CN214247529U - Aircraft engine core tail nozzle and aircraft engine core - Google Patents
Aircraft engine core tail nozzle and aircraft engine core Download PDFInfo
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- CN214247529U CN214247529U CN202120224562.XU CN202120224562U CN214247529U CN 214247529 U CN214247529 U CN 214247529U CN 202120224562 U CN202120224562 U CN 202120224562U CN 214247529 U CN214247529 U CN 214247529U
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- nozzle body
- tail
- aircraft engine
- jet nozzle
- engine core
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Abstract
The utility model discloses an aeroengine core tail spray tube and aeroengine core machine relates to the aeroengine field for improve the cooling to the tail spray tube. The aircraft engine core tail nozzle comprises a tail nozzle body, a cover body and a gas guide pipe. The jet nozzle body is configured to be annular and has a channel; the wall body of the tail nozzle body is provided with a gas film hole which is communicated with the channel. The cover body is arranged on the outer side of the tail nozzle body, and a cavity is defined between the inner wall of the cover body and the outer wall of the tail nozzle body. The bleed air duct is mounted to the cover and is in fluid communication with the cavity. The aircraft engine core aircraft tail nozzle that above-mentioned technical scheme provided for introduce cooling air from the test bench, cooling air forms the air film at the tail nozzle inner wall, thereby realizes the cooling of whole tail nozzle wall, and then realizes the cooling to the core aircraft nozzle, reduces or even has avoided the high temperature to bring the quick reduction of material mechanical properties, has improved the high temperature resistant ability of tail nozzle structure.
Description
Technical Field
The utility model relates to an aeroengine field, concretely relates to aeroengine core tail spray tube and aeroengine core machine.
Background
The core engine of the aircraft engine is test equipment used for verifying the performance of a high-pressure part, and high-pressure gas passing through a gas compressor, a combustion chamber and a turbine can be discharged in the test process.
The inventor finds that at least the following problems exist in the prior art: because the temperature of the airflow is too high, the temperature of the wall surface of the tail nozzle is higher due to the gas discharge, and simultaneously, larger thermal deformation thermal stress is caused, thus providing severe requirements on the high temperature resistance and other performances and the structural strength of the material of the tail nozzle.
SUMMERY OF THE UTILITY MODEL
The utility model provides an aeroengine core tail-nozzle and aeroengine core machine for improve the cooling to the tail-nozzle.
The embodiment of the utility model provides an aeroengine core tail spray tube, include:
a jet nozzle body configured to be annular and having a passageway; the wall body of the tail spray pipe body is provided with a gas film hole, and the gas film hole is communicated with the channel;
the cover body is arranged on the outer side of the tail nozzle body, and a cavity is defined between the inner wall of the cover body and the outer wall of the tail nozzle body; and
a bleed tube mounted to the cover and in fluid communication with the cavity.
In some embodiments, the film holes are disposed in rows in the jet nozzle body.
In some embodiments, the plurality of rows of the film holes are arranged along an axial direction of the jet nozzle body.
In some embodiments, a plurality of turns of the film holes are arranged along a circumferential direction of the jet nozzle body.
In some embodiments, the axis of each of the film holes is at a set angle to the wall in which the film hole is disposed.
In some embodiments, the set angle is 25 ° to 35 °.
In some embodiments, a plurality of the bleed air ducts are provided along a circumferential direction of the jet nozzle body.
In some embodiments, the cover is configured to be annular.
In some embodiments, the cover comprises:
a first plate extending radially of the jet nozzle body; one axial end of the tail nozzle body of the first plate is attached and fixed; and
the second plate extends along the axial direction of the jet nozzle body, the first plate is fixedly connected with the second plate, and one end, far away from the first plate, of the second plate is fixed with the axial middle part or the other end of the jet nozzle body.
The embodiment of the utility model provides a still provide an aeroengine core machine, include the utility model discloses the aeroengine core tail-nozzle that any technical scheme provided.
The aircraft engine core engine tail nozzle provided by the technical scheme is used for introducing cooling air flow from the test bed, discharging the cooling air flow through the densely distributed air film holes in the wall surface of the tail nozzle body, and enabling the air flow to converge into heat flow after passing through the air film holes. The cooling air flow forms the air film at the jet nozzle inner wall to realize the cooling of whole jet nozzle wall, and then realize the cooling to the core engine spray tube, reduce and avoided the high temperature even to bring the quick reduction of material mechanical properties, improved the high temperature resistant ability of jet nozzle structure.
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 undue limitation to the invention. In the drawings:
fig. 1 is a schematic perspective view of a core nozzle of an aircraft engine according to an embodiment of the present invention;
FIG. 2 is a schematic perspective view of a nozzle body of an aircraft engine core nozzle according to an embodiment of the present invention;
FIG. 3 is an enlarged view of part A of FIG. 2;
fig. 4 is a schematic partial cross-sectional view of an aircraft engine core nozzle according to an embodiment of the present invention.
Detailed Description
The technical solution provided by the present invention will be explained in more detail with reference to fig. 1 to 4.
Technical terms or nouns used in the embodiments of the present invention need to be interpreted.
A core machine: the test piece in the development process of the aero-engine is used for verifying the performance of high-voltage components, obtaining the working characteristics and matching performance of each component system and providing support for the design and test of the whole machine. The core engine does not do work during the test, so the temperature of the components of the core engine can be very high, which makes the temperature of the exhaust nozzle body at the tail end of the engine core engine extremely high.
The tail nozzle body: the exhaust structure is positioned at the tail end of the core engine and mainly has the function of exhausting high-temperature airflow generated by the core engine and providing certain thrust.
The embodiment of the utility model provides an aeroengine core tail nozzle, including tail nozzle body 1, the cover body 2 and bleed pipe 3.
Is configured as a ring and has a channel 11; the wall body of the tail nozzle body 1 is provided with a gas film hole 12, and the gas film hole 12 is communicated with the channel 11. The jet nozzle body 1 is annular, the passage 11 of which is intended to be passed through by a high-temperature gas flow. Referring to fig. 2 and 3, in some embodiments, the film holes 12 are arranged in rows in the jet nozzle body 1. The number of the air film holes 12 on the tail nozzle body 1 is large, so that a large amount of cooling air can enter the channel 11 of the tail nozzle body 1 through the air film holes 12, then a large amount of air films are formed on the surface of the inner wall of the tail nozzle body 1, the air films play a cooling and cooling role on the tail nozzle body 1, and the temperature of the tail nozzle body 1 is greatly reduced.
In some embodiments, multiple rows of film holes 12 are arranged along the axial direction of the jet nozzle body 1. The air inlet flow channel formed by the multiple rows of air film holes 12 has large area, large amount of cold air entering and good cooling effect.
In some embodiments, multiple circles of film holes 12 are arranged along the circumferential direction of the jet nozzle body 1. The air inlet flow channel formed by the multiple circles of air film holes 12 is large in area, large in cold air inlet quantity and good in cooling effect.
In some embodiments, the axis of each film hole 12 is at a set angle α to the wall in which the film hole 12 is disposed. The angle α is not preferably set to 90 ° because it is unlikely to form a gas film on the inner wall surface of the jet nozzle body 1. The included angle between each air film hole 12 and the wall body where the air film hole is located is a set included angle. The jet nozzle body 1 is not a standard cylinder, so that the angles of the film openings 12 located at different positions in the axial direction of the jet nozzle body 1 relative to the axis of the jet nozzle body 1 are also different.
In some embodiments, the included angle is set to 25 ° to 35 °. Specifically, for example, 25 °, 30 °, 35 °, etc. These angles enable a large amount of air film to be easily formed on the inner wall surface of the jet nozzle body 1, thereby ensuring the cooling effect.
Referring to fig. 1 and 4, the cover body 2 is disposed outside the jet nozzle body 1, and a cavity 20 is defined between an inner wall of the cover body 2 and an outer wall of the jet nozzle body 1. The cavity 20 plays a role in pressurization and pressure stabilization, so that airflow introduced by the subsequent air guide pipe 3 can smoothly enter the cavity 20 and then flows through the air film hole 12 to form an air film on the inner wall surface of the tail nozzle body 1. The size of the cavity 20 is greater than the sum of the flow areas of all the film holes 12. The shape of the cavity 20 is not required.
Referring to fig. 1 and 4, in some embodiments, the cover 2 is configured to be annular. The structure of the cover body 2 is substantially matched with that of the jet nozzle body 1. The jet nozzle body 1 is substantially in the shape of a frustum, and a bending edge is provided at an upstream end portion (i.e., one end in the axial direction and one end portion having a large inner diameter of the frustum) of the jet nozzle body 1.
Referring to fig. 4, in some embodiments, the enclosure 2 includes a first panel 21 and a second panel 22. The first plate 21 extends in the radial direction of the jet nozzle body 1. The axial one end laminating and the fixed of first board 21 tail nozzle body 1 specifically can welded fastening. The first plate 21 plays a role of fixed connection, and the length of the first plate 21 in the radial direction of the jet nozzle body 1 is smaller than the size of the bent edge of the jet nozzle body 1. The second plate 22 extends along the axial direction of the jet nozzle body 1, the first plate 21 and the second plate 22 are fixedly connected, and one end of the second plate 22, which is far away from the first plate 21, is fixed with the axial middle part or the other end of the jet nozzle body 1. The end of the second plate 22 remote from the first plate 21 is relatively thick, so that a substantial area of the second plate 22 is free from the jet body 1, which gap serves to form the cavity 20.
Referring to fig. 1 and 4, a bleed air duct 3 is mounted to the housing 2, the bleed air duct 3 being in fluid communication with the cavity 20. The bleed air duct 3 serves to introduce outside cooling air into the interior of the bleed air duct 3. Bleed pipe 3 is connected with the bleed device of core rack, acquires the air as cooling air flow from the rack, and cooling air flow passes through the intake pipe and gets into cavity 20 in, and a large amount of air film holes 12 on the rethread tail nozzle wall are discharged, and closely distributed air film holes 12 provide very big cooling area equivalent, and then have realized the cooling to the tail nozzle structure.
Referring to fig. 1, in some embodiments, a plurality of bleed air ducts 3 are provided along the circumferential direction of the jet nozzle body 1. A plurality of air guide pipes 3 are arranged in the circumferential direction of the jet nozzle body 1, so that cooling air can enter the cavity 20 more uniformly, and a large amount of cold air can flow into the air film holes 12 in different positions of the jet nozzle body 1 in the circumferential direction in time.
Referring to fig. 4, the flow direction of the air flow is generally described as follows: referring to fig. 4, air is used as the cooling air flow S1, the cooling air flow S1 enters the air through the air inlet pipe, then enters the air film hole 12, and finally flows into the channel 11 of the jet nozzle body 1, and an air film is formed on the inner wall surface of the jet nozzle body 1 to cool the jet nozzle body 1; and a portion of the cooling air flow S1 also mixes with the high temperature air flow S2 within the passageway 11 of the jet nozzle body 1 and exits the jet nozzle body 1 with the high temperature air flow S2. During the flow of this part of the cooling airflow, the inner wall surface of the jet nozzle body 1 is further cooled.
The embodiment of the utility model provides a still provide an aeroengine core machine, include the utility model discloses any technical scheme provides an aeroengine core tail spray tube.
In the description of the present invention, it should 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 the orientations or positional relationships shown in the drawings, and are only for convenience of description and simplicity of description, but 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 thus, should not be construed as limiting the scope of the present invention.
Finally, it should be noted that: the above embodiments are only used to illustrate the technical solution of the present invention, and not to limit it; although the present invention has been described in detail with reference to the foregoing embodiments, it should be understood by those skilled in the art that: it is to be understood that modifications may be made to the above-described arrangements in the embodiments or equivalents may be substituted for some of the features of the embodiments, but such modifications or substitutions do not depart from the spirit and scope of the present invention.
Claims (10)
1. An aircraft engine core jet nozzle, comprising:
a jet nozzle body (1) configured annularly and having a channel (11); the wall body of the tail nozzle body (1) is provided with a gas film hole (12), and the gas film hole (12) is communicated with the channel (11);
the cover body (2) is arranged on the outer side of the tail nozzle body (1), and a cavity (20) is enclosed between the inner wall of the cover body (2) and the outer wall of the tail nozzle body (1); and
a bleed conduit (3) mounted to the housing (2) and in fluid communication with the cavity (20).
2. The aircraft engine core jet nozzle according to claim 1, characterized in that the film holes (12) are arranged in rows in the jet nozzle body (1).
3. The aircraft engine core jet nozzle according to claim 1, characterized in that along the direction of the axis of the jet nozzle body (1) there are arranged a plurality of rows of said film holes (12).
4. The aircraft engine core jet nozzle according to claim 1, characterized in that along the circumferential direction of the jet nozzle body (1) there are arranged a plurality of turns of the film holes (12).
5. The aircraft engine core nozzle according to claim 1, wherein the axis of each film hole (12) forms a set angle with the wall body in which the film hole (12) is disposed.
6. The aircraft engine core nozzle according to claim 5, wherein the set angle is 25 ° to 35 °.
7. The aircraft engine core jet nozzle according to claim 1, characterized in that a plurality of the bleed air ducts (3) are provided along the circumferential direction of the jet nozzle body (1).
8. The aircraft engine core nozzle according to claim 1, characterized in that the hood (2) is configured annularly.
9. The aircraft engine core nozzle according to claim 1, characterized in that the shroud (2) comprises:
a first plate (21) extending in a radial direction of the jet nozzle body (1); one axial end of the tail nozzle body (1) of the first plate (21) is attached and fixed; and
the second plate (22) extends along the axial direction of the jet nozzle body (1), the first plate (21) and the second plate (22) are fixedly connected, and one end, far away from the first plate (21), of the second plate (22) is fixed with the axial middle part or the other end of the jet nozzle body (1).
10. An aircraft engine core engine comprising an aircraft engine core nozzle according to any one of claims 1 to 9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202120224562.XU CN214247529U (en) | 2021-01-27 | 2021-01-27 | Aircraft engine core tail nozzle and aircraft engine core |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202120224562.XU CN214247529U (en) | 2021-01-27 | 2021-01-27 | Aircraft engine core tail nozzle and aircraft engine core |
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| Publication Number | Publication Date |
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| CN214247529U true CN214247529U (en) | 2021-09-21 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN202120224562.XU Active CN214247529U (en) | 2021-01-27 | 2021-01-27 | Aircraft engine core tail nozzle and aircraft engine core |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113864081A (en) * | 2021-10-28 | 2021-12-31 | 中国航发沈阳发动机研究所 | Stealthy structure suitable for strong infrared suppression effect of binary vector spray tube |
| CN115013182A (en) * | 2022-06-17 | 2022-09-06 | 中国航发贵阳发动机设计研究所 | An efficient cooling jet engine |
-
2021
- 2021-01-27 CN CN202120224562.XU patent/CN214247529U/en active Active
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113864081A (en) * | 2021-10-28 | 2021-12-31 | 中国航发沈阳发动机研究所 | Stealthy structure suitable for strong infrared suppression effect of binary vector spray tube |
| CN115013182A (en) * | 2022-06-17 | 2022-09-06 | 中国航发贵阳发动机设计研究所 | An efficient cooling jet engine |
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