CN110749536B - Solid rocket engine thermal protection material ablation experimental device - Google Patents
Solid rocket engine thermal protection material ablation experimental device Download PDFInfo
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- CN110749536B CN110749536B CN201910982905.6A CN201910982905A CN110749536B CN 110749536 B CN110749536 B CN 110749536B CN 201910982905 A CN201910982905 A CN 201910982905A CN 110749536 B CN110749536 B CN 110749536B
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- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N17/00—Investigating resistance of materials to the weather, to corrosion, or to light
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Abstract
The invention discloses an ablation experimental device for a thermal protection material of a solid rocket engine, which comprises an end combustion charge unit, an ablation test piece unit, a test spray pipe unit and a main connector, wherein the ablation test piece unit comprises a first combustion chamber shell, a first blocking cover and an ablation test piece, the end combustion charge unit comprises a second combustion chamber shell, a second blocking cover and an end combustion column, the end combustion column is arranged in the second combustion chamber shell, the test spray pipe unit comprises a test spray pipe and a spray pipe connector, and the end combustion charge unit, the ablation test piece unit and the test spray pipe unit are mutually communicated through the main connector. The engine thermal protection material ablation test is to perform the ablation test on the material of the engine thermal protection structure on the premise of avoiding the whole engine test, select the proper thermal protection material according to the test result and reasonably design the thermal protection structure, and the test device has the advantages of simple structure, simplified process and lower cost.
Description
Technical Field
The invention relates to the technical field of thermal protection testing of solid rocket engines, in particular to an ablation experimental device for a thermal protection material of a solid rocket engine.
Background
The solid rocket engine is a propelling power device for rocket and missile flight, mainly comprises a combustion chamber, an igniter, propellant powder charge, a spray pipe and other components, and has the characteristics of simple structure, low cost, convenient and quick maintenance, strong quick response capability and the like. During the operation of the solid rocket engine, the heat exchange in the engine is carried out in a high-temperature, high-flow-rate and high-pressure gas environment, so that the heat protection of the shell, the combustion chamber, the nozzle and other parts of the solid rocket engine is particularly important.
Based on the requirements of modern national defense and scientific research, the application of a solid rocket engine in missile propulsion units and rocket weapons is more and more extensive, and in order to achieve the aims of enabling rocket projectiles to reach a farther range and obtain higher operational capability, reasonable energy distribution can be carried out by controlling conditions such as engine pulse time, charging and thrust, and the like, so that the purposes of improving the range and performance are achieved. With the widespread use of high energy propellants and the design of high specific impulse engines, the pressure load and temperature to which the engine casing, combustion chamber and nozzle are subjected will increase, thereby affecting the normal operation of the engine. Insufficient thickness or poor quality of the thermal protection structure can cause the engine to be at too high a temperature, resulting in damage to the engine structure, whereas too thick or too complex a design can result in the engine affecting flight performance due to too much negative mass.
In the prior art, the ablation research on the material of the engine thermal protection structure is generally to fix the whole engine on the ground and actually operate the engine for a preset time, and after the operation of the engine is finished, the required parts are taken down for the ablation research, so that the whole research process is complex, the cost is high, and the flexibility is poor.
Disclosure of Invention
The invention aims to provide an ablation experimental device for a thermal protection material of a solid rocket engine.
In order to achieve the purpose, the technical scheme adopted by the invention is as follows:
a solid rocket engine thermal protection material ablation experimental device comprises an end combustion charge unit, an ablation test piece unit, a test spray pipe unit and a main connector, wherein the ablation test piece unit comprises a first combustion chamber shell, a first blocking cover and an ablation test piece, a first combustion chamber is formed in the first combustion chamber shell, one end of the first combustion chamber shell is connected with the first blocking cover, the other end of the first combustion chamber shell is connected with one side of the main connector, a boss extending into the first combustion chamber is arranged on the first blocking cover, and the ablation test piece is fixedly arranged on the end face of the boss of the first blocking cover;
the end combustion explosive charging unit comprises a second combustion chamber shell, a second blocking cover and an end combustion explosive column, one end of the second combustion chamber shell is connected with the second blocking cover, the other end of the second combustion chamber shell is connected with the other side of the main connecting body, and the end combustion explosive column is filled in the second combustion chamber shell;
the test spray pipe unit comprises a test spray pipe and a spray pipe connecting body, one end of the spray pipe connecting body is connected with the upper end of the main connecting body, the test spray pipe is arranged on the end face of the other end of the spray pipe connecting body, and the end combustion charge unit, the ablation test piece unit and the test spray pipe unit are mutually communicated through the main connecting body.
Furthermore, the bottom end of the main connecting body is provided with an exhaust hole.
Furthermore, the main connecting body is a hollow cube, and three faces of the cube are provided with through holes respectively used for being connected with the first combustion chamber shell, the second combustion chamber shell and the spray pipe connecting body.
Further, first blanking cover includes lid and boss, the lid links to each other with first combustion chamber casing, and boss one end links to each other with the lid, and in the other end stretched into first combustion chamber, the terminal surface outside of the other end of boss stretched out to boss axial direction has a plurality of archs, form an installation space between a plurality of archs, the installation space is used for installing the ablation test piece.
Further, the ablation test piece unit further comprises a first combustion chamber liner, and the first combustion chamber liner is attached to the inner wall of the first combustion chamber shell.
Furthermore, the end combustion charge unit also comprises a second combustion chamber lining, and the second combustion chamber lining is positioned between the inner wall of the second combustion chamber shell and the end combustion charge column.
Further, the test nozzle is detachably mounted on the nozzle connection body.
Further, the first blocking cover, the first combustion chamber shell, the main connecting body, the nozzle connecting body, the second combustion chamber shell and the second blocking cover are made of 45 steel or 30CrMnSiA or 35 CrMnSiA.
Further, the material of the test nozzle is carbon/phenolic aldehyde.
Furthermore, the material of the first combustion chamber lining and/or the second combustion chamber lining adopts nitrile rubber or ethylene propylene diene monomer rubber or silicon rubber.
Compared with the prior art, the invention has the remarkable advantages that:
(1) the engine thermal protection material ablation test is to perform the ablation test on the material of the engine thermal protection structure on the premise of avoiding the whole engine test, select a proper thermal protection material according to the test result and reasonably design the thermal protection structure, and the test device has the advantages of simple structure, simplified process and lower cost;
(2) the experimental device can simulate the working environment of the spray pipe in the working process of the engine through the spray pipe part, can replace the test spray pipe according to the experimental requirements and measure the ablation condition of the spray pipe, and the end fuel column is adjustable, so that the ablation test piece is convenient to replace and has high controllability;
(3) the bottom of the connecting body part of the experimental device is provided with an exhaust hole, so that the danger caused by overhigh pressure in the cavity can be prevented;
(4) through this experimental apparatus, can simulate the operational environment and the condition of hot protective structure among the engine working process, compensatied hot protective structure and be difficult to the not enough of measuring in operational environment.
Drawings
FIG. 1 is a three-dimensional view of the solid rocket engine thermal protection material ablation experimental device.
FIG. 2 is a cross-sectional view of the solid rocket engine thermal protection material ablation experimental apparatus of the present invention.
FIG. 3 is a three-dimensional view of an ablation test piece unit of the solid rocket engine thermal protection material ablation experimental device.
FIG. 4 is a cross-sectional view of an ablation test piece unit of the solid rocket engine thermal protection material ablation experimental device.
FIG. 5 is a schematic diagram of an ablation test piece unit of the solid rocket engine thermal protection material ablation experimental device provided by the invention.
FIG. 6 is a three-dimensional view of an end-fire charging unit of the solid rocket engine thermal protection material ablation experimental device.
FIG. 7 is a cross-sectional view of an end-fire charging unit of the solid rocket engine thermal protection material ablation experimental device.
FIG. 8 is a three-dimensional view of a test nozzle unit of the solid rocket engine thermal protective material ablation experimental device.
FIG. 9 is a sectional view of a test nozzle unit of the solid rocket engine thermal protective material ablation experimental device.
FIG. 10 is a three-dimensional view of a main connecting body of the solid rocket engine thermal protection material ablation experimental device.
FIG. 11 is a sectional view of a main connecting body of the solid rocket engine thermal protection material ablation experimental device.
Detailed Description
The following describes the implementation of the present invention in detail with reference to specific embodiments.
As shown in fig. 1-11, an ablation experimental device for thermal protective materials of a solid rocket engine comprises an end-burning charge unit, an ablation test piece unit, a test nozzle unit and a main connector 5, wherein the ablation test piece unit comprises a first combustion chamber shell 4, a first blocking cover 1 and an ablation test piece 2, a first combustion chamber is formed in the first combustion chamber shell 4, one end of the first combustion chamber shell is connected with the first blocking cover 1, the other end of the first combustion chamber shell is connected with the left side of the main connector 5, a boss extending into the first combustion chamber is arranged on the first blocking cover 1, and the ablation test piece 2 is fixedly arranged on the boss of the first blocking cover; the end combustion explosive charging unit comprises a second combustion chamber shell 10, a second blocking cover 11 and an end combustion explosive column 8, one end of the second combustion chamber shell 10 is connected with the second blocking cover 11, the other end of the second combustion chamber shell 10 is connected with the right side of the main connecting body 5, and the end combustion explosive column 8 is arranged in the second combustion chamber shell 10; the test spray pipe unit comprises a test spray pipe 6 and a spray pipe connecting body 7, the bottom end of the spray pipe connecting body 7 is connected with the upper side of the main connecting body 5, the test spray pipe 6 is arranged at the upper end of the spray pipe connecting body 7, and the end combustion charge unit, the ablation test piece unit and the test spray pipe unit are communicated with each other through the main connecting body 5.
As shown in fig. 3-5, the first blocking cover 1 comprises a cover body and a boss, wherein the inner side of the cover body is provided with an internal thread for connecting with an external thread on the first combustion chamber shell 4; one end of the boss is connected with the cover body, the other end of the boss extends into the first combustion chamber, with reference to fig. 5, a plurality of protrusions 1-1 extend out of the outer side of the end face of the other end of the boss towards the axial direction of the boss, an installation space is formed among the protrusions 1-1 and used for installing the ablation test piece 2, the ablation test piece 2 is installed in the installation space during installation, the ablation test piece 2 is in close contact with the installation space, the installation mode is particularly convenient when the ablation test piece 2 is taken out, and the ablation test piece 2 only needs to be clamped at a gap between the protrusions 1-1 and taken out. The second closing cap 11 is provided with an internal thread inside for connection with one end of the second combustion chamber housing 10.
With reference to fig. 6-7, the ablation test piece unit further includes a first combustion chamber liner 3, and the first combustion chamber liner 3 is attached to the inner wall of the first combustion chamber housing 4, preferably by adhesion. The end combustion charge unit also comprises a second combustion chamber inner liner 9, and the second combustion chamber inner liner 9 is attached to the inner wall of a second combustion chamber shell 10, preferably in a bonding mode. The first combustion chamber lining 3 and the second combustion chamber lining 9 are made of rubber materials, such as nitrile rubber, ethylene propylene diene monomer rubber, silicon rubber and the like. The rubber material can protect the combustion chamber housing from overheating.
As shown in fig. 8-9, the nozzle connecting body 7 is a cylindrical structure with a hollow cavity, the lower end of the cylindrical structure is open, the upper end of the cylindrical structure is provided with a through hole, an internal thread is arranged inside the through hole, the test nozzle 6 is arranged in the through hole of the nozzle connecting body 7, and the front end of the contraction section of the test nozzle 6 is provided with an external thread for connecting with the internal thread of the through hole. The test nozzle 6 is made of a heat-resistant and corrosion-resistant material, such as carbon/phenolic aldehyde, so as to facilitate experimental tests.
The first combustion chamber shell 4 and the second combustion chamber shell 10 are both cylindrical cavities, and the first combustion chamber shell 4, the second combustion chamber shell 10, the spray pipe connecting body 7 and the main connecting body 5 are connected through threads. The first blocking cover 1, the ablation test piece 2, the first combustion chamber lining 3, the first combustion chamber shell 4, the test spray pipe 6, the spray pipe connecting body 7, the end fuel column 8, the second combustion chamber lining 9, the second combustion chamber shell 10 and the second blocking cover 11 are all of a single-shaft rotating body structure, and processing is convenient.
The first blocking cover 1, the first combustion chamber shell 4, the main connecting body 5, the nozzle connecting body 7, the second combustion chamber shell 10 and the second blocking cover 11 are all made of high-strength steel materials, such as 45 steel, 30CrMnSiA, 35CrMnSiA and the like, the high-strength steel is high in strength and good in toughness, and the structural integrity of the experimental device can be guaranteed under the conditions of long-time high pressure and high temperature.
As shown in fig. 10-11, the main connecting body 5 is a hollow cube, three sides of the cube are provided with through holes respectively for connecting with the first combustion chamber housing 4, the second combustion chamber housing 10 and the nozzle connecting body 7, the three through holes are connected with the hollow cavity of the cube, and the bottom of the main connecting body 5 is provided with an exhaust hole 5-1.
The working principle of the invention is as follows: simulating a heat transfer environment under the working condition of a solid rocket engine, researching the ablation condition of a thermal protection material, putting a corresponding igniter ignition end fuel column 8 into a groove 5-2 formed in one side of a connector 5, igniting the end fuel column 8, transmitting fuel gas to an ablation test piece 2 and a test spray pipe 6 through the connector 5 and a spray pipe part connector 7 respectively, waiting for the cooling of an experimental device after the end fuel column 8 works, detaching the ablation test piece 2 and the test spray pipe 6, recording the ablation condition of the test piece through corresponding observation and test means, and providing a reference value for the thermal protection of the solid rocket engine.
The foregoing illustrates and describes the principles, general features, and advantages of the present invention. It will be understood by those skilled in the art that the present invention is not limited to the embodiments described above, which are described in the specification and illustrated only to illustrate the principle of the present invention, but that various changes and modifications may be made therein without departing from the spirit and scope of the present invention, which fall within the scope of the invention as claimed. The scope of the invention is defined by the appended claims and equivalents thereof.
Claims (7)
1. The ablation experimental device for the thermal protection material of the solid rocket engine is characterized by comprising an end combustion charge unit, an ablation test piece unit, a test spray pipe unit and a main connector (5), wherein the ablation test piece unit comprises a first combustion chamber shell (4), a first blocking cover (1) and an ablation test piece (2), a first combustion chamber is formed in the first combustion chamber shell (4), one end of the first combustion chamber shell (4) is connected with the first blocking cover (1), the other end of the first combustion chamber shell (4) is connected with one side of the main connector (5), a boss extending into the first combustion chamber is arranged on the first blocking cover (1), and the ablation test piece (2) is fixedly arranged on the end face of the boss of the first blocking cover (1);
the end combustion explosive charging unit comprises a second combustion chamber shell (10), a second blocking cover (11) and an end combustion explosive column (8), one end of the second combustion chamber shell (10) is connected with the second blocking cover (11), the other end of the second combustion chamber shell (10) is connected with the other side of the main connecting body (5), and the end combustion explosive column (8) is filled in the second combustion chamber shell (10);
the test spray pipe unit comprises a test spray pipe (6) and a spray pipe connecting body (7), one end of the spray pipe connecting body (7) is connected with the upper end of the main connecting body (5), the test spray pipe (6) is arranged on the end face of the other end of the spray pipe connecting body (7), and the end combustion charge unit, the ablation test piece unit and the test spray pipe unit are mutually communicated through the main connecting body (5);
the bottom end of the main connecting body (5) is provided with an exhaust hole (5-1);
the main connecting body (5) is a hollow cube, and three sides of the cube are provided with through holes respectively used for being connected with the first combustion chamber shell (4), the second combustion chamber shell (10) and the spray pipe connecting body (7);
first blanking cover (1) includes lid and boss, the lid links to each other with first combustion chamber casing (4), and boss one end links to each other with the lid, and in the other end stretched into first combustion chamber, the terminal surface outside of the other end of boss stretched out to boss axial direction has a plurality of archs (1-1), form an installation space between a plurality of archs (1-1), the installation space is used for installing ablation test piece (2).
2. The solid rocket engine thermal protection material ablation experimental device according to claim 1, wherein the ablation test piece unit further comprises a first combustion chamber lining (3), and the first combustion chamber lining (3) is attached to the inner wall of the first combustion chamber shell (4).
3. The ablation experimental facility for the thermal protective material of the solid rocket engine as recited in claim 2, wherein the end-firing charge unit further comprises a second combustion chamber liner (9), and the second combustion chamber liner (9) is located between the inner wall of the second combustion chamber housing (10) and the end-firing charge column (8).
4. The solid rocket engine thermal protection material ablation experimental device according to claim 1, wherein the test nozzle (6) is detachably mounted on the nozzle connecting body (7).
5. The ablation experimental device for the thermal protective material of the solid rocket engine according to claim 1, wherein the first blocking cover (1), the first combustion chamber shell (4), the main connecting body (5), the nozzle connecting body (7), the second combustion chamber shell (10) and the second blocking cover (11) are made of 45 steel or 30CrMnSiA or 35 CrMnSiA.
6. The ablation experimental device for the thermal protective material of the solid rocket engine according to claim 1, wherein the material of the test nozzle (6) is carbon/phenolic aldehyde.
7. The ablation experimental device for the thermal protective material of the solid rocket engine according to claim 1, wherein the material of the first combustion chamber liner (3) and/or the second combustion chamber liner (9) is nitrile rubber, ethylene propylene diene monomer rubber or silicone rubber.
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| CN111927652B (en) * | 2020-07-29 | 2022-06-10 | 南京理工大学 | Double-pulse solid rocket engine interlayer ablation carbonization controllable experimental device |
| 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 |
| CN112485012A (en) * | 2020-11-13 | 2021-03-12 | 东北大学 | Solid rocket engine experiment table and stress testing method |
| CN113790335B (en) * | 2021-09-14 | 2023-04-07 | 湖北三江航天红林探控有限公司 | Space composite pipeline for engine gas output |
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