CN109296474B - Rocket engine - Google Patents
Rocket engine Download PDFInfo
- Publication number
- CN109296474B CN109296474B CN201811129627.1A CN201811129627A CN109296474B CN 109296474 B CN109296474 B CN 109296474B CN 201811129627 A CN201811129627 A CN 201811129627A CN 109296474 B CN109296474 B CN 109296474B
- Authority
- CN
- China
- Prior art keywords
- engine
- head
- injector
- throat
- fuel
- 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.)
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Links
- 239000000446 fuel Substances 0.000 claims abstract description 32
- 238000002485 combustion reaction Methods 0.000 claims abstract description 31
- 238000009792 diffusion process Methods 0.000 claims abstract description 22
- 238000001816 cooling Methods 0.000 claims abstract description 20
- 239000007800 oxidant agent Substances 0.000 claims abstract description 19
- 230000001590 oxidative effect Effects 0.000 claims abstract description 18
- 238000002347 injection Methods 0.000 abstract description 9
- 239000007924 injection Substances 0.000 abstract description 9
- 238000003754 machining Methods 0.000 description 8
- 238000010146 3D printing Methods 0.000 description 6
- 238000005516 engineering process Methods 0.000 description 6
- 230000004048 modification Effects 0.000 description 4
- 238000012986 modification Methods 0.000 description 4
- 239000003380 propellant Substances 0.000 description 3
- 239000007921 spray Substances 0.000 description 3
- 230000009286 beneficial effect Effects 0.000 description 2
- 238000000034 method Methods 0.000 description 2
- 238000006243 chemical reaction Methods 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 238000007789 sealing Methods 0.000 description 1
- 239000000243 solution Substances 0.000 description 1
Classifications
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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
-
- 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
-
- 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
- F02K9/60—Constructional parts; Details not otherwise provided for
- F02K9/62—Combustion or thrust chambers
- F02K9/64—Combustion or thrust chambers having cooling arrangements
-
- 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
- F02K9/97—Rocket nozzles
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Testing Of Engines (AREA)
Abstract
The invention is applicable to the technical field of aerospace, and provides a rocket engine, which comprises: the head injector, the oxidant inlet, the head flange, the fuel inlet, the fuel connecting pipeline and the fuel inlet are integrally formed; the oxidant inlet and the fuel inlet are formed on the head injector, the head flange is formed on the periphery of the head injector, the head injector is communicated with the engine combustion part, the engine combustion part is communicated with the engine throat, the engine throat is communicated with the engine diffusion section, and the engine body cooling flow passage is formed on the engine combustion part and is connected with the engine head injector through the fuel pipeline. The problems that the cooling flow channel is complex in shape, too many head parts, the injection flow channel of the injector is complex, and the processing and the forming are difficult are solved.
Description
Technical Field
The invention is suitable for the technical field of aerospace, and particularly relates to a rocket engine.
Background
The rocket and the space vehicle need the support of a power propulsion system in the track changing and posture adjusting process, and a rocket engine is the most critical component of the space propulsion power system.
The shape of the injector and combustion chamber of the rocket engine head is directly related to whether the engine fuel is fully mixed and fully combusted, and thus, the performance of the engine. However, due to the limitations of machining, complex runner forms are difficult to achieve for better propellant mixing. Meanwhile, due to the non-integrated design of the head part and the combustion chamber, the problem of connection between the head part and the combustion chamber and the problem of sealing between the head part and the combustion chamber are also involved in assembly, so that the problems of weight increase and reliability reduction of the engine are caused.
Disclosure of Invention
The invention provides a rocket engine, which aims to solve the problems that the rocket engine has more parts and great manufacturing difficulty and reduces the performance of the rocket engine.
The invention is realized in that a rocket engine comprises: the head injector, the oxidant inlet, the head flange, the fuel inlet, the engine throat, the engine diffusion section and the engine body cooling flow passage are integrally formed; the oxidant inlet and the fuel inlet are formed on the head injector, the head flange is formed on the periphery of the head injector, the head injector is communicated with the engine combustion part, the engine combustion part is communicated with the engine throat, the engine throat is communicated with the engine diffusion section, and the engine body cooling flow passage is formed on the engine combustion part and is connected with the engine head injector through the fuel pipeline.
Further, the head flange comprises four fixing portions, the four fixing portions are arranged at equal intervals along the circumferential direction of the head injector, and fixing holes are formed in the fixing portions.
Further, the engine diffuser is formed with a plurality of mounting holes, each of which is disposed at equal intervals in the circumferential direction of the engine diffuser.
Still further, the engine combustion part, the engine throat and the engine diffusion section are all arranged in a hollow cylinder shape, the diameter of the engine throat is smaller than that of the engine combustion part, and the inner space of the engine diffusion section is conical.
The rocket engine provided by the embodiment of the invention has the beneficial effects that: according to the rocket engine, the head injector, the oxidant inlet, the head flange, the fuel inlet, the engine throat, the engine diffusion section and the engine body cooling flow passage are integrally formed, and then manufactured through a 3D printing technology; the problems of complex cooling flow channel shape, complex flow channel fine machining process, too many head parts, complex injection flow channel of the injector, difficulty in machining and forming and the like are solved, meanwhile, parts of the rocket engine are reduced, and the performance of the rocket engine is greatly improved.
Drawings
Fig. 1 is a schematic diagram of the overall structure of a rocket engine provided by the invention.
Wherein, each reference sign in the figure:
1-a head injector; 11-oxidant inlet; 12-fuel inlet; 13-a fixing part; 131-fixing holes; 2-a head flange; 3-a fuel connecting line; 4-an engine combustion section; 5-engine throat; 6-an engine diffuser; 61-mounting holes.
Detailed Description
The present invention will be described in further detail with reference to the drawings and examples, in order to make the objects, technical solutions and advantages of the present invention more apparent. It should be understood that the specific embodiments described herein are for purposes of illustration only and are not intended to limit the scope of the invention.
The distinguishing features between the present invention and the prior art are that:
The head injector 1, the oxidant inlet 11, the head flange 2, the fuel inlet 12, the engine throat 5, the engine diffusion section 6 and the engine body cooling flow passage are integrally formed; the oxidant inlet 11 and the fuel inlet 12 are formed on the head injector 1, the head flange 2 is formed on the periphery of the head injector 1, the head injector 1 communicates with the engine combustion portion 4, the engine combustion portion 4 communicates with the engine throat portion 5, the engine throat portion 5 communicates with the engine diffuser 6, and an engine body cooling flow passage (not shown) is formed on the engine combustion portion 4 and connected with the engine head injector 1 through a fuel line.
The rocket engine provided by the embodiment of the invention has the beneficial effects that: the head injector 1, the oxidant inlet 11, the head flange 2, the fuel inlet 12, the engine throat 5, the engine diffusion section 6 and the engine body cooling flow passage of the rocket engine are integrally formed and designed, and then manufactured by a 3D printing technology; the problems of complex cooling flow channel shape, complex flow channel fine machining process, too many head parts, complex injection flow channel of the injector, difficulty in machining and forming and the like are solved, meanwhile, parts of the rocket engine are reduced, and the performance of the rocket engine is greatly improved.
Example 1
A rocket engine, comprising: the head injector 1, the oxidant inlet 11, the head flange 2, the fuel connecting pipeline 3, the fuel inlet 12, the engine combustion part 4, the engine throat 5, the engine diffusion section 6 and the engine body cooling flow passage are integrally formed; the oxidant inlet 11 and the fuel inlet 12 are formed on the head injector 1, the head flange 2 is formed on the periphery of the head injector 1, the head injector 1 communicates with the engine combustion portion 4, the engine combustion portion 4 communicates with the engine throat portion 5, the engine throat portion 5 communicates with the engine diffuser 6, and the engine body cooling flow passage is formed on the engine combustion portion 4 and connected with the engine head injector 1 through a fuel line. In the working process of the engine, oxidant directly enters the engine head injector 1 through the oxidant inlet 11, fuel enters the engine body cooling flow passage through the fuel inlet 12, the engine is cooled and then enters the engine head injector 1 through the fuel connecting pipe, the oxidant and the fuel are mixed in the combustion chamber and combusted, and high-temperature and high-pressure fuel is injected through the engine throat 5 and the engine diffusion section 6 to form thrust. The head injector 1, the oxidant inlet 11, the head flange 2, the fuel inlet 12, the engine throat 5, the engine diffuser 6 and the engine body cooling flow passage are integrally formed and designed, and then manufactured by a 3D printing technology; the problems of complex cooling flow channel shape, complex flow channel fine machining process, too many head parts, complex injection flow channel of the injector, difficulty in machining and forming and the like are solved, meanwhile, parts of the rocket engine are reduced, and the performance of the rocket engine is greatly improved.
Example two
The present embodiment is a modification of the first embodiment, in which the head flange 2 includes four fixing portions 13, the four fixing portions 13 being disposed at equal intervals along the circumferential direction of the head injector 1, and fixing holes 131 being formed in each of the fixing portions 13. The head flange 2 and the head injector 1 of the rocket engine are integrally formed through a 3D printing technology, so that the problems that parts of the head are too many, an injection runner of the injector is complex, and the injection runner of the injector is difficult to machine and form are solved. Four fixing portions 13 are formed at equal intervals around the head injector 1, and the head injector 1 can be well fixed to an external device through the fixing holes 131.
Example III
The present embodiment is a modification of the first embodiment in that a plurality of mounting holes 61 are formed in the engine diffusing section 6, and the mounting holes 61 are arranged at equal intervals in the circumferential direction of the engine diffusing section 6. The engine diffuser 6, i.e. the tail of the engine, is oriented to orient the jet of the engine and the plurality of mounting holes act to more firmly attach the long nozzle to effect the conversion of the short nozzle engine into a long nozzle engine.
Example IV
In this embodiment, according to a modification of the first embodiment, the engine combustion portion 4, the engine throat portion 5 and the engine diffuser 6 are all provided in a hollow cylindrical shape, the diameter of the engine throat portion 5 is smaller than the diameter of the engine combustion portion 4, and the internal space of the engine diffuser 6 is conical. The propellant is combusted in the engine combustion part 4 to generate high-temperature and high-pressure gas, the gas is accelerated through the engine throat 5 and then is ejected through the engine diffusion section 6, the internal space of the engine diffusion section 6 is conical, and the propulsion area of the high-temperature gas is increased and stable thrust is obtained by gradually expanding the diffusion section.
According to the rocket engine, the head injector, the oxidant inlet, the head flange, the fuel inlet, the engine throat, the engine diffusion section and the engine body cooling flow passage are integrally formed, and then manufactured through a 3D printing technology; the problems that the cooling flow channel is complex in shape, the flow channel fine machining process is complex, the injection flow channel of the injector is complex, the injection flow channel is difficult to machine and form and the like are solved, meanwhile, parts of the rocket engine are reduced, and the performance of the rocket engine is greatly improved. The head flange of the rocket engine and the head injector are integrally formed through a 3D printing technology, so that the problems that parts of the head are too many, an injection runner of the injector is complex, and the processing and the forming are difficult are solved. Four fixing parts are formed at the periphery of the head injector at equal intervals, and the head injector can be well fixed with external equipment through the fixing holes; the plurality of mounting holes are used for more firmly connecting the long spray pipes, so that the short spray pipe engine is converted into the long spray pipe engine. The propellant is burnt in the combustion part of the engine to generate high-temperature and high-pressure gas, the gas is accelerated through the throat part of the engine and then is sprayed out through the diffusion section of the engine, the internal space of the diffusion section of the engine is conical, and the propulsion area of the high-temperature gas is increased and stable thrust is obtained by gradually expanding the diffusion section.
The foregoing description of the preferred embodiments of the invention is not intended to be limiting, but rather is intended to cover all modifications, equivalents, and alternatives falling within the spirit and principles of the invention.
Claims (4)
1. A rocket engine, comprising: head injector, oxidant entry, head flange, fuel connecting line, fuel inlet, engine combustion part, engine throat, engine diffuser and engine fuselage cooling flow channel, its characterized in that: the head injector, the oxidant inlet, the head flange, the fuel inlet, the engine throat, the engine diffuser and the engine block cooling flow passage are integrally formed; the oxidant inlet and the fuel inlet are formed on the head injector, the head flange is formed on the periphery of the head injector, the head injector is communicated with the engine combustion part, the engine combustion part is communicated with the engine throat, the engine throat is communicated with the engine diffusion section, and the engine body cooling flow passage is formed on the engine combustion part and is connected with the engine head injector through the fuel pipeline.
2. A rocket engine as recited in claim 1, wherein: the head flange comprises four fixing portions, the four fixing portions are arranged at equal intervals along the circumferential direction of the head injector, and fixing holes are formed in the fixing portions.
3. A rocket engine as recited in claim 1, wherein: a plurality of mounting holes are formed in the engine diffusion section, and the mounting holes are arranged at equal intervals along the circumferential direction of the engine diffusion section.
4. A rocket engine as recited in claim 1, wherein: the engine combustion part, the engine throat part and the engine diffusion section are all in a hollow cylindrical shape, the diameter of the engine throat part is smaller than that of the engine combustion part, and the inner space of the engine diffusion section is conical.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811129627.1A CN109296474B (en) | 2018-09-27 | 2018-09-27 | Rocket engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811129627.1A CN109296474B (en) | 2018-09-27 | 2018-09-27 | Rocket engine |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN109296474A CN109296474A (en) | 2019-02-01 |
| CN109296474B true CN109296474B (en) | 2024-05-28 |
Family
ID=65164829
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201811129627.1A Active CN109296474B (en) | 2018-09-27 | 2018-09-27 | Rocket engine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN109296474B (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20220307450A1 (en) * | 2019-11-22 | 2022-09-29 | Aerojet Rocketdyne, Inc. | Catalytic thruster |
Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN209011970U (en) * | 2018-09-27 | 2019-06-21 | 宁波天擎航天科技有限公司 | A kind of rocket engine |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE102016212399B4 (en) * | 2016-07-07 | 2022-04-28 | Arianegroup Gmbh | rocket engine |
| JP2019534409A (en) * | 2016-09-01 | 2019-11-28 | アッディティブ ロケット コーポレーション | Additional manufactured combustion engine |
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2018
- 2018-09-27 CN CN201811129627.1A patent/CN109296474B/en active Active
Patent Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN209011970U (en) * | 2018-09-27 | 2019-06-21 | 宁波天擎航天科技有限公司 | A kind of rocket engine |
Also Published As
| Publication number | Publication date |
|---|---|
| CN109296474A (en) | 2019-02-01 |
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