CN215213719U

CN215213719U - Jet mechanism of aviation turbine engine

Info

Publication number: CN215213719U
Application number: CN202121625117.0U
Authority: CN
Inventors: 申红林
Original assignee: Nanjing Longpu Power Technology Co ltd
Current assignee: Nanjing Longpu Power Technology Co ltd
Priority date: 2021-07-16
Filing date: 2021-07-16
Publication date: 2021-12-17
Anticipated expiration: 2031-07-16

Abstract

The utility model relates to an engine insufflating technical field, and an aviation turbine engine's insufflating mechanism is disclosed, including the injector base member, the ignition spray tube has all been cup jointed in the inside slip of injector base member, the oxidant chamber has all been seted up to the upper end of injector base member and the both sides that are located the ignition spray tube, the bottom plate has been cup jointed in the lower extreme slip of injector base member, the bottom plate sets up with the internal surface lower extreme contact of injector base member, the fuel chamber has all been seted up to the inside both ends of injector base member and the top that is located the bottom plate, the intercommunication is equipped with the nozzle between oxidant chamber and the bottom plate, the fixed cover in upper end of injector base member has connect the bush, the bush is through the internal surface top contact fixed connection in first welding seam and fuel chamber, the outside of bush is equipped with sealing mechanism. The utility model discloses effectively prevent that the injector welding seam from taking place to reveal, improve the stability in use and the life-span of injector.

Description

Jet mechanism of aviation turbine engine

Technical Field

The utility model relates to an engine spouts technical field, especially relates to an aviation turbine engine's spouting mechanism.

Background

A turbine engine is a form of engine that uses rotating parts to extract kinetic energy from fluid passing through it, and is a type of internal combustion engine, while internal injectors are a major component of aviation worm engines.

When the existing injector is installed, the injector is connected with a flame spraying pipe in a welding mode, a base body of the injector is divided into a fuel cavity and an oxidant cavity, and the welding mode is adopted, so that welding seam leakage is easy to occur in the using process of the injector, and the fuel and the oxidant are subjected to cavity channeling.

SUMMERY OF THE UTILITY MODEL

The utility model aims at solving the problem that the welding connection is adopted inside the injector in the prior art, the welding line is easy to leak when in use, and the fuel and the oxidant lead to fleeing the cavity, and the injection mechanism of the aviation turbine engine is provided.

In order to achieve the above purpose, the utility model adopts the following technical scheme:

the injection mechanism of the aviation turbine engine comprises an injector base body, an ignition spray pipe is sleeved in the injector base body in a sliding mode, oxidant cavities are formed in the upper end of the injector base body and located on two sides of the ignition spray pipe, a bottom plate is sleeved at the lower end of the injector base body in a sliding mode, the bottom plate is arranged in contact with the lower end of the inner surface of the injector base body, fuel cavities are formed in the two ends of the inner portion of the injector base body and located above the bottom plate, a nozzle is communicated between the oxidant cavities and the bottom plate, a lining is fixedly sleeved at the upper end of the injector base body, the lining is fixedly connected with the top of the inner surface of the fuel cavity in a contacting mode through a first welding line, and a sealing mechanism is arranged outside the lining.

Preferably, the second welding seam arranged at the lower end of the outer side wall of the ignition spray pipe in a surrounding mode is fixedly connected with the bottom plate.

Preferably, both sides of the bottom plate are fixedly connected with both ends of the bottom of the injector base body through third welding seams.

Preferably, the sealing mechanism comprises a cover plate, the top of the inner surface of the fuel cavity is provided with a ring-shaped groove, the cover plate is in contact with the groove wall of the ring-shaped groove, the bottom of the cover plate is fixedly provided with a support plate, the support plate is in sliding sleeve connection with the bush, the bottom of the support plate is fixedly provided with an internal thread pipe, the lower end of the outer side wall of the bush is provided with a thread, and the internal thread pipe is in threaded fit with the outer side wall of the bush.

Preferably, sealing mechanism includes annular seal plate, the seal groove has been seted up at the internal surface top in fuel chamber, annular seal plate and seal groove contact setting, and the fixed lagging in bottom, the lagging slides with the bush and cup joints, and the fixed ejector sleeve that is equipped with in bottom, the equal screw thread of both sides lower extreme of ejector sleeve has cup jointed the bolt, the screw hole has been seted up to the surface both sides lower extreme of bush, the pore wall screw-thread fit of bolt and screw hole.

Preferably, the cover plate is arranged annularly.

Preferably, the sealing groove is an annular groove.

Compared with the prior art, the utility model provides an aviation turbine engine's spouting mechanism possesses following beneficial effect:

1. the injection mechanism of the aviation turbine engine can complete the mixing of the oxidant and the fuel and carry out ignition combustion operation when the engine is started through the oxidant cavity, the nozzle, the ignition spray pipe and the fuel cavity.

2. According to the injection mechanism of the aviation turbine engine, the welding line between the ignition nozzle and the injector base body is sealed through the arranged lining, the sealing mechanism and the ignition nozzle, and the phenomenon that oxidizing agents and fuels flee the cavity is prevented.

The part that does not relate to among the device all is the same with prior art or can adopt prior art to realize, the utility model discloses convenient operation can stabilize effectual parcel setting to the welding seam connection position of injector, prevents that the emergence welding seam from revealing, has improved the stability in use of injector.

Drawings

Fig. 1 is a schematic structural diagram of an injection mechanism of an aircraft turbine engine according to the present invention;

FIG. 2 is an enlarged view of a portion A of FIG. 1;

fig. 3 is a schematic structural diagram of another embodiment of an injection mechanism of an aircraft turbine engine according to the present invention;

fig. 4 is an enlarged view of a portion B of fig. 3.

In the figure: 1 injector base body, 2 ignition spray pipes, 3 oxidant cavities, 4 bottom plates, 5 fuel cavities, 6 nozzles, 7 bushings, 8 first welding seams, 9 second welding seams, 10 third welding seams, 11 cover plates, 12 supporting plates, 13 internal threaded pipes, 14 annular sealing plates, 15 sleeve plates, 16 push pipes, 17 bolts and 18 threaded holes.

Detailed Description

The technical solutions in the embodiments of the present invention will be described clearly and completely with reference to the accompanying drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only some embodiments of the present invention, not all embodiments.

In the description of the present invention, it is to be understood that the terms "upper", "lower", "front", "rear", "left", "right", "top", "bottom", "inner", "outer", and the like 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, 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 construed as limiting the present invention.

Example 1

Referring to fig. 1-2, the injection mechanism of the aircraft turbine engine comprises an injector base body 1, an ignition spray pipe 2 is slidably sleeved inside the injector base body 1, oxidant cavities 3 are formed in the upper end of the injector base body 1 and located on two sides of the ignition spray pipe 2, a bottom plate 4 is slidably sleeved at the lower end of the injector base body 1, the bottom plate 4 is in contact with the lower end of the inner surface of the injector base body 1, fuel cavities 5 are formed in the two ends of the inside of the injector base body 1 and located above the bottom plate 4, a nozzle 6 is communicated between the oxidant cavities 3 and the bottom plate 4, a lining 7 is fixedly sleeved at the upper end of the injector base body 1, the lining 7 is in contact and fixed connection with the top of the inner surface of the fuel cavity 5 through a first welding line 8, and a sealing mechanism is arranged outside the lining 7.

And a second welding seam 9 arranged at the lower end of the outer side wall of the ignition spray pipe 2 in a surrounding manner is fixedly connected with the bottom plate 4.

Both sides of the bottom plate 4 are fixedly connected with both ends of the bottom of the injector base body 1 through third welding seams 10.

Sealing mechanism has seted up the ring channel including apron 11, the internal surface top in fuel chamber 5, apron 11 sets up with the cell wall contact of ring channel, the fixed backup pad 12 that is equipped with in bottom of apron 11, backup pad 12 cup joints with 7 slip of bush, the fixed internal thread pipe 13 that is equipped with in bottom of backup pad 12, the lateral wall lower extreme of bush 7 is seted up threadedly, the external lateral wall screw-thread fit of internal thread pipe 13 and bush 7.

The cover plate 11 is arranged in an annular shape, and is convenient to contact and match with the groove wall of the annular groove to seal the first welding line 8.

In the utility model, when the injector matrix 1 is installed, the ignition spray pipe 2 is inserted into the injector matrix 1, then the bush 7 is fixed on the upper end of the ignition spray pipe 2, the ignition spray pipe 2 is connected with the injector matrix 1 through the first welding line 8, then the internal thread pipe 13 is rotated to push the support plate 12 and the cover plate 11 while moving up slowly, finally the cover plate 11 is contacted with the annular groove at the top of the inner surface of the fuel cavity 5, the first welding line 8 is wrapped to prevent the fuel cavity 5 and the oxidant cavity 3 from channeling, then the bottom plate 4 is connected with the ignition spray pipe 2 and the injector matrix 1 through the second welding line 9 and the third welding line 10, when the injector works, the oxidant and the fuel enter the ignition area through the oxidant cavity 3, the fuel cavity 5 and the nozzle 6 respectively, and cooperate with the ignition spray pipe 2 to perform combustion work, the start of the engine is completed.

Example 2

Referring to fig. 3-4, the injection mechanism of the aircraft turbine engine comprises an injector base body 1, an ignition spray pipe 2 is slidably sleeved inside the injector base body 1, oxidant cavities 3 are formed in the upper end of the injector base body 1 and located on two sides of the ignition spray pipe 2, a bottom plate 4 is slidably sleeved at the lower end of the injector base body 1, the bottom plate 4 is arranged in contact with the lower end of the inner surface of the injector base body 1, fuel cavities 5 are formed in the two ends of the inside of the injector base body 1 and located above the bottom plate 4, a nozzle 6 is communicated between the oxidant cavities 3 and the bottom plate 4, a lining 7 is fixedly sleeved at the upper end of the injector base body 1, the lining 7 is fixedly connected with the top of the inner surface of the fuel cavity 5 in a contact mode through a first welding line 8, and a sealing mechanism is arranged outside the lining 7.

Sealing mechanism includes annular seal plate 14, and the seal groove has been seted up at the internal surface top in fuel chamber 5, and annular seal plate 14 sets up with the seal groove contact, and the fixed lagging 15 that is equipped with in bottom, and lagging 15 and bush 7 slip cup joint, and the fixed ejector sleeve 16 that is equipped with in bottom, and the equal screw thread of both sides lower extreme of ejector sleeve 16 has cup jointed bolt 17, and threaded hole 18 is seted up to the surface both sides lower extreme of bush 7, the pore wall screw-thread fit of bolt 17 and threaded hole 18.

The seal groove is an annular groove, so that the seal groove is in contact fit with the groove wall of the seal groove, and the first welding seam 8 is sealed.

In the utility model, when the injector base body 1 is installed, the ignition spray pipe 2 is inserted into the injector base body 1, then the bush 7 is fixed on the upper end of the ignition spray pipe 2, the ignition spray pipe 2 is connected with the injector base body 1 through the first welding line 8, then the push pipe 16 is pushed, so that the ignition spray pipe 2 moves upwards, when the push pipe 16 moves upwards, the sleeve plate 15 at the top of the push pipe pushes the annular sealing plate 14 and slowly contacts with the sealing groove, then the two bolts 17 are in threaded connection with the push pipe 16 and finally in threaded fit with the threaded holes 18 on the side wall of the bush 7, the position determination of the push pipe 16 and the annular sealing plate 14 is completed, the first welding line 8 is wrapped, the fuel cavity 5 and the oxidant cavity 3 are prevented from channeling, and then the bottom plate 4 is connected with the ignition spray pipe 2 and the injector base body 1 through the second welding line 9 and the third welding line 10, when the injector works, oxidant and fuel enter an ignition region through the oxidant cavity 3, the fuel cavity 5 and the nozzle 6 respectively, and are matched with the ignition nozzle 2 to perform combustion work, so that the starting of an engine is completed.

The above, only be the concrete implementation of the preferred embodiment of the present invention, but the protection scope of the present invention is not limited thereto, and any person skilled in the art is in the technical scope of the present invention, according to the technical solution of the present invention and the utility model, the concept of which is equivalent to replace or change, should be covered within the protection scope of the present invention.

Claims

1. An injection mechanism of an aircraft turbine engine comprises an injector base body (1) and is characterized in that an ignition spray pipe (2) is sleeved in the injector base body (1) in a sliding mode, oxidant cavities (3) are formed in the upper end of the injector base body (1) and located on two sides of the ignition spray pipe (2), a bottom plate (4) is sleeved at the lower end of the injector base body (1) in a sliding mode, the bottom plate (4) is arranged in contact with the lower end of the inner surface of the injector base body (1), fuel cavities (5) are formed in the two inner ends of the injector base body (1) and located above the bottom plate (4), a nozzle (6) is communicated between the oxidant cavities (3) and the bottom plate (4), a lining (7) is fixedly connected to the upper end of the injector base body (1) in a fixed mode, and the lining (7) is fixedly connected with the top of the inner surface of the fuel cavities (5) through a first welding line (8), and a sealing mechanism is arranged outside the bushing (7).

2. Injection mechanism for an aircraft turbine engine according to claim 1, characterised in that the lower end of the outer side wall of the ignition lance (2) is fixedly connected to the base plate (4) by a second weld seam (9) arranged around it.

3. Injection mechanism of an aircraft turbine engine according to claim 1, characterised in that both sides of the bottom of the base plate (4) are fixedly connected to both ends of the bottom of the injector base body (1) by means of third welds (10).

4. An injection mechanism of an aircraft turbine engine according to claim 1, wherein the sealing mechanism comprises a cover plate (11), an annular groove is formed in the top of the inner surface of the fuel cavity (5), the cover plate (11) is arranged in contact with the groove wall of the annular groove, a support plate (12) is fixedly arranged at the bottom of the cover plate (11), the support plate (12) is slidably sleeved with the liner (7), an internally threaded pipe (13) is fixedly arranged at the bottom of the support plate (12), the lower end of the outer side wall of the liner (7) is provided with a thread, and the internally threaded pipe (13) is in thread fit with the outer side wall of the liner (7).

5. An injection mechanism of an aircraft turbine engine according to claim 1, wherein the sealing mechanism comprises an annular sealing plate (14), a sealing groove is formed in the top of the inner surface of the fuel cavity (5), the annular sealing plate (14) is in contact with the sealing groove, a sleeve plate (15) is fixedly arranged at the bottom of the annular sealing plate, the sleeve plate (15) is in sliding sleeve connection with the liner (7), a push pipe (16) is fixedly arranged at the bottom of the annular sealing plate, bolts (17) are sleeved at the lower ends of two sides of the push pipe (16) in a threaded manner, threaded holes (18) are formed in the lower ends of two sides of the outer surface of the liner (7), and the bolts (17) are in threaded fit with the hole walls of the threaded holes (18).

6. Injection mechanism for an aircraft turbine engine according to claim 4, characterised in that said cover plate (11) is annular.

7. An injector mechanism for an aircraft turbine engine according to claim 5, wherein said seal groove is an annular groove.