TANDEM ROTOR BLADES WITH COOLING FEATURES

Abstract

A gas turbine engine has a compressor section with a downstream most blade stage and a downstream vane row positioned downstream of the downstream most blade stage. The downstream most blade stage includes a plurality of blade pairs with the majority of blade pairs being circumferentially spaced apart from others of the blade pairs, each the blade pair being operatively connected to a rotor disk disposed radially inward from the blade pairs with each blade pair including a forward blade and an aft blade. The aft blades are configured to further condition airflow with respect to the forward blade. A tangential onboard injector (TOBI) is positioned radially inwardly of the downstream vane row.

Description

BACKGROUND OF THE INVENTION

This application relates to tandem compressor rotor blades having improved cooling structure.

Gas turbine engines are known, and typically include a fan delivering air into a bypass duct as propulsion air. The fan also delivers air into a compressor section in a core engine. The compressed air is delivered into a combustor where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors, driving them to rotate.

As can be appreciated, the components in the turbine section and downstream portions of the compressor section experience very high temperatures. As such, a good deal of design is put into cooling those areas.

Historically, the fan in the gas turbine engine rotated at the same speed as a fan drive turbine. However, more recently, a gear reduction has been placed between the two such that the fan can rotate at a slower speed

With this, design opportunities have been given to the designer of a gas turbine engine. One such opportunity is to increase the overall pressure ratio delivered by the compressor section. However, this, in turn, increases temperature and, in particular, the temperature at a downstream end of the compressor section.

SUMMARY OF THE INVENTION

In a featured embodiment, a gas turbine engine has a compressor section with a downstream most blade stage and a downstream vane row positioned downstream of the downstream most blade stage. The downstream most blade stage includes a plurality of blade pairs with the majority of blade pairs being circumferentially spaced apart from others of the blade pairs, each the blade pair being operatively connected to a rotor disk disposed radially inward from the blade pairs with each blade pair including a forward blade and an aft blade. The aft blades are configured to further condition airflow with respect to the forward blade. A tangential onboard injector (TOBI) is positioned radially inwardly of the downstream vane row.

In another embodiment according to the previous embodiment, there is no intervening stator vane stage between the forward blades and the aft blades.

In another embodiment according to any of the previous embodiments, at least 10% of the cooling air passing through the TOBI is compressor midstream air.

In another embodiment according to any of the previous embodiments, the compressor midstream air is cooled in a heat exchanger before passing into the TOBI.

In another embodiment according to any of the previous embodiments, air passing through the TOBI is air having been compressed by downstream most blade stage of the compressor.

In another embodiment according to any of the previous embodiments, the air passes through a heat exchanger before passing through the TOBI.

In another embodiment according to any of the previous embodiments, a seal is provided between a downstream end of the downstream most blade stage and an upstream end of the downstream vane row, and the seal defining a cavity extending radially inwardly along a hub of the disk.

In another embodiment according to any of the previous embodiments, the seal is provided on both the downstream end of the downstream most blade stage, and the upstream end of the downstream vane row.

In another embodiment according to any of the previous embodiments, a static structure radially inward of the downstream vane row includes a plate extending at an angle which is non-perpendicular relative to a center axis of the engine and having a component extending radially inwardly and the disk hub also having a portion extending in a direction which is non-perpendicular to the central axis and having a component which is radially inwardly, the plate causing the air having passed through the TOBI to be confined adjacent to the disk hub portion.

In another embodiment according to any of the previous embodiments, the plate and the disk hub portion each having a radially inner portion which extends closer to being parallel to the central axis than does a radially outer portion.

In another embodiment according to any of the previous embodiments, a mini disk is positioned radially inwardly of the radially inner portion of the hub.

In another embodiment according to any of the previous embodiments, the downstream most blade stage has air passages between the blades and the rotor disk such that air from the cavity can pass upstream toward a vane row upstream of the downstream most blade stage.

In another embodiment according to any of the previous embodiments, cooling air passing through the TOBI is compressor midstream air.

In another embodiment according to any of the previous embodiments, air passing through the TOBI is air having been compressed by the downstream most blade stage of the compressor.

In another embodiment according to any of the previous embodiments, a seal is provided between a downstream end of the downstream most blade stage and an upstream end of the downstream vane row, and the seal defining a cavity extending radially inwardly along a hub of the disk.

In another embodiment according to any of the previous embodiments, the seal is provided on both the downstream end of the downstream most blade stage, and the upstream end of the downstream vane row.

In another embodiment according to any of the previous embodiments, a static structure radially inward of the downstream vane row includes a plate extending at an angle which is non-perpendicular relative to a center axis of the engine and having a component extending radially inwardly and the disk hub also having a portion extending in a direction which is non-perpendicular to the central axis and having a component which is radially inwardly, the plate causing the air having passed through the TOBI to be confined adjacent to the disk hub portion.

In another embodiment according to any of the previous embodiments, a mini disk is positioned radially inwardly of the radially inner portion of the hub.

In another embodiment according to any of the previous embodiments, the downstream most blade stage has air passages between the blades and the rotor disk such that cavity air from the cavity can pass upstream toward a vane row upstream of the downstream most blade stage.

In another featured embodiment, a gas turbine engine has a compressor section with a rotor disk defined between a compressor case and a centerline axis. A plurality of stages is defined radially inward relative to the compressor case. The plurality of stages includes at least one tandem blade stage, wherein the at least one tandem blade stage includes. A plurality of blade pairs, each blade pair circumferentially spaced apart from the other blade pairs and operatively connected to the rotor disk. Each blade pair includes a forward blade and an aft blade, wherein the aft blade is configured to further condition air flow with respect to the forward blade without an intervening stator vane stage shrouded cavity therebetween.

These and other features may be best understood from the following drawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic cross-sectional side elevation view of an exemplary embodiment of a gas turbine engine constructed in accordance with the present disclosure, showing a location of a tandem blade stage;

FIG. 2 is an enlarged schematic side elevation view of a portion of the gas turbine engine of FIG. 1, showing the last stages of the HPC with the tandem blade stage forward of an exit guide vane stage;

FIG. 3 is a top perspective view of an exemplary embodiment of a tandem blade constructed in accordance with the present disclosure, showing a forward blade and an aft blade; and

FIG. 4 is a schematic side elevation view of a portion of another exemplary embodiment of a gas turbine engine, showing the last stages of the HPC with the tandem blade stage forward of a tandem stator vane stage, where the blades of the tandem blade stage do not overlap one another.

DETAILED DESCRIPTION

Reference will now be made to the drawings wherein like reference numerals identify similar structural features or aspects of the subject disclosure. For purposes of explanation and illustration, and not limitation, a cross-sectional view of an exemplary embodiment of the gas turbine engine 100 constructed in accordance with the disclosure is shown in FIG. 1 and is designated generally by reference character 10. Other embodiments of gas turbine engines constructed in accordance with the disclosure, or aspects thereof, are provided in FIGS. 2-4, as will be described.

As shown in FIG. 1, a gas turbine engine 10 defines a centerline axis A and includes a fan section 12, a compressor section 14, a combustor section 16 and a turbine section 18. Gas turbine engine 10 also includes a case 20. Compressor section 14 drives air along a gas path C for compression and communication into the combustor section 16 then expansion through the turbine section 18. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

Gas turbine engine 10 also includes an inner shaft 30 that interconnects a fan 32, a LPC 34 and a low pressure turbine 36. Inner shaft 30 is connected to fan 32 through a speed change mechanism, which in exemplary gas turbine engine 10 is illustrated as a geared architecture 38. An outer shaft 40 interconnects a HPC 42 and high pressure turbine 44. A combustor 46 is arranged between HPC 42 and high pressure turbine 44. The core airflow is compressed by LPC 34 then HPC 42, mixed and burned with fuel in combustor 46, then expanded over the high pressure turbine 44 and low pressure turbine 36.

With continued reference to FIG. 1, HPC 42 is aft of LPC 34. Gas path C is defined in HPC 42 between the compressor case, e.g. engine case 20, and a rotor disk 50. A plurality of stages 22 are defined in gas path C. Plurality of stages 22 includes at least one tandem blade stage 24. Gas turbine engine 10 includes an exit guide vane stage 26 aft of tandem blade stage 24. Exit guide vane stage 26 defines the end of compressor section 14. At least one forward stator vane stage 28 is disposed forward of tandem blade stage 24. Forward stator vane stage 28 and tandem blade stage 24 define the last two sequential stages before exit guide vane stage 26. While embodiments of the tandem blade stage are described herein with respect to a gas turbine engine, those skilled in the art will readily appreciate that embodiments of the tandem blade stage can be used in a variety of turbomachines and in a variety of locations throughout a turbomachine, for example the tandem blade stage can be used in the fan, LPC, low pressure turbine and high pressure turbine.

Tandem blade stage 24 combines two, typically discrete, blade stages into a single stage. For example, a traditional compressor configuration generally has the last stages in the pattern of stator stage, rotor stage, stator stage, rotor stage, and exit guide vane stage. Embodiments described herein have the pattern of stator stage 28, tandem rotor stage 24, and exit guide vane stage 26 or a tandem stator stage, described below. Tandem rotor stage 24 does more work than a traditional single blade stage, providing additional pressure-ratio and also reducing the need for a traditional stator vane stage that typically separates two traditional single blade stages. By removing one of the stator vane stages, respective shrouded cavities that are typically associated with each vane in the stator vane stage, are no longer needed. Shrouded cavities tend to increase metal temperatures because of the interface between a seal, typically a knife edge seal, and the rotor disk. The increased temperatures at the knife edge seal cause increased overall temperatures as part of windage heat-up. By removing one of the shrouded cavities, the windage heat-up is reduced and temperatures of other engine components in the last stages of the HPC are also reduced.

Those skilled in the art will readily appreciate that by reducing the temperatures, the component life can be improved. For example, by removing the intervening stator vane stage and its knife edge seal, the remaining knife edge seals can be approximately ten to fifteen percent of compressor discharge temperature cooler than they would be if the traditional intervening stator stage and knife edge seal was included. Not only does this potentially increase the life of the remaining seals, it also increases the life of the surrounding engine components due to the reduced windage heat-up temperature. On the other hand, the overall operating temperatures can be increased in order to increase the pressure ratio while still remaining within the traditional temperature tolerances of the engine components. Reducing the need for a traditional stator vane stage by using a tandem blade stage also reduces the length of the compressor since gaps between stages can be removed, and/or tandem rotor blades can overlap each other in the axial direction.

As shown in FIG. 2, tandem blade stage 24 includes a plurality of circumferentially disposed blade platforms 48, each having a blade pair 53. Each blade platform 48 is operatively connected to rotor disk 50 disposed radially inward from blade platforms 48. Blade pair 53 extends radially from each of blade platforms 48 and includes a forward blade 52 and an aft blade 54. Those skilled in the art will readily appreciate that each blade pair 53 can be integrally formed with a respective one of blade platforms 48. While tandem blade stage 24 is described herein as having a plurality of blade platforms 48, each with a respective blade pair 53, those skilled in the art will readily appreciate that blade platforms 48 can include multiple blade pairs 53 on a single platform and/or a first blade platform can have forward blade 52 and a second blade platform directly aft of the first blade platform can have aft blade 54, similar to a blade pair 124 described below. Forward stator vane stage 28 includes a plurality of circumferentially disposed stator vanes 64. Each stator vane 64 extends from a vane root 66 to a blade tip 68 along a respective vane axis B and can be operatively connected to a shrouded cavity 70 disposed radially between vane root 66 and rotor disk 50. Knife edge seals 72 are between rotor disk 50 and an inner diameter surface, or platform, 74 of shrouded cavity 70.

As shown in FIG. 3, forward blade 52 extends radially from blade platform 48 to an opposed forward blade tip 56 along a forward blade axis D. Aft blade 54 extends radially from blade platform 48 to an opposed aft blade tip 58 along an aft blade axis E. Aft blade 54 further directs air flow without an intervening stator vane stage shrouded cavity, e.g. a shrouded cavity similar to shrouded cavity 70. A leading edge 60 of aft blade 54 is defined forward of a trailing edge 62 of forward blade 52 with respect to centerline axis A, shown in FIG. 1. Those skilled in the art will readily appreciate that forward blade 52 and aft blade 54 do not need to overlap one another, for example, it is contemplated that leading edge 60 of aft blade 54 can be defined aft of trailing edge 62 of forward blade 52, similar to tandem blade stage 124, described below.

Now with reference to FIG. 4, another embodiment of a gas turbine engine 100 is shown. Gas turbine engine 100 differs from gas turbine engine 10 in that gas turbine engine 100 has a tandem stator vane stage 126 aft of tandem blade stage 124, instead of having an exit guide vane stage, e.g. exit guide vane stage 26. Tandem stator vane stage 126 includes a vane platform 127 radially inward of a compressor case, e.g. compressor case 20, shown in FIG. 1. A stator vane pair 129 extending radially from vane platform 127. Stator vane pair 129 includes a forward stator vane 131 and an aft stator vane 133. Forward stator vane 131 extends radially from the vane platform to an opposed forward stator vane tip 135 along a forward stator vane axis F. Aft stator vane 133 extends radially from vane platform 127 to an opposed aft stator vane tip 137 along an aft stator vane axis G. A leading edge 141 of aft stator vane 133 does not axially overlap a trailing edge 139 of forward stator vane 131. However, those skilled in the art will readily appreciate that leading edge 141 of aft stator vane 133 can be defined forward of trailing edge 139 of forward stator vane 131, similar to tandem blade stage 24, described above. Tandem stator vane stage 126 defines the end of compressor section 114 and tandem blade stage 124 and the tandem stator vane stage 126 define the last two sequential stages in compressor section 114.

With continued reference to FIG. 4, gas turbine engine 100 also differs from gas turbine engine 10 in that a trailing edge 162 of forward blade 152 does not overlap a leading edge 160 of aft blade 154. Further, instead of a single blade platform, e.g. blade platform 48, each respective blade pair 124 includes a respective blade platform 148 for each of blades 152 and 154. Those skilled in the art will readily appreciate that a similar platform configuration can be utilized for tandem stator stage 126. It is also contemplated that that leading edge 160 of aft blade 154 can be defined forward of trailing edge 162 of forward blade 152, similar to tandem blade stage 24, described above.

The methods and systems of the present disclosure, as described above and shown in the drawings, provide for gas turbine engines with superior properties including improved control over fluid flow properties through the engine and reduced windage heat up.

As mentioned above, designers would like to increase the temperature at this last compressor rotor stage 24. However, with this is a corresponding rise in the stress level and, in particular, the stress levels seen by a disk 50 of the compressor rotor. During high power operation, such as takeoff, the outer portions of the compressor rotor, such as the blades, will increase in temperature more rapidly than will the disk 50. This results in very high stresses, which require the compressor rotor to be able to withstand such stresses. Designers would like to further increase these temperatures.

Generally, two avenues exist for allowing the increase in these temperatures. One would be to increase the material properties of the compressor rotor and, in particular, the disk. The other is to improve cooling. As mentioned above, the disclosed arrangement reduces the cooling load. This application introduces additional cooling features which allow the designer to reach even higher temperatures at blade stage 24.

Returning to FIG. 2, a tangential onboard injector (TOBI) 202 is positioned radially inwardly of the vane stage 26. As shown schematically in the circle, the injector 202 has vanes which extend at an angle which is non-parallel to a center axis A of the engine. In a TOBI, small vanes spray onto an outer portion of a rotating disk rim. Thus, air is sprayed on a tangent to a disk of stage 24.

The air may be taken from a compressor midstream source 200 or from a high pressure source 201, such as air compressed by stage 26. Both airflows may be cooled in a heat exchanger 203. In one embodiment, at least 10% of the cooling air passing through the TOBI is compressor midstream air. The air downstream of the TOBI reaches a chamber 205 and is directed to cool rotor disk 50. Seals 204 and 206, such as honeycomb seals, may seal between the blade stage 24 and the vane stage 26 to direct the air from the TOBI 202 radially inwardly into a channel 210 formed between a plate 208 and a disk portion 212. In other embodiments, a knife edged seal may be provided on one of the blade stage 24 and vane stage 26 with a honeycomb abradable seal on the other. Alternatively, a brush seal could be provided on vane stage 26 sealing against a cylindrical surface on the blade stage 24. Any number of other seal embodiments could be utilized within the scope of this disclosure.

As illustrated, the plate may have an axially extending inner portion 213, which extends along an axially extending inner end 214 of the disk 212. An intermediate seal 215 may be utilized at this location to control the airflow. The plate 208 and portion 213 ensure that the air from the TOBI tends to adhere to the disk 50 and, in particular, inner end portions 212 and 214 are providing cooling air.

The plate 208 may be called a static structure radially inward of last blade stage 26 and includes a plate extending at an angle which is non-perpendicular relative to a center axis of said engine and having a component extending radially inwardly. The disk hub also has a portion 212 extending in a direction which is non-perpendicular to the central axis and having a component which is radially inwardly. Plate 208 causes the air having passed through said TOBIs to be confined to a channel 210 adjacent to hub portion 212.

The plate and the hub have a radially inner portion 213/214 which extend closer to being parallel to the central axis than radially outer portions.

In addition, a mini disk 216 may be provided to support a middle section of the hub of the disk 50. The mini disk 216 assists in the “fight” between the outer area 50 of the disk and the radially inner areas, which will tend to heat up less rapidly than the outer areas 50. Air passages 218 are formed between blades 52 and 54 and disk 50 to provide cooling air such as to the seals 72 and platform 74.

Similar structure is shown in the FIG. 4 embodiment.

Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the true scope and content of this disclosure.

Claims

1. A gas turbine engine comprising: a compressor section having a downstream most blade stage and a downstream vane row positioned downstream of said downstream most blade stage, said downstream most blade stage including a plurality of blade pairs with the majority of blade pairs being circumferentially spaced apart from others of said blade pairs, each said blade pair being operatively connected to a rotor disk disposed radially inward from said blade pairs with each said blade pair including a forward blade and an aft blade, said aft blades being configured to further condition airflow with respect to said forward blade; anda tangential onboard injector (TOBI) being positioned radially inwardly of said downstream vane row.

2. The gas turbine engine as set forth in claim 1, wherein there is no intervening stator vane stage between said forward blades and said aft blades.

3. The gas turbine engine as set forth in claim 2, wherein at least 10% of the cooling air passing through said TOBI is compressor midstream air.

4. The gas turbine engine as set forth in claim 3, wherein said compressor midstream air is cooled in a heat exchanger before passing into said TOBI.

5. The gas turbine engine as set forth in claim 2, wherein air passing through said TOBI is air having been compressed by downstream most blade stage of said compressor.

6. The gas turbine engine as set forth in claim 5, wherein said air passes through a heat exchanger before passing through said TOBI.

7. The gas turbine engine as set forth in claim 2, wherein a seal is provided between a downstream end of said downstream most blade stage and an upstream end of said downstream vane row, and said seal defining a cavity extending radially inwardly along a hub of said disk.

8. The gas turbine engine as set forth in claim 7, wherein said seal is provided on both said downstream end of said downstream most blade stage, and said upstream end of said downstream vane row.

9. The gas turbine engine as set forth in claim 7, wherein a static structure radially inward of said downstream vane row includes a plate extending at an angle which is non-perpendicular relative to a center axis of said engine and having a component extending radially inwardly and said disk hub also having a portion extending in a direction which is non-perpendicular to said central axis and having a component which is radially inwardly, said plate causing the air having passed through said TOBI to be confined adjacent to said disk hub portion.

10. The gas turbine engine as set forth in claim 9, wherein said plate and said disk hub portion each having a radially inner portion which extends closer to being parallel to said central axis than does a radially outer portion.

11. The gas turbine engine as set forth in claim 10, wherein a mini disk is positioned radially inwardly of said radially inner portion of said hub.

12. The gas turbine engine as set forth in claim 11, wherein said downstream most blade stage has air passages between said blades and said rotor disk such that air from said cavity can pass upstream toward a vane row upstream of said downstream most blade stage.

13. The gas turbine engine as set forth in claim 1, wherein cooling air passing through said TOBI is compressor midstream air.

14. The gas turbine engine as set forth in claim 1, wherein air passing through said TOBI is air having been compressed by said downstream most blade stage of said compressor.

15. The gas turbine engine as set forth in claim 1, wherein a seal is provided between a downstream end of said downstream most blade stage and an upstream end of said downstream vane row, and said seal defining a cavity extending radially inwardly along a hub of said disk.

16. The gas turbine engine as set forth in claim 15, wherein said seal is provided on both said downstream end of said downstream most blade stage, and said upstream end of said downstream vane row.

17. The gas turbine engine as set forth in claim 1, wherein a static structure radially inward of said downstream vane row includes a plate extending at an angle which is non-perpendicular relative to a center axis of said engine and having a component extending radially inwardly and said disk hub also having a portion extending in a direction which is non-perpendicular to said central axis and having a component which is radially inwardly, said plate causing the air having passed through said TOBI to be confined adjacent to said disk hub portion.

18. The gas turbine engine as set forth in claim 1, wherein a mini disk is positioned radially inwardly of said radially inner portion of said hub.

19. The gas turbine engine as set forth in claim 1, wherein said downstream most blade stage has air passages between said blades and said rotor disk such that cavity air from said cavity can pass upstream toward a vane row upstream of said downstream most blade stage.

20. A gas turbine engine, comprising: a compressor section having a rotor disk defined between a compressor case and a centerline axis; and