CN114810430A - Low-ablation rocket engine nozzle structure with active cooling throat insert and cooling method - Google Patents
Low-ablation rocket engine nozzle structure with active cooling throat insert and cooling method Download PDFInfo
- Publication number
- CN114810430A CN114810430A CN202210378851.4A CN202210378851A CN114810430A CN 114810430 A CN114810430 A CN 114810430A CN 202210378851 A CN202210378851 A CN 202210378851A CN 114810430 A CN114810430 A CN 114810430A
- Authority
- CN
- China
- Prior art keywords
- throat
- air
- pressure
- ablation
- low
- 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.)
- Pending
Links
- 238000001816 cooling Methods 0.000 title claims abstract description 58
- 238000002679 ablation Methods 0.000 title claims abstract description 34
- 239000007921 spray Substances 0.000 claims description 22
- 238000000034 method Methods 0.000 claims description 3
- 230000006835 compression Effects 0.000 claims description 2
- 238000007906 compression Methods 0.000 claims description 2
- 239000000284 extract Substances 0.000 claims description 2
- 239000002245 particle Substances 0.000 description 11
- 239000000112 cooling gas Substances 0.000 description 5
- 230000000694 effects Effects 0.000 description 5
- 239000007789 gas Substances 0.000 description 5
- 230000003628 erosive effect Effects 0.000 description 3
- 238000009413 insulation Methods 0.000 description 2
- 230000006978 adaptation Effects 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000002485 combustion reaction Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification 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
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/97—Rocket nozzles
- F02K9/972—Fluid cooling arrangements for nozzles
-
- 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)
- Jet Pumps And Other Pumps (AREA)
Abstract
The invention discloses a low-ablation rocket engine nozzle structure with an active cooling throat insert and a cooling method, wherein the nozzle structure comprises a nozzle body, an active cooling system and a control system; the nozzle body comprises a skirt part and a throat liner which are integrally formed along the axial direction; an annular common rail cavity is formed in the throat liner in the direction around the axis; the inner wall surface of the throat insert is provided with a plurality of air inlets connected with the common rail cavity; external low-pressure cooling air is extracted and compressed through a low-pressure air inlet pipe by an electric air pump, and high-pressure air is input into the common rail cavity through a high-pressure air outlet pipe connected with the common rail cavity to cool the throat liner; on the other hand, high-pressure air is sent into the nozzle body, a layer of cooling air film is formed on the inner wall surface of the nozzle body, the throat insert and the inner wall surface of the nozzle body are cooled, and the expansion speed of the throat diameter due to high-temperature ablation is reduced; the main nozzle structure provided by the invention can effectively ensure that the engine is in the optimal performance state for a longer time, and improve the specific impulse of the rocket engine.
Description
Technical Field
The invention relates to the technical field of rocket engine spray pipe design, in particular to a low-ablation rocket engine spray pipe structure with an active cooling throat insert and a cooling method.
Background
The nozzle of the rocket engine is a device for converting heat energy into kinetic energy, wherein the size of the throat part determines the working point of the rocket engine, the pressure and the gas flow rate of a combustion chamber can be controlled, and the performance and the working safety of the rocket engine are greatly influenced. The high-temperature and high-pressure heat flow and particle erosion which are circulated when the throat part of the spray pipe works can cause ablation of an inner profile, so that the profile moves back and the size changes, the throat diameter is enlarged, the performance of the engine is finally reduced, and the specific impulse of the engine is reduced. The spray pipe heat insulation structure in the prior art mainly adopts a mode of increasing the thickness of a heat insulation layer to improve the reliability of an engine, increases the passive quality and is not beneficial to improving the mass ratio of the engine. In addition, patent 2020116067168 discloses a web structure for guiding the cooling air to the rear position of the nozzle, and the cooling air is compressed by the profile, but this technique cannot control the pressure of the cooling air in real time, and cannot improve the robustness of the system.
Disclosure of Invention
The purpose of the invention is as follows: aiming at the problems in the prior art, the invention provides a low-ablation rocket engine nozzle structure capable of actively cooling a throat insert and a cooling method. In addition, when the throat diameter is enlarged due to serious ablation at the throat, the diameter of an interface formed by the cooling gas and hot air flow is reduced by increasing the pressure of the cooling gas, namely the effective diameter of the throat diameter is ensured, and the robustness of the nozzle is improved. Meanwhile, the cooling effect of the cooling air film on the inner profile of the spray pipe can reduce the design thickness of the heat insulating layer of the spray pipe and reduce the passive quality.
The technical scheme is as follows: in order to achieve the purpose, the invention adopts the technical scheme that:
a low ablation rocket engine nozzle structure of an active cooling throat insert comprises a nozzle body, an active cooling system and a control system; the nozzle body comprises a skirt part and a throat liner which are integrally formed along the axial direction; an annular common rail cavity is formed in the throat liner along the direction of an axis; the inner wall surface of the throat insert is provided with a plurality of air inlets connected with the common rail cavity; the active cooling system comprises an electric air pump, a low-pressure air inlet pipe and a high-pressure air outlet pipe; the electric air pump extracts external low-pressure cooling air through a low-pressure air inlet pipe and compresses the air, and high-pressure air is input into the common rail cavity through a high-pressure air outlet pipe connected with the common rail cavity; the control system is used for controlling the active cooling system to work.
Further, the electric air pump is mounted on the nozzle body through a bracket; the lower part of the spray pipe body is also provided with a power battery; and the power battery is connected with the electric air pump through a lead.
Furthermore, after the control system controls the rocket to ignite, the control system further controls the power battery to be communicated with the electric air pump, so that the electric air pump works.
Further, cold air flow is uniformly sprayed out of the air inlet, and the diameter of an interface between the cold air flow and the high-temperature high-pressure high-speed air flow at the position of the throat insert during working is the equivalent throat diameter; according to the actual nozzle ablation condition, the pressure is increased or reduced through the electric air pump, and the equivalent throat diameter can be ensured to be in the effective diameter range.
The active cooling method of the low ablation rocket engine nozzle structure adopting the active cooling throat insert comprises the following steps:
step S1, after the control system controls the rocket to ignite, high-temperature and high-pressure high-speed gas flows through the throat; at the moment, the control system controls the power battery to be communicated with the electric air pump, and the electric air pump starts to work;
step S2, the electric air pump sucks air through the low-pressure air inlet pipe, after compression, the compressed air is sent into the common rail cavity through the high-pressure air outlet pipe, and the throat liner is cooled by the high-pressure air in the annular common rail cavity; high-pressure air flows into the inner wall surface of the throat insert through an air inlet connected with the common rail cavity and is diffused to the skirt part along the flow direction of high-speed air; form one deck cooling air film along whole spray tube body internal profile, further cool off the spray tube body, effectively prevent the ablation of high temperature high-pressure gas to this internal profile of spray tube.
Has the advantages that:
(1) according to the rocket engine nozzle structure with the active cooling system, the annular common rail cavity is arranged, and the active cooling system can cool the throat insert by air with controllable pressure, so that the temperature of the throat insert is reduced.
(2) According to the rocket engine nozzle structure with the active cooling system, the cooling air at the throat is air, the airflow direction is perpendicular to the high-temperature high-pressure high-speed airflow, the particle speed is effectively reduced by the cooling airflow, the impact of particles on the throat is relieved, the particles are further oxidized by the air, and the smaller particles have smaller erosion effect on the throat.
(3) When the throat diameter is enlarged due to serious ablation of the throat, the invention increases the pressure of the cooling air flow by controlling the electric air pump, reduces the diameter of an interface formed by the cooling air and the hot air flow and ensures that the equivalent throat diameter is in an effective diameter range.
The low ablation rocket engine nozzle structure of the active cooling throat insert can slow down the increase of the throat diameter, simultaneously realizes the compensation of the ablation throat diameter by controlling the electric air pump, improves the robustness of the nozzle, can ensure the optimal performance of the engine more durably and improves the specific impulse of the rocket engine.
(4) The air film formed on the throat lining of the spray pipe can have a certain cooling effect on the throat lining and the skirt part of the spray pipe, so that the reliability is improved, and under the application of the same working condition, a thinner heat-insulating layer of the spray pipe can be adopted, so that the passive quality of the spray pipe is reduced.
Drawings
FIG. 1 is a schematic view of a low ablation rocket engine nozzle configuration for an actively cooled throat insert provided in accordance with the present invention;
FIG. 2 is a cross-sectional view of a low ablation rocket engine nozzle configuration for an actively cooled throat insert provided in accordance with the present invention;
FIG. 3 is a schematic view of the control principle of a low ablation rocket engine nozzle of an active cooling throat insert provided by the present invention;
FIG. 4 is a schematic view of the low ablation rocket engine nozzle structure with the actively cooling throat insert provided by the present invention.
Description of reference numerals:
1-a nozzle body; 1.1-skirt; 1.2-throat liner; 1.3-common rail cavity; 1.4-air intake; 2-an active cooling system; 2.1-electric air pump; 2.2-low pressure inlet pipe; 2.3-high pressure air outlet pipe; 2.4-scaffold; 3-a power battery; 3.1-wire; 4-a control system; 5, igniting the rocket; 6-cooling the air film; 6.1-cooling gas direction; 7-high-speed high-temperature high-pressure airflow; 7.1-hot gas flow particle direction.
Detailed Description
The invention will be further described with reference to the following description of an embodiment thereof, which is provided in connection with the accompanying drawings.
The invention provides a low-ablation rocket engine nozzle structure capable of actively cooling a throat insert, which is shown in figure 1 and comprises a nozzle body 1, an active cooling system 2, a power battery 3 and a control system 4. The nozzle body 1 comprises a truncated cone-shaped skirt 1.1 and a cylindrical throat insert 1.2 which are integrally formed in the axial direction. A circular common rail cavity 1.3 is arranged in the throat liner 1.2 along the direction of an axis. The inner wall surface of the throat insert 1.2 is provided with a plurality of air inlet holes 1.4 connected with the common rail cavity 1.3. In particular, in the present embodiment, both the skirt portion and the throat insert portion are made of an ablation-resistant structure.
The active cooling system 2 comprises an electric air pump 2.1, a low-pressure air inlet pipe 2.2, a high-pressure air outlet pipe 2.3 and a support 2.4. The electric air pump 2.1 is arranged at the upper end of the nozzle body 1 and is fixedly installed through a support 2.4. The electric air pump 2.1 is externally extended with a low pressure air inlet pipe 2.2, external low pressure cooling air is extracted and compressed through the low pressure air inlet pipe 2.2, and high pressure gas is input into the common rail cavity 1.3 through a high pressure air outlet pipe 2.3 connected with the common rail cavity 1.3, as shown in fig. 2.
The electric air pump controls the air inlet to uniformly spray cold air flow, and the diameter of the interface of the air inlet and the high-temperature, high-pressure and high-speed air flow at the position of the throat lining is the equivalent throat diameter when the electric air pump works; according to the actual nozzle ablation condition, the pressure is increased or reduced through the electric air pump, and the equivalent throat diameter can be ensured to be in the effective diameter range.
The power battery 3 is fixedly arranged at the lower part of the nozzle body 1, and connects the electric air pump 2.1 together through a lead 3.1 to supply power for the electric air pump 2.1.
The control method of the low ablation rocket engine nozzle structure of the active cooling throat insert is shown in figure 3, and after the control system 4 controls the rocket ignition 5 to work, the power battery 3 is controlled to be communicated with the electric air pump 2.1 through a circuit, so that the electric air pump 2.1 is controlled to work normally. The electric air pump 2.1 pressurizes air sucked from the low-pressure air inlet pipe 2.2, the high-pressure air is sent into the common rail cavity 1.3 through the high-pressure air outlet pipe 2.3, the throat liner 1.2 is cooled by the high-pressure air in the high-pressure common rail cavity 1.3, meanwhile, the high-pressure air is sent into the nozzle body 1 through the air inlet 1.4 in the throat liner 1.2, a layer of cooling air film 6 is formed on the inner profile of the nozzle body 1 under the action of high-speed high-temperature high-pressure airflow 7 on the inner wall surface of the nozzle body 1, the throat liner 1.2 and the inner wall surface of the nozzle body 1 are cooled, and the expansion speed of the throat diameter due to high-temperature ablation is reduced, and the electric air pump is particularly shown in fig. 4.
Wherein, the cooling gas of the throat insert 1.2 is air and the direction 6.1 of the cooling gas is vertical to the direction 7.1 of hot air flow particles, the air flow reduces the particle speed, the impact of the particles on the throat insert 1.2 is slowed down, simultaneously, the air further oxidizes the particles, and the smaller particles have smaller erosion effect on the throat insert 1.2.
In addition, when the throat diameter is enlarged due to serious ablation of the throat liner 1.2, the control system 4 controls the electric air pump 1.2 to improve the pressure of the cooling air, reduce the diameter of an interface formed by the cooling air and hot air flow, ensure that the equivalent throat diameter is always maintained in an effective throat diameter range, slow down the increase of the throat diameter, simultaneously realize the compensation of the ablation throat diameter by controlling the electric air pump, obviously improve the robustness of the spray pipe, effectively ensure that the engine is in an optimal performance state for a longer time, and improve the specific impulse of the rocket engine. The cooling effect of the cooling air film on the inner profile of the spray pipe can reduce the design thickness of the heat insulating layer of the spray pipe and reduce the passive quality.
The above description is only of the preferred embodiments of the present invention, and it should be noted that: it will be apparent to those skilled in the art that various modifications and adaptations can be made without departing from the principles of the invention and these are intended to be within the scope of the invention.
Claims (5)
1. A low ablation rocket engine nozzle structure of an active cooling throat insert is characterized by comprising a nozzle body, an active cooling system and a control system; the nozzle body comprises a skirt part and a throat liner which are integrally formed along the axial direction; an annular common rail cavity is formed in the throat liner along the direction of an axis; the inner wall surface of the throat insert is provided with a plurality of air inlets connected with the common rail cavity; the active cooling system comprises an electric air pump, a low-pressure air inlet pipe and a high-pressure air outlet pipe; the electric air pump extracts external low-pressure cooling air through a low-pressure air inlet pipe and compresses the air, and high-pressure air is input into the common rail cavity through a high-pressure air outlet pipe connected with the common rail cavity; the control system is used for controlling the active cooling system to work.
2. The actively-cooled throat insert low ablation rocket engine nozzle structure as claimed in claim 1 wherein said electric air pump is mounted on the nozzle body by a bracket; the lower part of the spray pipe body is also provided with a power battery; and the power battery is connected with the electric air pump through a lead.
3. The low ablation rocket engine nozzle structure of active cooling throat insert of claim 1 wherein said control system further controls the power battery to communicate with the electric air pump to operate the electric air pump after controlling rocket ignition.
4. The low ablation rocket engine nozzle structure of claim 3 wherein the cold air flow is ejected uniformly from the inlet orifice, and the diameter of the interface between the cold air flow and the high temperature, high pressure and high velocity air flow at the throat insert position during operation is the equivalent throat diameter; according to the actual nozzle ablation condition, the pressure is increased or reduced through the electric air pump, and the equivalent throat diameter can be ensured to be in the effective diameter range.
5. A method of actively cooling a low ablation rocket engine nozzle structure based on an actively cooled throat insert according to any one of claims 1-4, comprising the steps of:
step S1, after the control system controls the rocket to ignite, high-temperature and high-pressure high-speed gas flows through the throat; at the moment, the control system controls the power battery to be communicated with the electric air pump, and the electric air pump starts to work;
step S2, the electric air pump sucks air through the low-pressure air inlet pipe, after compression, the compressed air is sent into the common rail cavity through the high-pressure air outlet pipe, and the throat liner is cooled by the high-pressure air in the annular common rail cavity; high-pressure air flows into the inner wall surface of the throat insert through an air inlet connected with the common rail cavity and is diffused to the skirt part along the flow direction of high-speed air; a layer of cooling air film is formed along the inner profile of the whole spray pipe body, the spray pipe body is further cooled, and ablation of high-temperature and high-pressure air on the inner profile of the spray pipe body is prevented;
and step S3, when the actual throat diameter is increased due to ablation of the throat liner, the control system controls the electric air pump to increase the pressure of the cooling air flow, reduce the equivalent throat diameter, ensure that the equivalent throat diameter is always maintained in the range of the effective throat diameter, and slow down the increase of the throat diameter.
Priority Applications (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210378851.4A CN114810430A (en) | 2022-04-12 | 2022-04-12 | Low-ablation rocket engine nozzle structure with active cooling throat insert and cooling method |
| ZA2022/12593A ZA202212593B (en) | 2022-04-12 | 2022-11-21 | A structure and cooling method of a low-ablation rocket engine nozzle with active-cooling throat-insert |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210378851.4A CN114810430A (en) | 2022-04-12 | 2022-04-12 | Low-ablation rocket engine nozzle structure with active cooling throat insert and cooling method |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN114810430A true CN114810430A (en) | 2022-07-29 |
Family
ID=82534578
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202210378851.4A Pending CN114810430A (en) | 2022-04-12 | 2022-04-12 | Low-ablation rocket engine nozzle structure with active cooling throat insert and cooling method |
Country Status (2)
| Country | Link |
|---|---|
| CN (1) | CN114810430A (en) |
| ZA (1) | ZA202212593B (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115434829A (en) * | 2022-10-17 | 2022-12-06 | 西安探火航天技术有限公司 | Reusable rocket engine nozzle assembly with variable expansion ratio |
| CN115434828A (en) * | 2022-10-17 | 2022-12-06 | 西安探火航天技术有限公司 | Variable expansion ratio rocket engine jet pipe |
Citations (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5159809A (en) * | 1989-12-21 | 1992-11-03 | Societe Europeenne De Propulsion | Highly adaptable combined propulsion engine for an aircraft or a space-going airplane |
| US5996936A (en) * | 1997-09-29 | 1999-12-07 | General Electric Company | Fluidic throat exhaust nozzle |
| US20080141656A1 (en) * | 2006-08-01 | 2008-06-19 | Snecma | Bypass turbomachine with artificial variation of its throat section |
| JP2011027120A (en) * | 2010-11-11 | 2011-02-10 | Mitsubishi Heavy Ind Ltd | Rocket nozzle and control method for rocket engine combustion gas flow |
| US20140331682A1 (en) * | 2012-11-08 | 2014-11-13 | Mark Bovankovich | High-speed-launch ramjet booster |
| JP2016061263A (en) * | 2014-09-19 | 2016-04-25 | 三菱重工メカトロシステムズ株式会社 | Rocket engine high altitude combustion test equipment, and method of operating rocket engine high altitude combustion test equipment |
| CN111122767A (en) * | 2019-11-29 | 2020-05-08 | 南京理工大学 | Detachable solid rocket engine jet pipe throat lining ablation test device |
| US20200338494A1 (en) * | 2019-04-24 | 2020-10-29 | Norman A. Pullen | Jet engine and method of operation |
| CN112672454A (en) * | 2020-12-30 | 2021-04-16 | 中国航天空气动力技术研究院 | Inner wall air film of electric arc heater |
| CN112682954A (en) * | 2020-12-30 | 2021-04-20 | 中国航天空气动力技术研究院 | Electrode of electric arc heater |
| CN213360272U (en) * | 2020-10-10 | 2021-06-04 | 江西洪都航空工业集团有限责任公司 | Pneumatic type spray pipe throat diameter adjusting device |
| CN113107611A (en) * | 2021-04-22 | 2021-07-13 | 南京航空航天大学 | Turbine blade trailing edge cooling structure based on double-throat pneumatic vectoring nozzle and trailing path control method thereof |
-
2022
- 2022-04-12 CN CN202210378851.4A patent/CN114810430A/en active Pending
- 2022-11-21 ZA ZA2022/12593A patent/ZA202212593B/en unknown
Patent Citations (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5159809A (en) * | 1989-12-21 | 1992-11-03 | Societe Europeenne De Propulsion | Highly adaptable combined propulsion engine for an aircraft or a space-going airplane |
| US5996936A (en) * | 1997-09-29 | 1999-12-07 | General Electric Company | Fluidic throat exhaust nozzle |
| US20080141656A1 (en) * | 2006-08-01 | 2008-06-19 | Snecma | Bypass turbomachine with artificial variation of its throat section |
| JP2011027120A (en) * | 2010-11-11 | 2011-02-10 | Mitsubishi Heavy Ind Ltd | Rocket nozzle and control method for rocket engine combustion gas flow |
| US20140331682A1 (en) * | 2012-11-08 | 2014-11-13 | Mark Bovankovich | High-speed-launch ramjet booster |
| JP2016061263A (en) * | 2014-09-19 | 2016-04-25 | 三菱重工メカトロシステムズ株式会社 | Rocket engine high altitude combustion test equipment, and method of operating rocket engine high altitude combustion test equipment |
| US20200338494A1 (en) * | 2019-04-24 | 2020-10-29 | Norman A. Pullen | Jet engine and method of operation |
| CN111122767A (en) * | 2019-11-29 | 2020-05-08 | 南京理工大学 | Detachable solid rocket engine jet pipe throat lining ablation test device |
| CN213360272U (en) * | 2020-10-10 | 2021-06-04 | 江西洪都航空工业集团有限责任公司 | Pneumatic type spray pipe throat diameter adjusting device |
| CN112672454A (en) * | 2020-12-30 | 2021-04-16 | 中国航天空气动力技术研究院 | Inner wall air film of electric arc heater |
| CN112682954A (en) * | 2020-12-30 | 2021-04-20 | 中国航天空气动力技术研究院 | Electrode of electric arc heater |
| CN113107611A (en) * | 2021-04-22 | 2021-07-13 | 南京航空航天大学 | Turbine blade trailing edge cooling structure based on double-throat pneumatic vectoring nozzle and trailing path control method thereof |
Non-Patent Citations (3)
| Title |
|---|
| ZHAOHUI YAO等: "Performance study of a pre-cooled turbo-rocket combined engine under a wide Mach number of 0~5", 《CASE STUDIES IN THERMAL ENGINEERING》 * |
| 孙瑞斌等: "高超声速风洞喷管中气膜的防热效应和流场影响", 《兵器装备工程学报》 * |
| 王立武等: "固体火箭发动机喷管喉衬烧蚀研究进展", 《固体火箭技术》 * |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115434829A (en) * | 2022-10-17 | 2022-12-06 | 西安探火航天技术有限公司 | Reusable rocket engine nozzle assembly with variable expansion ratio |
| CN115434828A (en) * | 2022-10-17 | 2022-12-06 | 西安探火航天技术有限公司 | Variable expansion ratio rocket engine jet pipe |
| CN115434828B (en) * | 2022-10-17 | 2023-08-29 | 西安探火航天技术有限公司 | Rocket engine spray pipe with variable expansion ratio |
Also Published As
| Publication number | Publication date |
|---|---|
| ZA202212593B (en) | 2023-03-29 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN114810430A (en) | Low-ablation rocket engine nozzle structure with active cooling throat insert and cooling method | |
| CN101067372B (en) | High-pressure mixed gas generator used for petroleum thermal recovery gas injection machine | |
| CN101975122B (en) | Stabilized knocking engine with magnetic fluid energy bypath system | |
| CN109538375B (en) | Injection rocket applied to rocket-based combined cycle engine | |
| CN113137634A (en) | Variable-structure bimodal stamping combustion chamber | |
| CN115726904A (en) | Thrust device of liquid rocket engine in composite cooling form | |
| CN108825405A (en) | A kind of full runner of axially symmetric structure RBCC using multistep rocket | |
| CN102893009B (en) | Device for reducing the noise emitted by the jet of an aircraft propulsion engine | |
| CN112211753A (en) | Novel rocket engine tail nozzle throat structure | |
| CN112594737B (en) | Oblique detonation wave stationary control method and variable-geometry combustion chamber | |
| CN110822475A (en) | On-duty flame stabilizer using air cooling structure to match ignition position and design method | |
| CN101825041B (en) | Gas piston pulse engine | |
| CN107100740A (en) | Air-flow compresses directional cumulation plasma arc ignition burner | |
| CN104832318A (en) | Ramjet engine | |
| CN216691282U (en) | Combined propulsion system of separate-exhaust air turbine rocket | |
| CN113048516B (en) | Detonation combustion chamber, scramjet engine and hypersonic aircraft | |
| CN211399862U (en) | Energy-saving consumption-reducing high-efficiency combustion chamber | |
| CN214533288U (en) | Long tail nozzle structure of segmentation of solid rocket engine | |
| CN115450798A (en) | Long-time working solid rocket engine and long tail jet pipe thereof | |
| CN101929406A (en) | Vortex cold vacuum aircraft engine | |
| CN114876673A (en) | Low-cost ablation-resistant embedded spray pipe and machining method thereof | |
| CN108087153A (en) | A kind of Ducted rocket with cooling combination | |
| CN210861777U (en) | Refrigeration cycle system for aircraft and aircraft | |
| CN110107402B (en) | Ceramic turbine engine | |
| CN109252979B (en) | A kind of cyclone air suction type aero-engine and its driving method |
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 | ||
| RJ01 | Rejection of invention patent application after publication | ||
| RJ01 | Rejection of invention patent application after publication |
Application publication date: 20220729 |