
0. 40. A turbine nozzle in accordance with
0. 41. A turbine nozzle in accordance with
0. 42. A turbine nozzle in accordance with
0. 44. An impingement baffle in accordance with
0. 45. An impingement baffle in accordance with
0. 46. An impingement baffle in accordance with
0. 47. An impingement baffle in accordance with
0. 49. A turbine nozzle in accordance with
0. 50. A turbine nozzle in accordance with
0. 51. A turbine nozzle in accordance with
0. 52. A turbine nozzle in accordance with
0. 53. A turbine nozzle in accordance with
0. 54. A turbine nozzle in accordance with
0. 55. A turbine nozzle in accordance with
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The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. Energy is extracted from the gases in corresponding turbine stages which power the compressor and produce useful work, such as powering a fan in a turbofan engine for propelling an aircraft in flight, for example.
Since the turbines are bathed in the hot combustion gases during operation, they must be suitably cooled which is typically accomplished by bleeding a portion of the pressurized air from the compressor and channeling it through the turbine components.
A high pressure turbine directly receives gases from the combustor and includes a stator nozzle and a corresponding first stage rotor having a plurality of rotor blades extending radially outwardly from a supporting disk. A second stage nozzle then directs the combustion gases through a corresponding row of rotor blades extending from another rotor disk. The second stage nozzle receives lower temperature combustion gases than the first stage nozzle and therefore has different cooling requirements, which are typically effected in a different manner than that for the first stage nozzle.
Turbine nozzles are designed for durability with extensive lives measured in thousands of hours or thousands of cycles of operation. Such extended life is difficult to achieve since the nozzles are subject to various differential temperatures during operation which create thermal loads and stress therefrom. And, temperature distributions and heat transfer coefficients of the combustion gases channeled through the nozzle vary significantly and increase the complexity of providing corresponding cooling. Suitable nozzle cooling is required to limit thermal stresses and ensure a useful life.
A typical turbine nozzle includes a row of stator vanes joined at radially opposite ends to corresponding outer and inner bands. The bands are typically segmented in the circumferential direction, and include two or more vanes in corresponding sectors. The vane sectors permit differential movement during combustion gas temperature changes for reducing undesirable thermal stress during operation.
The individual vanes are hollow and typically include an impingement baffle therein which is a perforated sheet metal sleeve spaced from the inner surface of the vane cavity for channeling cooling air in impingement jets there against.
This type of turbine nozzle specifically configured for a second stage turbine has enjoyed many years of commercial service in this country. However, these nozzles are beginning to experience distress at high cycle operation which may require their replacement prior to their expected useful life. Nozzle distress is caused by locally high heat transfer coefficients in different regions of the nozzle at which corresponding cooling is limited. Thermal gradients lead to thermal stress, which adversely affect the useful life of the nozzle.
Accordingly, it is desired to uncover the source of high cycle turbine nozzle distress, and improve the nozzle design for increasing nozzle durability and corresponding life.
A turbine nozzle includes a plurality of vanes joined at opposite ends to outer and inner bands. The inner band has a forward hook which is segmented to reduce thermal mismatch. And, in additional embodiments the vane includes an impingement baffle having preferential cooling.
The invention, in accordance with preferred and exemplary embodiments, together with further objects and advantages thereof, is more particularly described in the following detailed description taken in conjunction with the accompanying drawings in which:
Illustrated in
The engine includes a high pressure turbine 20 having a first stage stator nozzle 22 followed in turn by a row of first stage turbine rotor blades 24 extending radially outwardly from a supporting disk. The combustion gases 18 are channeled through the nozzle vanes 22 and blades 24 for powering the compressor in a conventional manner.
Disposed immediately downstream from the first stage blades 24 is a second stage turbine stator or nozzle 26 which in turn channels the combustion gases to a downstream row of second stage turbine rotor blades (not shown) which power the fan in a conventional manner.
But for the improved second stage turbine nozzle 26 illustrated in
More specifically, the second stage turbine nozzle 26 illustrated in
The inner band 32 is relatively thin and is locally enlarged at its forward and aft ends for integrally including a forward hook 36 and an aft flange 38 which extend radially inwardly for supporting a honeycomb rotor seal 40 in a conventional manner.
As illustrated in more particularly in
The forward hook 36 includes an arcuate inner lip 36a which is circumferentially continuous between its opposite ends, and is spaced radially inwardly from a plurality of circumferentially spaced apart outer lips 36b spaced radially outwardly from the inner lip to define a corresponding retention slot 42 therein. The outer lips 36b are better illustrated in
The seal 40 illustrated in
Instead of being circumferentially continuous like the lower lip 36a, the outer lip 36b is segmented for removing substantial thermal mass from the inner band to reduce or eliminate thermal mismatch between the forward hook and the inner band during transient operation. During such operation, the combustion gases 18 flow over the inner band 32 causing heating thereof. Since the forward hook 36 is hidden below the inner band it is isolated from the combustion gases and therefore has a delayed thermal response. By segmenting only the forward hook outer lip 36b, a significant reduction in thermal mass may be obtained without compromising the performance of the forward hook for supporting the rotor seal 40 in a sealed fit therewith.
As shown in
Since the individual vanes 28 are integrally joined to both the outer and inner bands 30,32 as illustrated in
As shown in more detail in
As shown in
As shown in
As shown in
As shown in
Impingement baffles in turbine nozzles are well known in commercial use and typically include impingement holes in uniform patterns on the concave and convex sides thereof. However, in accordance with another embodiment of the present invention, the pattern of the impingement holes 50 on the concave side 48a of the baffle as shown in
In the preferred embodiment illustrated in
The baffles 48 preferably include imperforate zones or regions 54 at the radially outer and inner opposite span ends thereof that generally converge toward the baffle midspan between the corresponding leading and trailing edges 48c,d thereof. In this way, the improved baffle 48 may use the same amount of cooling air found in the previous baffle used in commerce, but preferentially distributes the cooling air to the thermally distressed areas near the midspan of the vanes.
As shown in FIGS. 5,8 and 11, the baffles 48 preferably also include a row of larger impingement holes 50b extending along the leading edges 48c thereof to preferentially cool each vane behind its leading edge. The large impingement holes 50b illustrated in
The impingement holes 50,50b are preferably arranged in patterns having different flow density or flow per unit area for preferentially impingement cooling the different regions of the vanes. As shown in
And, in accordance with another embodiment of the present invention, the large impingement holes 50b along the baffle leading edge have a greater flow density for preferentially cooling the inside of the vane leading edge, as shown in
It is noted that a given amount of cooling air 14 is provided for each vane and corresponding baffle which must be suitably distributed inside the different regions of the vane. The high density holes on the baffle concave side 48a provide more cooling of the vane pressure side than the lower density impingement holes in the baffle convex side 48b on the vane suction side.
Correspondingly, the high density impingement holes 50b along the baffle leading edge 48c concentrate cooling along the back of the vane leading edge. The increased amount of impingement cooling air provided along the vane leading edge and pressure side is at the expense of a reduced amount on the suction side.
However, by introducing the imperforate regions 54 along both sides of the baffle near the outer and inner ends thereof, additional cooling air is provided for the remaining impingement holes by eliminating impingement cooling in the imperforate regions 54.
As illustrated in
Since the combustion gases 18 stagnate at the vane leading edge during operation, they effect a correspondingly high external heat transfer coefficient along the vane leading edge. By increasing the radius of the leading edge, and correspondingly increasing the radius of the baffle leading edge 48c and introducing the large impingement holes 50b therein, a significant increase in the ratio of the cooling area behind the vane leading edge to the heated area outside the vane leading edge is provided, with a corresponding reduction in temperature of the vane leading edge.
As shown in
As shown in
These various improvements described above provide tailored and preferential cooling of the different portions of the nozzle vanes 28 themselves for reducing thermal distress and improving nozzle durability and life. Furthermore, the improved forward hook 36 of the inner band 32 and the compound fillet 44 at the outer band 30 provide significant reductions in local thermal stress and mismatch which further improves the durability and life of the nozzle. The nozzle therefore enjoys decreased metal temperature during operation, a more balanced thermal design, and reduced peak stresses which all directly contribute to increased durability of the nozzle and enhanced life.
While there have been described herein what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein, and it is, therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention.
Nichols, Glenn H., Tressler, Judd Dodge
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