CN116291970A - Observable solid rocket engine ignition test platform - Google Patents
Observable solid rocket engine ignition test platform Download PDFInfo
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- CN116291970A CN116291970A CN202310105185.1A CN202310105185A CN116291970A CN 116291970 A CN116291970 A CN 116291970A CN 202310105185 A CN202310105185 A CN 202310105185A CN 116291970 A CN116291970 A CN 116291970A
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- quartz glass
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- 239000007787 solid Substances 0.000 title claims abstract description 34
- 238000012360 testing method Methods 0.000 title claims abstract description 29
- 238000002485 combustion reaction Methods 0.000 claims abstract description 50
- VYPSYNLAJGMNEJ-UHFFFAOYSA-N Silicium dioxide Chemical compound O=[Si]=O VYPSYNLAJGMNEJ-UHFFFAOYSA-N 0.000 claims abstract description 48
- 239000007921 spray Substances 0.000 claims abstract description 20
- RYGMFSIKBFXOCR-UHFFFAOYSA-N Copper Chemical compound [Cu] RYGMFSIKBFXOCR-UHFFFAOYSA-N 0.000 claims description 16
- 229910052802 copper Inorganic materials 0.000 claims description 16
- 239000010949 copper Substances 0.000 claims description 16
- 229910052751 metal Inorganic materials 0.000 claims description 7
- 239000002184 metal Substances 0.000 claims description 7
- 238000009413 insulation Methods 0.000 claims description 4
- 238000007405 data analysis Methods 0.000 claims description 3
- 239000011248 coating agent Substances 0.000 claims description 2
- 238000000576 coating method Methods 0.000 claims description 2
- 238000000034 method Methods 0.000 abstract description 24
- 230000008569 process Effects 0.000 abstract description 20
- 230000001052 transient effect Effects 0.000 abstract description 15
- 239000002360 explosive Substances 0.000 description 12
- 239000003380 propellant Substances 0.000 description 9
- 239000010410 layer Substances 0.000 description 8
- 238000011160 research Methods 0.000 description 5
- 229910000851 Alloy steel Inorganic materials 0.000 description 3
- 229910000629 Rh alloy Inorganic materials 0.000 description 3
- 230000008859 change Effects 0.000 description 3
- 239000002737 fuel gas Substances 0.000 description 3
- 239000002245 particle Substances 0.000 description 3
- PXXKQOPKNFECSZ-UHFFFAOYSA-N platinum rhodium Chemical compound [Rh].[Pt] PXXKQOPKNFECSZ-UHFFFAOYSA-N 0.000 description 3
- 239000000843 powder Substances 0.000 description 3
- 230000000903 blocking effect Effects 0.000 description 2
- 239000011247 coating layer Substances 0.000 description 2
- 230000001808 coupling effect Effects 0.000 description 2
- 230000007123 defense Effects 0.000 description 2
- 238000013461 design Methods 0.000 description 2
- 239000007789 gas Substances 0.000 description 2
- 239000004449 solid propellant Substances 0.000 description 2
- 238000012546 transfer Methods 0.000 description 2
- 244000003416 Asparagus officinalis Species 0.000 description 1
- 235000005340 Asparagus officinalis Nutrition 0.000 description 1
- AZKVWQKMDGGDSV-BCMRRPTOSA-N Genipin Chemical group COC(=O)C1=CO[C@@H](O)[C@@H]2C(CO)=CC[C@H]12 AZKVWQKMDGGDSV-BCMRRPTOSA-N 0.000 description 1
- 230000009471 action Effects 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000008878 coupling Effects 0.000 description 1
- 238000010168 coupling process Methods 0.000 description 1
- 238000005859 coupling reaction Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000010304 firing Methods 0.000 description 1
- 230000000977 initiatory effect Effects 0.000 description 1
- 238000012423 maintenance Methods 0.000 description 1
- 230000007246 mechanism Effects 0.000 description 1
- 238000012544 monitoring process Methods 0.000 description 1
- 238000011056 performance test Methods 0.000 description 1
- 238000004080 punching Methods 0.000 description 1
- 238000004088 simulation Methods 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
- 238000010998 test method Methods 0.000 description 1
- 238000013519 translation Methods 0.000 description 1
Images
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
- F02K9/96—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by specially adapted arrangements for testing or measuring
-
- 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
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/10—Shape or structure of solid propellant charges
-
- 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
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/32—Constructional parts; Details not otherwise provided for
-
- 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
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/32—Constructional parts; Details not otherwise provided for
- F02K9/34—Casings; Combustion chambers; Liners thereof
-
- 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
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/32—Constructional parts; Details not otherwise provided for
- F02K9/34—Casings; Combustion chambers; Liners thereof
- F02K9/343—Joints, connections, seals therefor
-
- 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
- F02K9/95—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by starting or ignition means or arrangements
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Testing Of Engines (AREA)
- Investigating Or Analyzing Materials Using Thermal Means (AREA)
Abstract
The invention relates to an observable solid rocket engine ignition test platform, which comprises an igniter, a measuring section and a spray pipe, wherein the igniter is arranged at the front end of the measuring section, and the spray pipe is arranged at the rear end of the measuring section; the measuring section comprises a Fang Xingqiang body, a quartz glass cover plate, quartz glass, a grain section, a temperature sensor, a pressure sensor, a combustion chamber and a high-speed camera. The invention aims to solve the problem that the ignition transient process of the engine cannot be directly observed and the problem that the temperature/pressure in the combustion chamber cannot be directly monitored.
Description
Technical Field
The invention relates to an observable solid rocket engine ignition test platform.
Background
The solid rocket engine has the characteristics of simple structure, excellent performance, good maneuverability, low cost, easy maintenance and the like, and is widely applied to the fields of national defense, military and aerospace. The ignition transient process of the solid engine is a coupling action process between the jet flow of the igniter and the propellant grain, namely, when the jet flow of the igniter acts on the surface of the grain, the temperature of the surface of the propellant rises, when the temperature of a certain position of the grain reaches the ignition point of the grain, the grain is ignited at the position, flame is formed and spread in the combustion chamber, and the newly generated hot fuel gas continues to ignite the propellant at other positions together with the jet flow of the igniter, so that a large amount of high-temperature and high-pressure fuel gas is generated, and then the high-temperature and high-pressure fuel gas is discharged through the Laval nozzle in an accelerating way, so that thrust is provided for a solid rocket.
The solid rocket engine is usually ignited in several milliseconds to hundreds of milliseconds, the action time is short, the physical quantity is changed severely, and the data indicate that the faults caused by the ignition stage account for more than 10% of the total faults of the engine [1]. Therefore, the research on the matching among igniter parameters, propellant types and tail nozzle structures in the ignition transient process is carried out, and technical support can be provided for the optimal design of the solid rocket engine.
The ignition transient process of the solid rocket engine is a complex strong transient and unsteady process, relates to the problems of gas-solid two-phase combustion reaction and fluid-solid coupling heat and mass transfer, and adopts a theoretical method and a numerical simulation method to research the transient process to have lower precision, so that the related research work is carried out by a test method in the present stage. However, as the solid rocket engine test device is mostly composed of totally-enclosed high-temperature-resistant alloy steel, the coupling action process between the jet flow of the igniter in the combustion chamber and the propellant grains cannot be observed [2]; meanwhile, as the traditional solid propellant burns high, the temperature sensor cannot directly measure the temperature inside the combustion chamber, only can measure the temperature of the outer wall surface of the combustion chamber, and then the temperature inside the combustion chamber is calculated through a heat transfer formula, so that the accuracy of test data is reduced; in addition, most solid rocket engine ignition test devices are only provided with pressure sensors in a punching mode at the position of a spray pipe, the pressure change at the position of the spray pipe is used for replacing the pressure change [3] in the combustion chamber, the evolution process of the pressure in the combustion chamber is ignored, the pressure distribution in the combustion chamber of the solid rocket engine with large length-diameter ratio has large difference, and the pressure at the position of the spray pipe cannot be used for representing the pressure of a flow field in the ignition transient process.
Based on the problems, the current transient ignition test of the solid rocket engine has the problems of incapability of observation, lower test data precision and the like, and the test data is corrected according to engineering experience when the actual engine is designed, so that the optimal design of the solid rocket engine has redundancy and even can cause the working failure of the engine.
Reference to the literature
[1] Williams F.A, huang N.C, barrel M. Basic problem of solid propellant rocket engines [ M ]. Genipin group translation.
National defense industry Press, 1977:125:382-384.
[2] Li Jian, wu Feichun, asparagus Ji Yang, etc. small-sized launching-stage solid rocket engine ignition performance test research [ J ]. Initiating explosive device,
2018(3):6-8.
[3] zhao Ruyan, huang Zhiyong, zhou Gongmei, et al solid engine single firing test study [ J ]. Projectile arrow and guidance, 2011,31 (6): 122-125.
Disclosure of Invention
Aiming at the defects of the prior art, the invention aims to provide an observable solid rocket engine ignition test platform, and aims to solve the problems that the ignition transient process of an engine cannot be directly observed and the temperature/pressure in a combustion chamber cannot be directly monitored.
The specific effects are as follows:
(1) And quartz glass is arranged on the upper half part of the combustion chamber and is sealed with high-temperature resistant alloy steel on the lower half part through bolts and red copper gaskets, so that the combustion chamber with the transparent window is formed. Then a high-speed camera is placed above the combustion chamber cavity, so that the ignition transient process of the solid rocket engine can be recorded through high-speed shooting, and the problem that the ignition test of the engine cannot be directly observed is solved.
(2) The test process adopts a platinum rhodium alloy temperature/pressure sensor, can work under 1900K temperature environment, can be directly installed in a combustion chamber to monitor the temperature of an internal flow field, avoids monitoring the temperature of the outer wall surface of the combustion chamber, and then calculates the data error caused by the temperature in the combustion chamber through an empirical formula.
(3) The temperature/pressure sensor is embedded into the propellant grains in the charging process, and the temperature/pressure change process in the combustion chamber in the ignition transient process of the solid rocket engine is directly measured, so that the temperature/pressure difference at different positions in the combustion chamber can be obtained, and the method is particularly suitable for engines with large length-diameter ratio, and is more beneficial to revealing the flame propagation mechanism in the combustion chamber.
The invention is realized by the following technical scheme.
An observable solid rocket engine ignition test platform comprises an igniter, a measuring section and a spray pipe, wherein the igniter is arranged at the front end of the measuring section, and the spray pipe is arranged at the rear end of the measuring section;
the measuring section comprises a Fang Xingqiang body, a quartz glass cover plate, quartz glass, a grain section, a temperature sensor, a pressure sensor, a combustion chamber and a high-speed camera, wherein the combustion chamber is arranged below the square cavity through a red copper gasket and a bolt, the quartz glass is arranged above the square cavity through the quartz glass cover plate, the red copper gasket and the bolt, a cylindrical combustion chamber cavity is formed between the quartz glass and the combustion chamber, the quartz glass comprises thick quartz glass and thin quartz glass, the thin quartz glass clings to the inner wall of the thick quartz glass, the grain section is of a complete grain structure, the grain section is sequentially provided with an ignition grain, a grain coating layer, a grain heat insulation layer and a grain metal layer from inside to outside, the grain metal layer clings to the inner wall of the combustion chamber, through holes are formed in the grain section and the combustion chamber, the through holes are used for setting the temperature sensor and the pressure sensor, the high-speed camera is arranged above the quartz glass, and the temperature sensor and the pressure sensor acquire data and enter a data analysis system.
Further, the igniter is arranged at the front end of the square cavity body through an igniter cover plate, a red copper gasket and bolts.
Further, the spray pipe is installed at the rear end of the measuring section through a red copper gasket and a bolt, and an engine rear plug cover is arranged at the rear end of the spray pipe.
Further, nine temperature sensors and six pressure sensors are buried in the grain section, three rows are provided in total along the axial direction, each row comprises three temperature sensors and two pressure sensors, and the temperature sensors and the pressure sensors are alternately distributed at 15 DEG intervals in the circumferential direction.
Compared with the prior art, the invention has the advantages that:
(1) By changing the cavity structure of the combustion chamber, quartz glass is additionally arranged above the explosive column, and the ignition process of the engine is recorded through a high-speed camera, so that the problem that the ignition transient process of the current solid rocket engine cannot be observed is solved.
(2) The temperature/pressure sensor of the high-temperature-resistant platinum-rhodium alloy is embedded in the propellant grains, so that the pressure/temperature evolution law in the combustion chamber can be directly monitored, and the accuracy of test data is ensured.
(3) The fire section, the test section and the spray pipe section are connected through the bolts, the cover plate and the red copper gasket, so that igniters and spray pipes with different configurations can be replaced conveniently, and the matching research among igniter parameters, propellant types and tail spray pipe structures in the ignition transient process can be carried out.
(4) Because a large amount of black solid particles can be generated in the solid engine ignition test, the solid particles can be attached to the surface of the quartz glass after the test is finished, the bottom of the thick quartz glass is padded with a thin quartz glass sheet, and the thin quartz glass sheet is replaced after each test, so that the test platform can be reused;
(5) The square cavity body enables the test section to be fixed more easily, and the stability of the device is prevented from being influenced by jet flow of the igniter.
(6) The explosive column section is of a complete explosive column structure, and the temperature sensor and the pressure sensor are directly buried in the explosive column, so that the ignition transient process of the solid rocket engine under the real condition can be better tested
Drawings
FIG. 1 is a standard isometric view of the invention
FIG. 2 is a block diagram of the overall split of the apparatus of the present invention
FIG. 3 is a schematic view showing the connection of the igniting powder column and the sensor of the device
FIG. 4 is a schematic view of the ignition powder column of the device of the present invention
FIG. 5 is a yz cross-sectional view of the present invention
FIG. 6 is an integrated view of a solid rocket engine ignition test platform system according to the present invention
FIG. 7 is a flow chart of the present invention;
in the figure: 1. an igniter; 2. an igniter cover plate; 3. a red copper gasket; 4. a quartz glass cover plate; 5. thick quartz glass; 6. thin quartz glass; 7. a grain section; 8. a temperature sensor; 9. a pressure sensor; 10. a combustion chamber; 11. a combustion chamber cavity; 12. a spray pipe; 13. a rear plug cover of the engine; 14. igniting the powder column; 15. a grain coating layer; 16. a grain insulation layer; 17. a grain metal layer; 18. a high-speed camera; 19. a pressure/temperature data acquisition system; 20. and a square cavity.
Detailed Description
The invention is further described below with reference to the drawings and specific examples, which are not intended to be limiting.
As shown in fig. 1 to 7, an observable solid rocket engine ignition test platform comprises an igniter 1, a measuring section and a spray pipe 12, wherein the igniter 1 is arranged at the front end of the measuring section, and the spray pipe 12 is arranged at the rear end of the measuring section;
the measuring section comprises a Fang Xingqiang body 20, a quartz glass cover plate 4, quartz glass, a grain section 7, a temperature sensor 8, a pressure sensor 9, a combustion chamber 10 and a high-speed camera 18, wherein the combustion chamber 10 is arranged below the square cavity 20 through a red copper gasket and a bolt, the quartz glass is arranged above the square cavity 20 through the quartz glass cover plate 4, the red copper gasket and the bolt, a cylindrical combustion chamber cavity 11 is formed between the quartz glass and the combustion chamber 10, the quartz glass comprises a thick quartz glass 5 and a thin quartz glass 6, the thin quartz glass 6 is tightly attached to the inner wall of the thick quartz glass 5, the grain section 7 is of a complete grain structure, the grain section 7 is sequentially provided with a igniting grain 14, a grain coating 15, a grain heat insulation layer 16 and a grain metal layer 17 from inside to outside, the grain metal layer 17 is tightly attached to the inner wall of the combustion chamber 10, through holes are formed in the grain section 7 and the combustion chamber 10, the through holes are used for setting the temperature sensor 8 and the temperature sensor 9, the high-speed camera 6 is tightly attached to the inner wall of the combustion chamber 10, and the temperature sensor 8 is arranged above the high-speed camera and the temperature sensor 8, and the temperature sensor 8 is arranged at the terminal end of the pressure sensor and the pressure sensor is used for data acquisition.
Further, the igniter 1 is mounted at the front end of the square cavity 20 through an igniter cover plate 2, a red copper gasket and bolts.
Further, the spray pipe 12 is installed at the rear end of the measuring section through a red copper gasket and a bolt, and an engine rear plug 13 is arranged at the rear end of the spray pipe 12.
Further, nine temperature sensors and six pressure sensors are buried in the column section 7 in total, and three rows are provided in the axial direction, each row including three temperature sensors and two pressure sensors alternately distributed at 15 ° intervals in the circumferential direction.
The data analysis system terminal is a pressure/temperature data acquisition system.
The combustion chamber 10 is made of high temperature alloy steel.
The temperature sensor 8 and the pressure sensor 9 are platinum rhodium alloy temperature/pressure sensors.
The implementation method of the invention is as follows: after the explosive column section is installed in the combustion chamber, the igniter, the quartz glass, the combustion chamber, the spray pipe and the rear engine blocking cover are hermetically connected through a red copper gasket and a bolt, a pressure/temperature sensor is connected, the sensor is connected with a data acquisition system, and a nodding type high-speed photographic instrument is installed above the device. After the igniter ignites, the gas jet flows impact the explosive column at the ignition impact point on the surface of the explosive column, the temperature in the combustion chamber rises sharply after the explosive column ignites, and the propellant at other positions is ignited after the ignition point of the explosive column is reached. The explosive column burns to increase the pressure inside the combustion chamber, so that the nozzle blocking cover is broken, and high-temperature and high-pressure gas is accelerated to be sprayed out through the Laval nozzle. The entire ignition transient is recorded by a high-speed camera through quartz glass and pressure/temperature changes at different locations are measured by pressure/temperature sensors. After the test is finished, the thin quartz glass can be replaced to prevent black particles generated by burning of the explosive column from adhering to the quartz glass, so that the repeated use of the device is affected.
The foregoing description is only illustrative of the preferred embodiments of the present invention and is not to be construed as limiting the scope of the invention, and it will be appreciated by those skilled in the art that equivalent substitutions and obvious variations may be made using the description and illustrations of the present invention, and are intended to be included within the scope of the present invention.
Claims (4)
1. The observable solid rocket engine ignition test platform is characterized by comprising an igniter (1), a measuring section and a spray pipe (12), wherein the igniter (1) is arranged at the front end of the measuring section, and the spray pipe (12) is arranged at the rear end of the measuring section;
the measuring section comprises a Fang Xingqiang body (20), a quartz glass cover plate (4), quartz glass, a grain section (7), a temperature sensor (8), a pressure sensor (9), a combustion chamber (10) and a high-speed camera (18), wherein the combustion chamber (10) is arranged below the square cavity (20) through a red copper gasket and a bolt, the quartz glass is arranged above the square cavity (20) through the quartz glass cover plate (4), the red copper gasket and the bolt, a cylindrical combustion chamber cavity (11) is formed between the quartz glass and the combustion chamber (10), the quartz glass comprises a thick quartz glass (5) and a thin quartz glass (6), the thin quartz glass (6) is clung to the inner wall of the thick quartz glass (5), the grain section (7) is of a complete grain structure, the grain section (7) is sequentially provided with a grain (14), a grain coating (15), a grain heat insulation layer (16), a grain metal layer (17), the inner wall of the combustion chamber (10) is clung to the inner wall of the column metal layer (17), the combustion chamber (7) and the combustion chamber (10) is provided with a through hole (18), the temperature sensor (8) is arranged on the grain section, the temperature sensor (9) is arranged on the high-speed camera, and data acquired by the temperature sensor (8) and the pressure sensor (9) enter a data analysis system terminal.
2. The observable solid rocket engine ignition test platform of claim 1, characterized in that the igniter (1) is mounted at the front end of the square cavity (20) through an igniter cover plate (2), a red copper gasket and bolts.
3. The observable solid rocket engine ignition test platform of claim 1, characterized in that the jet pipe (12) is mounted at the rear end of the measuring section through a red copper gasket and a bolt, and the rear end of the jet pipe (12) is provided with an engine rear plug cover (13).
4. An observable solid rocket engine ignition test platform according to claim 1, characterized in that the grain section (7) has nine temperature sensors and six pressure sensors embedded therein, and three rows of three temperature sensors and two pressure sensors are arranged in total along the axial direction, each row comprising three temperature sensors and two pressure sensors alternately distributed at 15 ° intervals in the circumferential direction.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310105185.1A CN116291970A (en) | 2023-02-13 | 2023-02-13 | Observable solid rocket engine ignition test platform |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310105185.1A CN116291970A (en) | 2023-02-13 | 2023-02-13 | Observable solid rocket engine ignition test platform |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN116291970A true CN116291970A (en) | 2023-06-23 |
Family
ID=86826586
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202310105185.1A Pending CN116291970A (en) | 2023-02-13 | 2023-02-13 | Observable solid rocket engine ignition test platform |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN116291970A (en) |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114810429A (en) * | 2022-04-11 | 2022-07-29 | 北京航空航天大学 | Device and method for measuring charge burning rate of solid-liquid rocket engine |
| CN116877300A (en) * | 2023-09-05 | 2023-10-13 | 南京理工大学 | Solid rocket engine propellant erosion function measuring device and measuring method |
| CN117250006A (en) * | 2023-11-20 | 2023-12-19 | 中国空气动力研究与发展中心设备设计与测试技术研究所 | Rocket-based combined cycle model engine combustion chamber tester |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5379699A (en) * | 1993-08-02 | 1995-01-10 | The United States Of America As Represented By The Secretary Of The Navy | Active spray rocket propellant ignition controller |
| CN108981503A (en) * | 2018-09-05 | 2018-12-11 | 西安近代化学研究所 | A kind of condensed phase high explosive detonation property multi-parameter method for synchronously measuring |
| CN112485006A (en) * | 2020-11-13 | 2021-03-12 | 东北大学 | Solid rocket engine experiment table and combustion chamber heat insulation layer ablation thickness measuring method |
| CN114810429A (en) * | 2022-04-11 | 2022-07-29 | 北京航空航天大学 | Device and method for measuring charge burning rate of solid-liquid rocket engine |
| CN115112376A (en) * | 2022-07-05 | 2022-09-27 | 南昌航空大学 | Gas injection and propellant coupling effect observation test device |
-
2023
- 2023-02-13 CN CN202310105185.1A patent/CN116291970A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5379699A (en) * | 1993-08-02 | 1995-01-10 | The United States Of America As Represented By The Secretary Of The Navy | Active spray rocket propellant ignition controller |
| CN108981503A (en) * | 2018-09-05 | 2018-12-11 | 西安近代化学研究所 | A kind of condensed phase high explosive detonation property multi-parameter method for synchronously measuring |
| CN112485006A (en) * | 2020-11-13 | 2021-03-12 | 东北大学 | Solid rocket engine experiment table and combustion chamber heat insulation layer ablation thickness measuring method |
| CN114810429A (en) * | 2022-04-11 | 2022-07-29 | 北京航空航天大学 | Device and method for measuring charge burning rate of solid-liquid rocket engine |
| CN115112376A (en) * | 2022-07-05 | 2022-09-27 | 南昌航空大学 | Gas injection and propellant coupling effect observation test device |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114810429A (en) * | 2022-04-11 | 2022-07-29 | 北京航空航天大学 | Device and method for measuring charge burning rate of solid-liquid rocket engine |
| CN114810429B (en) * | 2022-04-11 | 2024-01-19 | 北京航空航天大学 | Device and method for measuring explosive loading burning speed of solid-liquid rocket engine |
| CN116877300A (en) * | 2023-09-05 | 2023-10-13 | 南京理工大学 | Solid rocket engine propellant erosion function measuring device and measuring method |
| CN116877300B (en) * | 2023-09-05 | 2023-12-15 | 南京理工大学 | Solid rocket engine propellant erosion function measuring device and measuring method |
| CN117250006A (en) * | 2023-11-20 | 2023-12-19 | 中国空气动力研究与发展中心设备设计与测试技术研究所 | Rocket-based combined cycle model engine combustion chamber tester |
| CN117250006B (en) * | 2023-11-20 | 2024-01-23 | 中国空气动力研究与发展中心设备设计与测试技术研究所 | Rocket-based combined cycle model engine combustion chamber tester |
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