WO2015094990A1 - Adjustable clearance control system for airfoil tip in gas turbine engine - Google Patents

Adjustable clearance control system for airfoil tip in gas turbine engine Download PDF

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Publication number
WO2015094990A1
WO2015094990A1 PCT/US2014/070190 US2014070190W WO2015094990A1 WO 2015094990 A1 WO2015094990 A1 WO 2015094990A1 US 2014070190 W US2014070190 W US 2014070190W WO 2015094990 A1 WO2015094990 A1 WO 2015094990A1
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WO
WIPO (PCT)
Prior art keywords
airfoil
ring segment
adjustable ring
axially
axially adjustable
Prior art date
Application number
PCT/US2014/070190
Other languages
French (fr)
Inventor
John A. Orosa
Original Assignee
Siemens Aktiengesellschaft
Siemens Energy, Inc.
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Siemens Aktiengesellschaft, Siemens Energy, Inc. filed Critical Siemens Aktiengesellschaft
Publication of WO2015094990A1 publication Critical patent/WO2015094990A1/en

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/20—Actively adjusting tip-clearance
    • F01D11/22—Actively adjusting tip-clearance by mechanically actuating the stator or rotor components, e.g. moving shroud sections relative to the rotor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F05D2220/321—Application in turbines in gas turbines for a special turbine stage
    • F05D2220/3215—Application in turbines in gas turbines for a special turbine stage the last stage of the turbine
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/55—Seals
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/28—Three-dimensional patterned
    • F05D2250/283—Three-dimensional patterned honeycomb

Definitions

  • the components After the turbine engine has been operating at full load conditions for a period of time, the components reach a maximum operating condition at which maximum thermal expansion occurs. In this state, it is desirable that the gap between the blade tips and the casing of the turbine engine and the gap between the vanes and rotatable blade and disc assembly be as small as possible to limit leakage past the tips of the airfoils.
  • An airfoil system for use in a gas turbine engine having an adjustable clearance control system including an axially adjustable ring segment releasably coupled to the stationary turbine component whereby the axially adjustable ring segment may be controlled independently of other airfoil stages is disclosed.
  • the adjustable clearance control system may thus control the flow of hot gases passing one particular airfoil stage while the flow passing other airfoil stages within the component of the turbine engine remains unchanged.
  • the adjustable clearance control system may control the size of the gap between the axially adjustable ring segment and the tip of an airfoil through axial movement of the axially adjustable ring segment.
  • the axially adjustable ring segment may include a radially inward contact surface that is positioned nonparallel and nonorthogonal relative to a direction of movement of the axial adjustable ring segment to adjust the size of the gap.
  • the adjustable clearance control system may address the need to open the clearance on the last blade tip without opening the clearance on the other blades.
  • turbine engines are operated with the tightest possible clearance on every blade.
  • diffuser performance can be significantly improved by increasing the last blade tip leakage jet. Even though opening the gap reduces the blade efficiency, that loss can be more than offset by the improvement in the downstream diffuser.
  • One particular situation in which diffuser performance can be significantly improved is at baseload, which is maximum power standard day, if the turbine has a hub strong velocity profile. This situation occurs when turbine engines are operated at higher power levels, which can occur by increasing mass flow, without changing the turbine design. This situation can also occur in well designed turbines having a flat velocity profile when they operate on cold days.
  • the airfoil system may include a rotor assembly having a plurality of airfoils extending radially therefrom and aligned axially to form a circumferentially extending row of airfoils forming an airfoil stage.
  • the airfoil system may also include a stationary turbine component positioned radially outward from a tip of the airfoil.
  • the airfoil system may include an adjustable clearance control system including an axially adjustable ring segment releasably coupled to the stationary turbine component, wherein the axially adjustable ring segment is adjustable axially.
  • the axially adjustable ring segment may be adjustable axially when positioned radially outward from a single row of turbine blades, thereby enabling the axially adjustable ring segment to be moved axially to change a gap size between the tip of the airfoil and the axially adjustable ring segment independently of other airfoil stages.
  • the axially adjustable ring segment may include a radially inward contact surface having at least a portion aligned with the tip of the airfoil.
  • the axially adjustable ring segment may include a honeycomb seal land coupled to the axially adjustable ring segment.
  • a radially inward contact surface of the honeycomb seal land may be positioned nonparallel and nonorthogonal relative to a direction of movement of the axial adjustable ring segment.
  • One or more ribs may extend radially outward from the tip of the turbine blade and may terminate before
  • the adjustable clearance control system may enable the axially adjustable ring segment to be moved axially such that the gap may be reduced or increased. For instance, during startup of the gas turbine engine, the adjustable clearance control system may move the axially adjustable ring segment axially such that the gap may be increased to prevent tip rubbing and possible damage. Once at steady state operating conditions, the adjustable clearance control system may move the axially adjustable ring segment to reduce the gap
  • Increasing the last stage blade tip gap can inject high velocity air at the OD to keep the OD flow healthy in the exhaust diffuser.
  • an advantage of the adjustable clearance control system is that the system may control the size of a gap between airfoil tips and a radially outward ring segment by adjusting the axially adjustable ring segment axially independently of other airfoil stages.
  • the size of the gap between the airfoil tips and a radially outward ring segment may be adjusted without adjusting or interfering with gap sizes between airfoil tips and ring segments of other airfoil stages within the same gas turbine engine.
  • the system may include a honeycomb seal land having a radially inward contact surface configured to absorb contact from airfoil tips without damaging the airfoils.
  • Figure 3 is a perspective view of a honeycomb seal land coupled to the axially adjustable ring segment.
  • Figure 4 is a perspective view of a shrouded turbine blade usable with the adjustable clearance control system.
  • the airfoil system 10 may include a rotor assembly 32 having a plurality of airfoils 26 extending radially therefrom and aligned axially to form a circumferentially extending row 34 of airfoils 26 forming an airfoil stage 20.
  • the airfoils 26 may have any appropriate shape or configuration.
  • the airfoil system 10 may also include a stationary turbine component 36 positioned radially outward from the tip 24 of the airfoil 26.
  • the stationary component 36 may be a turbine ring segment.
  • the stationary component 36 may be other components that remain stationary relative to the rotor assembly 32.
  • the axially adjustable ring segment 16 may include a radially inward contact surface 28 having at least a portion aligned with the tip 24 of the airfoil 26. In other embodiments, the radially inward contact surface 28 may be misaligned with the tip 24 of the airfoil 26.
  • the axially adjustable ring segment 16 may include a honeycomb seal land 40 coupled to the axially adjustable ring segment 16.
  • the honeycomb seal land 40 may be formed from a plurality of hollow cavities 42 with walls 44 taking the shape of a honeycomb shape.
  • the honeycomb shaped walls 44 may have any appropriate shape.
  • the honeycomb seal land 40 may be formed from any appropriate material capable of withstanding the
  • a radially inward contact surface 28 of the honeycomb seal land 40 may be positioned nonparallel and nonorthogonal relative to the direction of movement 30 of the axial adjustable ring segment 16.
  • the adjustable clearance control system 14 may include one or more ribs 46 extending radially outward from the tip 24 of the turbine blade 38 and terminating before contacting a radially inward contact surface 28 of the axially adjustable ring segment 16.
  • the rib 46 may be formed from a tip shroud, as shown in Figure 4.
  • the rib 46 may have any appropriate configuration and shape.
  • the rib 46 may also may be formed from any appropriate material capable of withstanding the environment within a hot gas path in a gas turbine engine 12.
  • the adjustable clearance control system 14 may enable the axially adjustable ring segment 16 to be moved axially such that the gap 22 may be reduced or increased.
  • the adjustable clearance control system 14 may move the axially adjustable ring segment 16 to increase the gap 22 independently of other stages 18 to increase tip jet flow when diffuser inlet conditions cause high OD loading.
  • High OD loading occurs where a hub strong velocity profile entrains or pulls flow away from the OD and toward the hub. This causes more diffusion to occur near the OD wall. Hence higher loading or diffusion which can lead to flow
  • the adjustable clearance control system 14 may move the axially adjustable ring segment 16 axially such that the gap 22 may be increased to prevent tip rubbing and possible damage. Once at steady state operating conditions, the adjustable clearance control system 14 may move the axially adjustable ring segment 16 to reduce the gap 22 independently of other stages 18. During shutdown of the gas turbine engine 12, the adjustable clearance control system 14 may move the axially adjustable ring segment 16 to increase the gap 22 independently of other stages 18 to prevent tip rubbing and possible damage. In addition, increasing the clearance of the last stage blade 20 on cold days or on turbine engines 12 that have a hub strong exit velocity profile may improve engine performance.
  • the exhaust diffuser 48 may be overloaded on the outer diamater (OD) due to higher velocities on the hub pulling flow away. Increasing the last stage blade 20 tip gap 22 can inject high velocity air at the OD to keep the OD flow healthy in the exhaust diffuser 48.
  • the adjustable clearance control system 14 may address the need to open the clearance on the last blade tip 24 without opening the clearance on the other blades.
  • turbine engines 12 are operated with the tightest possible clearance on every blade.
  • diffuser performance can be significantly improved by increasing the last blade tip leakage jet. Even though opening the gap reduces the blade efficiency, that loss can be more than offset by the improvement in the downstream diffuser.
  • One particular situation in which diffuser performance can be significantly improved is at baseload, which is maximum power standard day, if the turbine has a hub strong velocity profile. This situation occurs when turbine engines 12 are operated at higher power levels, which can occur by increasing mass flow, without changing the turbine design. This situation can also occur in well designed turbines having a flat velocity profile when they operate on cold days.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

An airfoil system (10) for use in a gas turbine engine (12) having an adjustable clearance control system (14) including an axially adjustable ring segment (16) releasably coupled to the stationary turbine component (36) whereby the axially adjustable ring segment (16) may be controlled independently of other airfoil stages (18) is disclosed. The adjustable clearance control system (14) may thus control the flow of hot gases passing one particular airfoil stage (20) while the flow passing other airfoil stages (18) within the component of the turbine engine (12) remains unchanged. The adjustable clearance control system (14) may control the size of the gap (22) between the axially adjustable ring segment (16) and the tip (24) of an airfoil (26) through axial movement of the axially adjustable ring segment (16). The axially adjustable ring segment (16) may include a radially inward contact surface (28) that is positioned nonparallel and nonorthogonal relative to a direction of movement of the axial adjustable ring segment (16).

Description

ADJUSTABLE CLEARANCE CONTROL SYSTEM FOR
AIRFOIL TIP IN GAS TURBINE ENGINE
FIELD OF THE INVENTION
This invention is directed generally to turbine engines, and more particularly to systems for controlling gaps between airfoil tips and radially adjacent components in turbine engines so as to control the flow of fluids downstream based upon diffuser inlet conditions.
BACKGROUND
Typically, gas turbine engines are formed from a compressor positioned upstream from a turbine blade assembly. The compressor and turbine are formed from a plurality of blade stages coupled to discs that are capable of rotating about a longitudinal axis. Each blade stage is formed from a plurality of blades extending radially about the circumference of the disc.
The tips of the blades are located in close proximity to an inner surface of the casing of the turbine engine. There typically exists a gap between the blade tips and the casing of the turbine engine so that the blades may rotate without striking the casing. Likewise, for nonshrouded vanes, there typically exists a gap between the vane tips and an internal rotatable blade and disc assembly so that the rotatable blade and disc assembly may rotate without the vanes contacting the rotatable blade and disc assembly. During operation, gases pass the blades and vanes and compress to high temperature and pressure. These gases also heat the casing, blades, vanes and discs causing each to expand due to thermal expansion. After the turbine engine has been operating at full load conditions for a period of time, the components reach a maximum operating condition at which maximum thermal expansion occurs. In this state, it is desirable that the gap between the blade tips and the casing of the turbine engine and the gap between the vanes and rotatable blade and disc assembly be as small as possible to limit leakage past the tips of the airfoils.
However, reducing the gap cannot be accomplished by simply positioning the components so that the gap is minimal under full load conditions because the configuration of the components forming the gap must account for warm restart conditions in which the casing and the compressor vane carriers, having less mass than the blade and disc assembly, cools faster than the blade and disc assembly. During a warm restart, the discs expand due to centrifugal forces and the clearances tighten before the casing begins to heat up and expand. Therefore, unless the components have been positioned so that a sufficient gap has been established between the blades and the casing and between the vanes and the rotatable blade and disc assembly under operating conditions, the airfoils may strike the casing or the rotatable blade and disc assembly because the diameter of components forming the casing have not heated up and expanded yet. Collision between the blades and the casing or compressor vanes and the rotatable blade and disc assembly often causes severe airfoil tip rubs and may result in damage. There exist systems that reducing gaps between every blade tip and a casing simultaneously. However, at certain conditions, diffuser performance can be significantly improved by increasing the last blade tip leakage jet. Thus, a need exists for an improved system for regulating the last blade tip leakage jet.
SUMMARY OF THE INVENTION
An airfoil system for use in a gas turbine engine having an adjustable clearance control system including an axially adjustable ring segment releasably coupled to the stationary turbine component whereby the axially adjustable ring segment may be controlled independently of other airfoil stages is disclosed. The adjustable clearance control system may thus control the flow of hot gases passing one particular airfoil stage while the flow passing other airfoil stages within the component of the turbine engine remains unchanged. The adjustable clearance control system may control the size of the gap between the axially adjustable ring segment and the tip of an airfoil through axial movement of the axially adjustable ring segment. The axially adjustable ring segment may include a radially inward contact surface that is positioned nonparallel and nonorthogonal relative to a direction of movement of the axial adjustable ring segment to adjust the size of the gap.
The adjustable clearance control system may address the need to open the clearance on the last blade tip without opening the clearance on the other blades. Normally, turbine engines are operated with the tightest possible clearance on every blade. However, at certain conditions, diffuser performance can be significantly improved by increasing the last blade tip leakage jet. Even though opening the gap reduces the blade efficiency, that loss can be more than offset by the improvement in the downstream diffuser. One particular situation in which diffuser performance can be significantly improved is at baseload, which is maximum power standard day, if the turbine has a hub strong velocity profile. This situation occurs when turbine engines are operated at higher power levels, which can occur by increasing mass flow, without changing the turbine design. This situation can also occur in well designed turbines having a flat velocity profile when they operate on cold days.
The airfoil system may include a rotor assembly having a plurality of airfoils extending radially therefrom and aligned axially to form a circumferentially extending row of airfoils forming an airfoil stage. The airfoil system may also include a stationary turbine component positioned radially outward from a tip of the airfoil. The airfoil system may include an adjustable clearance control system including an axially adjustable ring segment releasably coupled to the stationary turbine component, wherein the axially adjustable ring segment is adjustable axially. The axially adjustable ring segment may be adjustable axially when positioned radially outward from a single row of turbine blades, thereby enabling the axially adjustable ring segment to be moved axially to change a gap size between the tip of the airfoil and the axially adjustable ring segment independently of other airfoil stages.
The axially adjustable ring segment may include a radially inward contact surface having at least a portion aligned with the tip of the airfoil. The axially adjustable ring segment may include a honeycomb seal land coupled to the axially adjustable ring segment. A radially inward contact surface of the honeycomb seal land may be positioned nonparallel and nonorthogonal relative to a direction of movement of the axial adjustable ring segment. One or more ribs may extend radially outward from the tip of the turbine blade and may terminate before
contacting a radially inward contact surface of the axially adjustable ring segment or of the honeycomb seal land. In at least one embodiment, the rib may be formed from a tip shroud. The plurality of airfoils forming the circumferentially extending row of airfoils forming the airfoil stage may form an airfoil stage closest to an exhaust diffuser. The axially adjustable ring segment of the adjustable clearance control system may extend at least partially circumferentially about the airfoil stage.
During operation, the adjustable clearance control system may enable the axially adjustable ring segment to be moved axially such that the gap may be reduced or increased. For instance, during startup of the gas turbine engine, the adjustable clearance control system may move the axially adjustable ring segment axially such that the gap may be increased to prevent tip rubbing and possible damage. Once at steady state operating conditions, the adjustable clearance control system may move the axially adjustable ring segment to reduce the gap
independently of other stages. During shutdown of the gas turbine engine, the adjustable clearance control system may move the axially adjustable ring segment to increase the gap independently of other stages to prevent tip rubbing and possible damage. In addition, increasing the clearance of the last stage blade on cold days or on turbine engines that have a hub strong exit velocity profile may improve engine performance. Under such conditions, the exhaust diffuser may be overloaded on the outer diamater (OD) due to higher velocities on the hub pulling flow away.
Increasing the last stage blade tip gap can inject high velocity air at the OD to keep the OD flow healthy in the exhaust diffuser.
An advantage of the adjustable clearance control system is that the system may control the size of a gap between airfoil tips and a radially outward ring segment by adjusting the axially adjustable ring segment axially independently of other airfoil stages. Thus, the size of the gap between the airfoil tips and a radially outward ring segment may be adjusted without adjusting or interfering with gap sizes between airfoil tips and ring segments of other airfoil stages within the same gas turbine engine.
Another advantage of the adjustable clearance control system is that the system may include a honeycomb seal land having a radially inward contact surface configured to absorb contact from airfoil tips without damaging the airfoils.
These and other embodiments are described in more detail below. BRIEF DESCRIPTION OF THE DRAWINGS
The accompanying drawings, which are incorporated in and form a part of the specification, illustrate embodiments of the presently disclosed invention and, together with the description, disclose the principles of the invention.
Figure 1 is a cross-sectional perspective view of a turbine engine with an adjustable clearance control system.
Figure 2 is a cross-sectional detail view of the adjustable clearance control system positioned in a gas turbine engine taken at detail 2-2 in Figure 1 .
Figure 3 is a perspective view of a honeycomb seal land coupled to the axially adjustable ring segment.
Figure 4 is a perspective view of a shrouded turbine blade usable with the adjustable clearance control system.
DETAILED DESCRIPTION OF THE INVENTION
As shown in Figures 1 -4, an airfoil system 10 for use in a gas turbine engine 12 having an adjustable clearance control system 14 including an axially adjustable ring segment 16 releasably coupled to the stationary turbine component 36 whereby the axially adjustable ring segment 16 may be controlled independently of other airfoil stages 18 is disclosed. The adjustable clearance control system 14 may thus control the flow of hot gases passing one particular airfoil stage 20 while the flow passing other airfoil stages 18 within the component of the turbine engine 12 remains unchanged. The adjustable clearance control system 14 may control the size of the gap 22 between the axially adjustable ring segment 16 and a tip 24 of an airfoil 26 through axial movement of the axially adjustable ring segment 16. The axially adjustable ring segment 16 may include a radially inward segment contact surface 27 that is positioned nonparallel and nonorthogonal relative to a direction of movement 30 of the axial adjustable ring segment 16.
The airfoil system 10 may include a rotor assembly 32 having a plurality of airfoils 26 extending radially therefrom and aligned axially to form a circumferentially extending row 34 of airfoils 26 forming an airfoil stage 20. The airfoils 26 may have any appropriate shape or configuration. The airfoil system 10 may also include a stationary turbine component 36 positioned radially outward from the tip 24 of the airfoil 26. In at least one embodiment, the stationary component 36 may be a turbine ring segment. In other embodiments, the stationary component 36 may be other components that remain stationary relative to the rotor assembly 32.
The airfoil system 10 may also include an adjustable clearance control system 14 including an axially adjustable ring segment 16 releasably coupled to the stationary turbine component 36, whereby the axially adjustable ring segment 16 may be adjustable axially. In particular, the axially adjustable ring segment 16 may be adjustable axially when positioned radially outward from a single row 34 of turbine blades 38 forming an airfoil stage 20, thereby enabling the axially adjustable ring segment 16 to be moved axially to change a gap size 22 between the tip 24 of the airfoil 26 and the axially adjustable ring segment 16 independently of other airfoil stages 18. In at least one embodiment, the plurality of airfoils 26 forming the circumferentially extending row 34 of airfoils forming the airfoil stage 20 forms an airfoil stage 20 closest to an exhaust diffuser 48. As such, the airfoil stage 20 may be the last turbine blade stage in a turbine blade assembly in reference to moving in a downstream direction. The exhaust diffuser 48 may be positioned aft of the turbine. The axially adjustable ring segment 16 may be used to control the flow of fluid through the gap 22 based on the diffuser inlet conditions. The axially adjustable ring segment 16 of the adjustable clearance control system 14 may extend at least partially circumferentially about the airfoil stage 20. In one embodiment, the axially adjustable ring segment 16 may extend circumferentially around the airfoil stage 20.
The axially adjustable ring segment 16 may include a radially inward contact surface 28 having at least a portion aligned with the tip 24 of the airfoil 26. In other embodiments, the radially inward contact surface 28 may be misaligned with the tip 24 of the airfoil 26. In at least one embodiment, the axially adjustable ring segment 16 may include a honeycomb seal land 40 coupled to the axially adjustable ring segment 16. The honeycomb seal land 40 may be formed from a plurality of hollow cavities 42 with walls 44 taking the shape of a honeycomb shape. The honeycomb shaped walls 44 may have any appropriate shape. The honeycomb seal land 40 may be formed from any appropriate material capable of withstanding the
environment within a hot gas path in a gas turbine engine 12. A radially inward contact surface 28 of the honeycomb seal land 40 may be positioned nonparallel and nonorthogonal relative to the direction of movement 30 of the axial adjustable ring segment 16.
In at least one embodiment, the adjustable clearance control system 14 may include one or more ribs 46 extending radially outward from the tip 24 of the turbine blade 38 and terminating before contacting a radially inward contact surface 28 of the axially adjustable ring segment 16. In at least one embodiment, the rib 46 may be formed from a tip shroud, as shown in Figure 4. The rib 46 may have any appropriate configuration and shape. The rib 46 may also may be formed from any appropriate material capable of withstanding the environment within a hot gas path in a gas turbine engine 12.
During operation, the adjustable clearance control system 14 may enable the axially adjustable ring segment 16 to be moved axially such that the gap 22 may be reduced or increased. For instance, the adjustable clearance control system 14 may move the axially adjustable ring segment 16 to increase the gap 22 independently of other stages 18 to increase tip jet flow when diffuser inlet conditions cause high OD loading. High OD loading occurs where a hub strong velocity profile entrains or pulls flow away from the OD and toward the hub. This causes more diffusion to occur near the OD wall. Hence higher loading or diffusion which can lead to flow
separation off the OD wall.
Also, during startup of the gas turbine engine 12, the adjustable clearance control system 14 may move the axially adjustable ring segment 16 axially such that the gap 22 may be increased to prevent tip rubbing and possible damage. Once at steady state operating conditions, the adjustable clearance control system 14 may move the axially adjustable ring segment 16 to reduce the gap 22 independently of other stages 18. During shutdown of the gas turbine engine 12, the adjustable clearance control system 14 may move the axially adjustable ring segment 16 to increase the gap 22 independently of other stages 18 to prevent tip rubbing and possible damage. In addition, increasing the clearance of the last stage blade 20 on cold days or on turbine engines 12 that have a hub strong exit velocity profile may improve engine performance. Under such conditions, the exhaust diffuser 48 may be overloaded on the outer diamater (OD) due to higher velocities on the hub pulling flow away. Increasing the last stage blade 20 tip gap 22 can inject high velocity air at the OD to keep the OD flow healthy in the exhaust diffuser 48.
The adjustable clearance control system 14 may address the need to open the clearance on the last blade tip 24 without opening the clearance on the other blades. Normally, turbine engines 12 are operated with the tightest possible clearance on every blade. However, at certain conditions, diffuser performance can be significantly improved by increasing the last blade tip leakage jet. Even though opening the gap reduces the blade efficiency, that loss can be more than offset by the improvement in the downstream diffuser. One particular situation in which diffuser performance can be significantly improved is at baseload, which is maximum power standard day, if the turbine has a hub strong velocity profile. This situation occurs when turbine engines 12 are operated at higher power levels, which can occur by increasing mass flow, without changing the turbine design. This situation can also occur in well designed turbines having a flat velocity profile when they operate on cold days.
The foregoing is provided for purposes of illustrating, explaining, and describing embodiments of this invention. Modifications and adaptations to these embodiments will be apparent to those skilled in the art and may be made without departing from the scope or spirit of this invention.

Claims

CLAIMS I claim:
1 . An airfoil system (10) for use in a gas turbine engine (12),
characterized in that:
a rotor assembly (32) having a plurality of airfoils (26) extending radially therefrom and aligned axially to form a circumferentially extending row (34) of airfoils (26) forming an airfoil stage (20);
a stationary turbine component (36) positioned radially outward from a tip (24) of at least one airfoil (26);
an adjustable clearance control system (14) including an axially adjustable ring segment (16) releasably coupled to the stationary turbine component (36), wherein the axially adjustable ring segment (16) is adjustable axially; and
wherein the axially adjustable ring segment (16) is adjustable axially when positioned radially outward from a single row (34) of turbine blades, thereby enabling the axially adjustable ring segment (16) to be moved axially to change a gap (22) size between the tip (24) of the airfoil (26) and the axially adjustable ring segment (16) independently of other airfoil stages (18).
2. The airfoil system (10) of claim 1 , characterized in that the axially adjustable ring segment (16) includes a radially inward segment contact surface (28) having at least a portion aligned with the tip (24) of the airfoil (26).
3. The airfoil system (10) of claim 1 , characterized in that the axially adjustable ring segment (16) includes a honeycomb seal land (40) coupled to the axially adjustable ring segment (16).
4. The airfoil system (10) of claim 3, characterized in that a radially inward contact surface of the honeycomb seal land (40) is positioned nonparallel and nonorthogonal relative to a direction of movement of the axial adjustable ring segment (16).
5. The airfoil system (10) of claim 1 , further characterized in that at least one rib (46) extending radially outward from the tip (24) of the turbine blade and terminating before contacting a radially inward contact surface (28) of the axially adjustable ring segment (16).
6. The airfoil system (10) of claim 1 , characterized in that the plurality of airfoils (26) forming the circumferentially extending row (34) of airfoils (26) forming the airfoil stage (20) forms an airfoil stage (20) closest to an exhaust diffuser (48).
7. The airfoil system (10) of claim 1 , characterized in that the axially adjustable ring segment (16) of the adjustable clearance control system (14) extends at least partially circumferentially about the airfoil stage (20).
PCT/US2014/070190 2013-12-18 2014-12-13 Adjustable clearance control system for airfoil tip in gas turbine engine WO2015094990A1 (en)

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US14/132,024 US20150167488A1 (en) 2013-12-18 2013-12-18 Adjustable clearance control system for airfoil tip in gas turbine engine
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