CN112196697A - Integrated structure injector for rocket engine - Google Patents
Integrated structure injector for rocket engine Download PDFInfo
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- CN112196697A CN112196697A CN202011120831.4A CN202011120831A CN112196697A CN 112196697 A CN112196697 A CN 112196697A CN 202011120831 A CN202011120831 A CN 202011120831A CN 112196697 A CN112196697 A CN 112196697A
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- fuel
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- injector
- injection holes
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/52—Injectors
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- Engineering & Computer Science (AREA)
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- Combustion & Propulsion (AREA)
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- General Engineering & Computer Science (AREA)
- Fuel-Injection Apparatus (AREA)
Abstract
The invention provides an integrated structure injector for a rocket engine, which comprises an oxidant inlet, an oxidant liquid collecting cavity, an oxidant distribution channel, an oxidant annular cavity and an oxidant injection hole, wherein the oxidant inlet is communicated with the oxidant liquid collecting cavity; the fuel injector comprises a fuel inlet, a fuel collecting cavity, a fuel conveying channel, a fuel distributing channel, a fuel annular cavity and a fuel injection hole, wherein each propellant flows through the respective channel, enters the respective annular cavity and enters the corresponding injection hole. According to the processing method, the number of parts is effectively reduced, the processing flow is simplified, the manufacturing period of the injector is accelerated, the processing cost is reduced, and the reliability of the product is improved. The oxidant annular cavity and the fuel annular cavity are arranged in a circle and respectively correspond to the injection ports, so that the oxidant and the fuel are respectively injected; the injection holes form injection pairs, the oxidant fog fans and the fuel fog fans are formed by mutual impact, the fog fans are mixed with each other, the oxidant and the fuel can be uniformly mixed according to the proportion, and the combustion efficiency is improved.
Description
Technical Field
The invention relates to the technical field of rocket engines, in particular to an integrated structure injector for a rocket engine.
Background
The development of rocket engines is directed towards simplifying the structure to improve engine reliability while ensuring engine performance. The injector of the traditional liquid rocket engine is used for supplying and mixing the tissue oxidant and the fuel, and the structure of the injector is more complex along with the increase of the thrust of the injector. The injector of the general liquid rocket engine is composed of a plurality of parts, and then the parts are connected together by a plurality of welding methods such as brazing, argon arc welding and the like, so that the manufacturing period is long, the cost is high, and the quick verification of the engine scheme and the reduction of the engine product cost are restricted.
CN106735829A discloses a welding method for a multi-part structure of a liquid rocket engine thrust chamber head injector. The processing technology for forming the whole injector in a welding mode is complex, the nozzles of the oxidant and the fuel are sleeved with each other, the liquids are directly mixed, and the permeability between the liquids is poor.
Disclosure of Invention
Aiming at the problems in the prior art, the invention provides an integrated structure injector for a rocket engine, which adopts an integrated structure and simplifies the processing flow; the injection holes form injection pairs, the oxidant fog fans and the fuel fog fans are formed by mutual impact, the fog fans are mixed with each other, the oxidant and the fuel can be uniformly mixed according to the proportion, and the combustion efficiency is improved.
In order to achieve the purpose, the invention provides an integrated structure injector for a rocket engine, which comprises an oxidant inlet, an oxidant liquid collecting cavity, an oxidant distributing channel, an oxidant annular cavity, an oxidant injecting hole, a fuel inlet, a fuel liquid collecting cavity, an oblique fuel conveying channel, a transverse fuel conveying channel, a fuel distributing channel, a fuel annular cavity and a fuel injecting hole;
the oxidant enters the oxidant liquid collecting cavity through the oxidant inlet, enters the corresponding oxidant annular cavity through the oxidant distribution channels and is sprayed out of the oxidant injection hole;
the fuel enters the fuel collecting cavity through the fuel inlet, then enters the plurality of fuel distribution channels through the fuel conveying channel, enters the corresponding fuel annular cavity through the fuel distribution channels, and is sprayed out through the fuel injection holes;
the oxidant annulus is disposed between two of the fuel annuli.
Furthermore, a plurality of oxidant injection pairs are formed by the oxidant injection holes, and the oxidant injected by two oxidant injection holes in each oxidant injection pair mutually attacks to form an oxidant fog fan;
the fuel injection holes form a plurality of fuel injection pairs, and the fuel injected by two oxygen fuel injection holes in each fuel injection pair mutually impacts to form a fuel fog fan;
the oxidant fog fan and the fuel fog fan are mixed with each other.
Furthermore, the axes of the two oxidant injection holes in each oxidant injection pair form an included angle of 45-85 degrees; the axes of the two fuel injection holes in each fuel injection pair form an included angle of 50-90 degrees.
Further, according to the proportion requirement of the oxidant and the fuel, the diameters and the number of the oxidant injection holes and the fuel injection holes are determined, wherein the pressure drop of the injection holes is 15% -25% of the pressure of the combustion chamber.
Further, the number of the oxidant ring cavities is 6, and the number of the fuel ring cavities is 8; the number of the oxidant injection holes is the same as that of the fuel injection holes, and the diameter of each oxidant injection hole is larger than that of each fuel injection hole.
Furthermore, the injector is a conical sleeve, the top of the injector is an oxidant inlet, and an oxidant liquid collecting cavity is arranged inside the sleeve; the fuel conveying channel comprises a transverse fuel conveying channel and an oblique fuel conveying channel; the side wall of the injector is provided with a fuel inlet, a fuel liquid collecting cavity and an inclined fuel conveying channel arranged along the circumferential direction of the side wall; and a plurality of transverse fuel conveying channels distributed along the radius direction are uniformly arranged at the bottom of the injector and are connected to the inclined fuel conveying channel.
Furthermore, the oxidant ring cavity and the fuel ring cavity are positioned on the same layer at the bottom and positioned on the lower layer of the transverse fuel conveying channel; the oxidant distribution channels are distributed along the oxidant ring cavity and avoid the transverse fuel conveying channel, and the oxidant distribution channels are communicated to the oxidant ring cavity from the top end of the bottom;
the oxidant injection hole is communicated to the bottom end of the bottom part from the oxidant annular cavity; the fuel injection holes are communicated to the bottom end of the bottom part through the fuel ring cavity.
Furthermore, the two oxidant injection holes of the oxidant pair are positioned in the same circle, and one circle or two circles of the oxidant pair are correspondingly arranged in each circle of the oxidant annular cavity;
the two fuel injection holes of the fuel pairs are positioned in the same circle, and one circle or two circles of the fuel pairs are correspondingly arranged in each circle of the fuel ring cavity.
Furthermore, the injector is of an integrated structure and is formed by 3D printing and processing of metal powder; the metal powder comprises stainless steel, titanium alloy or high-temperature alloy GH 4169.
Further, the oxidant is dinitrogen tetroxide, and the fuel is hydrazines; or the oxidant is liquid oxygen and the fuel is liquid hydrogen; or the oxidant is liquid oxygen, and the fuel is liquid hydrocarbon fuel; or the oxidant is liquid nitrous oxide, and the fuel is liquid hydrocarbon fuel.
The technical scheme of the invention has the following beneficial technical effects:
(1) according to the processing method, the flow channel and the spray holes of the injector are flexibly designed, the number of parts is effectively reduced, welding is not needed, the processing flow is simplified, the manufacturing period of the injector is accelerated, the processing cost is reduced, and the reliability of products is improved.
(2) The oxidant annular cavity and the fuel annular cavity are arranged in a circle and respectively correspond to the injection ports, so that the oxidant and the fuel are respectively injected; the injection holes form injection pairs, the oxidant fog fans and the fuel fog fans are formed by mutual impact, the fog fans are mixed with each other, the oxidant and the fuel can be uniformly mixed according to the proportion, and the combustion efficiency is improved.
(3) The injector has an integrated structure, is simple in structure and high in precision, and is convenient for batch production; the injector has high structural strength and can be repeatedly used.
Drawings
FIG. 1 is an axial view of an injector of the present invention for a rocket engine;
FIG. 2 is a schematic cross-sectional view taken along line A of FIG. 1;
FIG. 3 is a schematic cross-sectional view of a transverse fuel transport channel;
FIG. 4 is a partial schematic view of FIG. 1 taken along line C and line D;
FIG. 5 is a schematic of injection face injection versus impingement blending.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention will be described in further detail with reference to the accompanying drawings in conjunction with the following detailed description. It should be understood that the description is intended to be exemplary only, and is not intended to limit the scope of the present invention. Moreover, in the following description, descriptions of well-known structures and techniques are omitted so as to not unnecessarily obscure the concepts of the present invention.
The invention provides an integrated structure injector for a rocket engine, which comprises the following components in sequence of fluid flow: an oxidant inlet 101, an oxidant collection chamber 102, an oxidant distribution channel 103, an oxidant annular chamber 104 and an oxidant injection hole 105; a fuel inlet 201, a fuel collection cavity 202, an inclined fuel delivery channel 203, a transverse fuel delivery channel 204, a fuel distribution channel 205, a fuel annulus 206 and a fuel injector hole 207. Is formed by adopting an integrated structure.
As shown in fig. 1-3, the injector is a conical sleeve, the top of the injector is a flat top and serves as an oxidant inlet 101, and a conical chamber inside the sleeve is an oxidant liquid collecting cavity 102; the fuel delivery passages include a lateral fuel delivery passage 204 and an oblique fuel delivery passage 203; the side wall of the injector is provided with a fuel inlet 201, a fuel liquid collecting cavity 202 and an inclined fuel conveying channel 203, the inclined fuel conveying channel 203 is an annular cavity formed along the circumferential direction of the side wall, and the bottom of the inclined fuel conveying channel 203 is communicated with a transverse fuel conveying channel 204; the bottom of the injector is uniformly provided with a plurality of transverse fuel conveying channels 204 distributed along the radius direction and connected to the inclined fuel conveying channels 203.
The bottom of the injector is of a multilayer structure, and the oxidant annular cavity 104 and the fuel annular cavity 206 are positioned at the same layer of the bottom and positioned at the lower layer of the transverse fuel conveying channel 204; the oxidant distribution channels 103 are arranged in a vertical direction, distributed along the oxidant ring 104 and avoiding the transverse fuel conveying channels 204, and the oxidant distribution channels 103 are communicated to the oxidant ring 104 from the top end at the bottom. The oxidant injection hole 105 is communicated to the bottom end of the bottom part by the oxidant annular cavity 104; the fuel injection holes 207 are connected by the fuel annulus 206 to the bottom end of the bottom.
The oxidant ring cavities 104 are arranged annularly and alternately with the fuel ring cavities 206, and the oxidant ring cavities 104 are arranged between the two fuel ring cavities 206. The two oxidant injection holes 105 of the oxidant pair are positioned in the same circle, and one circle or two circles of the oxidant pair are correspondingly arranged in each circle of the oxidant annular cavity 104; also, the two fuel injection holes 207 of the fuel pair are located at the same circle, and one or two circles of the oxidizer pairs are correspondingly arranged at each circle of the fuel ring cavity 206.
The oxidant of the injector enters the oxidant liquid collecting cavity 102 after entering from the oxidant inlet 101, then enters the oxidant annular cavity 104 through a plurality of oxidant distribution channels 103 which are communicated with the oxidant liquid collecting cavity 102 and the oxidant annular cavity 104 in the vertical direction, and is sprayed out through oxidant injection holes 105 below the oxidant annular cavity 104.
The injector fuel enters the fuel collecting cavity 202 after entering from the lateral fuel inlet 201, then passes through a plurality of oblique fuel conveying channels 203 which form a certain angle with the vertical direction, turns and enters a plurality of transverse fuel conveying channels 204 at the lower part, then enters the fuel annular cavity 206 through a plurality of fuel distributing channels 205 which are communicated with the transverse fuel conveying channels 204 and the fuel annular cavity 206 in the vertical direction, and finally is ejected through the fuel injection holes 207 at the lower part of the fuel annular cavity 206.
Combining 4, wherein the axis of each oxidant injection hole 105 forms an included angle of 45-85 degrees with the axis of the two oxidant injection holes 105 in each oxidant injection pair; the axes of the two fuel injection holes 207 in each fuel injection pair form an included angle of 50-90 degrees.
Referring to fig. 5, the injector oxidant injection holes and fuel injection holes are in a unit of a single injection pair, wherein the injection pair is in a self-tapping form. Every two injection holes in the same circle form a certain angle, and the sprayed liquid mutually hits. The single injection pairs are circumferentially distributed on the distribution circle according to the structure; in the radial direction, the oxidant injection holes and the fuel injection holes are arranged in a plurality of circles according to the annular cavity. The sprayer is characterized in that a fog fan formed by self-impact of the same propellant and the same propellant on the same distribution circle is mixed with a fog fan formed by self-impact of the different propellant adjacent to the distribution circle.
The size of the injection holes is calculated according to the propellant flow, the selected injection pressure drop, the flow coefficient and the number of the injection holes, wherein the injection pressure drop of each propellant flow is 15% -25% of the room pressure according to the thrust of the engine, the flow coefficient of the injection holes is generally 0.6-0.8, and the number of the injection holes is determined according to the diameter of the injection panel and the number of the arranged injection rings.
Further, the processing method of the injector comprises 3D printing and powder metallurgy.
Further, the material of the injector comprises stainless steel, titanium alloy or high-temperature alloy GH 4169.
Further, the injector is suitable for propellants, including dinitrogen tetroxide/hydrazine propellants, oxygen/liquid hydrocarbon fuel propellants, liquid oxygen/liquid hydrogen propellants and liquid nitrous oxide/liquid hydrocarbon fuel propellants.
Examples
Aiming at a 30-ton liquid rocket engine thrust chamber head injector, the maximum diameter is phi 393mm, the height is 188mm, an oxidant (O) is nitrous oxide, a fuel (F) is ethylene, and the mixing ratio of a propellant is 6.
As shown in fig. 1 to 5, the present invention provides an integrated structure injector, including, in a fluid flow sequence:
an oxidant inlet 101, an oxidant collection chamber 102, an oxidant distribution channel 103, an oxidant annular chamber 104 and an oxidant injection hole 105;
a fuel inlet 201, a fuel plenum 202, a fuel delivery channel 203, a fuel delivery channel 204, a fuel dispensing orifice 205, a fuel annulus 206, and a fuel injector orifice 207.
Oxidant enters an oxidant collecting cavity 102 from an oxidant inlet 101, enters six oxidant annular cavities 104 after passing through an oxidant distribution channel 103 which is axially arranged, and is finally uniformly sprayed out from 928 oxidant spraying holes 105, wherein two circles of circumferentially arranged spraying pairs are arranged below each annular cavity, the spraying pairs are gradually increased from inside to outside, two spraying holes in each circle form self-impact pairs at 60 degrees, and a fog fan is formed after the self-impact pairs.
The fuel enters the fuel collecting cavity 202 from the fuel inlet 201, passes through the fuel delivery channel 203, enters the fuel delivery channel 204, passes through the fuel distribution channel 205, enters the eight fuel annular cavities 206, and is finally uniformly sprayed out from the 928 fuel spray holes 207, namely the eight fuel annular cavities marked by F in the figure 2, and the eight fuel annular cavities are the first fuel annular cavity to the eighth fuel annular cavity from inside to outside in sequence. Referring to fig. 1, the circle is the oxidizer ring 104 and the black dot is the fuel ring 206. Except the first, the third and the fourth fuel ring cavities, two circles of injection pairs are circumferentially arranged below the rest fuel ring cavities, one circle of injection pairs are circumferentially arranged below the first, the third and the fourth ring cavities, the injection pairs are gradually increased from inside to outside, two injection holes of each circle are self-impacted by 60 degrees to form self-impacting pairs, and a fog fan is formed after self-impacting.
The oxidant self-impingement formed mist fan blends with an adjacent fuel self-impingement fan.
The flow rates of the oxidant injection holes and the fuel injection holes are 91.5 kg/s and 15.3kg/s respectively, the injection pressure drop is 1.5MPa and 1.2MPa respectively, the number of the injection holes is 928, and the diameters of the oxidant injection holes and the fuel injection holes are calculated to be phi 1.85mm and phi 1.02mm respectively.
In this embodiment, the material of the injector is 316L. In the embodiment, the processing method of the injector is 3D printing, and the 3D printing is followed by heat treatment.
According to the processing method, the number of parts is effectively reduced, the processing flow is simplified, the manufacturing period of the injector is accelerated, the processing cost is reduced, and the reliability of the product is improved.
In conclusion, the invention provides an integrated structure injector for a rocket engine, which comprises an oxidant inlet, an oxidant liquid collecting cavity, an oxidant distribution channel, an oxidant annular cavity and an oxidant injection hole, wherein the oxidant inlet is communicated with the oxidant liquid collecting cavity; the fuel injector comprises a fuel inlet, a fuel collecting cavity, a fuel conveying channel, a fuel distributing channel, a fuel annular cavity and a fuel injection hole, wherein each propellant flows through the respective channel, enters the respective annular cavity and enters the corresponding injection hole. . According to the processing method, the number of parts is effectively reduced, the processing flow is simplified, the manufacturing period of the injector is accelerated, the processing cost is reduced, and the reliability of the product is improved. The oxidant annular cavity and the fuel annular cavity are arranged in a circle and respectively correspond to the injection ports, so that the oxidant and the fuel are respectively injected; the injection holes form injection pairs, the oxidant fog fans and the fuel fog fans are formed by mutual impact, the fog fans are mixed with each other, the oxidant and the fuel can be uniformly mixed according to the proportion, and the combustion efficiency is improved.
It is to be understood that the above-described embodiments of the present invention are merely illustrative of or explaining the principles of the invention and are not to be construed as limiting the invention. Therefore, any modification, equivalent replacement, improvement and the like made without departing from the spirit and scope of the present invention should be included in the protection scope of the present invention. Further, it is intended that the appended claims cover all such variations and modifications as fall within the scope and boundaries of the appended claims or the equivalents of such scope and boundaries.
Claims (10)
1. An integrated structure injector for a rocket engine is characterized by comprising an oxidant inlet (101), an oxidant liquid collecting cavity (102), an oxidant distribution channel (103), an oxidant annular cavity (104), oxidant injection holes (105), a fuel inlet (201), a fuel liquid collecting cavity (202), an inclined fuel conveying channel (203), a transverse fuel conveying channel (204), a fuel distribution channel (205), a fuel annular cavity (206) and fuel injection holes (207);
the oxidant enters an oxidant liquid collecting cavity (102) through an oxidant inlet (101), enters a corresponding oxidant annular cavity (104) through a plurality of oxidant distribution channels (103), and is sprayed out through an oxidant spraying hole (105);
the fuel enters the fuel collecting cavity (202) through the fuel inlet (201), then enters the plurality of fuel distribution channels (205) through the fuel conveying channel, enters the corresponding fuel annular cavity (206) through the fuel distribution channels (205), and is sprayed out through the fuel injection holes (207);
the oxidant annulus (104) is disposed between two of the fuel annuli (206).
2. The integrated structure injector for rocket engines according to claim 1, wherein the oxidant injection holes (105) form a plurality of oxidant injection pairs, and the oxidant injected from two oxidant injection holes (105) in each oxidant injection pair mutually hits to form an oxidant fog fan;
the fuel injection holes (207) form a plurality of fuel injection pairs, and the fuel injected by two oxygen fuel injection holes (207) in each fuel injection pair mutually hits to form a fuel fog fan;
the oxidant fog fan and the fuel fog fan are mixed with each other.
3. The integrated structure injector for rocket engines according to claim 2, wherein the axes of two oxidant injection holes (105) in each oxidant injection pair form an included angle of 45-85 °; the axes of the two fuel injection holes (207) in each fuel injection pair form an included angle of 50-90 degrees.
4. The integrated structural injector for rocket engines according to claim 3, wherein the diameters and the number of the oxidant injection holes (105) and the fuel injection holes (207) are determined according to the proportion requirement of the oxidant and the fuel, wherein the injection hole pressure drop is 15-25% of the combustion chamber pressure.
5. An integrated structural injector for rocket engines according to claim 4 wherein said oxidizer ring cavities (104) are 6, said fuel ring cavities (206) are 8; the number of the oxidant injection holes (105) and the number of the fuel injection holes (207) are the same, and the diameter of the oxidant injection holes (105) is larger than that of the fuel injection holes (207).
6. An integrated structural injector for rocket engines according to claim 3, characterized in that said injector is a conical sleeve, the top is the oxidant inlet (101) and the inside of the sleeve is the oxidant plenum (102); the fuel delivery channels comprise a transverse fuel delivery channel (204) and an oblique fuel delivery channel (203); the side wall of the injector is provided with a fuel inlet (201), a fuel liquid collecting cavity (202) and an inclined fuel conveying channel (203) which is arranged along the circumferential direction of the side wall; and a plurality of transverse fuel conveying channels (204) distributed along the radius direction are uniformly arranged at the bottom of the injector and are connected to the inclined fuel conveying channel (203).
7. An integrated structural injector for rocket engines according to claim 6, wherein said oxidizer ring (104) and said fuel ring (206) are located at the same level of the bottom, at the lower level of the transverse fuel conveying channels (204); the oxidant distribution channels (103) are distributed along the oxidant ring cavity (104) and avoid the transverse fuel conveying channels (204), and the oxidant distribution channels (103) are communicated to the oxidant ring cavity (104) from the bottom top end;
the oxidant injection hole (105) is communicated to the bottom end of the bottom part by the oxidant annular cavity (104); the fuel injection holes (207) are connected to the bottom end of the bottom by the fuel annulus (206).
8. An integrated structural injector for rocket engines according to claim 7, characterized in that said two oxidant injection holes (105) of said pair are located in the same circle, one or two circles of said pair of oxidants being provided for each circle of said oxidant annulus (104);
the two fuel injection holes (207) of the fuel pair are positioned in the same circle, and one circle or two circles of the fuel pair are correspondingly arranged in each circle of the fuel ring cavity (206).
9. An injector of an integrated structure for a rocket engine according to claim 1 or 2, wherein said injector is of an integrated structure and is formed by 3D printing and processing of metal powder; the metal powder comprises stainless steel, titanium alloy or high-temperature alloy GH 4169.
10. An integrated structural injector for rocket engines according to claim 1 or 2 wherein said oxidizer is dinitrogen tetroxide and said fuel is a hydrazine; or the oxidant is liquid oxygen and the fuel is liquid hydrogen; or the oxidant is liquid oxygen, and the fuel is liquid hydrocarbon fuel; or the oxidant is liquid nitrous oxide, and the fuel is liquid hydrocarbon fuel.
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|---|---|---|---|
| CN202011120831.4A CN112196697A (en) | 2020-10-19 | 2020-10-19 | Integrated structure injector for rocket engine |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011120831.4A CN112196697A (en) | 2020-10-19 | 2020-10-19 | Integrated structure injector for rocket engine |
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| CN112196697A true CN112196697A (en) | 2021-01-08 |
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| CN202011120831.4A Pending CN112196697A (en) | 2020-10-19 | 2020-10-19 | Integrated structure injector for rocket engine |
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Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113339159A (en) * | 2021-07-06 | 2021-09-03 | 西安航天动力研究所 | Coaxial double-centrifugal injector based on 3D printing and liquid oxygen kerosene rocket engine |
| CN114382614A (en) * | 2022-03-02 | 2022-04-22 | 北京星际荣耀科技有限责任公司 | Injector, engine and aircraft |
| CN114412660A (en) * | 2021-12-22 | 2022-04-29 | 首都航天机械有限公司 | Integrated coaxial nozzle of liquid rocket engine and design and manufacturing method thereof |
| CN114412664A (en) * | 2021-12-23 | 2022-04-29 | 北京航天动力研究所 | Thrust chamber head structure for reducing ignition impact |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| RU2124647C1 (en) * | 1993-05-11 | 1999-01-10 | Сосьете Оропеен де Пропюльсьон | Nozzle and injection member |
| CN101737198A (en) * | 2008-11-11 | 2010-06-16 | 北京航空航天大学 | Gas-gas nozzle with constriction section |
| CN101782030A (en) * | 2009-01-19 | 2010-07-21 | 北京航空航天大学 | Multiunit coaxial injector component for test and manufacturing method |
| CN106134392B (en) * | 2009-12-03 | 2013-12-11 | 上海空间推进研究所 | A kind of method of combination of ejector filler spray unit of double elements liquid-propellant rocket engine |
| CN108915899A (en) * | 2018-06-29 | 2018-11-30 | 北京航天动力研究所 | A kind of three chamber ejector filler of four bottom |
-
2020
- 2020-10-19 CN CN202011120831.4A patent/CN112196697A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| RU2124647C1 (en) * | 1993-05-11 | 1999-01-10 | Сосьете Оропеен де Пропюльсьон | Nozzle and injection member |
| CN101737198A (en) * | 2008-11-11 | 2010-06-16 | 北京航空航天大学 | Gas-gas nozzle with constriction section |
| CN101782030A (en) * | 2009-01-19 | 2010-07-21 | 北京航空航天大学 | Multiunit coaxial injector component for test and manufacturing method |
| CN106134392B (en) * | 2009-12-03 | 2013-12-11 | 上海空间推进研究所 | A kind of method of combination of ejector filler spray unit of double elements liquid-propellant rocket engine |
| CN108915899A (en) * | 2018-06-29 | 2018-11-30 | 北京航天动力研究所 | A kind of three chamber ejector filler of four bottom |
Non-Patent Citations (2)
| Title |
|---|
| 孙宏明: "直流式喷注器设计", 《火箭推进》 * |
| 李子阳: "火箭喷注器的流体特性研究及结构优化", 《中国优秀硕士学位论文全文数据局 工程科技Ⅱ辑》 * |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113339159A (en) * | 2021-07-06 | 2021-09-03 | 西安航天动力研究所 | Coaxial double-centrifugal injector based on 3D printing and liquid oxygen kerosene rocket engine |
| CN114412660A (en) * | 2021-12-22 | 2022-04-29 | 首都航天机械有限公司 | Integrated coaxial nozzle of liquid rocket engine and design and manufacturing method thereof |
| CN114412660B (en) * | 2021-12-22 | 2023-12-29 | 首都航天机械有限公司 | Integrated coaxial nozzle of liquid rocket engine and design and manufacturing method thereof |
| CN114412664A (en) * | 2021-12-23 | 2022-04-29 | 北京航天动力研究所 | Thrust chamber head structure for reducing ignition impact |
| CN114412664B (en) * | 2021-12-23 | 2023-08-29 | 北京航天动力研究所 | Thrust chamber head structure for reducing ignition impact |
| CN114382614A (en) * | 2022-03-02 | 2022-04-22 | 北京星际荣耀科技有限责任公司 | Injector, engine and aircraft |
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Application publication date: 20210108 |