GB2042086A - Gas turbine engine seal - Google Patents
Gas turbine engine seal Download PDFInfo
- Publication number
- GB2042086A GB2042086A GB7933920A GB7933920A GB2042086A GB 2042086 A GB2042086 A GB 2042086A GB 7933920 A GB7933920 A GB 7933920A GB 7933920 A GB7933920 A GB 7933920A GB 2042086 A GB2042086 A GB 2042086A
- Authority
- GB
- United Kingdom
- Prior art keywords
- seal
- compressor
- cooling air
- air
- cooling
- 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.)
- Granted
Links
- 238000001816 cooling Methods 0.000 claims abstract description 57
- 238000010926 purge Methods 0.000 claims abstract description 8
- 238000000034 method Methods 0.000 claims description 10
- 238000011144 upstream manufacturing Methods 0.000 claims description 5
- 238000012986 modification Methods 0.000 claims description 2
- 230000004048 modification Effects 0.000 claims description 2
- 238000010438 heat treatment Methods 0.000 abstract 2
- 238000007789 sealing Methods 0.000 description 20
- 208000028659 discharge Diseases 0.000 description 9
- 230000007423 decrease Effects 0.000 description 3
- 230000003247 decreasing effect Effects 0.000 description 3
- 239000000463 material Substances 0.000 description 3
- 238000013021 overheating Methods 0.000 description 3
- 238000002485 combustion reaction Methods 0.000 description 2
- 230000003292 diminished effect Effects 0.000 description 2
- 230000003068 static effect Effects 0.000 description 2
- 208000027418 Wounds and injury Diseases 0.000 description 1
- 230000015556 catabolic process Effects 0.000 description 1
- 230000002301 combined effect Effects 0.000 description 1
- 230000006378 damage Effects 0.000 description 1
- 238000006731 degradation reaction Methods 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 238000002347 injection Methods 0.000 description 1
- 239000007924 injection Substances 0.000 description 1
- 208000014674 injury Diseases 0.000 description 1
- 230000001141 propulsive effect Effects 0.000 description 1
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
The seal is cooled by injecting air from the outlet of compressor diffuser 15 into a region between labyrinth teeth 20 at an angle in the direction of rotation of the seal teeth to minimize frictional heating of the injected air. Additionally, an outlet 36 for the cooling air is provided to a cavity 35 forward of the seal to purge this cavity and prevent stagnant air from heating the seal structure and other turbine components. <IMAGE>
Description
SPECIFICATION
Seal cooling method and apparatus
This invention pertains to seal cooling structures and methods in gas turbine engines and, more particularly, to seal cooling structures and methods for labyrinth-type seals.
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 compressor 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.
While preventing excessive air leakage, 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 1 100'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 labyrinth 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 between the sealing teeth and the sealing surface. This is recognized, and to some extent alleviated, by widespread use of honeycomb material or other abradable, 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 minimum clearance available when the sealing surface is at its maximum growth position, and the sealing teeth are at their minimum growth position.
To minimize thermal growth, methods and structures have been developed for cooling labyrinth seals by directing cooling air along the outer surface of the sealing structure, as shown in U.S. Patent No. 3,527,053 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. Patent No. 3,989,410 also assigned to the same assignee as the present 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 discharge 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 portion of the sealing structure in the region where the cooling air is directed 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 100"F (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, therefore, 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 air 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 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 structures.
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.
Breifly, in the method and apparatus of the present invention, 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 sufficiently compressed for use in this region. This air is cooler than boundary layer air extracted from the base of 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 the seal into a space between first and 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 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 allows 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 prevent possible overheating of the seal components and turbine rotor structure.
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 of 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 Fig.
2.
Referring now to Fig. 1, a compressor discharge pressure seal 10 is shown in its usual location within a typical gas turbine engine.
This seal 10 is generally located between compressor 11 and a combustor 1 6 in serial flow relation. In the gas turbine engine, a compressor section compresses engine 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 Fig. 1, an aft compressor blade 1 2 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 of rotation of the turbine engine and then directed through an outlet guide vane 14 and a compressor dump diffuser 1 5 to diffuse the compressed air and direct the air into the combustor 16. In the combustor section of the engine, the 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 engine.
If the reader desires an explanation of the operations within a gas turbine engine that affect a compressor discharge pressure seal, the reader is referred to U.S. Patent No.
3,527,053, the disclosure of which is incorporated herein by reference.
The compressor discharge pressure seal 10 is provided to prevent compressed air from escaping into central regions 1 9 of the gas turbine engine while, at the same time, permitting rotation of a compressor rotor 1 8 in iespect to the outlet guide vane 14 and combustor 16, which do not rotate. Compressor blades 12, one of which is shown in Fig.
1, are attached to the compressor rotor 18, 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 removes a component of rotational velocity of the compressed air before it enters the combustor.
The compressor dump diffuser 1 5 diffuses the air, causing a decrease in flow velocity and an increase in pressure.
The compressor discharge seal 10 is comprised of a series of circumferential labyrinth teeth 20 contiguous with a 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 close fit against the stator 22. Upon 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 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 this labyrinth-type seal and thereby maintain a closer fit between the teeth 20 and 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 teeth and seal stator to maintain the seal components at lower, more consistent temperatures. In the prior art system shown in Fig. 1, boundary layer air from the compressor discharge at the base of the aft compressor blade 1 2 is directed radially inward and axially aft along the compressor rotor 1 8 to the region of the compressor discharge pressure seal 1 0. Some of the compressed air then leaks through the seal along the path depicted by a wavy arrow, and continues to flow aft into a central region 1 9 of the gas turbine engine.
The method and apparatus of the present invention is shown in Figs. 2 and 3. In Fig. 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 1 5. In one embodiment of this invention, air in this region is approximately 100"F (55.55on) cooler than the compressor boundary layer air used in prior art systems, such as the system shown in Fig. 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 a 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 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 from the engine axis to impart a tangential component of velocity in the direction of rotor rotation. The direction of annular orientation is shown in Fig. 3 wherein it 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 1 8 for rotation therewith.Thus, the labyrinth teeth 20 rotate during engine operation, while the seal stator 22 does not. By orienting the 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, 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 1 5 is at a higher static pressure than the central regions 1 9 of the gas turbine beyond the compressor discharge seal 1 0. 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 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 stator 22 because of diminished thermal expansion and diminished differential thermal 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 1 8 as a result of zero throughflow.
As can be seen in Fig. 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 1 5 is higher than at the exit of the compressor, upstream of the guide vane 14. These holes 36 are cut at an angle to impart a tangential velocity component in the direction of rotor rotation, similar to the manner in which a tangential velocity component is imparted by passages 34.Tangential injection reduces the amount of frictional drag between the injected air and the rotating compressor rotor 1 8. 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 another advantage over the previous seal cooling system, shown in Fig. 1, wherein a flow of air is created through the forward cavity 35 because of 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.
Cavity 35 flow-through can approach zero, and overheating of the rotor 1 8 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 compressor discharge pressure seal. Therefore, if the seal clearance diminishes causing a temporary drop in seal leakage flow, the forward cavity 35 remains purged, and the affected portions of the rotor 1 8 do not overheat.The combined effects of the cooling airflow through passages 34 and inlet holes 36 serve to maintain both the compressor rotor 1 8 and the compressor dis charge 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 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 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 (11)
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.
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.
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 claims 1, 2, 3, or 4, and further including:
a cavity forward of said seal in flow communication 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.
6. The apparatus recited in claim 5 wherein said means to inject a portion of said cooling air comprises one 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.
7. An improved turbine engine comprising a compressor an$ rotating parts rotating about an axis of rotation of said turbine engine and a compressed cooling air circuit deriving cooling air from air discharged from the compressor diffuser for cooling said rotating parts, wherein the improvement comprises:
a passage or passages to direct cooling air from said cooling air circuit against said rotating parts wherein said passage or passages are angled in respect to a radius from said axis to impart a tangential velocity component in the direction of rotation of said parts.
8. An improved method of cooling a pressure seal in a gas turbine engine having a compressor, a compressor diffuser, and a combustor, all in serial flow relation, comprising the step of:
directing a portion of air discharged from the compressor diffuser into the pressure seal for cooling said seal.
9. An improved method of cooling a pressure seal in a gas turbine engine with an axis of rotation and having a compressor, a compressor diffuser, and a combustor, all in serial flow relation and wherein said seal has a section rotating about said engine axis, comprising the steps of:
directing a portion of air discharged from the compressor diffuser to the pressure seal for cooling said seal,
directing the cooling air into the seal at an angle in respect to a radius from the engine axis and in the direction of rotation of the rotating seal section to minimize frictional drag work done on the cooling air.
10. The method recited in claims 8 or 9 and further including the step of:
injecting a portion of air discharged from the compressor diffuser to a cavity forward of said seal to purge the cavity and avoid zero airflow conditions during engine operation.
11. A method as claimed in Claim 8 and substantially as hereinbefore described.
1 2. A turbomachine substantially in accordance with any embodiment (or modification thereof) of the invention claimed in Claim 1 or
Claim 7 and described and/or illustrated herein.
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US1557879A | 1979-02-26 | 1979-02-26 |
Publications (2)
Publication Number | Publication Date |
---|---|
GB2042086A true GB2042086A (en) | 1980-09-17 |
GB2042086B GB2042086B (en) | 1983-10-12 |
Family
ID=21772249
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
GB7933920A Expired GB2042086B (en) | 1979-02-26 | 1979-10-01 | Gas turbine engine seal |
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) |
Cited By (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB2158166A (en) * | 1984-04-27 | 1985-11-06 | Gen Electric | Labyrinth type rotary seal |
FR2570764A1 (en) * | 1984-09-27 | 1986-03-28 | Snecma | DEVICE FOR AUTOMATICALLY CONTROLLING THE PLAY OF A TURBOMACHINE LABYRINTH SEAL |
GB2263945A (en) * | 1992-02-05 | 1993-08-11 | Gen Electric | Diffuser clean air bleed assembly for gas turbine engines. |
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 |
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 |
Families Citing this family (4)
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 |
DE3627306A1 (en) * | 1986-02-28 | 1987-09-03 | Mtu Muenchen Gmbh | DEVICE FOR VENTILATING ROTOR COMPONENTS FOR COMPRESSORS OF GAS TURBINE ENGINE PLANTS |
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 |
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
Cited By (11)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB2158166A (en) * | 1984-04-27 | 1985-11-06 | Gen Electric | Labyrinth type rotary seal |
DE3514392A1 (en) * | 1984-04-27 | 1985-11-07 | General Electric Co., Schenectady, N.Y. | ROTARY SEAL |
FR2570764A1 (en) * | 1984-09-27 | 1986-03-28 | Snecma | DEVICE FOR AUTOMATICALLY CONTROLLING THE PLAY OF A TURBOMACHINE LABYRINTH SEAL |
EP0176447A1 (en) * | 1984-09-27 | 1986-04-02 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." | Apparatus for the automatic control of the play of a labyrinth seal of a turbo machine |
US4662821A (en) * | 1984-09-27 | 1987-05-05 | Societe Nationale D'etude Et De Construction De Moteur D'aviation S.N.E.C.M.A. | Automatic control device of a labyrinth seal clearance in a turbo jet engine |
GB2263945A (en) * | 1992-02-05 | 1993-08-11 | Gen Electric | Diffuser clean air bleed assembly for gas turbine engines. |
GB2263945B (en) * | 1992-02-05 | 1995-06-07 | Gen Electric | 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 |
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 |
Also Published As
Publication number | Publication date |
---|---|
DE2947439A1 (en) | 1980-08-28 |
GB2042086B (en) | 1983-10-12 |
FR2449789B1 (en) | 1984-10-19 |
JPS631452B2 (en) | 1988-01-12 |
IT7927210A0 (en) | 1979-11-12 |
FR2449789A1 (en) | 1980-09-19 |
CA1140474A (en) | 1983-02-01 |
DE2947439C2 (en) | 1994-09-22 |
IT1125667B (en) | 1986-05-14 |
JPS55114852A (en) | 1980-09-04 |
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Legal Events
Date | Code | Title | Description |
---|---|---|---|
746 | Register noted 'licences of right' (sect. 46/1977) |