CN112483281B - High-burning-speed solid propellant burning speed testing device - Google Patents
High-burning-speed solid propellant burning speed testing device Download PDFInfo
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- CN112483281B CN112483281B CN202011181249.9A CN202011181249A CN112483281B CN 112483281 B CN112483281 B CN 112483281B CN 202011181249 A CN202011181249 A CN 202011181249A CN 112483281 B CN112483281 B CN 112483281B
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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
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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
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/24—Charging rocket engines with solid propellants; Methods or apparatus specially adapted for working solid propellant charges
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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
- F02K9/08—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
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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
- F02K9/95—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by starting or ignition means or arrangements
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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
- F02K9/97—Rocket nozzles
Abstract
A high-burning-speed solid propellant burning speed testing device belongs to the technical field of solid propellant testing, and comprises a top cover (1), a combustion chamber shell (3), propellant grains (4), a spray pipe base (10), a throat liner (12), an ignition device (9) and a spray pipe plugging cover (11); the combustion chamber shell (3) is of a circular tube type structure; the propellant grain (4) is a solid end-burning grain and is arranged in the combustion chamber shell (3), and the inner cavity of the combustion chamber shell (3) is a combustion chamber (5); the combustion chamber shell (3) is provided with a pressure measuring hole (6), and the pressure measuring hole (6) is communicated with the combustion chamber (5); one end of the combustion chamber shell (3) is connected with the top cover (1), and the other end of the combustion chamber shell is connected with the spray pipe base (10); the throat liner (12) is arranged in the nozzle base (10), and the throat liner (12) and the nozzle base (10) form a nozzle; and the ignition device (9) and the spray pipe blocking cover (11) are sequentially arranged into the throat part of the spray pipe.
Description
Technical Field
The invention relates to a burning rate testing device for a high-burning-rate solid propellant, and belongs to the technical field of burning rate testing of solid propellants.
Background
The high-burning-speed solid propellant is taken as an important branch of the solid propellant, and along with the one-by-one solution of the use bottleneck problem, the application range of the solid propellant is wider and wider, and the application requirement is higher and higher. The burning rate is used as a main parameter for characterizing the performance of the high-burning-rate solid propellant, and the test accuracy of the solid propellant has a critical effect on guiding the design of a propellant formula or the detection of the charging performance of related products.
In the process of developing and producing the propellant, the burning rate test method of the solid propellant widely applied at present mainly comprises two main types: one type is a burning rate instrument method represented by a target line method and an underwater acoustic emission method; the other is an engine method represented by various small-sized standard test engine test runs. The burning rate instrument method is influenced by pressure fluctuation in the propellant burning process, the testing precision is not high, particularly the testing precision of the high burning rate solid propellant is poorer, and the national military standard GJB770B-2005 clearly stipulates that the application ranges of the two burning rate testing methods of the target line method and the underwater acoustic emission method are the solid propellant with the burning rate not more than 50 mm/s. The instrument method for determining the fuel ratio and the burning rate of the propellant by the engine method is closer to the reality, the test precision is higher, the burning rate of the propellant is tested by utilizing a small-sized standard test engine, the real working condition of the propellant burning in a full-size engine can be accurately simulated, the burning characteristic of the propellant can be accurately reflected, and the instrument method is an effective means for developing the propellant formula and monitoring the production process. However, the existing various standard test engines adopt a tubular side combustion grain charging structure, and when a propellant is charged, core inserting operation is required, so that the eccentricity problem is easy to occur, and the test precision is influenced; the propellant grains need to be demoulded and reshaped after being solidified and formed, the process is complicated, and particularly for high-burning-speed solid propellants, potential safety hazards exist; meanwhile, the combustion surface of the tubular charge structure is large, when the low-working-pressure combustion speed test is carried out on the high-combustion-speed solid propellant, the throat diameter of a spray pipe is required to be selected and used and often exceeds the design upper limit of various test engines under the existing structural condition, and abnormal combustion phenomena such as difficult ignition, flameout, low-frequency oscillation, surge and the like are easy to occur. Therefore, in order to effectively represent the burning rate level of the high-burning-rate solid propellant in a wide pressure range, especially the burning rate level under lower working pressure, a safe and reliable burning rate testing device for the high-burning-rate solid propellant under lower working pressure is urgently needed.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: overcome prior art's not enough, a high speed of combustion solid propellant burning rate testing arrangement is provided, the device includes top cap, combustion chamber casing, propellant grain, spray tube base, throat lining, ignition, the spray tube blanking cover, the combustion chamber casing is equipped with the pressure measurement hole, with the combustion chamber intercommunication, the propellant grain is for holding to fire solid grain, the combined type structure of spray tube base and throat lining constitutes the spray tube, be clearance fit between combustion chamber casing and top cap, the spray tube, adopt bolt and nut connection to fix as an organic whole, and seal with O type sealing washer, ignition sets up in the inboard middle zone of throat, be connected with ignition power supply after the spray tube blanking cover through the ignition line. The device avoids the problems that the existing burning rate instrument method is not high in testing precision, the existing small-sized standard test engine grain charging process is complex and complex, the safety is poor, and the burning rate of the high-burning-rate solid propellant can not be tested under lower working pressure, and has the advantages of simple, convenient and safe charging process, economy and practicability, high testing precision and capability of meeting the burning rate testing requirement of the high-burning-rate solid propellant under lower working pressure.
The purpose of the invention is realized by the following technical scheme:
a high-burning-speed solid propellant burning speed testing device comprises a top cover, a combustion chamber shell, a propellant grain, a spray pipe base, a throat liner, an ignition device and a spray pipe plugging cover;
the combustion chamber shell is of a circular tube type structure; the propellant grain is a solid grain with a burning end and is arranged in the combustion chamber shell, and the inner cavity of the combustion chamber shell is a combustion chamber; the combustion chamber shell is provided with a pressure measuring hole, and the pressure measuring hole is communicated with the combustion chamber;
one end of the combustion chamber shell is connected with the top cover, and the other end of the combustion chamber shell is connected with the spray pipe base; the throat insert is arranged in the nozzle base, and the throat insert and the nozzle base form a nozzle; the ignition device and the spray pipe plugging cover are sequentially installed in the throat part of the spray pipe.
Above-mentioned high burning rate solid propellant burning rate testing arrangement, preferred, the both ends of combustion chamber casing all adopt flange formula connection structure to be connected with top cap and spray tube base respectively.
Preferably, two ends of the combustion chamber shell are respectively connected with the top cover and the nozzle base in a sealing manner; wherein, clearance fit is formed between the combustion chamber shell and the top cover as well as between the combustion chamber shell and the nozzle base, and then sealing is carried out by utilizing the sealing ring.
Preferably, the two ends of the combustion chamber shell are connected with the top cover and the nozzle base through bolts and nuts.
Preferably, the ignition device is located in the middle area inside the throat part.
Preferably, the ignition device is connected with an external ignition power supply through an ignition wire after passing through the nozzle blocking cover.
Above-mentioned high burning rate solid propellant burning rate testing arrangement, it is preferred, in first face and the second face of back on the protecgulum, first face is circular plane, is equipped with circular boss on the second face, the boss with the terminal surface adaptation of combustion chamber casing.
Preferably, the combustion speed testing device for the high-combustion-speed solid propellant has the advantage that the fit clearance distance between the combustion chamber shell and the top cover is 0-0.02 mm.
Preferably, the propellant grains are directly formed after being subjected to vacuum pouring and curing of propellant grains and are directly attached to the inner surface of the combustion chamber shell and the circular boss on the second surface of the top cover after roughening treatment.
Preferably, the number of the pressure measuring holes is 1 or 2.
Compared with the prior art, the invention has the following beneficial effects:
the invention provides a high-burning-rate solid propellant burning rate testing device, which is based on the existing standard test engine, according to the design principle of a rocket engine, by changing the propellant grain charging structure design and the local design of a combustion chamber shell and a top cover matched with the propellant grain charging structure design, under the premise of not changing the main appearance structure and the matched components of the existing standard test engine, the lower limit of the working pressure of the high-burning-rate solid propellant tested by the test engine is effectively widened, the testing requirements of the high-burning-rate solid propellant on low-working-pressure burning rate and combustion characteristics are met, the propellant burning rate is only determined by the thickness of the propellant grain and the pressure of the combustion chamber, and is not influenced by the eccentricity of the propellant grain and the size change of the inner diameter of the combustion chamber shell, and the device has the characteristics of less interference factors, high testing precision, strong universality, high reuse rate of the weight of the combustion chamber shell and the like; meanwhile, the propellant grain is directly formed after being poured and cured in vacuum by adopting the propellant grain slurry, so that the manufacturing process is simplified, the complex operations of inserting cores, demoulding, shaping and the like during the manufacturing of the existing small-sized standard test engine propellant grain are avoided, and the operation is simple, convenient, efficient and safe.
Drawings
Fig. 1 is a schematic view of a combustion rate testing device for a high-combustion-rate solid propellant according to an embodiment of the present invention.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, embodiments of the present invention will be described in detail with reference to the accompanying drawings.
A high-burning-speed solid propellant burning speed testing device comprises a top cover 1, a combustion chamber shell 3, a propellant grain 4, a nozzle base 10, a throat liner 12, an ignition device 9 and a nozzle blocking cover 11;
the combustion chamber shell 3 is of a circular tube type structure; the propellant grain 4 is a solid end-burning grain and is arranged in the combustion chamber shell 3, and the inner cavity of the combustion chamber shell 3 is a combustion chamber 5; the combustion chamber shell 3 is provided with a pressure measuring hole 6, and the pressure measuring hole 6 is communicated with the combustion chamber 5;
one end of the combustion chamber shell 3 is connected with the top cover 1, and the other end of the combustion chamber shell is connected with the spray pipe base 10; the throat insert 12 is arranged in the nozzle base 10, and the throat insert 12 and the nozzle base 10 form a nozzle; the ignition device 9 and the nozzle blocking cover 11 are sequentially arranged in the throat part of the nozzle.
As a preferred scheme of the present invention, both ends of the combustion chamber housing 3 are connected to the top cover 1 and the nozzle base 10 respectively by using a flange type connecting structure.
As a preferred scheme of the present invention, two ends of the combustion chamber housing 3 are respectively connected with the top cover 1 and the nozzle base 10 in a sealing manner; wherein, the combustion chamber shell 3 is in clearance fit with the top cover 1 and the nozzle base 10, and then is sealed by the seal ring 2; and two ends of the combustion chamber shell 3 are connected with the top cover 1 and the spray pipe base 10 by bolts and nuts.
As a preferable mode of the present invention, the ignition device 9 is located in the intermediate region inside the throat portion; the ignition device 9 is connected with an external ignition power supply through an ignition wire 13 after passing through the spray pipe blocking cover 11.
As a preferable scheme of the present invention, in the first surface 101 and the second surface 102 of the top cover 1, which are opposite to each other, the first surface 101 is a circular plane, and the second surface 102 is provided with a circular boss, and the boss is adapted to the end surface of the combustion chamber housing 3.
In a preferred embodiment of the present invention, a fitting clearance distance between the combustion chamber housing 3 and the top cover 1 is 0 to 0.02 mm.
In a preferred embodiment of the present invention, the propellant grains 4 are directly formed by vacuum casting and curing propellant grains, and directly attached to the inner surface of the combustion chamber housing 3 and the circular bosses of the second surface 102 of the top cover 1 after being roughened.
In a preferred embodiment of the present invention, the number of the pressure measuring holes 6 is 1 or 2.
Example (b):
as shown in fig. 1, an embodiment of the present invention provides a high-burning-rate solid propellant burning rate testing device, which includes a top cover 1, a combustion chamber shell 3, a propellant grain 4, a nozzle base 10, a throat liner 12, an ignition device 9, and a nozzle blocking cover 11, wherein the combustion chamber shell 3 is a circular tube structure, adopts a flange type connection structure, and is provided with a pressure measuring hole 6, and is communicated with a combustion chamber 5, the propellant grain 4 is an end-burning solid grain, the nozzle base 10 and the throat liner 12 form a nozzle in a composite structure, the combustion chamber shell 3, the top cover 1 and the nozzle are in clearance fit, are connected and fixed into a whole by bolts 8 and nuts 7, and are sealed by an O-shaped sealing ring 2, the ignition device 9 is arranged in a middle area inside the throat, and is connected with an ignition power supply through an ignition wire 13 via the nozzle blocking cover 11.
According to the high-burning-rate solid propellant burning rate testing device provided by the embodiment of the invention, before 4 propellant columns are filled with 4 propellant, the inner surface of a combustion chamber shell 3 is roughened; then the top cover 1 and the combustion chamber shell 3 are connected and fixed into a whole by bolts 8 and nuts 7 and are sealed by O-shaped sealing rings 2 to form a propellant grain forming die; then pouring the mixed propellant slurry into a propellant grain forming die through the opening end of the die cavity in a quantitative pouring mode by adopting a vacuum pouring method; after pouring, removing the boiling layer slurry on the inner surface of the combustion chamber shell 3, putting the boiling layer slurry into a drying oven for curing, keeping the surface of the slurry horizontal in the curing process, forming a propellant grain 4 after curing, and measuring and calculating the height of the grain, namely the thickness of the grain. The ignition device 9 is fixed in the middle area of the inner side of the throat part of the spray pipe assembled by a spray pipe base 10 and a throat liner 12 and is led out through a spray pipe blanking cover 11 through an ignition wire 13; then connecting and fixing the spray pipe and the other end face of the combustion chamber shell 3 provided with the propellant grain 4 into a whole by adopting a bolt 8 and a nut 7, and sealing by using an O-shaped sealing ring 2 to obtain the high-burning-rate solid propellant burning rate testing device; and finally, testing the burning rate of the propellant according to the existing standard test engine burning rate test method and steps.
In an optional embodiment, the top cover comprises a first surface 101 and a second surface 102 which are opposite to each other, the first surface 101 is a circular flat plate structure and is used for being attached to a horizontal plane of a curing oven so as to ensure the levelness of the propellant grain 4, and the second surface 102 is provided with a circular boss which is matched with the end surface of the combustion chamber shell 3.
In an optional embodiment, the matching clearance distance between the combustion chamber shell 3 and the top cover 1 is 0-0.02 mm, so that the two can be tightly attached to each other, and propellant slurry is prevented from entering a matching gap between the two.
In an optional embodiment, the pressure taps 6 of the combustion chamber shell 3 are arranged at the end, close to the nozzle, of the combustion chamber shell 3 and do not affect the nozzle assembly position, and the number of the pressure taps 6 is 1 or 2, so that the test run data of the whole combustion process of the propellant grain 4 can be normally and completely acquired, and the mounting and dismounting operations of the testing device are not affected.
In an alternative embodiment, the combustion chamber housing 3 is roughened on its inner surface before charging, which facilitates good engagement between the propellant grains 4 and the inner surface of the combustion chamber housing 3 and increases the mechanical bonding force therebetween.
In an optional embodiment, the propellant grain 4 is an end-fired solid grain which is directly formed after being poured and cured in vacuum by adopting propellant grain slurry, so that the manufacturing process is simplified, the complex operations of inserting cores, demoulding, shaping and the like during the manufacturing of the propellant grain of the existing standard test engine are avoided, and the operation is simple, convenient, efficient and safe; meanwhile, according to the equilibrium pressure formula of the engine combustion chamber
Pressure P of combustion chambercWith engine parameters (combustion area A)bNozzle throat area At) And propellant characteristics (characteristic velocity C)*Density rhoPCombustion rate coefficient a, pressure index n), combustion chamber pressure P for conventional solid propellants (n is less than 1)cAnd the combustion area AbIn a positive correlation relationship. After replacing the tubular side combustion grain charge structure of the existing standard test engine with the end combustion grain charge structure, estimating the combustion area A of the propellant grain according to the structure size of each standard test engine specified in the national military standard GJB96A-2001bDifferent from 1/6 of a test engine with different specifications and standards, the lower limit of the limiting pressure of the device can be obviously reduced under the same test state, the charging quality is greatly reduced under the condition of the same propellant grain thickness, the manufacturing cost is saved, and the device is very suitable for developing a small-dose new formula; in addition, the burning speed of the propellant of the device is only determined by the thickness of the propellant grain and the pressure of the combustion chamber, is not influenced by the eccentricity of the propellant grain and the size change of the inner diameter of the combustion chamber shell, and has the advantages of few interference factors, high testing precision, high weight reuse rate of the combustion chamber shell and the like.
In an alternative embodiment, the ignition device 9 is arranged in the middle area of the inner side of the throat part to ensure that the initial combustion surface of the propellant grain 4 is consistent with the contact state of high-temperature gas generated by ignition, so that partial ignition of the combustion surface of the propellant grain is avoided.
In an alternative embodiment, the nozzle, the ignition device 9, the nozzle plug cover 11, the O-shaped sealing ring 2, the bolt 8 and the nut 7 are all the same components of the existing standard test engine, so that the universality and the applicability of the device are improved.
The test shows that: the burning rate testing device for the high-burning-rate solid propellant can meet the testing requirement of the high-burning-rate solid propellant with the burning rate of 7MPa of 82mm/s in the pressure intensity range of 1 MPa-20 MPa, the testing process is normal, the problems of pressure relief and fire blowby are avoided, the testing P-t curve is stable, and the abnormal fluctuation phenomenon is avoided.
Those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.
Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make variations and modifications of the present invention without departing from the spirit and scope of the present invention by using the methods and technical contents disclosed above.
Claims (7)
1. A high-burning-speed solid propellant burning speed testing device is characterized by comprising a top cover (1), a combustion chamber shell (3), a propellant grain (4), a nozzle base (10), a throat liner (12), an ignition device (9) and a nozzle plugging cover (11);
the combustion chamber shell (3) is of a circular tube type structure; the propellant grain (4) is a solid end-burning grain and is arranged in the combustion chamber shell (3), and the inner cavity of the combustion chamber shell (3) is a combustion chamber (5); the combustion chamber shell (3) is provided with a pressure measuring hole (6), and the pressure measuring hole (6) is communicated with the combustion chamber (5);
one end of the combustion chamber shell (3) is connected with the top cover (1), and the other end of the combustion chamber shell is connected with the spray pipe base (10); the throat liner (12) is arranged in the nozzle base (10), and the throat liner (12) and the nozzle base (10) form a nozzle; the ignition device (9) and the spray pipe blanking cover (11) are sequentially arranged into the throat part of the spray pipe;
both ends of the combustion chamber shell (3) are respectively connected with the top cover (1) and the spray pipe base (10) by adopting flange type connecting structures;
the ignition device (9) is positioned in the middle area of the inner side of the throat part;
the propellant grain (4) is directly formed after being subjected to vacuum pouring and curing of propellant grain slurry, and is directly attached to the inner surface of the combustion chamber shell (3) after being subjected to roughening treatment and the circular boss of the second surface (102) of the top cover (1).
2. The high-burning-rate solid propellant burning rate testing device of claim 1, wherein two ends of the combustion chamber shell (3) are respectively connected with the top cover (1) and the nozzle base (10) in a sealing way; wherein, clearance fit is formed between the combustion chamber shell (3) and the top cover (1) and between the combustion chamber shell and the nozzle base (10), and then the combustion chamber shell is sealed by the sealing ring (2).
3. The high-burning-rate solid propellant burning rate testing device as claimed in claim 2, wherein both ends of the combustion chamber shell (3) are connected with the top cover (1) and the nozzle base (10) by bolts and nuts.
4. The high-burning-rate solid propellant burning rate testing device as claimed in any one of claims 1 to 3, wherein the ignition device (9) is connected with an external ignition power source through an ignition wire (13) via the nozzle cap (11).
5. The high-burning-rate solid propellant burning rate testing device as claimed in any one of claims 1 to 3, wherein, of the first surface (101) and the second surface (102) of the top cover (1) which are opposite to each other, the first surface (101) is a circular plane, and the second surface (102) is provided with a circular boss which is matched with the end surface of the combustion chamber shell (3).
6. The burning rate testing device of the solid propellant with high burning rate as claimed in any one of claims 1 to 3, wherein the fit clearance distance between the combustion chamber shell (3) and the top cover (1) is 0 to 0.02 mm.
7. The high-burning-rate solid propellant burning rate testing device as claimed in any one of claims 1 to 3, wherein the number of the pressure taps (6) is 1 or 2.
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| 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 |
| CN113417763B (en) * | 2021-07-27 | 2022-10-14 | 西北工业大学 | Solid rocket engine combustion surface retreating calculation method |
| CN114233517B (en) * | 2021-12-21 | 2024-03-29 | 西安零壹空间科技有限公司 | Solid rocket engine for high overload and drug pattern design method thereof |
| CN114412666A (en) * | 2022-01-04 | 2022-04-29 | 湖北三江航天江河化工科技有限公司 | Engine for testing pressure index of solid propellant and testing method thereof |
| CN114876667B (en) * | 2022-04-27 | 2023-09-12 | 西安零壹空间科技有限公司 | Composite solid rocket engine, charging method and filling equipment |
| CN115112376A (en) * | 2022-07-05 | 2022-09-27 | 南昌航空大学 | Gas injection and propellant coupling effect observation test device |
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| AU2003304725A1 (en) * | 2002-01-16 | 2008-05-29 | W.E. Research, Llc | Electrically controlled extinguishable solid propellant motors |
| CN104975985A (en) * | 2015-07-09 | 2015-10-14 | 南京理工大学 | Solid rocket engine igniting test device |
| CN110953090B (en) * | 2019-10-17 | 2021-06-04 | 西北工业大学 | Test device for simulating two-phase flow ablation environment in isobaric multichannel rocket engine |
| CN110886669A (en) * | 2019-11-25 | 2020-03-17 | 湖北三江航天江河化工科技有限公司 | Test engine for testing solid rocket propellant |
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