CN110608892A - High accuracy burning rate testing arrangement suitable for high pressure - Google Patents
High accuracy burning rate testing arrangement suitable for high pressure Download PDFInfo
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- CN110608892A CN110608892A CN201910969119.2A CN201910969119A CN110608892A CN 110608892 A CN110608892 A CN 110608892A CN 201910969119 A CN201910969119 A CN 201910969119A CN 110608892 A CN110608892 A CN 110608892A
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- combustion chamber
- shell
- burning rate
- spray pipe
- high pressure
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M15/00—Testing of engines
- G01M15/02—Details or accessories of testing apparatus
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M15/00—Testing of engines
- G01M15/04—Testing internal-combustion engines
- G01M15/10—Testing internal-combustion engines by monitoring exhaust gases or combustion flame
Abstract
The invention relates to the field of solid rocket engine casings, in particular to a high-precision burning rate testing device suitable for high pressure, which comprises a top cover, a combustion chamber and a spray pipe, wherein the top cover and the combustion chamber casing are respectively connected with each other through a thread structure, the combustion chamber consists of a combustion chamber casing, a combustion limiting layer and a grain, the combustion limiting layer is arranged on the front end face of the grain, the spray pipe consists of a spray pipe casing, a heat insulation sleeve and a throat liner, the spray pipe casing and the heat insulation polyphenyl ether sleeve are sealed through supporting silicon rubber dynamic and static sealing putty, and the throat liner and the heat insulation sleeve are bonded through HX98-1 adhesive. The material of the nozzle shell, the material of the top cover and the material of the combustion chamber shell are all made of 30 CrMnSiA. The connecting structure adopts a threaded connecting structure, is convenient to disassemble and assemble, can be repeatedly used for a plurality of times, and improves the efficiency; the ignition mode of the invention adopts a medicine bag ignition mode, the installation is convenient, and the cost is saved.
Description
Technical Field
The invention relates to the technical field of solid rocket engine testing, in particular to a high-precision burning rate testing device suitable for high pressure.
Background
The demonstration of a certain project verifies that the specific impulse performance of a primary engine (code: Gf 0-119) is improved by improving the working pressure, and the average pressure reaches more than 12 MPa. In the process of debugging the formula of the propellant in the early period, the burning rate result (8.25 mm/s, 12 MPa) tested by a phi 108 burning rate test engine in the prior art is lower than the real burning rate of the propellant; meanwhile, the combustion speed test has larger deviation due to the difference of the design state, the forming condition and the test run state of the engine tested by different phi 108 combustion speeds. According to the test result (9.43 mm/s, 13.26 MPa) of the Gf0-119 engine, the burning rate correlation coefficient of the burning rate test result (8.6 mm/s, 12 MPa) reaches 1.065 according to the burning rate conversion formula, and is obviously improved compared with other BHT four-component propellant engine models. Due to the fact that the propellant burning rate test result has large deviation, the predicted inner trajectory of the Gf0-119 engine has large difference from the measured result.
Considering that the dynamic burning rate of the propellant is directly related to the internal ballistic design, prediction precision and working reliability of the engine, the dynamic burning rate of the propellant, particularly the burning rate characteristic under the high-pressure condition, needs to be accurately tested and obtained. Therefore, the invention provides a high-precision burning rate testing device which is not only suitable for low pressure intensity but also suitable for high pressure intensity.
The invention is improved on the basis of a phi 165 burning speed test engine specified in GJB 96A. The structure of the phi 165 burning rate test engine specified in GJB 96A is shown in figure 1.
The engine structure and material comparison of the invention and the GJB 96A phi 165 standard test is shown in Table 1, and the working parameter and size parameter comparison is shown in Table 2.
TABLE 1 Engine Structure and materials
TABLE 2 Engine operating parameters and dimensional parameters
The GJB 96A phi 165 standard test engine throat insert material is a graphite material, and the graphite material has poor ablation resistance, especially under the condition of high pressure. The propellant has the condition of pressure index mutation under the condition of high pressure, and if the graphite throat liner material is adopted under the condition of high pressure, the pressure maintaining capability is lower, the pressure index mutation and the burning rate test result are inaccurate.
The phi 165 standard test engine in GJB 96A has a thickness of 14 or 18 mm. When the burning rate of the propellant is tested to be high, the working time is short, the influence degree of the data processing deviation of the burning time on the burning rate analysis is increased, and the test result is inaccurate.
The connecting structure of the combustion chamber and the front top cover of the phi 165 standard test engine in the GJB 96A and the connecting structure of the combustion chamber and the spray pipe adopt a flange bolt structure. Because the burning rate test number of times is more, flange bolted connection dismouting is wasted time and energy.
The phi 165 standard test engine ignition device in the GJB 96A adopts an ignition mode of an electric ignition head and an ignition head, so that the cost is high, and the installation steps are complicated.
Disclosure of Invention
The technical problems solved by the invention are as follows:
the invention provides a high-precision burning rate testing device suitable for high pressure, and aims to solve the problems that the deviation between a predicted internal trajectory and an actually-measured internal trajectory is large and the burning rate test is inaccurate under the high pressure condition of a solid rocket engine.
The technical scheme adopted by the invention is as follows:
the utility model provides a high accuracy burning rate testing arrangement suitable for high pressure, includes top cap, combustion chamber, spray tube, the top cap adopts threaded structure to be connected respectively with combustion chamber casing, spray tube and combustion chamber casing, the combustion chamber comprises combustion chamber casing, flame retardant layer and powder column, the flame retardant layer sets up terminal surface before the powder column, the spray tube comprises spray tube casing, adiabatic cover and throat lining, the spray tube casing is used with adiabatic cover polyphenyl ether to prop silicon rubber and move, the static seal putty is sealed, bonds with HX98-1 adhesive between throat lining and the adiabatic cover.
Further, the thread structure adopts a thread specification of M160 x 3.
3. The high-precision burning rate testing device suitable for high pressure according to claim 1, wherein sealing structures are further arranged at the connection positions of the threaded structures of the top cover and the combustion chamber shell, and the threaded structures of the nozzle and the combustion chamber shell, the sealing structures are sealing rings made of M204 materials with high Shore hardness, the sealing rings are designed to have a compression rate of 14.36 ~ 22.14.14% and a design area ratio of 1.110 ~ 1.346.
Further, the top cover material is made of a 30CrMnSiA material with high strength.
Further, the thickness of the combustion limiting layer is 2mm, the length of the grain is 300mm, the thickness of the grain is 28mm, the mass of the grain is 5.0kg, the shell material of the combustion chamber is 30CrMnSiA, and the wall thickness of the shell is designed to be 7 mm.
Further, the material of the nozzle shell is 30CrMnSiA, the material of the heat insulating sleeve is a high-silica phenolic resin molded product, and the material of the throat lining is a tungsten copper-infiltrated material with an ablation rate of approximately 0 mm/s.
The invention has the beneficial effects that:
the ballistic pressure curve is straight in a short time of the burning rate test, the pressure index has no problem of mutation, and the burning rate test is more accurate; the thickness of the drug type is increased, the burning time is prolonged, the influence degree of the data processing deviation of the burning time on the burning rate analysis is greatly reduced, and the burning rate test precision is increased; the connecting structure adopts a threaded connecting structure, is convenient to disassemble and assemble, can be repeatedly used for a plurality of times, and improves the efficiency; the ignition mode of the invention adopts a medicine bag ignition mode, the installation is convenient, and the cost is saved.
Drawings
FIG. 1: a phi 165 standard engine structure diagram in a GJB 96A;
FIG. 2: a structure schematic diagram of the burning speed testing device;
FIG. 3: a combustion chamber structure diagram;
FIG. 4: the structural diagram of the spray pipe is simplified;
FIG. 5: pressure-time curve (7.5 mm/s, 10 MPa).
Wherein: 1. the burner comprises a pressing cap, a sealing plug, a top cover, a heat insulation plate, a sealing ring, a bolt, a nut, a gasket, a powder column, a combustion chamber shell, an ignition device, a convergence section, a nozzle base, a nozzle throat lining, a nozzle cover pressing ring, a nozzle blocking cover, a nozzle, a combustion chamber, a nozzle, a combustion limiting layer, a nozzle shell and a heat insulation sleeve, wherein the pressing cap is 2, the sealing plug is 3, the top cover is 4, the heat insulation plate is 5, the bolt is 6, the nut is 7, the gasket is 8, the powder column is 9, the combustion chamber shell is 10.
Detailed Description
The following detailed description is made with reference to the accompanying drawings and specific embodiments:
as shown in figure 2, the combustion speed testing device disclosed by the invention consists of a top cover, a combustion chamber and a spray pipe.
a) Considering that the number of times of testing the burning speed engine is large, the top cover and the combustion chamber shell as well as the spray pipe and the combustion chamber shell adopt a threaded connection structure which is convenient to assemble and disassemble, and the thread specification is M160 multiplied by 3;
b) sealing structures of the top cover and the combustion chamber shell as well as the spray pipe and the combustion chamber shell are sealing rings made of M204 materials with high Shore hardness, the design compression rate of the sealing rings is 14.36 ~ 22.14.14%, and the design area ratio is 1.110 ~ 1.346;
c) the top cover material is made of 30CrMnSiA material with higher strength.
As shown in figure 3, the combustion chamber consists of a combustion-limiting layer, a grain and a combustion chamber shell.
a) The function of the combustion limiting layer is mainly to limit the combustion of the front end face of the explosive column, and the front end face of the explosive column adopts the combustion limiting layer with the thickness of 2 mm;
b) according to the calculation result of the inner trajectory, the length of the grain is 300mm, the thickness of the meat is 28mm, and the design mass of the grain is 5.0 kg;
c) the combustor shell is made of a 30CrMnSiA material with high strength, and the wall thickness of the shell is designed to be 7 mm.
As shown in figure 4, the nozzle consists of a nozzle shell, a heat insulation sleeve and a throat insert.
a) The material of the spray pipe shell is a 30CrMnSiA material with higher strength;
b) the heat insulating sleeve material is a high-silica phenolic resin molded product;
c) the throat lining material is a tungsten copper infiltration material with ablation rate of approximately 0 mm/s;
d) the jet pipe shell and the heat insulating sleeve are sealed by polyphenyl ether-silicon rubber dynamic and static sealing putty, and the throat lining and the heat insulating sleeve are bonded by HX98-1 adhesive.
Claims (6)
1. The utility model provides a high accuracy burning rate testing arrangement suitable for high pressure which characterized in that: the improved combustion chamber comprises a top cover, a combustion chamber and a spray pipe, wherein the top cover is connected with a combustion chamber shell, the spray pipe is connected with the combustion chamber shell through a threaded structure, the combustion chamber is composed of the combustion chamber shell, a combustion limiting layer and a grain, the combustion limiting layer is arranged on the front end face of the grain, the spray pipe is composed of a spray pipe shell, an insulating sleeve and a throat liner, the spray pipe shell and the insulating sleeve are sealed through polyphenyl ether-supported silicon rubber dynamic and static sealing putty, and the throat liner and the insulating sleeve are bonded through HX98-1 adhesive.
2. The high-precision burning rate testing device suitable for high pressure according to claim 1, wherein: the thread structure adopts a thread specification of M160 multiplied by 3.
3. The high-precision burning rate testing device suitable for high pressure according to claim 1, wherein sealing structures are further arranged at the connection positions of the threaded structures of the top cover and the combustion chamber shell, and the threaded structures of the nozzle and the combustion chamber shell, the sealing structures are sealing rings made of M204 materials with high Shore hardness, the sealing rings are designed to have a compression rate of 14.36 ~ 22.14.14% and a design area ratio of 1.110 ~ 1.346.
4. The high-precision burning rate testing device suitable for high pressure according to claim 1, wherein: the top cover material is made of a 30CrMnSiA material with high strength.
5. The high-precision burning rate testing device suitable for high pressure according to claim 1, wherein: the thickness of the combustion limiting layer is 2mm, the length of the grain is 300mm, the thickness of the grain is 28mm, the mass of the grain is 5.0kg, the shell material of the combustion chamber is 30CrMnSiA, and the wall thickness of the shell is designed to be 7 mm.
6. The high-precision burning rate testing device suitable for high pressure according to claim 1, wherein: the jet pipe shell material is a 30CrMnSiA material, the heat insulation sleeve material is a high-silica phenolic resin molded product, and the throat lining material is a tungsten copper infiltration material with an ablation rate of approximately 0 mm/s.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910969119.2A CN110608892A (en) | 2019-10-12 | 2019-10-12 | High accuracy burning rate testing arrangement suitable for high pressure |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910969119.2A CN110608892A (en) | 2019-10-12 | 2019-10-12 | High accuracy burning rate testing arrangement suitable for high pressure |
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| CN201910969119.2A Pending CN110608892A (en) | 2019-10-12 | 2019-10-12 | High accuracy burning rate testing arrangement suitable for high pressure |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111237089A (en) * | 2020-01-19 | 2020-06-05 | 北京中科宇航探索技术有限公司 | Solid engine with adjustable thrust and center adjusting cone thereof |
| CN112555058A (en) * | 2021-02-19 | 2021-03-26 | 星河动力(北京)空间科技有限公司 | Test device for nozzle throat lining of rocket engine |
| CN112555057A (en) * | 2021-02-19 | 2021-03-26 | 星河动力(北京)空间科技有限公司 | Test device for nozzle throat lining of rocket engine |
-
2019
- 2019-10-12 CN CN201910969119.2A patent/CN110608892A/en active Pending
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111237089A (en) * | 2020-01-19 | 2020-06-05 | 北京中科宇航探索技术有限公司 | Solid engine with adjustable thrust and center adjusting cone thereof |
| CN112555058A (en) * | 2021-02-19 | 2021-03-26 | 星河动力(北京)空间科技有限公司 | Test device for nozzle throat lining of rocket engine |
| CN112555057A (en) * | 2021-02-19 | 2021-03-26 | 星河动力(北京)空间科技有限公司 | Test device for nozzle throat lining of rocket engine |
| CN112555057B (en) * | 2021-02-19 | 2021-05-07 | 星河动力(北京)空间科技有限公司 | Test device for nozzle throat lining of rocket engine |
| CN112555058B (en) * | 2021-02-19 | 2021-05-07 | 星河动力(北京)空间科技有限公司 | Test device for nozzle throat lining of rocket engine |
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