CN113250856B - Aircraft engine spray tube expansion section cooling structure - Google Patents
Aircraft engine spray tube expansion section cooling structure Download PDFInfo
- Publication number
- CN113250856B CN113250856B CN202110491756.0A CN202110491756A CN113250856B CN 113250856 B CN113250856 B CN 113250856B CN 202110491756 A CN202110491756 A CN 202110491756A CN 113250856 B CN113250856 B CN 113250856B
- Authority
- CN
- China
- Prior art keywords
- cooling
- heat insulation
- aircraft engine
- plate
- expansion section
- 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.)
- Active
Links
- 238000001816 cooling Methods 0.000 title claims abstract description 103
- 239000007921 spray Substances 0.000 title claims description 6
- 238000009413 insulation Methods 0.000 claims abstract description 61
- 239000007789 gas Substances 0.000 description 15
- 230000003068 static effect Effects 0.000 description 12
- 239000000112 cooling gas Substances 0.000 description 11
- 230000007423 decrease Effects 0.000 description 4
- 230000006837 decompression Effects 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 238000006467 substitution reaction Methods 0.000 description 2
- 230000007547 defect Effects 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 230000007717 exclusion Effects 0.000 description 1
- 230000000750 progressive effect Effects 0.000 description 1
- 230000005855 radiation Effects 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/78—Other construction of jet pipes
- F02K1/82—Jet pipe walls, e.g. liners
- F02K1/822—Heat insulating structures or liners, cooling arrangements, e.g. post combustion liners; Infrared radiation suppressors
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/78—Other construction of jet pipes
- F02K1/82—Jet pipe walls, e.g. liners
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The application belongs to the technical field of the cooling design of jet pipe expansion section of aircraft engine, concretely relates to jet pipe expansion section cooling structure of aircraft engine, include: an aircraft engine nozzle expansion section; the heat insulation plate is arranged on the inner side of the jet pipe expansion section of the aircraft engine and forms a cooling channel with the jet pipe expansion section of the aircraft engine; and the corrugated pressure reducing plate is arranged in the cooling channel, is connected to the expansion section of the jet pipe of the aircraft engine, has a gap with the heat insulating plate, and has an upper peak-trough extending from the inlet end of the cooling channel to the outlet end.
Description
Technical Field
The application belongs to the technical field of cooling design of an expansion section of an aircraft engine spray pipe, and particularly relates to a cooling structure of the expansion section of the aircraft engine spray pipe.
Background
The jet pipe of the aero-engine is used for exhausting engine tail gas, and comprises a jet pipe connecting cylinder, a convergence section and an expansion section, wherein the inlet end of the jet pipe connecting cylinder is connected to an exhaust port of the aero-engine, the convergence section is hinged to the outlet end of the jet pipe connecting cylinder, the expansion section is hinged to the outlet end of the convergence section, and the radial size of the convergence section gradually shrinks from the inlet end to the outlet end; the radial dimension of the expanding section gradually expands from the inlet end to the outlet end.
The expansion section in some types of aircraft engine spray pipes bears higher heat load, in order to reduce infrared radiation and ensure stealth performance, the inner side of the expansion section is mostly provided with a heat insulation plate at present, and cooling air is introduced into a cooling channel formed between the expansion section and the heat insulation plate to cool the expansion section, and the scheme has better cooling effect on the expansion section, but has the following defects:
when tail gas in the expansion section of the jet pipe of the aircraft engine reaches supersonic speed, the tail gas flows on the inner side of the heat insulation plate, the static pressure of the tail gas is sharply reduced along the flow direction, and the static pressure of cooling gas flowing on the outer side of the heat insulation plate is not obviously reduced along the flow direction, so that huge unbalanced pressure is generated on two sides of the heat insulation plate, and the heat insulation plate bears huge unbalanced force and is easy to seriously deform so as to be damaged.
The present application has been made in view of the above-mentioned technical drawbacks.
It should be noted that the above background disclosure is only for the purpose of assisting understanding of the inventive concept and technical solutions of the present invention, and does not necessarily belong to the prior art of the present patent application, and the above background disclosure should not be used for evaluating the novelty and inventive step of the present application without explicit evidence to suggest that the above content is already disclosed at the filing date of the present application.
Disclosure of Invention
It is an object of the present application to provide an aircraft engine nozzle flare cooling arrangement that overcomes or mitigates at least one of the known disadvantages.
The technical scheme of the application is as follows:
an aircraft engine nozzle flare cooling structure comprising:
an aircraft engine nozzle expansion section;
the heat insulation plate is arranged on the inner side of the jet pipe expansion section of the aircraft engine and forms a cooling channel with the jet pipe expansion section of the aircraft engine;
and the corrugated pressure reducing plate is arranged in the cooling channel, is connected to the expansion section of the jet pipe of the aircraft engine, has a gap with the heat insulating plate, and has wave crests-wave troughs extending from the inlet end of the cooling channel to the outlet end.
According to at least one embodiment of the application, in the aircraft engine nozzle expansion segment cooling structure, the gap between each wave trough part of the corrugated pressure reducing plate and the heat insulation plate is gradually reduced from the inlet end to the outlet end of the cooling channel.
According to at least one embodiment of the application, in the above-mentioned aircraft engine nozzle extension cooling structure, L2 ═ a · L1;
wherein,
l1 is the distance between the midline of the corrugated pressure reducing plate and the inlet end of the heat shield at the cooling channel;
l2 is the distance between the midline of the corrugated pressure reducing plate and the heat insulating plate at the outlet end of the cooling channel;
a is a reduction coefficient and is within 0.3-0.7.
According to at least one embodiment of the application, in the cooling structure of the nozzle expanding section of the aircraft engine, the heat insulation plate is provided with a plurality of heat insulation cooling holes; the heat insulation cooling holes are distributed in a direction from the inlet end to the outlet end of the cooling channel.
According to at least one embodiment of the present application, the above-mentioned cooling structure for the nozzle extension section of an aircraft engine further comprises:
the impact plate is arranged on the inner side of the heat insulation plate, a heat insulation gap is formed between the impact plate and the heat insulation plate, and a plurality of impact cooling holes are formed in the impact plate; the impingement cooling holes are distributed in a direction from the inlet end to the outlet end of the cooling channel.
According to at least one embodiment of the application, in the cooling structure of the nozzle expanding section of the aircraft engine, the heat insulation cooling holes and the impingement cooling holes are arranged in a staggered mode.
Drawings
FIG. 1 is a schematic illustration of an aircraft engine nozzle flare cooling configuration provided in an embodiment of the present application;
FIG. 2 is a schematic diagram of a cooling structure of an expansion section of an aircraft engine nozzle provided by an embodiment of the application;
FIG. 3 is a schematic view of the pressure distribution on both sides of a heat shield plate based on a conventional nozzle divergent section cooling structure of an aircraft engine;
FIG. 4 is a schematic diagram illustrating pressure distribution on two sides of a heat insulation plate of a cooling structure of a nozzle expansion section of an aircraft engine provided according to an embodiment of the application;
wherein:
1-an aircraft engine nozzle expansion section; 2-insulating board; 3-corrugated step-down plate; 4-an impact plate;
H1-H5 is the clearance between each wave trough part of the corrugated decompression plate and the heat insulation plate.
For the purpose of better illustrating the present embodiments, certain elements of the drawings may be omitted, enlarged or reduced, and do not represent the size of an actual product; furthermore, the drawings are for illustrative purposes, and the terms used to describe positional relationships are merely exemplary in nature and are not to be construed as limiting the present patent.
Detailed Description
In order to make the technical solutions and advantages of the present application clearer, the technical solutions of the present application will be described in detail with reference to the accompanying drawings, and it should be understood that the specific embodiments described herein are only some of the embodiments of the present application, and are used for explaining the present application and not limiting the present application. It should be noted that, for convenience of description, only the parts related to the present application are shown in the drawings, other related parts may refer to general designs, and the embodiments and technical features in the embodiments in the present application may be combined with each other to obtain a new embodiment without conflict.
In addition, unless otherwise defined, technical or scientific terms used in the description of the present application shall have the ordinary meaning as understood by one of ordinary skill in the art to which the present application belongs. The terms "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inner", "outer", and the like used in the description of the present application, which indicate orientations, are used only to indicate relative directions or positional relationships, and do not imply that devices or elements must have specific orientations, be constructed and operated in specific orientations, and that when the absolute position of an object to be described is changed, the relative positional relationships may be changed accordingly, and thus, should not be construed as limiting the present application. The use of "first," "second," "third," and the like in the description of the present application is for descriptive purposes only to distinguish between different components and is not to be construed as indicating or implying relative importance. The use of the terms "a," "an," or "the" and similar referents in the context of describing the application is not to be construed as an absolute limitation on the number, but rather as the presence of at least one. The word "comprising" or "comprises", and the like, when used in this description, is intended to specify the presence of stated elements or items, but not the exclusion of any other elements or items.
Further, it is noted that, unless expressly stated or limited otherwise, the terms "mounted," "connected," and the like are used in the description of the invention in a generic sense, e.g., connected as either a fixed connection or a removable connection or integrally connected; can be mechanically or electrically connected; they may be directly connected or indirectly connected through an intermediate medium, or they may be connected through the inside of two elements, and those skilled in the art can understand their specific meaning in this application according to the specific situation.
The present application will be described in further detail with reference to fig. 1 to 4.
An aircraft engine nozzle flare cooling structure comprising:
an aircraft engine nozzle expansion section 1;
the heat insulation plate 2 is arranged on the inner side of the jet pipe expansion section 1 of the aircraft engine and forms a cooling channel with the jet pipe expansion section 1 of the aircraft engine;
and the corrugated pressure reducing plate 3 is arranged in the cooling channel, is connected to the nozzle expansion section 1 of the aircraft engine, has a gap with the heat insulation plate 2, and extends from the inlet end to the outlet end of the cooling channel along the wave crest-wave trough.
For the cooling structure of the expansion section of the jet pipe of the aircraft engine disclosed in the above embodiment, as can be understood by those skilled in the art, the heat insulation plate 2 is arranged on the inner side of the expansion section 1 of the jet pipe of the aircraft engine, so that the heat transfer rate of tail gas of the aircraft engine to the expansion section 1 of the jet pipe of the aircraft engine can be reduced, and cooling gas is introduced into a cooling channel between the heat insulation plate 2 and the expansion section 1 of the jet pipe of the aircraft engine, so that the heat transfer rate of tail gas of the aircraft engine to the expansion section 1 of the jet pipe of the aircraft engine can be further reduced, and heat on the expansion section 1 of the jet pipe of the aircraft engine can be taken away, thereby effectively cooling the expansion section 1 of the jet pipe of the aircraft engine.
For the cooling structure of the nozzle expansion section of the aircraft engine disclosed in the above embodiment, it can be further understood by those skilled in the art that the cooler introduced into the cooling channel flows between the gap between the heat insulating plate 2 and the corrugated pressure-reducing plate 3, and during the flow of the cooling gas from the inlet end to the outlet end of the cooling channel, a significant pressure-expanding loss is generated when the cooling gas flows from the trough portion to the peak portion of each of the corrugated pressure-reducing plates 3, so that the static pressure of the cooling gas can be sharply reduced along the flow direction, and in the specific application, the structural parameters of the corrugated pressure-reducing plate 3 are designed for the actual situation that the tail gas in the nozzle expansion section of the aircraft engine reaches supersonic speed and the static pressure of the tail gas flowing inside the heat insulating plate is sharply reduced along the flow direction, so that the degree of reduction of the static pressure of the cooling gas outside the heat insulating plate 2 along the flow direction is matched with the degree of reduction of the static pressure of the tail gas inside the heat insulating plate 2 along the flow direction, therefore, the pressure difference between the two sides of the heat insulation plate 2 can be maintained, and the heat insulation plate 2 is prevented from being seriously deformed and damaged.
In some alternative embodiments, in the above-mentioned aircraft engine nozzle divergent section cooling structure, gaps between each of the wave trough portions of the corrugated decompression plate 3 and the heat insulating plate 2 gradually decrease from the inlet end to the outlet end of the cooling channel, as shown in fig. 2H 1> H2> H3> H4> H5, specific values may be designed by those skilled in the relevant art in a specific application according to a practical situation that static pressure of the exhaust gas flowing inside the heat insulating plate sharply decreases along the flow direction, so as to effectively control a decrease degree of the static pressure of the cooling gas outside the heat insulating plate 2 along the flow direction to match a decrease degree of the static pressure of the exhaust gas inside the heat insulating plate 2 along the flow direction.
In some alternative embodiments, in the above-mentioned nozzle diverging section cooling structure of an aircraft engine, L2 is a · L1;
wherein,
l1 is the distance between the central line of the corrugated pressure reducing plate 3 and the inlet end of the heat insulation plate 2 in the cooling channel;
l2 is the distance between the central line of the corrugated pressure reducing plate 3 and the outlet end of the heat insulation plate 2 in the cooling channel;
and a is a reduction coefficient and is used for restricting the change degree of the gap between the wave trough part of the corrugated pressure reducing plate 3 positioned at the outlet end of the cooling channel and the heat insulation plate 2 relative to the gap between the wave trough part of the corrugated pressure reducing plate 3 positioned at the inlet end of the cooling channel and the heat insulation plate 2, and the value of the change degree is limited within 0.3-0.7, so that the flow stability of cooling airflow can be kept while the reduction of the cooling gas static pressure outside the heat insulation plate 2 along the flow direction is effectively controlled.
In some alternative embodiments, in the above-mentioned nozzle divergent section cooling structure of an aircraft engine, the heat insulation plate 2 has a plurality of heat insulation cooling holes; the heat insulation cooling holes are distributed in a direction from the inlet end to the outlet end of the cooling channel.
With regard to the cooling structure of the nozzle divergent section of the aircraft engine disclosed in the above embodiment, it can be understood by those skilled in the art that the cooling air flow introduced into the cooling channel may flow out from each heat insulation cooling hole, and the cooling air flowing out from each heat insulation cooling hole may form an air pattern inside the heat insulation plate 2, so that the heat transfer rate of the aircraft engine exhaust to the nozzle divergent section 1 of the aircraft engine may be effectively reduced.
In some optional embodiments, the above-mentioned aircraft engine nozzle diverging section cooling structure further comprises:
an impingement plate 4 disposed inside the insulation plate 2 with an insulation gap formed therebetween and having a plurality of impingement cooling holes; the impingement cooling holes are arranged from the inlet end to the outlet end of the cooling channel.
For the cooling structure of the nozzle expansion section of the aircraft engine disclosed in the above embodiment, it can be understood by those skilled in the art that the tail gas exhausted from the aircraft engine flows outwards in the impingement plate 4, the cooling gas flow introduced into the cooling channel can flow out from each heat insulation cooling hole, the cooling gas flowing out from each heat insulation cooling hole can enter the heat insulation plate 2 and the impingement plate 4 to form a heat insulation gap, the flow in the heat insulation gap can reduce the heat transfer rate of the tail gas of the aircraft engine to the nozzle expansion section 1 of the aircraft engine, in addition, the cooling gas entering the heat insulation gap can flow out from each impingement cooling hole, and an air mold can be formed inside the impingement plate 4, so that the heat transfer rate of the tail gas of the aircraft engine to the nozzle expansion section 1 of the aircraft engine can be effectively reduced.
In some optional embodiments, in the above-mentioned nozzle divergent cooling structure of an aircraft engine, the respective heat-insulating cooling holes and the impingement cooling holes are arranged in a staggered manner, so as to prevent the cooling air entering the heat-insulating gap from the respective heat-insulating cooling holes from directly flowing out of the respective impingement cooling holes.
In a specific example, the pressure distribution measured on the two sides of the heat insulation plate based on the existing cooling structure for the nozzle extension section of the aircraft engine is shown in fig. 3, and the pressure distribution measured on the two sides of the heat insulation plate based on the cooling structure for the nozzle extension section of the aircraft engine provided in the embodiment of the present application is shown in fig. 4, by contrast, the cooling structure for the nozzle extension section of the aircraft engine provided in the embodiment of the present application can make the degree of reduction of the cooling static pressure outside the heat insulation plate 2 in the flow direction match with the degree of reduction of the static pressure of the tail gas inside the heat insulation plate 2 in the flow direction, so that the pressure difference on the two sides of the heat insulation plate 2 can be maintained, and the heat insulation plate 2 is prevented from being seriously deformed and damaged.
The embodiments are described in a progressive manner in the specification, each embodiment focuses on differences from other embodiments, and the same and similar parts among the embodiments are referred to each other.
Having thus described the present invention in connection with the preferred embodiments illustrated in the accompanying drawings, it will be understood by those skilled in the art that the scope of the present invention is not limited to those specific embodiments, and that equivalent changes or substitutions of the related technical features may be made by those skilled in the art without departing from the principle of the present invention, and those technical aspects after such changes or substitutions will fall within the scope of the present invention.
Claims (5)
1. An aircraft engine nozzle flare cooling structure, comprising:
an aircraft engine nozzle expansion section (1);
the heat insulation plate (2) is arranged on the inner side of the jet pipe expansion section (1) of the aircraft engine and forms a cooling channel with the jet pipe expansion section (1) of the aircraft engine;
the corrugated pressure reducing plate (3) is arranged in the cooling channel, is connected to the aircraft engine spray pipe expansion section (1), has a gap with the heat insulation plate (2), and has wave crests and wave troughs extending from the inlet end to the outlet end of the cooling channel;
gaps between the wave trough parts of the corrugated pressure reducing plate (3) and the heat insulation plate (2) are gradually reduced from the inlet end to the outlet end of the cooling channel.
2. The aero engine nozzle diverging section cooling arrangement of claim 1,
L2=a•L1；
wherein,
l1 is the distance between the center line of the corrugated pressure reducing plate (3) and the inlet end of the heat insulation plate (2) in the cooling passage;
l2 is the distance between the center line of the corrugated pressure reducing plate (3) and the heat insulation plate (2) at the outlet end of the cooling channel;
a is a reduction coefficient and is between 0.3 and 0.7.
3. The aero engine nozzle diverging section cooling arrangement of claim 1,
the heat insulation plate (2) is provided with a plurality of heat insulation cooling holes; and all the heat insulation cooling holes are distributed from the inlet end to the outlet end of the cooling channel in an arrayed manner.
4. The aero engine nozzle diverging section cooling arrangement of claim 3,
further comprising:
the impingement plate (4) is arranged on the inner side of the heat insulation plate (2), an insulation gap is formed between the impingement plate and the heat insulation plate (2), and a plurality of impingement cooling holes are formed in the impingement plate; and all the impingement cooling holes are distributed in an array from the inlet end to the outlet end of the cooling channel.
5. The aero engine nozzle flare cooling structure of claim 4,
and all the heat insulation cooling holes and the impingement cooling holes are distributed in a staggered manner.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110491756.0A CN113250856B (en) | 2021-05-06 | 2021-05-06 | Aircraft engine spray tube expansion section cooling structure |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110491756.0A CN113250856B (en) | 2021-05-06 | 2021-05-06 | Aircraft engine spray tube expansion section cooling structure |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN113250856A CN113250856A (en) | 2021-08-13 |
| CN113250856B true CN113250856B (en) | 2022-07-15 |
Family
ID=77223750
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202110491756.0A Active CN113250856B (en) | 2021-05-06 | 2021-05-06 | Aircraft engine spray tube expansion section cooling structure |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN113250856B (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114687888B (en) * | 2022-04-14 | 2024-01-30 | 中国航发沈阳发动机研究所 | Binary vector spray pipe cooling structure |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH01253555A (en) * | 1987-04-14 | 1989-10-09 | United Technol Corp <Utc> | Cooling type liner assembly for exhaust nozzle of gas turbine |
| CN1637247A (en) * | 2004-01-09 | 2005-07-13 | 联合工艺公司 | Extended impingement cooling device and method |
| CN104863750A (en) * | 2015-05-07 | 2015-08-26 | 南京航空航天大学 | Impingement and air-film cooling structure adopting variable-hole array pitches used for wall surface of jet tube |
| CN109779782A (en) * | 2019-03-08 | 2019-05-21 | 西北工业大学 | The double wall cooling structure with longitudinal ripple impact orifice plate for vector spray |
| CN111577481A (en) * | 2020-05-26 | 2020-08-25 | 中国航发沈阳发动机研究所 | Cooling channel structure suitable for binary stealthy spray tube |
-
2021
- 2021-05-06 CN CN202110491756.0A patent/CN113250856B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH01253555A (en) * | 1987-04-14 | 1989-10-09 | United Technol Corp <Utc> | Cooling type liner assembly for exhaust nozzle of gas turbine |
| CN1637247A (en) * | 2004-01-09 | 2005-07-13 | 联合工艺公司 | Extended impingement cooling device and method |
| CN104863750A (en) * | 2015-05-07 | 2015-08-26 | 南京航空航天大学 | Impingement and air-film cooling structure adopting variable-hole array pitches used for wall surface of jet tube |
| CN109779782A (en) * | 2019-03-08 | 2019-05-21 | 西北工业大学 | The double wall cooling structure with longitudinal ripple impact orifice plate for vector spray |
| CN111577481A (en) * | 2020-05-26 | 2020-08-25 | 中国航发沈阳发动机研究所 | Cooling channel structure suitable for binary stealthy spray tube |
Also Published As
| Publication number | Publication date |
|---|---|
| CN113250856A (en) | 2021-08-13 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| EP1990507B1 (en) | Impingement cooling structure | |
| US20240159151A1 (en) | Airfoil for a turbine engine | |
| EP2702250B1 (en) | A method of forming a multi-panel outer wall of a component for use in a gas turbine engine | |
| JP5579011B2 (en) | Gas turbine combustor | |
| JP5080076B2 (en) | Thermal control of gas turbine engine ring for active clearance control | |
| EP3719432B1 (en) | Heat exchanger fractal splitter | |
| EP2685048B1 (en) | Gas turbine rotor blade, and gas turbine | |
| US7464554B2 (en) | Gas turbine combustor heat shield panel or exhaust panel including a cooling device | |
| US5711650A (en) | Gas turbine airfoil cooling | |
| EP3190271A1 (en) | Heat exchanger for embedded engine applications | |
| US8668438B2 (en) | Turbine casing cooling | |
| US20100303610A1 (en) | Cooled gas turbine stator assembly | |
| US9145789B2 (en) | Impingement plate for damping and cooling shroud assembly inter segment seals | |
| EP2196634B1 (en) | Cavity ventilation | |
| US20190085705A1 (en) | Component for a turbine engine with a film-hole | |
| JP2007162698A (en) | System and method to exhaust spent cooling air of gas turbine engine active clearance control | |
| EP3128130B1 (en) | Partial cavity baffles for airfoils in gas turbine engines | |
| WO2015134005A1 (en) | Turbine airfoil | |
| JP2017129116A (en) | Heat exchanger for embedded engine applications: transduct segments | |
| CN113250856B (en) | Aircraft engine spray tube expansion section cooling structure | |
| KR20150110367A (en) | Turbine vane with cooled fillet | |
| US20140056719A1 (en) | Gas Turbine, Gas Turbine Blade, and Manufacturing Method of Gas Turbine Blade | |
| US20090238695A1 (en) | Full coverage trailing edge microcircuit with alternating converging exits | |
| CN111578310A (en) | Air film cooling hole structure for turboshaft engine | |
| US7481059B2 (en) | Method and apparatus for providing an afterburner fuel-feed arrangement |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |