CA1140474A - Seal cooling method and apparatus - Google Patents
Seal cooling method and apparatusInfo
- Publication number
- CA1140474A CA1140474A CA000344942A CA344942A CA1140474A CA 1140474 A CA1140474 A CA 1140474A CA 000344942 A CA000344942 A CA 000344942A CA 344942 A CA344942 A CA 344942A CA 1140474 A CA1140474 A CA 1140474A
- Authority
- CA
- Canada
- Prior art keywords
- seal
- compressor
- air
- cooling air
- rotor
- Prior art date
- Legal status (The legal status 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 status listed.)
- Expired
Links
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
- F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
- F01D11/04—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
SEAL COOLING METHOD AND APPARATUS
ABSTRACT OF THE DISCLOSURE
An improved seal cooling method and apparatus is provided for use in a gas turbine engine. In a particular embodiment, this method and apparatus is used to minimize leakage in a labyrinth-type pressure seal of a gas turbine engine by using tighter seal clearance made possible through lower seal operating temperatures.
The seal operating temperature is reduced by injecting relatively cool air from a compressor drump diffuser outlet into an interior region between labyrinth teeth of the seal at an angle in the direction of rotation of injected air. Additionally, an outlet for the cooling air is provided to a cavity forward of the seal to purge this cavity and prevent stagnant air from heating the seal structure and other turbine components.
ABSTRACT OF THE DISCLOSURE
An improved seal cooling method and apparatus is provided for use in a gas turbine engine. In a particular embodiment, this method and apparatus is used to minimize leakage in a labyrinth-type pressure seal of a gas turbine engine by using tighter seal clearance made possible through lower seal operating temperatures.
The seal operating temperature is reduced by injecting relatively cool air from a compressor drump diffuser outlet into an interior region between labyrinth teeth of the seal at an angle in the direction of rotation of injected air. Additionally, an outlet for the cooling air is provided to a cavity forward of the seal to purge this cavity and prevent stagnant air from heating the seal structure and other turbine components.
Description
~L14~
SEAL'COOLING METHOD AND APPARATUS
Field of the Invention This invention pertains to seal cooling structures and methods in gas turbine engines and, more particul-arly, to seal cooling structures and methods for labyrinth-type seals.
De'script'ion o`f''the'Pr'ior Axt A big factor in the performance of a gas turbine engine is the effectiveness of seal structures that must operate over a wide range of operational conditions.
In particular, a compressor discharge pressure seal must operate effectively at high pressures and usually at relatively high operating temperatures.
The compressor discharge pressure seal is employed to prevent compressed air from leaking between a rotating co~pres`sor section of an engine and a nonrotating combustor section. To optimize engine performance, tight clearance on this seal is highly desirable to minimize air leakage. Any air that leaks through the seal does not pass through the combustion cycle of the engine and, therefore, does not contribute to the power produced by the products of combustion.
7'~
While preventing excessive air leakageJ the compressor discharge pressure seal must also permit relative rotation between upper and lower sections of the seal and must operate in a region of the engine near the compressor discharge outlet where temperatures reach over 1100 F (593 C).
Current practice is to employ labyrinth-type seals in this region. A labyrinth-type seal comprises one or more circumferential teeth that are contiguous with a circumferential sealing surface wherein the teeth and sealing surface are relatively rotatable. Labyrinth seals can provide a high restriction to gas flow, and while there is some leakage, labyrinth seals do permit free rotation between upper and lower sections of the seal. This type of seal has many other well known advantages and is widely used at various sealing locations in gas turbine engines.
The effectiveness of labyrirlth seals is a function of the clearance between the sealing teeth and the contiguous sealing surface. While engine parts can be accurately machined to obtain minimum clearances and a highly effective seal, practical operation of the engine results in seal clearance degradation due to differential thermal growth betweerl the sealing teeth ~nd the sealing surface. This is recognized, arid to some extent alle~ated, by widespread use of honeycomb material or other abradahle, readily deformable materials to form the sealing surface with which the labyrinth teeth coact. By this approach, if the seaLing teeth grow at a faster rate than the sealing surface, the sealing surface will be deformed without injury to the sealing teeth. This will automatically establish the minirnum clearance available when the sealing surface is at its maximum growth position, and the sealing teeth are at their minimum growth position.
To minimi~e thermal growth, methods and structures have been developed for cooling labyrinth seals by directing cooling 347~L
air along the outer surface of the sealing structure, as shown in U.S. Pat.No.3,527,053 to Horn dated September 8, 1970 assigned to the same assignee as the present invention. Other systems have been developed which directly inject air into the space between the teeth of the labyrinth seal, as shown in U.S. Pat.No.
3,989,~10 to Ferrari dated Nov. 2,1976 also assigned to the same assignee as the presen~ invention. The use of cooling, air, as disclosed in these systems, greatly improves the efficiency of sealing structures by decreasing thermal expansion, thereby making it possible to maintain tighter seal clearance to cut down on seal leakage.
However, prior art systems involving compressor discharse pressure seals do not fully utilize the cooling air available in the most effective possible manner. In addition, problems have resulted because of friction-induced work done on the cooling air by the rotating portlon of the sealing structure in the region where the cooling air is dîrected against rapidly rotating portions of the seal. This increases the cooling air temperature and, therefore, the seal temperature. Also, in prior art systems, cooling air has been derived from boundary layer air at the base of the last blade of the compressor, and this boundary layer air can be as much as 100F
(55.5 C) warmer than nonboundary layer air. Finally, no matter where the cooling air comes from, the flow pattern must circulate through the entire seal and around support structures surrounding the seal. Any regions within the seal or surrounding the seal with zero airflow will undergo insufficient cooling, increasing the risk of material failure due to unnecessarily high temperatures.
It is, thereforer an object of the present invention to derive the cooling air for the compressor discharge pressure seal from a sufficiently pressurized source with - the lowest practical ~ir temperature.
It is another object of the present invention to inject the cooling air into the seal in a more efficient manner which decreases the amount of work done on the cooling air by the internal seal structure and, therefore, which results in cooler 5 air temperature .
It is another object of the present invention to distribute the cooling air in and around the compressor discharge pressure seal with a flow pattern that does not create "dead spots" with zero airflow in the seal or in the surrounding support 10 strUcture s .
These and other objects will be more fully understood from the drawings and from the following description and example, all of which are intended to be representative of rather than in any way limiting on the scope of the present invention.
Briefly, in the method and apparatus of the present in~ention, cooling air for a compressor discharge pressure seal is derived from compressed air discharged from a compressor dump diffuser. Air from this source is the coolest available source of air that is suficiently compressed for use in this region.
20 This air is cooler than boundary layer air extracted frvm the base OI a compressor blade, as is done in previous compressor seal cooling systems.
In a specific form, the cooling air is injected through passages in an outer ring of She seal into a space between first and 25 second teeth of the seal, taken from the upstream direction. These passages are angled to direct the cooling air with a tangential velocity component in the direction of rotation of the teeth and downstream portion of the seal structure. The tangential velocity component reduces the amount of frictional drag work done by the 30 teeth, seal rotor, and associated seal structure on the cooling air injected into the seal. This ultimately reduces the increase in cooling air temperature caused by friction and allol,vs the cooling air to extract more heat energy from the seal structure.
The cooling airflow path is also provided with a forward purge outlet into an open cavity in the support structures located forward of the seal. This forward purge outlet is provided to avoid zero airflow conditions from occurring in this forward cavity and S prevent possihle overheating of the seal components and turbine rotor structure.
DESCRIPTION OF THE DE~AWINGS
, ~
While the specification concludes with claims distinctly claiming and particularly pointing out the invention described herein, it is believed that the invention will be more clearly understood by reference to the discussion below in conjunction with the foLlowing drawings:
FIGURE 1 is a vertical cross-sectional view of a prior art seal cooling structure;
FIGURE 2 is a vertical cross-sectional view of seal cooling structure o~' the present invention and the surrounding components of a gas turbine engine; and FIGURE 3 is a cross-sectional view of the present invention taken along line 3-3 of Figure 2.
DESC~IPTION OF THE PREFERRED E~IBODIMENT
Referring now to Figure 1, a compressor discharge pressure seal 10 is shown in its u~ual location within a typical gas turbine engine. This seal 10 is generally located between compressor 11 and a combustor 16 in serial flow relation. In the gas turbine engine, a compressor section cornpresses engi~e intake air and a compressor discharge pressure seal retains this compressed air in the thrust-producing flowpath of the engine while permitting relative rotation along this flowpath of compressor parts in relation to the nonrotating combustor 16.
In Figure 1, an aft compressor blade 12 of the compressor section 11 is shown forward of the compressor discharge seal 10.
Intake air is compressed by compressor blades rotating about a central axis oE rotation of the turbine engine and then directed through an out~et guide vane 14 and a compressor dump diffuser lS to diffuse the compressed air and direct the air into the combustor 16. In the combustor section of the engine, the 5 compressed air is combined with fuel and ignited to form a thrust-producing propulsive gas flowstream. For the purpose of simplifying the description of this invention, a complete gas turbine engine is not shown. It is believed that the reader will fully appreciate this invention without a description of an entire 10 engine. If the reader desires an e~fplanation of the operations within a gas turbine engine that affect ~ compressor discharge pressure seal, the reader is referred to United Stat~s ~atent Numl~er 3, 527, 053 to Horn dated September 8, 1~70 .
The compressor discharge pressure seal 10 is provided to prevent compressed air Erom escaping into central regions 19 of the gas turbine engine while, at the same time, permitting rotation of a compressor rotor 18 in respect to the outlet guide vane 14 and combustor 16, which do not rotate.
20 Compressor blades 12, one of which is shown in Figure 1, are attached to the compressor rotor I8, and the rotor rotates the compressor blades to compress intake air passing through the compressor section 11 of the engine. The outlet guide vane 14 does not rotate and rerno~es a component of rotational velocity 25 of the compressed air before it enters the combustor. The compressor dump diffuser 15 diffuses the air, causing a decrease in flow velocity and an increase in pressure.
The compressor discharge seal 10 is compris2d of a series of circumferential labyrinth teeth 20 contiguous with a 30 seal outer stator 22 that defines a sealing surface. Outer edges 24 of the teeth 20 are initially assembled so as to form a very t74L
close fit against the stator 22. IJpon rotation of the compressor rotor 18 and the attached labyrinth teeth 20 about the engine axis, the outer edges 24 of the teeth create a slight groove in the inner surface of the seal stator 22. The very close fit between the 5 teeth 20 and the seal stator 22 inside these grooves provides a high degree of restriction to gas flow between the rotation teeth 20 and the stationary seal outer stator 22.
An object of this invention is to minimize differential thermal growth between the interacting portions of th~s labyrinth-10 type seal and thereby maintain a closer fit between the teeth 20and the seal stator 22 to improve seal effectiveness under operating conditions .
In prior art systems, such thermal growth has been decreased by passing compressor discharge air between the seal 15 teeth and seal stator to maintain the seal components at lower, more consistent temperatures. In the prlor art system shown in Figure 1, boundary layer air from the compressor discharge at the base of the aft compressor blade 12 is directed radially inward and axially aft along the compressor rotor 18 to the region of the 20 compressor discharge pressure seal 10. Some of the compressed air then leaks through the seal along the path depicted by a wavy arrow, and caltinues to Ilow a~t into a central region 19 of the gas turbine engine.
The method and apparatus of the present invention is 25 shown in Figures 2 and 3. In Figure 2, the flowpath of compressed air used to cool the seal structure 10 is shown with multiple arrows.
This air first is extracted downstream of the compressor dump diffuser 15. In one embodiment of this invention, air in this region is approximately 100 F (55. 55 C) cooler than the compressor 30 boundary layer air used in prior art systems, such as the system shown in Figure 1. The cooling air is directed through an inlet 30 to an open region 32 radially surrounding the seal stator 22.
From this open region 32, the air is directed through passages 34 in ;q seal bracket 23 and across an open slot 21 in the seal stator 22 into the space between the first and second labyrinth teeth 20 of the compressor discharge seal. The first and second 5 teeth are furthest upstream in respect to airflow through the turbine. The passages 34 are uniquely oriented on an angle in respect to a radius fromthe engine axis to impart a tangential component of velocity in the direction of rotor rotation. The direction of annular orientation is shown in Figure 3 wherein it 10 can be readily appreciated that the passages 34 cause the cooling air to be injected into the seal in the direction of rotor rotation.
The labyrinth teeth 20 are attached to the rotor 18 for rotation therewith. Thus, the labyrinth teeth 20 rotate during en~ne operation, while the seal stator 22 does not. By orienting the 15 passages 34 on an angle, the cooling air is injected in the direction of rotation of the teeth 20, thereby decreasing the frictional drag between the injected air and the teeth. The tangential component of velocity provided by this form of the present invention reduces the work done by the frictional drag on the cooling air and, 20 therefore, decreases the resulting increase in the temperature of the cooling air, Ultimately, the internal seal structure is maintained at a lower temperature.
The cooling air tends to flow through passages 34 because the region downstream of the compressor dump diffuser 25 15 is at a higher static pressure than the central regions 1~ of the gas turbine beyond the compressor discharge seal 10. The cooling air tends to leak in the aft direction across the region between the teeth outer edges 24 and the seal stator 22. A
small but continuous flow of leakage air is sufficient to cool 30 the seal components and maintain the internal seal structure at a relatively low temperature. This allows the seal to maintain a closer fit between the outer edges 24 of the seal teeth and seal 3~'~4 g stator 22 because of diminished thermal expansion and diminished differential therrnal growth.
Another unique feature of this invention is the manner in which a cavity 35 forward of the compressor discharge seal 10 is purged with air to avoid overheating of the rotor structure 18 as a result of zero throughflow. As can be seen in Figure 2, an annular series of inlet holes 36, one of which is shown, is provided to inject a small quantity of air into this forward cavity 35, Air will flow in this direction because the static pressure downstream of the diffuser 15 is higher than at the exit of the compressor, upstream of the guide vane 14. These holes 36 are cut at an an~le to impart a tangential velocity component in the direction of rotor rotation, similar to the manner in which a tangential velocity COInpOnent is imparted by passages 34. Tan~ential injection reduces the amount of frictional drag between the injected air and the rotating compressor rotor 18. This reduces the amount of work done on the injected air, the resulting increase in temperature of the air and, consequently, the rotor 18, and minimizes the quantity of air required for cavity purge~
This forward cavity purge apparatus offers anothe~
advantage over the previous seal cooling system, shown in ~igure 1, wherein a flow of air is created through the forl,vard cavity 35 because o~ leakage flow through the compressor discharge pressure seal. In the prior art system, when thermal expansion of the labyrinth teeth 20 of the seal is such as to cause a lesser clearance between the outer edges 24 of the teeth and the seal stator 22, the leakage flow of air is substantially reduced. Ca~ity 35 flow-through can approach zero, and overheating of the rotor 18 can result. In the present invention, because inlet holes 36 are provided, the amount of air injected into cavity 35 for cavity purge remains relatively constant and is not affected by clearance change in the '79L
compressor discharge pressure seal. Therefore, if the seal clearance diminishes causing a temporary drGp in seal leakage flow, the forward cavity 35 remains pur~ed, and the affected portions of the rotor 18 do not overheat. The combined effects 5 of the cooling airflow through passages 34 and inlet holes 36 serve to maintain both the compressor rotor 18 an~ the compressor discharge seal 10 at reasonable temperatures, thereby improving the performance of the high pressure turbine seal and rotor cooling circuit.
While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the scope of the invention as recited in the appended claims. For example, while the invention has been described in 15 conjunction with a labyrinth-type compressor discharge pressure seal in a gas turbine engine, it will be appreciated that various aspects of this invention are applicable to other sealing regions in a gas turbine engine and can be applied to sealing structures other than labyrinth-type seals. The methods and apparatus of the 20 present invention can be used to increase the performance of various seals in any type of turbomachinery. The scope of the invention, therefore, is to be derived from the following claims.
SEAL'COOLING METHOD AND APPARATUS
Field of the Invention This invention pertains to seal cooling structures and methods in gas turbine engines and, more particul-arly, to seal cooling structures and methods for labyrinth-type seals.
De'script'ion o`f''the'Pr'ior Axt A big factor in the performance of a gas turbine engine is the effectiveness of seal structures that must operate over a wide range of operational conditions.
In particular, a compressor discharge pressure seal must operate effectively at high pressures and usually at relatively high operating temperatures.
The compressor discharge pressure seal is employed to prevent compressed air from leaking between a rotating co~pres`sor section of an engine and a nonrotating combustor section. To optimize engine performance, tight clearance on this seal is highly desirable to minimize air leakage. Any air that leaks through the seal does not pass through the combustion cycle of the engine and, therefore, does not contribute to the power produced by the products of combustion.
7'~
While preventing excessive air leakageJ the compressor discharge pressure seal must also permit relative rotation between upper and lower sections of the seal and must operate in a region of the engine near the compressor discharge outlet where temperatures reach over 1100 F (593 C).
Current practice is to employ labyrinth-type seals in this region. A labyrinth-type seal comprises one or more circumferential teeth that are contiguous with a circumferential sealing surface wherein the teeth and sealing surface are relatively rotatable. Labyrinth seals can provide a high restriction to gas flow, and while there is some leakage, labyrinth seals do permit free rotation between upper and lower sections of the seal. This type of seal has many other well known advantages and is widely used at various sealing locations in gas turbine engines.
The effectiveness of labyrirlth seals is a function of the clearance between the sealing teeth and the contiguous sealing surface. While engine parts can be accurately machined to obtain minimum clearances and a highly effective seal, practical operation of the engine results in seal clearance degradation due to differential thermal growth betweerl the sealing teeth ~nd the sealing surface. This is recognized, arid to some extent alle~ated, by widespread use of honeycomb material or other abradahle, readily deformable materials to form the sealing surface with which the labyrinth teeth coact. By this approach, if the seaLing teeth grow at a faster rate than the sealing surface, the sealing surface will be deformed without injury to the sealing teeth. This will automatically establish the minirnum clearance available when the sealing surface is at its maximum growth position, and the sealing teeth are at their minimum growth position.
To minimi~e thermal growth, methods and structures have been developed for cooling labyrinth seals by directing cooling 347~L
air along the outer surface of the sealing structure, as shown in U.S. Pat.No.3,527,053 to Horn dated September 8, 1970 assigned to the same assignee as the present invention. Other systems have been developed which directly inject air into the space between the teeth of the labyrinth seal, as shown in U.S. Pat.No.
3,989,~10 to Ferrari dated Nov. 2,1976 also assigned to the same assignee as the presen~ invention. The use of cooling, air, as disclosed in these systems, greatly improves the efficiency of sealing structures by decreasing thermal expansion, thereby making it possible to maintain tighter seal clearance to cut down on seal leakage.
However, prior art systems involving compressor discharse pressure seals do not fully utilize the cooling air available in the most effective possible manner. In addition, problems have resulted because of friction-induced work done on the cooling air by the rotating portlon of the sealing structure in the region where the cooling air is dîrected against rapidly rotating portions of the seal. This increases the cooling air temperature and, therefore, the seal temperature. Also, in prior art systems, cooling air has been derived from boundary layer air at the base of the last blade of the compressor, and this boundary layer air can be as much as 100F
(55.5 C) warmer than nonboundary layer air. Finally, no matter where the cooling air comes from, the flow pattern must circulate through the entire seal and around support structures surrounding the seal. Any regions within the seal or surrounding the seal with zero airflow will undergo insufficient cooling, increasing the risk of material failure due to unnecessarily high temperatures.
It is, thereforer an object of the present invention to derive the cooling air for the compressor discharge pressure seal from a sufficiently pressurized source with - the lowest practical ~ir temperature.
It is another object of the present invention to inject the cooling air into the seal in a more efficient manner which decreases the amount of work done on the cooling air by the internal seal structure and, therefore, which results in cooler 5 air temperature .
It is another object of the present invention to distribute the cooling air in and around the compressor discharge pressure seal with a flow pattern that does not create "dead spots" with zero airflow in the seal or in the surrounding support 10 strUcture s .
These and other objects will be more fully understood from the drawings and from the following description and example, all of which are intended to be representative of rather than in any way limiting on the scope of the present invention.
Briefly, in the method and apparatus of the present in~ention, cooling air for a compressor discharge pressure seal is derived from compressed air discharged from a compressor dump diffuser. Air from this source is the coolest available source of air that is suficiently compressed for use in this region.
20 This air is cooler than boundary layer air extracted frvm the base OI a compressor blade, as is done in previous compressor seal cooling systems.
In a specific form, the cooling air is injected through passages in an outer ring of She seal into a space between first and 25 second teeth of the seal, taken from the upstream direction. These passages are angled to direct the cooling air with a tangential velocity component in the direction of rotation of the teeth and downstream portion of the seal structure. The tangential velocity component reduces the amount of frictional drag work done by the 30 teeth, seal rotor, and associated seal structure on the cooling air injected into the seal. This ultimately reduces the increase in cooling air temperature caused by friction and allol,vs the cooling air to extract more heat energy from the seal structure.
The cooling airflow path is also provided with a forward purge outlet into an open cavity in the support structures located forward of the seal. This forward purge outlet is provided to avoid zero airflow conditions from occurring in this forward cavity and S prevent possihle overheating of the seal components and turbine rotor structure.
DESCRIPTION OF THE DE~AWINGS
, ~
While the specification concludes with claims distinctly claiming and particularly pointing out the invention described herein, it is believed that the invention will be more clearly understood by reference to the discussion below in conjunction with the foLlowing drawings:
FIGURE 1 is a vertical cross-sectional view of a prior art seal cooling structure;
FIGURE 2 is a vertical cross-sectional view of seal cooling structure o~' the present invention and the surrounding components of a gas turbine engine; and FIGURE 3 is a cross-sectional view of the present invention taken along line 3-3 of Figure 2.
DESC~IPTION OF THE PREFERRED E~IBODIMENT
Referring now to Figure 1, a compressor discharge pressure seal 10 is shown in its u~ual location within a typical gas turbine engine. This seal 10 is generally located between compressor 11 and a combustor 16 in serial flow relation. In the gas turbine engine, a compressor section cornpresses engi~e intake air and a compressor discharge pressure seal retains this compressed air in the thrust-producing flowpath of the engine while permitting relative rotation along this flowpath of compressor parts in relation to the nonrotating combustor 16.
In Figure 1, an aft compressor blade 12 of the compressor section 11 is shown forward of the compressor discharge seal 10.
Intake air is compressed by compressor blades rotating about a central axis oE rotation of the turbine engine and then directed through an out~et guide vane 14 and a compressor dump diffuser lS to diffuse the compressed air and direct the air into the combustor 16. In the combustor section of the engine, the 5 compressed air is combined with fuel and ignited to form a thrust-producing propulsive gas flowstream. For the purpose of simplifying the description of this invention, a complete gas turbine engine is not shown. It is believed that the reader will fully appreciate this invention without a description of an entire 10 engine. If the reader desires an e~fplanation of the operations within a gas turbine engine that affect ~ compressor discharge pressure seal, the reader is referred to United Stat~s ~atent Numl~er 3, 527, 053 to Horn dated September 8, 1~70 .
The compressor discharge pressure seal 10 is provided to prevent compressed air Erom escaping into central regions 19 of the gas turbine engine while, at the same time, permitting rotation of a compressor rotor 18 in respect to the outlet guide vane 14 and combustor 16, which do not rotate.
20 Compressor blades 12, one of which is shown in Figure 1, are attached to the compressor rotor I8, and the rotor rotates the compressor blades to compress intake air passing through the compressor section 11 of the engine. The outlet guide vane 14 does not rotate and rerno~es a component of rotational velocity 25 of the compressed air before it enters the combustor. The compressor dump diffuser 15 diffuses the air, causing a decrease in flow velocity and an increase in pressure.
The compressor discharge seal 10 is compris2d of a series of circumferential labyrinth teeth 20 contiguous with a 30 seal outer stator 22 that defines a sealing surface. Outer edges 24 of the teeth 20 are initially assembled so as to form a very t74L
close fit against the stator 22. IJpon rotation of the compressor rotor 18 and the attached labyrinth teeth 20 about the engine axis, the outer edges 24 of the teeth create a slight groove in the inner surface of the seal stator 22. The very close fit between the 5 teeth 20 and the seal stator 22 inside these grooves provides a high degree of restriction to gas flow between the rotation teeth 20 and the stationary seal outer stator 22.
An object of this invention is to minimize differential thermal growth between the interacting portions of th~s labyrinth-10 type seal and thereby maintain a closer fit between the teeth 20and the seal stator 22 to improve seal effectiveness under operating conditions .
In prior art systems, such thermal growth has been decreased by passing compressor discharge air between the seal 15 teeth and seal stator to maintain the seal components at lower, more consistent temperatures. In the prlor art system shown in Figure 1, boundary layer air from the compressor discharge at the base of the aft compressor blade 12 is directed radially inward and axially aft along the compressor rotor 18 to the region of the 20 compressor discharge pressure seal 10. Some of the compressed air then leaks through the seal along the path depicted by a wavy arrow, and caltinues to Ilow a~t into a central region 19 of the gas turbine engine.
The method and apparatus of the present invention is 25 shown in Figures 2 and 3. In Figure 2, the flowpath of compressed air used to cool the seal structure 10 is shown with multiple arrows.
This air first is extracted downstream of the compressor dump diffuser 15. In one embodiment of this invention, air in this region is approximately 100 F (55. 55 C) cooler than the compressor 30 boundary layer air used in prior art systems, such as the system shown in Figure 1. The cooling air is directed through an inlet 30 to an open region 32 radially surrounding the seal stator 22.
From this open region 32, the air is directed through passages 34 in ;q seal bracket 23 and across an open slot 21 in the seal stator 22 into the space between the first and second labyrinth teeth 20 of the compressor discharge seal. The first and second 5 teeth are furthest upstream in respect to airflow through the turbine. The passages 34 are uniquely oriented on an angle in respect to a radius fromthe engine axis to impart a tangential component of velocity in the direction of rotor rotation. The direction of annular orientation is shown in Figure 3 wherein it 10 can be readily appreciated that the passages 34 cause the cooling air to be injected into the seal in the direction of rotor rotation.
The labyrinth teeth 20 are attached to the rotor 18 for rotation therewith. Thus, the labyrinth teeth 20 rotate during en~ne operation, while the seal stator 22 does not. By orienting the 15 passages 34 on an angle, the cooling air is injected in the direction of rotation of the teeth 20, thereby decreasing the frictional drag between the injected air and the teeth. The tangential component of velocity provided by this form of the present invention reduces the work done by the frictional drag on the cooling air and, 20 therefore, decreases the resulting increase in the temperature of the cooling air, Ultimately, the internal seal structure is maintained at a lower temperature.
The cooling air tends to flow through passages 34 because the region downstream of the compressor dump diffuser 25 15 is at a higher static pressure than the central regions 1~ of the gas turbine beyond the compressor discharge seal 10. The cooling air tends to leak in the aft direction across the region between the teeth outer edges 24 and the seal stator 22. A
small but continuous flow of leakage air is sufficient to cool 30 the seal components and maintain the internal seal structure at a relatively low temperature. This allows the seal to maintain a closer fit between the outer edges 24 of the seal teeth and seal 3~'~4 g stator 22 because of diminished thermal expansion and diminished differential therrnal growth.
Another unique feature of this invention is the manner in which a cavity 35 forward of the compressor discharge seal 10 is purged with air to avoid overheating of the rotor structure 18 as a result of zero throughflow. As can be seen in Figure 2, an annular series of inlet holes 36, one of which is shown, is provided to inject a small quantity of air into this forward cavity 35, Air will flow in this direction because the static pressure downstream of the diffuser 15 is higher than at the exit of the compressor, upstream of the guide vane 14. These holes 36 are cut at an an~le to impart a tangential velocity component in the direction of rotor rotation, similar to the manner in which a tangential velocity COInpOnent is imparted by passages 34. Tan~ential injection reduces the amount of frictional drag between the injected air and the rotating compressor rotor 18. This reduces the amount of work done on the injected air, the resulting increase in temperature of the air and, consequently, the rotor 18, and minimizes the quantity of air required for cavity purge~
This forward cavity purge apparatus offers anothe~
advantage over the previous seal cooling system, shown in ~igure 1, wherein a flow of air is created through the forl,vard cavity 35 because o~ leakage flow through the compressor discharge pressure seal. In the prior art system, when thermal expansion of the labyrinth teeth 20 of the seal is such as to cause a lesser clearance between the outer edges 24 of the teeth and the seal stator 22, the leakage flow of air is substantially reduced. Ca~ity 35 flow-through can approach zero, and overheating of the rotor 18 can result. In the present invention, because inlet holes 36 are provided, the amount of air injected into cavity 35 for cavity purge remains relatively constant and is not affected by clearance change in the '79L
compressor discharge pressure seal. Therefore, if the seal clearance diminishes causing a temporary drGp in seal leakage flow, the forward cavity 35 remains pur~ed, and the affected portions of the rotor 18 do not overheat. The combined effects 5 of the cooling airflow through passages 34 and inlet holes 36 serve to maintain both the compressor rotor 18 an~ the compressor discharge seal 10 at reasonable temperatures, thereby improving the performance of the high pressure turbine seal and rotor cooling circuit.
While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the scope of the invention as recited in the appended claims. For example, while the invention has been described in 15 conjunction with a labyrinth-type compressor discharge pressure seal in a gas turbine engine, it will be appreciated that various aspects of this invention are applicable to other sealing regions in a gas turbine engine and can be applied to sealing structures other than labyrinth-type seals. The methods and apparatus of the 20 present invention can be used to increase the performance of various seals in any type of turbomachinery. The scope of the invention, therefore, is to be derived from the following claims.
Claims (5)
1. An improved turbomachinery comprising a compressor, a compressor diffuser, and a combustor all in serial flow relation, and a pressure seal downstream of said compressor, wherein the improvement comprises:
means for deriving cooling air from air discharged from the compressor diffuser and directing that cooling air into said seal for cooling said seal;
a cavity forward of said seal in flow communica-tion with compressor discharge air and radially surrounding said compressor rotor; and means to inject a portion of said cooling air into said cavity to purge the region and prevent zero airflow conditions during engine operation.
means for deriving cooling air from air discharged from the compressor diffuser and directing that cooling air into said seal for cooling said seal;
a cavity forward of said seal in flow communica-tion with compressor discharge air and radially surrounding said compressor rotor; and means to inject a portion of said cooling air into said cavity to purge the region and prevent zero airflow conditions during engine operation.
2. The apparatus recited in claim 1 wherein said pressure seal is a labyrinth-type compressor discharge pressure seal with a seal stator and a seal rotor having two or more teeth, and further including means to direct said cooling air directly into a region between the teeth of the labyrinth seal.
3. The apparatus recited in claim 2 wherein said means to direct the cooling air into the seal comprises:
one or more passages through the seal stator into a region between a furthest upstream tooth and a tooth adjoining said upstream tooth of said seal rotor.
one or more passages through the seal stator into a region between a furthest upstream tooth and a tooth adjoining said upstream tooth of said seal rotor.
4. The apparatus recited in claim 3 wherein said gas turbine engine has an axis of rotation around which said seal rotor rotates and wherein said passages are angled in respect to a radius from said axis of the gas turbine engine to impart a tangential velocity component in the direction of seal rotor rotation to minimize the frictional drag work done on the cooling air.
5. The apparatus recited in claim 1 wherein said means to inject a portion of said cooling air comprises one
5. The apparatus recited in claim 1 wherein said means to inject a portion of said cooling air comprises one
Claim 5 continued:
or more passages and wherein said passage or passages are oriented at an angle in respect to a radius from an axis of rotation of said compressor rotor to impart a tangential velocity component in the direction of compressor rotor rotation.
or more passages and wherein said passage or passages are oriented at an angle in respect to a radius from an axis of rotation of said compressor rotor to impart a tangential velocity component in the direction of compressor rotor rotation.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US1557879A | 1979-02-26 | 1979-02-26 | |
| US15,578 | 1979-02-26 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1140474A true CA1140474A (en) | 1983-02-01 |
Family
ID=21772249
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000344942A Expired CA1140474A (en) | 1979-02-26 | 1980-02-01 | Seal cooling method and apparatus |
Country Status (6)
| Country | Link |
|---|---|
| JP (1) | JPS55114852A (en) |
| CA (1) | CA1140474A (en) |
| DE (1) | DE2947439C2 (en) |
| FR (1) | FR2449789A1 (en) |
| GB (1) | GB2042086B (en) |
| IT (1) | IT1125667B (en) |
Families Citing this family (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4397471A (en) * | 1981-09-02 | 1983-08-09 | General Electric Company | Rotary pressure seal structure and method for reducing thermal stresses therein |
| US4513975A (en) * | 1984-04-27 | 1985-04-30 | General Electric Company | Thermally responsive labyrinth seal |
| FR2570764B1 (en) * | 1984-09-27 | 1986-11-28 | Snecma | DEVICE FOR AUTOMATICALLY CONTROLLING THE PLAY OF A TURBOMACHINE LABYRINTH SEAL |
| DE3627306A1 (en) * | 1986-02-28 | 1987-09-03 | Mtu Muenchen Gmbh | DEVICE FOR VENTILATING ROTOR COMPONENTS FOR COMPRESSORS OF GAS TURBINE ENGINE PLANTS |
| US5211003A (en) * | 1992-02-05 | 1993-05-18 | General Electric Company | Diffuser clean air bleed assembly |
| US7234918B2 (en) | 2004-12-16 | 2007-06-26 | Siemens Power Generation, Inc. | Gap control system for turbine engines |
| US8167547B2 (en) * | 2007-03-05 | 2012-05-01 | United Technologies Corporation | Gas turbine engine with canted pocket and canted knife edge seal |
| DE102012206090A1 (en) * | 2012-04-13 | 2013-10-17 | Rolls-Royce Deutschland Ltd & Co Kg | Axial compressor for turbomachine, has circulating cover plate to protect material of rotor body against hot compressor air, where cover plate has sealing unit, which prevents passing of compressor air in cooling channel |
| DE102013217504A1 (en) | 2013-09-03 | 2015-03-05 | MTU Aero Engines AG | flow machine |
| EP3540180A1 (en) * | 2018-03-14 | 2019-09-18 | General Electric Company | Inter-stage cavity purge ducts |
| US11293295B2 (en) | 2019-09-13 | 2022-04-05 | Pratt & Whitney Canada Corp. | Labyrinth seal with angled fins |
Family Cites Families (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3043561A (en) * | 1958-12-29 | 1962-07-10 | Gen Electric | Turbine rotor ventilation system |
| US3303997A (en) * | 1965-04-21 | 1967-02-14 | United Aircraft Corp | Compressor air seal |
| US3365172A (en) * | 1966-11-02 | 1968-01-23 | Gen Electric | Air cooled shroud seal |
| CA939521A (en) * | 1970-04-28 | 1974-01-08 | Bruce R. Branstrom | Turbine coolant flow system |
| CA982828A (en) * | 1972-06-01 | 1976-02-03 | General Electric Company | Combustor casing cooling structure |
| US3989410A (en) * | 1974-11-27 | 1976-11-02 | General Electric Company | Labyrinth seal system |
| US3986720A (en) * | 1975-04-14 | 1976-10-19 | General Electric Company | Turbine shroud structure |
| US4103899A (en) * | 1975-10-01 | 1978-08-01 | United Technologies Corporation | Rotary seal with pressurized air directed at fluid approaching the seal |
| US4113406A (en) * | 1976-11-17 | 1978-09-12 | Westinghouse Electric Corp. | Cooling system for a gas turbine engine |
-
1979
- 1979-10-01 GB GB7933920A patent/GB2042086B/en not_active Expired
- 1979-11-12 IT IT27210/79A patent/IT1125667B/en active
- 1979-11-22 JP JP15077179A patent/JPS55114852A/en active Granted
- 1979-11-24 DE DE2947439A patent/DE2947439C2/en not_active Expired - Lifetime
- 1979-11-26 FR FR7929044A patent/FR2449789A1/en active Granted
-
1980
- 1980-02-01 CA CA000344942A patent/CA1140474A/en not_active Expired
Also Published As
| Publication number | Publication date |
|---|---|
| JPS631452B2 (en) | 1988-01-12 |
| FR2449789B1 (en) | 1984-10-19 |
| IT7927210A0 (en) | 1979-11-12 |
| FR2449789A1 (en) | 1980-09-19 |
| GB2042086B (en) | 1983-10-12 |
| DE2947439C2 (en) | 1994-09-22 |
| IT1125667B (en) | 1986-05-14 |
| JPS55114852A (en) | 1980-09-04 |
| DE2947439A1 (en) | 1980-08-28 |
| GB2042086A (en) | 1980-09-17 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| MKEX | Expiry |