Title:
Bullnose seal turbine stage
Kind Code:
A1


Abstract:
A turbine stage includes a stator nozzle cooperating with a row of turbine blades. The nozzle includes a row of vanes mounted between inner and outer bands, with the inner band adjoining the blades at inner platforms thereof at a rotary seal. The inner band includes a stator bullnose at the trailing edge thereof cooperating with the rotary seal for reducing aerodynamic losses thereat.



Inventors:
Klasing, Kevin Samuel (Springboro, OH, US)
Lee, Ching-pang (Cincinnati, OH, US)
Hunter, Scott David (Fairfield, OH, US)
Application Number:
11/642002
Publication Date:
06/19/2008
Filing Date:
12/19/2006
Assignee:
General Electric Company
Primary Class:
Other Classes:
415/191, 415/208.2
International Classes:
F04D29/08; F01D9/02
View Patent Images:



Primary Examiner:
VERDIER, CHRISTOPHER M
Attorney, Agent or Firm:
GENERAL ELECTRIC COMPANY (Huntersville, NC, US)
Claims:
Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims in which we claim:

1. A turbine stage comprising: an annular stator nozzle including a row of vanes mounted between inner and outer bands, and said inner band includes a radially inwardly extending mounting flange; a row of blades mounted to the perimeter of a rotor disk and spaced aft from said mounting flange to define a forward cavity for channeling purge air therethrough; each blade including an inner platform adjoining said inner band at a rotary seal therewith disposed atop said forward cavity for discharging said purge air; and said inner band having a stator bullnose at the trailing edge thereof cooperating with said rotary seal for reducing aerodynamic losses thereat.

2. A stage according to claim 1 wherein said blade platform includes a rotor bullnose joined to a blunt forward face extending radially inwardly therefrom to a rotor wing extending axially forwardly from said blunt face and spaced below said stator bullnose to define said rotary seal therewith.

3. A stage according to claim 2 wherein said stator bullnose axially overlaps said rotor wing and converges inwardly toward said rotor bullnose and said blunt face.

4. A stage according to claim 3 wherein said stator bullnose is convex radially inwardly past said rotor bullnose and terminates above said rotor wing at about mid-elevation of said blunt face.

5. A stage according to claim 4 wherein said inner band and platform are coextensive in elevation at said rotary seal, and converge radially inwardly between said stator and rotor bullnoses.

6. A stage according to claim 5 wherein: said stator bullnose commences at the trailing edges of said vanes, and extends aft toward said rotor bullnose; and said rotor bullnose commences forward from a leading edge of said blade and terminates in a radially aligned blunt face.

7. A stage according to claim 6 wherein said stator and rotor bullnoses are fully convex in radial elevation.

8. A stage according to claim 6 wherein said stator bullnose has a larger radius of curvature than said rotor bullnose.

9. A stage according to claim 6 wherein said rotary seal further comprises a stator wing extending aft from said mounting flange below said rotor wing.

10. A stage according to claim 6 wherein said inner band further includes a lower surface under said stator bullnose disposed generally parallel to and axially overlapping said rotor wing.

11. A turbine stage comprising: a stator nozzle including a row of vanes mounted between inner and outer bands; a row of blades mounted to the perimeter of a rotor disk, and each blade having an inner platform adjoining said inner band at a rotary seal therewith; and said inner band has a substantially full height bullnose at the trailing edge thereof cooperating with said rotary seal for reducing aerodynamic losses thereat.

12. A stage according to claim 11 wherein said blade platform includes a blunt forward face extending radially inwardly to a rotor wing extending axially forwardly therefrom and spaced below said bullnose to define said rotary seal therewith.

13. A stage according to claim 12 wherein said bullnose axially overlaps said rotor wing and converges inwardly toward said blunt face.

14. A stage according to claim 13 wherein said inner band further includes a lower surface under said bullnose disposed generally parallel to and axially overlapping said rotor wing.

15. A stage according to claim 14 wherein said bullnose is convex from the trailing edges of said vanes to said lower surface.

16. A stage according to claim 13 wherein said inner band and platform are coextensive in elevation at said rotary seal, and converge radially inwardly between said bullnose and blunt face.

17. A stage according to claim 16 wherein: said stator bullnose commences at the trailing edges of said vanes, and extends aft toward said platform; and said platform extends forward from a leading edge of said blade and joins said blunt face at a rotor bullnose.

18. A stage according to claim 17 wherein said stator bullnose has a larger radius of curvature than said rotor bullnose.

19. A stage according to claim 13 wherein said inner band further comprises a mounting flange, and said rotary seal further comprises a stator wing extending aft from said mounting flange below said rotor wing.

20. A stage according to claim 19 further comprising a forward cavity disposed between said nozzle and rotor disk in flow communication with said rotary seal for discharging purge air therethrough.

Description:

BACKGROUND OF THE INVENTION

The present invention relates generally to gas turbine engines, and, more specifically, to turbine efficiency therein.

In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the combustion gases in turbine stages which power the compressor through one drive shaft, and produce additional work for powering an upstream fan in a turbofan aircraft engine application, or driving an external drive shaft for marine and industrial (M&I) applications.

The basic core engine typically includes a multistage axial compressor having rows of compressor blades and corresponding guide vanes which pressurize ambient air in stages and correspondingly increase the temperature thereof. The air discharged from the aft end of the compressor has the highest pressure, commonly referred to as compressor discharge pressure (CDP), and a correspondingly high temperature.

In an exemplary configuration, the compressor may have seven stages for increasing air pressure many times atmospheric pressure along with many hundreds of degrees of temperature increase due to the compression cycle. A fewer or greater number of compression stages may be used as desired for the specific design of the gas turbine engine and its intended use.

A majority of the CDP air discharged from the compressor is mixed with fuel in the combustor for generating hot combustion gases. These combustion gases then undergo an expansion cycle in the several turbine stages for extracting energy therefrom which correspondingly reduces the pressure of the combustion gases and the temperature thereof. A high pressure turbine (HPT) immediately follows the combustor and is used to power the compressor blades in the core engine.

A low pressure turbine (LPT) follows the HPT and drives the second shaft for powering the upstream fan in the turbofan engine application, or driving an external drive shaft for M&I applications.

The overall efficiency of the gas turbine engine is dependent on the efficiency of air compression, efficiency of combustion, and efficiency of combustion gas expansion in the turbine stages.

Each turbine stage typically includes an upstream turbine nozzle or stator having a row of nozzle vanes which direct the combustion gases downstream through a corresponding row of turbine rotor blades. The blades are typically mounted to the perimeter of a supporting rotor disk in corresponding dovetail slots formed therein.

The turbine blades and vanes are typically hollow airfoils with corresponding internal cooling channels therein which receive compressor discharge air for cooling thereof during operation. The hollow blades and vanes typically include various rows of film cooling and other discharge holes through the pressure and suction sidewalls thereof for discharging the spent internal cooling air in corresponding external films for further protecting the airfoils.

The main turbine flowpath is designed to confine the combustion gases as they flow through the engine and decrease in temperature and pressure from the combustor. The various cooling circuits for the turbine components are independent from the main flowpath and must be provided with cooling air at sufficient pressure to prevent ingestion of the hot combustion gases therein during operation.

For example, suitable rotary seals are provided between the stationary turbine nozzles and the rotating turbine blades to prevent ingestion or backflow of the hot combustion gases into the cooling circuits.

Since the airfoils of the nozzle vanes and turbine blades typically include rows of cooling air outlet holes, the cooling air must also have sufficient pressure greater than that of the external combustion gases to provide a suitable backflow margin to prevent ingestion of the hot combustion gases into the turbine airfoils themselves.

Since the combustion gases and cooling air are channeled through corresponding flowpaths or flow circuits in the engine, they are subject to various aerodynamic losses which further decrease engine efficiency. Fluid flow is subject to friction or drag losses, flow separation losses, and mixing losses all of which reduce pressure and decrease efficiency.

The rotary seal between the first stage turbine nozzle and first stage turbine rotor blades is at one critical site which significantly affects turbine efficiency. The nozzle including its inner band is a stationary or stator component immediately followed downstream by the rotating turbine blades and their corresponding rotating platforms which form the radially inner flowpath boundary for the hot combustion gases being channeled through the first stage turbine.

The combustion gases discharged from the turbine nozzle must necessarily flow over the axial gap and rotor seal therebetween to reach the turbine blades. The rotary seal cooperates with the internal pressurized purge air being channeled through the axial gap to prevent backflow of the hot combustion gases into the purge cooling circuit.

Accordingly, the smooth flow of the hot combustion gases is interrupted at the inner band-platform axial gap, and the purge air mixes with the combustion gases at this site and further degrades smooth flow.

Flow separation of the combustion gases occurs at this rotary seal and the combustion gases mix with the different velocity purge air for collectively further reducing efficiency.

Efficiency losses at this location decrease the total pressure of the combustion gases available at the turbine blades, which in turn corresponding reduces turbine efficiency. And, the disrupted flow at the beginning of the blade platforms can lead to increased heating thereof.

Accordingly, it is desired to provide a turbine stage having an improved rotary seal between the stator nozzle and rotor blades for improving turbine efficiency.

BRIEF DESCRIPTION OF THE INVENTION

A turbine stage includes a stator nozzle cooperating with a row of turbine blades. The nozzle includes a row of vanes mounted between inner and outer bands, with the inner band adjoining the blades at inner platforms thereof at a rotary seal. The inner band includes a stator bullnose at the trailing edge thereof cooperating with the rotary seal for reducing aerodynamic losses thereat.

BRIEF DESCRIPTION OF THE DRAWINGS

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:

FIG. 1 is a partly sectional, axial schematic view of a turbofan gas turbine engine.

FIG. 2 is an enlarged, axial sectional view of the high pressure turbine illustrated in FIG. 1.

FIG. 3 is a further enlarged, axial sectional view of the rotary seal between the stator nozzle and rotor blades shown in FIG. 3.

DETAILED DESCRIPTION OF THE INVENTION

Illustrated schematically in FIG. 1 is an exemplary turbofan aircraft gas turbine engine 10. The engine is axisymmetrical about a longitudinal or axial centerline axis 12 and is suitably mounted to the wing or a fuselage of an aircraft (not shown) for powering an aircraft in flight in an exemplary application.

The engine includes in serial flow communication a fan 14, a low pressure or booster compressor 16, a high pressure (HP) compressor 18, an annular combustor 20, a high pressure turbine (HPT) 22, and a low pressure turbine (LPT) 24.

An annular nacelle 26 surrounds the fan 14 and defines an annular bypass duct 28 extending aft around the booster compressor 16. A first drive shaft 30 joins the HPT 22 to the HP compressor 18, and a second drive shaft 32 joins the LPT 24 to the fan 14 and booster compressor 16. The two drive shafts are suitably mounted in bearings in corresponding frames within the engine in a conventional configuration of the various engine components described above.

During operation, ambient air 34 enters the inlet of the engine and is pressurized in part by the fan 14 and discharged through the bypass duct 28 for providing a majority of propulsion thrust. Some of the air 34 passing the fan enters the booster compressor 16 and undergoes a further compression cycle in the multiple axial stages thereof, with additional compression also being provided in the HP compressor 18 in the multiple axial stages thereof.

The pressurized air 34 is discharged from the compressor and suitably mixed with fuel 36 in the combustor 20 for generating hot combustion gases 38. Energy is extracted from the combustion gases 38 in the HPT 22 to drive the first shaft 30 and power the HP compressor 18. Additional energy is extracted from the combustion gases in the LPT 24 to drive the second shaft 32 and power the fan 14 and booster compressor 16.

The engine as described above is conventional in configuration and operation and includes multiple compression stages and multiple turbine stages. For example, the booster compressor 16 may have four axial stages including four rows of compressor blades alternating axially with four rows of inlet guide vanes.

The high pressure compressor 18 may include seven axial stages for example, having seven rows of compressor blades alternating axially with corresponding rows of inlet guide vanes, and discharging the CDP air 34 through a conventional diffuser.

The HPT 22 is a single stage turbine followed in turn by an exemplary five stage LPT 24.

FIG. 2 illustrates in more detail the HPT 22 disposed in serial flow communication between the outlet of the combustor 20 and the inlet of the LPT 24.

The HPT 22 includes a first stage or HP turbine nozzle 40 having a row of stator vanes 42 suitably mounted in inner and outer bands 44,46.

Typically, two nozzle vanes 42 are integrally joined together in corresponding arcuate segments of the bands in a nozzle doublet in a unitary configuration. And, a full row of the doublets completes the annular nozzle.

The inner band 44 typically includes a radially inwardly extending mounting flange 48 suitably fixedly attached to a corresponding flange on a conical nozzle support 50, by a row of fastening bolts for example.

Following the vanes is a single row of HP turbine blades 52 removably mounted to the perimeter or rim of a first stage or HP rotor disk 54. The disk 54 is fixedly joined to the first drive shaft 30 which in turn is fixedly joined to the rotor disks supporting the compressor blades of the high pressure compressor 18.

Each nozzle vane 42 has the typical airfoil shape with generally concave pressure side and generally convex opposite suction side extending axially in chord between the upstream leading edge and the downstream trailing edge. Correspondingly, the airfoil portion of the first stage turbine blades 52 has the generally concave pressure side and generally convex opposite suction side extending axially between the leading and trailing edges thereof.

The vanes 42 and blades 52 may have any suitable internal cooling configuration. FIG. 2 illustrates forward and aft cooling cavities for the vanes. The blades have a pair of three-pass serpentine cooling channels over the midchord region, with corresponding inlets through the dovetail. And, the blade leading and trailing edges have dedicated cooling channels therefor.

In this way, pressurized air is bled from the compressor and channeled through the several internal cooling circuits of the vanes 42 and blade 52 for providing internal cooling thereof in any conventional manner, with the spent air then being discharged through various rows of outlet holes found in the pressure and suction sides of the airfoils from the leading edge to the trailing edge.

Since the turbine nozzle 40 is a stationary component and the turbine blades 52 and supporting disk 54 rotate during operation, an annular forward plenum or cavity 56 is disposed axially therebetween for receiving a portion of the pressurized CDP air 34 for use in cooling and purging this stator-rotor interface or junction.

Each turbine blade airfoil extends radially in span from a radially inner platform 58 to a radially outer tip spaced closely adjacent to a surrounding turbine shroud. The blade platform is integrally joined to a conventional axial-entry mounting dovetail retained in a complementary dovetail slot in the perimeter of the rotor disk 54.

The annular nozzle inner band 44 illustrated in FIG. 2 terminates aft in radial position or elevation with the forward end of the row of blade platforms 58 extending circumferentially together in another full annulus. The inner band and blade platforms are therefore axially coextensive at similar radial elevation to provide a substantially continuous inner flowpath or boundary for the hot combustion gases 38 being channeled through the main flowpath between the nozzle vanes and turbine blades.

The blade platforms 58 closely adjoin the inner band 44 at a small axial gap having a rotary seal 60 extending therebetween to control the discharge of the internal purge air 34 from the forward cavity 56 into the main turbine flowpath, and thereby prevent backflow or ingestion of the hot combustion gases radially inwardly into the forward cavity. The purge air has sufficient pressure locally greater than the pressure of the combustion gases to ensure discharge or outflow of the purge air through the rotary seal 60 and into the main flowpath along the junction of the nozzle inner band and the blade platforms.

As indicated above, this axial gap between the stator nozzle 40 and the rotor blades 52 provides a local interruption in the otherwise smooth inner flowpath for the hot combustion gases. Furthermore, the purge air 34 being discharged from the rotary seal necessarily mixes with the local combustion gases and provides additional aerodynamic losses therewith.

Accordingly, the inner band 44 illustrated in more detail in FIG. 3 is modified to include a substantially full radial height stator bullnose 62 at the aft or trailing edge thereof which cooperates with the rotary seal 60 for significantly reducing aerodynamic losses thereat.

Correspondingly, each blade platform 58 includes a rotor bullnose 64 at the leading edge or forward edge thereof which in turn is joined to a flat or blunt forward vertical face 66 extending radially inwardly therefrom.

The stator and rotor bullnoses 62,64 are convex outwardly in axial profile and are uniform circumferentially around the inner band and blade platforms. Both bullnoses 62,64 extend radially inwardly and converge together, with the blunt forward face 66 being disposed substantially vertically or radially without inclination in the exemplary embodiment.

The platform forward face 66 extends radially inwardly to integrally join a seal or rotor wing 68 extending axially forwardly therefrom and circumferentially along the full extent of each blade platform. The seal wing 68, commonly referred to as an angel wing, is spaced below the stator bullnose 62 to define the rotary seal 60 therewith, commonly also referred to as a labyrinth seal.

In these rotary seals, the cooperating rotary and stator components do not contact each other but are spaced closely adjacent to each other for imparting a substantial flow restriction against the free flow of pressurized fluid therepast.

Accordingly, as the pressurized purge air 34 flows radially upwardly, its free flow is restrained by the rotary seal 60, which nevertheless permits discharge of the purge air through the axial gap defined between the two bullnoses 62,64. At this location, the discharged purge air meets and mixes with the combustion gases 38 flowing downstream through the main turbine flowpath.

The introduction of the relatively large stator bullnose 62 delays the onset of flow separation of the combustion gases as they flow downstream over the rotary seal, which permits the resultant flow separation to occur within the recessed space or pocket provided between the stator bullnose 62 and the forward face 66 of the blade platform. In this protected region, the purge air 34 is discharged radially outwardly to meet and mix with the separated combustion gas flow.

Accordingly, the aerodynamic losses due to flow separation of the combustion gases, mixing of the different velocity combustion gases and discharged purge air, and interruption in the smooth continuity of the combustion gases over the rotary seal are reduced for significantly increasing aerodynamic efficiency of the turbine stage.

Computational fluid dynamics (CFD) analysis predicts a substantial improvement in aerodynamic efficiency of the turbine stage employing the cooperating stator bullnose interface of, for example, up to about a fraction of one percent.

As shown in FIG. 3, the stator bullnose 62 axially aft overlaps the forward or leading edge portion of the rotor wing 68, and converges radially inwardly toward the blunt face 66. The convex bullnose 62 uses the Coanda effect to draw the local combustion gas boundary flow radially inwardly where it meets the radially outwardly flowing purge air 34.

The stator bullnose 62 defines the aft outer surface of the inner band 44 at its trailing edge, and the inner band further includes a lower surface 70 under the aft bullnose 62 which is disposed generally parallel to and axially overlaps the rotor wing 68 for cooperating therewith. The purge air 34 must turn sharply as it flows around the angel wing 68, and the lower surface 70 of the stator bullnose 62 confines the purge flow aft to meet the diverted combustion gases in the recessed pocket between the inner band and blade platform.

The lower surface 70 of the stator bullnose 62 may be in the form of a local step at the trailing edge of the inner band, or it may extend upstream therefrom without a step as illustrated in phantom line in FIG. 3.

The stator bullnose 62 is convex axially and preferably extends radially inwardly past the convex rotor bullnose 64, and terminates above the rotor wing 68 at about mid-elevation in the blunt platform face 66. The stator bullnose 62 is preferably convex from the trailing edges of the nozzle vanes 42 and is fully arcuate and convex in radial elevation to its junction with the lower surface 70 of the inner band.

In this way, the stator bullnose 62 is completely arcuate and convex for the entire trailing edge of the inner band 44 aft from the vane trailing edges, over at least a majority of the thickness of the inner band at the trailing edge, and preferably over the full radial thickness or extent thereof. The convex bullnose 62 thusly provides a radially continuous and uninterrupted smooth aerodynamic surface along which the combustion gases may closely adhere with a smooth boundary layer to delay as long as possible flow separation therefrom.

Discontinuity along the trailing edge region of the inner band should be minimized or avoided to prevent the undesirable early separation of the combustion gas flow in the aft direction over the rotary seal.

The inner band 44 and blade platform 58 illustrated in FIG. 3 are preferably aligned axially and coextensive in radial elevation at the rotary seal 60 for providing an aerodynamically smooth inner boundary for the combustion gases. And, the inner band and platform converge radially inwardly between the stator bullnose 62 and the radial blunt face 66.

In this way, the small pocket formed therebetween is recessed in common height from the inner boundary of combustion flowpath for permitting premixing of the purge air with the separated combustion gases in the pocket, and thereby reduce aerodynamic losses associated therewith.

As indicated above, the stator bullnose 62 preferably commences at the trailing edges of the nozzle vanes 42 and extends aft toward the rotor bullnose 64 at the forward end of the blade platform 58.

In contrast, the rotor bullnose 64 commences forward from or spaced forward from the leading edge of the blade 52, and then terminates radially inwardly in radial alignment with the blunt face 66 before termination of the stator bullnose 62, which terminates at the mid-elevation of the blunt face 66. In this way, the combustion gases are drawn radially inwardly toward the middle of the outlet pocket above the rotor wing 68 where they meet and mix with the discharging purge air.

The configuration of the turbine blades 52 and their platforms 58 are conventional in the exemplary embodiment illustrated in FIG. 3, including a relatively small rotor bullnose 64 providing a sharp corner transition between the axial platform 58 and the radial blunt face 66.

Correspondingly, the stator bullnose 62 is specifically configured for improved aerodynamic performance therewith, and includes a larger radius of curvature than for the rotor bullnose 64.

Since the flow fields of the discharging purge air 34 and the combustion gases 38 in the main flowpath are quite complex through and around the rotary seal 60, the precise configuration of convex stator bullnose 62 must complement the configuration of the forward region of the blade platform 58 to minimize aerodynamic losses for increasing turbine efficiency.

For example, the typical profile of the trailing edge of a nozzle inner band is relatively square, with significant, blunt aft faces. The blunt aft face of the inner band promotes flow separation, and a corresponding decrease in turbine efficiency.

In contrast, the substantially fully convex stator bullnose 62 eliminates, or substantially eliminates, any blunt or straight aft face at which flow separation would be prevalent.

In the exemplary embodiment illustrated in FIG. 3, the rotary seal 60 also includes a stator angel wing 72 disposed radially below the rotor wing 68, and extending axially aft to partially overlap that wing and increase turning in the serpentine flowpath defined by the rotary seal for further restraining free flow of the discharged purge air.

In this way, the purge air is directed radially outwardly from the forward cavity 56, and then follows the circuitous flowpath around the rotor wing 68. The stator wing 72 first causes the purge air to flow axially forward and then radially outwardly and axially aft in reverse direction over the top of the rotor wing 68.

Here the purge air joins the combustion gases drawn radially inwardly by the Coanda effect of the convex stator bullnose 62 for delaying flow separation of the combustion gases and allowing premixing thereof with the purge air being discharged from the rotary seal.

The bullnose rotary seal 60 effected in FIG. 3 improves flow attachment of the combustion gases as they are discharged from the nozzle inner band 44, reduces flow separation thereof, and further reduces mixing losses with the different velocity purge air being discharged from the forward purge cavity 56.

In this interface region between the discharged purge air and diverted combustion gases, the intermixed flow is subject to smaller circulation and less mixing, and permits the cooler purge air to thermally protect the forward portion of the blade platform 58, including the rotor bullnose 64.

This configuration not only increases the total pressure of the combustion gases at the leading edges of the turbine blades 52 for increasing turbine efficiency, but also reduces the temperature of the blade platform along its leading edge. A cooler operating platform reduces thermal stresses therein and may increase the life thereof, or may be used for reducing any cooling air specifically provided for internally cooling the blade platform itself.

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.