CN104612858A - Adjustable double-pulse solid rocket engine jet pipe testing device - Google Patents
Adjustable double-pulse solid rocket engine jet pipe testing device Download PDFInfo
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- CN104612858A CN104612858A CN201510048129.4A CN201510048129A CN104612858A CN 104612858 A CN104612858 A CN 104612858A CN 201510048129 A CN201510048129 A CN 201510048129A CN 104612858 A CN104612858 A CN 104612858A
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- jet pipe
- safety pin
- nozzle block
- solid rocket
- pin
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Abstract
The invention discloses an adjustable double-pulse solid rocket engine jet pipe testing device. The adjustable double-pulse solid rocket engine jet pipe testing device comprises a jet pipe base, a throat lining piece, an expanding section component, a safety pin, an upper shearing sleeve and a lower shearing sleeve, wherein the throat lining piece is arranged at the fuel gas inlet end of the jet pipe base, the expanding section component is arranged at the fuel gas outlet end of the jet pipe base, the throat lining piece is matched with an annular step surface in the jet pipe base in a positioned mode through an annular step surface in the axial direction of the throat lining piece, the safety pin penetrates through a pin hole in the wall of the jet pipe base to stretch into a pin hole in the expanding section component, the upper shearing sleeve is arranged between the safety pin and the jet pipe base, the lower shearing sleeve is arranged between the safety pin and the expanding section component, and matched threads are arranged at corresponding positions of the ends, close to the outer wall of the jet pipe base, of the safety pin and the jet pipe base. By the adoption of the adjustable double-pulse solid rocket engine jet pipe testing device, testing safety can be improved, and testing cost can be reduced.
Description
Technical field
The invention belongs to Solid Rocket Engine Test research field, particularly a kind of adjustable Double pulse solid rocket motor nozzle test device.
Background technique
Along with the development of military science and technology, in order to improve existence and the penetration ability of guided missile, emphasize stand-off firing, trajectory variability.Pulsed motor utilizes isolation mounting that firing chamber is divided into some parts, by repeatedly shutting down and lighting a fire, reasonable distribution thrust and each burst length, realizes the optimum management of trajectory optimal control and engine power.For Double pulse solid rocket motor, have compared with traditional single thrust or Design of Single-chamber motor that range is farther, mobility is stronger, precision is higher.
Ground performance experimental study is the important means in development process.Because double pulsing motor barrier technique awaits further raising, the second pulsed motor accidental ignition is unsuccessfully caused for preventing interlayer work in test, chamber pressure is uprushed, housing is caused to burst, usually relief hole is set at head of combustion chamber or housing side, when internal pressure of combustion chamber be greater than burning chamber shell can bear pressure maximum time, relief hole is started working, and avoids the danger of blasting.But the method has certain defect.When being arranged on the relief hole work of head or side, firing chamber, the high-temperature high-pressure fuel gas of ejection may damage thrust pickup and other the data signal line of head.In addition, when carrying out many group tests, when throat's ablation has not seriously reached design objective, usually need change overall nozzle component, this method is uneconomical.Therefore, need to design a kind of novel adjustable jet pipe.
Summary of the invention
The object of the present invention is to provide a kind of adjustable Double pulse solid rocket motor nozzle test device that can improve experimental safe, save experimentation cost.
The technical solution realizing the object of the invention is:
A kind of adjustable Double pulse solid rocket motor nozzle test device, comprise nozzle block, larynx backing member, extending section parts, safety pin, upper shearing sleeve and down cut sleeve, larynx backing member is arranged on the fuel gas inlet end in nozzle block, extending section parts are arranged on the gas outlet end in nozzle block, larynx backing member is by the annular table terrace location fit in its annular table terrace axially and nozzle block, safety pin stretches into the pin-and-hole on extending section parts through the pin-and-hole on nozzle block wall, shearing sleeve is provided with between safety pin and nozzle block, down cut sleeve is set between safety pin and extending section parts, safety pin and nozzle block pin-and-hole arrange the screw thread matched at corresponding position near one end of nozzle block outer wall.
The present invention compared with prior art, its remarkable advantage:
(1) part of the present invention is few, and structure is simple, easy to assembly.
(2) the present invention is directed to different design objectives, under the prerequisite ensureing nozzle block size constancy, by changing the size of larynx backing member and expansion segment, can adapt to the designing requirement such as different divergence ratios, convergence half-angle, there is certain versatility.
(3) safety pin of the present invention and upper and lower shearing sleeve with the use of, diameter of safety pin and upper and lower shearing sleeve diameter can be regulated according to test demand, the convenient shear strength regulating safety pin.
(4) present invention, avoiding the defect needing integral replacing nozzle component when profile ablation is serious in jet pipe, only need larynx backing member serious for ablation to change, nozzle block still can continue to use, and effectively can save experimentation cost.
Below in conjunction with accompanying drawing, the present invention is described in further detail.
Accompanying drawing explanation
Fig. 1 be the present invention's adjustable Double pulse solid rocket motor nozzle test device partly cut open structural representation.
Fig. 2 is the axial sectional view of the present invention's adjustable Double pulse solid rocket motor nozzle test device.
Fig. 3 is the contrast scheme of installation of the present invention's adjustable Double pulse solid rocket motor nozzle test device different size safety pin; A () is that thicker safety pin coordinates thinner shearing sleeve, (b) is that thinner safety pin coordinates thicker shearing sleeve.
Embodiment
Composition graphs 1 ~ Fig. 3:
A kind of adjustable Double pulse solid rocket motor nozzle test device of the present invention, comprise nozzle block 1, larynx backing member 2, extending section parts 3, safety pin 4, upper shearing sleeve 5 and down cut sleeve 6, larynx backing member 2 is arranged on the fuel gas inlet end in nozzle block 1, extending section parts 3 are arranged on the gas outlet end in nozzle block 1, larynx backing member 2 is by the annular table terrace location fit in its annular table terrace axially and nozzle block 1, safety pin 4 stretches into the pin-and-hole on extending section parts 3 through the pin-and-hole on nozzle block 1 wall, shearing sleeve 5 is provided with between safety pin 4 and nozzle block 1, down cut sleeve 6 is set between safety pin 4 and extending section parts 3, safety pin 4 and nozzle block 1 pin-and-hole arrange the screw thread matched at corresponding position near one end of nozzle block 1 outer wall.
The internal diameter of larynx backing member 2 is less than the internal diameter of extending section parts 3, and seamlessly transits between the two.
Safety pin 4 is symmetricly set on the circumference of nozzle block 1.
Standard: according to design objective, loads nozzle block by the larynx backing member 2 meeting designing requirement successively with extending section parts 3, first puts into down cut sleeve 6, and then put into shearing sleeve 5 from the pin-and-hole of nozzle block 1, note precedence.Finally to screw on safety pin 4.After confirming that assembling is errorless, nozzle block 1 is screwed into firing chamber, screws, can test.
Working principle: under the prerequisite ensureing nozzle block 1 size constancy, by changing the size of larynx backing member 2 and extending section parts 3, can adapt to the designing requirement such as different divergence ratios, convergence half-angle; Regulate the internal diameter (as Fig. 3) of safety pin 4 diameter and upper shearing sleeve 5 and down cut sleeve 6 according to test demand, convenient adjustment safety pin 4 shear strength, makes design have certain versatility.
In test, if interlayer work failure under the first pulsed motor high-temperature high-pressure fuel gas effect, cause the second pulse accidental ignition, chamber pressure is caused sharply to increase, under high-pressure gas effect, safety pin 4 is sheared disconnected, larynx backing member 2 and extending section parts 3 are released from nozzle block 1, makes blast tube area become large, thus reach the effect of safety relief.
Claims (3)
1. an adjustable Double pulse solid rocket motor nozzle test device, it is characterized in that: comprise nozzle block (1), larynx backing member (2), extending section parts (3), safety pin (4), upper shearing sleeve (5) and down cut sleeve (6), larynx backing member (2) is arranged on the fuel gas inlet end in nozzle block (1), extending section parts (3) are arranged on the gas outlet end in nozzle block (1), larynx backing member (2) is by the annular table terrace location fit in its annular table terrace axially and nozzle block (1), safety pin (4) stretches into the pin-and-hole on extending section parts (3) through the pin-and-hole on nozzle block (1) wall, shearing sleeve (5) is provided with between safety pin (4) and nozzle block (1), down cut sleeve (6) is set between safety pin (4) and extending section parts (3), safety pin (4) and nozzle block (1) pin-and-hole arrange the screw thread matched at corresponding position near one end of nozzle block (1) outer wall.
2. adjustable Double pulse solid rocket motor nozzle test device according to claim 1, is characterized in that: the internal diameter of described larynx backing member (2) is less than the internal diameter of extending section parts (3), and seamlessly transits between the two.
3. adjustable Double pulse solid rocket motor nozzle test device according to claim 1 and 2, is characterized in that: described safety pin (4) is symmetricly set on the circumference of nozzle block (1).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201510048129.4A CN104612858B (en) | 2015-01-29 | 2015-01-29 | Adjustable Double pulse solid rocket motor nozzle test device |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201510048129.4A CN104612858B (en) | 2015-01-29 | 2015-01-29 | Adjustable Double pulse solid rocket motor nozzle test device |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN104612858A true CN104612858A (en) | 2015-05-13 |
| CN104612858B CN104612858B (en) | 2016-09-21 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN201510048129.4A Expired - Fee Related CN104612858B (en) | 2015-01-29 | 2015-01-29 | Adjustable Double pulse solid rocket motor nozzle test device |
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| CN (1) | CN104612858B (en) |
Cited By (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105089853A (en) * | 2015-08-25 | 2015-11-25 | 南京理工大学 | Combined exhaust pipe for solid rocket engine |
| CN105971769A (en) * | 2016-07-13 | 2016-09-28 | 西安航天动力测控技术研究所 | Fuel gas guide device for vertical type test run of small solid rocket engine |
| CN106050477A (en) * | 2016-07-28 | 2016-10-26 | 湖北航天技术研究院总体设计所 | Combined throat liner spraying pipe of solid rocket engine and manufacturing method |
| CN106762229A (en) * | 2016-11-04 | 2017-05-31 | 上海新力动力设备研究所 | A kind of dipulse missile propulsive plant anti-clogging pressure measurement structure |
| CN109252978A (en) * | 2018-08-31 | 2019-01-22 | 西安航天动力技术研究所 | A kind of control bar type change propulsive solid engines |
| CN110749536A (en) * | 2019-10-16 | 2020-02-04 | 南京理工大学 | Solid rocket engine thermal protection material ablation experimental device |
| CN112012852A (en) * | 2020-09-02 | 2020-12-01 | 西安航天动力测控技术研究所 | Reverse-injection protection and collection device and method for solid rocket engine |
| CN114109644A (en) * | 2021-10-09 | 2022-03-01 | 中国航发贵阳发动机设计研究所 | Method for adjusting area of exit throat of special-shaped fixed spray pipe of aircraft engine |
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| DE3428469A1 (en) * | 1984-08-02 | 1986-02-13 | Bayern-Chemie Gesellschaft für flugchemische Antriebe mbH, 8261 Aschau | Thrust nozzle for solid-fuel rocket motors |
| US4729512A (en) * | 1985-11-21 | 1988-03-08 | Laing Johannes L N | Rocket nozzles in layered construction |
| US20050198940A1 (en) * | 2004-03-10 | 2005-09-15 | Koshoffer John M. | Ablative afterburner |
| CN203285567U (en) * | 2013-05-20 | 2013-11-13 | 宜宾北方川安化工有限公司 | Combined type spray pipe structure |
| CN204003155U (en) * | 2014-04-30 | 2014-12-10 | 山西北方兴安化学工业有限公司 | A kind of rocket motor with decompression protection device |
| CN204419396U (en) * | 2015-01-29 | 2015-06-24 | 南京理工大学 | Adjustable Double pulse solid rocket motor nozzle test device |
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2015
- 2015-01-29 CN CN201510048129.4A patent/CN104612858B/en not_active Expired - Fee Related
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE3428469A1 (en) * | 1984-08-02 | 1986-02-13 | Bayern-Chemie Gesellschaft für flugchemische Antriebe mbH, 8261 Aschau | Thrust nozzle for solid-fuel rocket motors |
| US4729512A (en) * | 1985-11-21 | 1988-03-08 | Laing Johannes L N | Rocket nozzles in layered construction |
| US20050198940A1 (en) * | 2004-03-10 | 2005-09-15 | Koshoffer John M. | Ablative afterburner |
| CN203285567U (en) * | 2013-05-20 | 2013-11-13 | 宜宾北方川安化工有限公司 | Combined type spray pipe structure |
| CN204003155U (en) * | 2014-04-30 | 2014-12-10 | 山西北方兴安化学工业有限公司 | A kind of rocket motor with decompression protection device |
| CN204419396U (en) * | 2015-01-29 | 2015-06-24 | 南京理工大学 | Adjustable Double pulse solid rocket motor nozzle test device |
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105089853A (en) * | 2015-08-25 | 2015-11-25 | 南京理工大学 | Combined exhaust pipe for solid rocket engine |
| CN105971769A (en) * | 2016-07-13 | 2016-09-28 | 西安航天动力测控技术研究所 | Fuel gas guide device for vertical type test run of small solid rocket engine |
| CN106050477A (en) * | 2016-07-28 | 2016-10-26 | 湖北航天技术研究院总体设计所 | Combined throat liner spraying pipe of solid rocket engine and manufacturing method |
| CN106762229A (en) * | 2016-11-04 | 2017-05-31 | 上海新力动力设备研究所 | A kind of dipulse missile propulsive plant anti-clogging pressure measurement structure |
| CN109252978A (en) * | 2018-08-31 | 2019-01-22 | 西安航天动力技术研究所 | A kind of control bar type change propulsive solid engines |
| CN110749536A (en) * | 2019-10-16 | 2020-02-04 | 南京理工大学 | Solid rocket engine thermal protection material ablation experimental device |
| CN112012852A (en) * | 2020-09-02 | 2020-12-01 | 西安航天动力测控技术研究所 | Reverse-injection protection and collection device and method for solid rocket engine |
| CN112012852B (en) * | 2020-09-02 | 2021-06-18 | 西安航天动力测控技术研究所 | Reverse-injection protection and collection device and method for solid rocket engine |
| CN114109644A (en) * | 2021-10-09 | 2022-03-01 | 中国航发贵阳发动机设计研究所 | Method for adjusting area of exit throat of special-shaped fixed spray pipe of aircraft engine |
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| Publication number | Publication date |
|---|---|
| CN104612858B (en) | 2016-09-21 |
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Granted publication date: 20160921 Termination date: 20180129 |