CN114961870A - Aeroengine rim sealing system and aeroengine - Google Patents
Aeroengine rim sealing system and aeroengine Download PDFInfo
- Publication number
- CN114961870A CN114961870A CN202110205800.7A CN202110205800A CN114961870A CN 114961870 A CN114961870 A CN 114961870A CN 202110205800 A CN202110205800 A CN 202110205800A CN 114961870 A CN114961870 A CN 114961870A
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- Prior art keywords
- blade
- aircraft engine
- rim
- flange
- sealing system
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- 238000007789 sealing Methods 0.000 title claims abstract description 64
- 238000001816 cooling Methods 0.000 claims description 2
- 238000002156 mixing Methods 0.000 abstract description 8
- 230000000694 effects Effects 0.000 abstract description 7
- 239000007789 gas Substances 0.000 description 40
- 230000009545 invasion Effects 0.000 description 12
- 238000004378 air conditioning Methods 0.000 description 7
- 239000002737 fuel gas Substances 0.000 description 5
- 244000126211 Hericium coralloides Species 0.000 description 3
- 238000002485 combustion reaction Methods 0.000 description 3
- 230000003068 static effect Effects 0.000 description 3
- 230000009286 beneficial effect Effects 0.000 description 2
- 239000012141 concentrate Substances 0.000 description 2
- 230000000903 blocking effect Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
Images
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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
-
- 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
-
- 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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
-
- 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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Architecture (AREA)
- Gasket Seals (AREA)
Abstract
The invention relates to an aeroengine rim sealing system and an aeroengine. Wherein, aeroengine rim system of obturaging includes: the first blade is provided with a first edge plate; the second blade is arranged adjacent to the first blade and provided with a second flange, and a sealing cavity is formed between the root of the second blade and the root of the first blade; the second flange plate is close to the central axis of the aircraft engine relative to the first flange plate, the second flange plate and the first flange plate are provided with mutually overlapped parts, and a gap is formed between the mutually overlapped parts; a flow guide groove is formed in one side, away from the first flange, of the second flange, and a first end of the flow guide groove in the length extension direction is configured to guide part of the cold air in the sealing cavity to a second end of the flow guide groove in the length extension direction, so that part of the cold air flows out through a gap between the mutually overlapped parts; the groove width of the first end of the diversion groove is smaller than that of the second end of the diversion groove. The invention can reduce the mixing temperature in the sealing cavity of the wheel rim and improve the sealing effect of the wheel rim.
Description
Technical Field
The invention relates to the field of aerospace equipment, in particular to an aeroengine rim sealing system and an aeroengine.
Background
The gas turbine engine consists of a gas compressor, a combustion chamber and a turbine, wherein the gas compressor compresses gas, the combustion chamber provides energy through combustion, and the turbine expands to do work to drive the gas compressor. The turbine has several stages, each of which is composed of stator blades fixed to the turbine casing and rotor blades fixed to the turbine disk. Turbine blade is in engine sprue gas environment, and the temperature resistant level is higher, and the spare part except for the blade, especially the turbine dish, the temperature resistant level is lower, needs to adopt cooling air to cool off to design the wheel rim structure of obturating at the rim of turbine blade root and prevent that the gas of sprue from getting into inside parts such as arousing turbine dish overtemperature. In the related art, cold air flowing at the rim flows outwards from the rotor side, and fuel gas in the main flow channel flows inwards from the stator side to be mixed, so that the temperature inside the rim seal is high.
Disclosure of Invention
Some embodiments of the invention provide an aircraft engine rim sealing system and an aircraft engine, which are used for relieving the problem of high temperature inside rim sealing.
Some embodiments of the present invention provide an aircraft engine rim sealing system, comprising:
the first blade is provided with a first edge plate; and
the second blade is arranged adjacent to the first blade and provided with a second flange, and a sealing cavity is formed between the root of the second blade and the root of the first blade; the second flange is close to the central axis of the aircraft engine relative to the first flange, the second flange and the first flange are provided with mutually overlapped parts, and a gap is formed between the mutually overlapped parts; a flow guide groove is formed in one side, away from the first flange, of the second flange, and a first end of the flow guide groove in the length extending direction is configured to guide part of the cold air in the sealed cavity to a second end of the flow guide groove in the length extending direction, so that the part of the cold air flows out through a gap between the mutually overlapped parts; the groove width of the first end of the diversion groove is smaller than that of the second end of the diversion groove.
In some embodiments, the guide groove is configured to guide a portion of the cool air in the sealed cavity, so that the portion of the cool air flows around a central axis of the aircraft engine in a rotating manner and flows out through a gap between the mutually overlapped portions.
In some embodiments, a length extension direction of the flow guide groove is perpendicular to a central axis of the aircraft engine and a length extension direction of the second blade.
In some embodiments, the sidewalls of the channels are configured to be linear.
In some embodiments, the sidewalls of the channels are configured in an arcuate pattern.
In some embodiments, the channel width of the channel gradually increases from the first end to the second end along the length extension direction of the channel.
In some embodiments, the aeroengine rim sealing system further comprises a gas collection chamber and an exhaust hole arranged at the root of the second blade, the exhaust hole is communicated with the gas collection chamber and the sealing chamber, and the gas collection chamber is configured to provide cold air to the sealing chamber through the exhaust hole.
In some embodiments, the exhaust direction of the exhaust hole faces the diversion trench, or the exhaust direction of the exhaust hole is inclined to the central axis of the aircraft engine.
In some embodiments, the first flange and the second flange are arranged in a circle around a central axis of the aircraft engine, the second flange is provided with a plurality of the guiding grooves at intervals in the circumferential direction, the root of the second blade is provided with a plurality of the exhaust holes at intervals, each guiding groove in the plurality of guiding grooves is provided with at least one exhaust hole in the plurality of exhaust holes, and the exhaust holes are configured to guide the cold air in the air collecting cavity to the guiding grooves.
Some embodiments of the invention provide an aircraft engine comprising a turbine assembly and the aircraft engine rim sealing system, wherein the turbine assembly comprises a turbine disc and a turbine casing, the first blades are arranged on the turbine disc, and the second blades are arranged on the turbine casing.
Based on the technical scheme, the invention at least has the following beneficial effects:
in some embodiments, an aircraft engine rim seal system includes a first blade and a second blade; the second blade is arranged adjacent to the first blade, and a sealing cavity is formed between the root of the second blade and the root of the first blade; the second flange plate of the second blade and the first flange plate of the first blade are provided with mutually overlapped parts, and a gap is formed between the mutually overlapped parts; a diversion trench is arranged on one side, far away from the first edge plate, of the second edge plate; the first end of the length extension direction of the guide groove is constructed to lead part of the cold air in the sealing cavity to the second end of the length extension direction of the guide groove, so that part of the cold air flows out through the gap between the mutually overlapped parts to form sealing; the groove width of the first end of the diversion groove is smaller than that of the second end of the diversion groove; the guiding gutter is flaring type structure, and static pressure risees when making air conditioning expand along the guiding gutter, strengthens cold air flow in the part, resists the gas inflow of rim seal stator side better, resists the invasion that the inhomogeneous high temperature main gas of circumference brought, reduces the mixing temperature of the intracavity that seals of rim, improves the rim effect of sealing.
Drawings
The accompanying drawings, which are included to provide a further understanding of the invention and are incorporated in and constitute a part of this application, illustrate embodiment(s) of the invention and together with the description serve to explain the invention without limiting the invention. In the drawings:
FIG. 1 is a schematic view of an aircraft engine rim sealing system provided in accordance with some embodiments of the present invention;
FIG. 2 is a schematic illustration of the resultant velocity of high temperature mainstream fuel gas provided in accordance with some embodiments of the invention;
FIG. 3 is a schematic bottom view of a second platform according to a first embodiment of the present invention;
fig. 4 is a schematic bottom view of a second platform according to a second embodiment of the present invention.
The reference numbers in the drawings illustrate the following:
1-a first blade; 11-a first flange;
2-a second blade; 21-a second flange; 22-a diversion trench; 23-a vent hole;
3-a gas collection cavity;
4-sealing the cavity;
5-a honeycomb structure;
6-a grid structure.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments. It is to be understood that the described embodiments are merely a few embodiments of the invention, and not all embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments of the present invention without inventive step, shall fall within the scope of protection of the present invention.
In the description of the present invention, it is to be understood that the terms "center", "longitudinal", "lateral", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate orientations or positional relationships based on those shown in the drawings, and are used merely for convenience in describing the present invention and for simplifying the description, and do not indicate or imply that the device or element being referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore, should not be taken as limiting the scope of the present invention.
At the sealing position of a rim of an aircraft engine, cold air flowing at the rim flows outwards from the rotor side, and gas in a main flow channel flows inwards from the stator side to be mixed, so that the temperature of an inner cavity of the sealing of the rim is higher. Particularly, the main runner gas has a certain circumferential speed, the total pressure is high in the combined speed direction, flowing cold air at the wheel edge flows out from the stator side along the radial direction, the invasion blocking effect on the circumferential gas in the combined speed direction is poor, and the mixing temperature of the inner cavity sealed by the wheel rim is high; and the pressure of the main flow gas is periodically distributed in the circumferential direction, and the pressure in certain directions is high, so that the invasion of the stator side gas is more serious.
Based on this, some embodiments of this disclosure provide an aeroengine rim system of obturating for reduce the gas invasion of rim obturating stator side, resist the invasion that the inhomogeneous gas in circumference brought, reduce the mixing temperature in the inside chamber of rim obturating, improve the rim effect of obturating.
As shown in FIG. 1, in some embodiments, an aircraft engine rim sealing system includes a first blade 1 and a second blade 2.
The first blade 1 is provided with a first platform 11.
The second blade 2 is arranged adjacent to the first blade 1, the second blade 2 is provided with a second flange 21, and a sealing cavity 4 is formed between the root of the second blade 2 and the root of the first blade 1; the second flange plate 21 is close to the central axis of the aircraft engine relative to the first flange plate 11, the second flange plate 21 and the first flange plate 11 are provided with mutually overlapped parts, a gap is formed between the mutually overlapped parts, and the gap is sealed by air flow; the side of the second flange 21 away from the first flange 11 is provided with a diversion trench 22. The sealing cavity 4 is arranged on one side of the first flange 11 and the second flange 21 close to the central axis of the aircraft engine. The second flange 21 is one of the sidewalls of the sealing cavity 4, and the guiding groove 22 is located in the sealing cavity 4.
The first end of the guide groove 22 in the length extending direction is configured to guide part of the cold air in the sealing cavity 4 to the second end of the guide groove 22 in the length extending direction, so that part of the cold air flows out through the gap between the mutually overlapped parts to form rim sealing; the channel width of the first end of channel 22 is smaller than the channel width of the second end of channel 22. The diversion trench 22 is a flaring structure, the corresponding cold air flow area is increased, the airflow static pressure is increased, more cold air can be concentrated to mainly resist the invasion of the main flow gas high-pressure area and the high-pressure area gas, the mixing temperature of the inner cavity of the rim sealing is reduced, and the rim sealing effect is improved.
In the rim sealing system, cold air in the rim sealing cavity 4 is discharged outwards along the rotor, and high-temperature main stream gas flows towards the rim sealing cavity 4 along the stator. As shown in fig. 2, the velocity direction of the high-temperature main flow gas has a substantial circumferential component in addition to a radial component, resulting in a resultant velocity that is inclined.
In some embodiments, the diversion trench 22 is configured to guide and guide part of the cold air in the sealing cavity 4, so that part of the cold air flows around the central axis of the aircraft engine in a rotating manner and flows out through the gap between the mutually overlapped parts, that is, the diversion trench 22 enables the cold air to flow in a rotating manner and directly face the inflow direction of the high-temperature mainstream fuel gas, so that the invasion of the fuel gas can be positively resisted, the invasion caused by the high-temperature mainstream fuel gas which is circumferentially uneven can be resisted, the blending temperature of the rim sealing cavity 4 can be reduced, and the rim sealing effect can be improved.
In some embodiments, the length extension of channels 22 is perpendicular to the median axis of the aircraft engine, and the length extension of channels 22 is perpendicular to the length extension of second blade 2. Here, the following. The length extension direction of the second blade 2 is the direction from the root to the tip of the second blade 2.
In some embodiments, as shown in fig. 3, the lengthwise extending sidewalls of channels 22 are configured to be linear.
In some embodiments, as shown in fig. 4, the lengthwise extending sidewalls of channels 22 are configured in an arcuate line.
In some embodiments, the channel width of channel 22 increases from the first end to the second end along the length of channel 22.
In some embodiments, the aeroengine rim sealing system further comprises a gas collecting cavity 3 and an exhaust hole 23 which are arranged at the root part of the second blade 2, the exhaust hole 23 is communicated with the gas collecting cavity 3 and the sealing cavity 4, and the gas collecting cavity 3 is configured to provide cold air to the sealing cavity 4 through the exhaust hole 23.
In some embodiments, as shown in fig. 1, the exhaust direction of the exhaust hole 23 faces the diversion trench 22, and the flow area of the cold air corresponding to the diversion trench 22 is increased, so that the loss of the cold air entering the diversion trench 22 can be reduced, the airflow pressure in the diversion trench 22 is increased, and it is beneficial to concentrate more cold air to resist the gas invasion; the guiding groove 22 guides the cold air to enable the cold air to flow around the central axis of the aircraft engine in a rotating mode, the front side of the cold air can resist gas invasion, the mixing temperature of the rim sealing cavity 4 is reduced, and the rim sealing effect is improved.
In other embodiments, the exhaust direction of the exhaust port 23 is inclined toward the central axis of the aircraft engine.
In some embodiments, the first flange 11 and the second flange 21 are both arranged in a circle around the central axis of the aircraft engine, the second flange 21 is provided with a plurality of guide grooves 22 at intervals in the circumferential direction, the root of the second blade 2 is provided with a plurality of exhaust holes 23 at intervals, each guide groove 22 of the plurality of guide grooves 22 is provided with at least one exhaust hole 23 of the plurality of exhaust holes 23, the exhaust hole 23 is configured to guide the cold air in the air collecting cavity 3 to the sealed cavity 4, the guide groove 22 is located in the sealed cavity 4, and the exhaust hole 23 directly guides the cold air to the guide groove 22.
Be equipped with a plurality of guiding gutters 22 in second flange 21's circumference, a plurality of guiding gutters 22 lead to the air conditioning in the chamber 4 of obturating, guiding gutter 22 is flaring type structure, the air conditioning flow area who corresponds increases, the air current static pressure risees, can concentrate more air conditioning and counter the regional and high-pressure area gas invasion of mainstream gas high pressure, and can make air conditioning have the rotation angle around aeroengine axis runner, openly resist the gas invasion, reduce the mixing temperature in the inside chamber of rim obturating, improve the rim effect of obturating.
In some embodiments, as shown in fig. 1, the root of the second blade 2 is provided with a honeycomb structure 5, the honeycomb structure 5 is located below the gas collecting cavity 3, that is, the honeycomb structure 5 is located on one side of the gas collecting cavity 3 close to the central axis of the aircraft engine, a comb tooth structure 6 adapted to the honeycomb structure 5 is located below the honeycomb structure 5, that is, the comb tooth structure 6 is close to the central axis of the aircraft engine relative to the honeycomb structure 5, and the comb tooth structure 6 and the first blade 1 are fixed on the same rotor member.
And a part of cold air in the sealing cavity 4 flows backwards through the honeycomb structure 5 and the labyrinth structure 6 and is supplied to the rim behind the second blade 2 for sealing.
In some embodiments, one side of the first blade 1 may be provided with two first flanges 11, and the second flange 21 of the second blade 2 is located between the two first flanges 11.
Some embodiments also provide an aircraft engine including the aircraft engine rim sealing system described above.
In some embodiments, the aircraft engine further comprises a turbine assembly comprising a turbine disc and a turbine casing, the first blade 1 being provided on the turbine disc and the second blade 2 being provided on the turbine casing.
In some embodiments, the first blade 1 is a moving blade, and the first blade 1 is provided with a first flange 11 on both sides. The second blade 2 is a guide vane, and both sides of the second blade 2 are also provided with second flanges 21.
The part of the second blade 2 above the second flange 21 is hollow to form a cavity, the part of the second blade 2 below the second flange 21 is provided with a gas collecting cavity 3 communicated with the cavity, and the second blade 2 is provided with an exhaust hole 23 communicated with the gas collecting cavity 3. The cold air enters the sealing cavity 4 between the first blade 1 and the second blade 2 from the cavity of the second blade 2 through the air collecting cavity 3 and the air exhaust hole 23.
The first flange 11 and the second flange 21 between the first blade 1 and the second blade 2 have axial mutual overlapping parts, certain gaps are arranged between the mutual overlapping parts in the radial direction, and the formed gaps are subjected to rim sealing through air flow.
The portion above the first edge plate 11 of the first vane 1 and the portion above the second edge plate 21 of the second vane 2 are in the main flow gas atmosphere of the main flow passage and have high temperatures. Air conditioning gets into the chamber 4 of obturating between first blade 1 and the second blade 2 through gas collecting chamber 3 and exhaust hole 23 by the cavity of second blade 2, the chamber 4 of obturating is located the first flange 11 of first blade 1 and the below of the second flange 21 of second blade 2, partly air conditioning in the chamber 4 of obturating discharges the sprue through the clearance between the overlapping position each other between first flange 11 and the second flange 21, prevent that high temperature mainstream gas from getting into in the chamber 4 of obturating, and then prevent that high temperature mainstream gas from getting into inside the engine, form the rim and seal.
Based on the embodiments of the invention described above, the technical features of one of the embodiments can be advantageously combined with one or more other embodiments without explicit negatives.
In the description of the present invention, it should be understood that the terms "first", "second", "third", etc. are used to define the components, and are used only for the convenience of distinguishing the components, and if not otherwise stated, the terms have no special meaning, and thus, should not be construed as limiting the scope of the present invention.
Finally, it should be noted that the above examples are only used to illustrate the technical solutions of the present invention and not to limit the same; although the present invention has been described in detail with reference to preferred embodiments, those skilled in the art will understand that: modifications to the specific embodiments of the invention or equivalent substitutions for parts of the technical features may be made; without departing from the spirit of the present invention, it is intended to cover all aspects of the invention as defined by the appended claims.
Claims (10)
1. An aircraft engine rim sealing system, comprising:
a first blade (1) provided with a first edge plate (11); and
the second blade (2) is arranged adjacent to the first blade (1), the second blade (2) is provided with a second flange plate (21), and a sealing cavity (4) is formed between the root of the second blade (2) and the root of the first blade (1); the second platform (21) is close to the central axis of the aircraft engine relative to the first platform (11), the second platform (21) and the first platform (11) have mutual overlapping positions, and a gap is formed between the mutual overlapping positions; a flow guide groove (22) is formed in one side, away from the first flange (11), of the second flange (21), and a first end of the flow guide groove (22) in the length extension direction is configured to guide part of the cold air in the sealing cavity (4) to a second end of the flow guide groove (22) in the length extension direction, so that the part of the cold air flows out through a gap between the mutually overlapped parts; the groove width of the first end of the flow guide groove (22) is smaller than that of the second end of the flow guide groove (22).
2. An aircraft engine rim sealing system according to claim 1, characterised in that the flow guide channel (22) is configured to guide a portion of the cooling air in the sealing chamber (4) in such a way that it flows in a rotating manner around the central axis of the aircraft engine and out through the gap between the mutually overlapping points.
3. An aircraft engine rim sealing system according to claim 1, characterised in that the length extension of the channels (22) is perpendicular to the central axis of the aircraft engine and to the length extension of the second blades (2).
4. An aircraft engine rim sealing system according to claim 1, characterised in that the side walls of the channels (22) are configured as rectilinear.
5. An aircraft engine rim sealing system according to claim 1, characterised in that the side walls of the channels (22) are configured as arcs.
6. An aircraft engine rim sealing system according to any one of claims 1 to 5, characterised in that the channel width of the channel (22) increases progressively from the first end to the second end along the length extension of the channel (22).
7. An aircraft engine rim sealing system according to claim 1, further comprising a gas collection chamber (3) and a gas discharge hole (23) provided at the root of the second blade (2), the gas discharge hole (23) communicating the gas collection chamber (3) and the sealing chamber (4), the gas collection chamber (3) being configured to provide cold gas to the sealing chamber (4) through the gas discharge hole (23).
8. An aircraft engine rim sealing system according to claim 7, characterized in that the exhaust direction of the exhaust hole (23) is towards the diversion trench (22), or the exhaust direction of the exhaust hole (23) is inclined towards the central axis of the aircraft engine.
9. An aircraft engine rim sealing system according to claim 7, wherein the first rim plate (11) and the second rim plate (21) are arranged one turn around the central axis of the aircraft engine, the second rim plate (21) is provided with a plurality of the guiding grooves (22) at intervals in the circumferential direction, the root of the second blade (2) is provided with a plurality of the exhaust holes (23) at intervals, each guiding groove (22) in the plurality of guiding grooves (22) is provided with at least one exhaust hole (23) in the plurality of exhaust holes (23), and the exhaust hole (23) is configured to guide the cold air in the air collection chamber (3) to the guiding groove (22).
10. An aircraft engine, characterized in that it comprises a turbine assembly and an aircraft engine rim sealing system according to any one of claims 1 to 9, said turbine assembly comprising a turbine disk and a turbine casing, said first blades (1) being provided on said turbine disk and said second blades (2) being provided on said turbine casing.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110205800.7A CN114961870A (en) | 2021-02-24 | 2021-02-24 | Aeroengine rim sealing system and aeroengine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110205800.7A CN114961870A (en) | 2021-02-24 | 2021-02-24 | Aeroengine rim sealing system and aeroengine |
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| Publication Number | Publication Date |
|---|---|
| CN114961870A true CN114961870A (en) | 2022-08-30 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202110205800.7A Pending CN114961870A (en) | 2021-02-24 | 2021-02-24 | Aeroengine rim sealing system and aeroengine |
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| CN (1) | CN114961870A (en) |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1120694A (en) * | 1965-08-02 | 1968-07-24 | Snecma | Improvements in cooled turbine blades |
| CN1755063A (en) * | 2004-09-28 | 2006-04-05 | 斯奈克玛公司 | Turbine overspeed limiting device |
| CN104196574A (en) * | 2014-07-15 | 2014-12-10 | 西北工业大学 | Gas turbine cooling blade |
| CN105971674A (en) * | 2016-07-29 | 2016-09-28 | 上海电气燃气轮机有限公司 | Gas turbine rim sealing structure and method |
| CN107869362A (en) * | 2016-09-26 | 2018-04-03 | 中国航发商用航空发动机有限责任公司 | Rim sealing structure, turbine and gas turbine |
| CN109296402A (en) * | 2017-07-25 | 2019-02-01 | 中国航发商用航空发动机有限责任公司 | Labyrinth gas seals structure and aero-engine |
| CN109779696A (en) * | 2019-02-12 | 2019-05-21 | 中国科学院工程热物理研究所 | A kind of aperture rim sealing structure with fluidal texture adaptability |
-
2021
- 2021-02-24 CN CN202110205800.7A patent/CN114961870A/en active Pending
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1120694A (en) * | 1965-08-02 | 1968-07-24 | Snecma | Improvements in cooled turbine blades |
| CN1755063A (en) * | 2004-09-28 | 2006-04-05 | 斯奈克玛公司 | Turbine overspeed limiting device |
| CN104196574A (en) * | 2014-07-15 | 2014-12-10 | 西北工业大学 | Gas turbine cooling blade |
| CN105971674A (en) * | 2016-07-29 | 2016-09-28 | 上海电气燃气轮机有限公司 | Gas turbine rim sealing structure and method |
| CN107869362A (en) * | 2016-09-26 | 2018-04-03 | 中国航发商用航空发动机有限责任公司 | Rim sealing structure, turbine and gas turbine |
| CN109296402A (en) * | 2017-07-25 | 2019-02-01 | 中国航发商用航空发动机有限责任公司 | Labyrinth gas seals structure and aero-engine |
| CN109779696A (en) * | 2019-02-12 | 2019-05-21 | 中国科学院工程热物理研究所 | A kind of aperture rim sealing structure with fluidal texture adaptability |
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