CN111810318B - Single-chamber double-thrust solid rocket engine and rocket - Google Patents
Single-chamber double-thrust solid rocket engine and rocket Download PDFInfo
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- CN111810318B CN111810318B CN202010594698.XA CN202010594698A CN111810318B CN 111810318 B CN111810318 B CN 111810318B CN 202010594698 A CN202010594698 A CN 202010594698A CN 111810318 B CN111810318 B CN 111810318B
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- combustion chamber
- wing groove
- thrust
- rocket engine
- propellant grain
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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
- F02K9/10—Shape or structure of solid propellant charges
- F02K9/12—Shape or structure of solid propellant charges made of two or more portions burning at different rates or having different characteristics
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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
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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
- F02K9/346—Liners, e.g. inhibitors
-
- 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
-
- 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
-
- 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
- F02K9/974—Nozzle- linings; Ablative coatings
-
- 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
- F02K9/978—Closures for nozzles; Nozzles comprising ejectable or discardable elements
Abstract
The application provides a single-chamber double-thrust solid rocket engine and a rocket, wherein the engine comprises a combustion chamber shell, a heat insulating layer attached to the inner surface of the combustion chamber shell, a propellant grain positioned in the combustion chamber shell, an igniter arranged at one end of the combustion chamber shell and a nozzle assembly arranged at the other end of the combustion chamber shell; the propellant grain is provided with inner holes penetrating through two ends of the propellant grain; the propellant grain is uniformly provided with a front wing groove extending in the radial direction at one end of the inner hole close to the igniter, and a tail wing groove extending in the radial direction at one end of the inner hole close to the spray pipe component; the ratio of the wing groove depth of the front wing groove and the tail wing groove to the maximum thickness of the propellant grain is more than or equal to 0.81; the gradient range of the axial slope of the rear wing of the grain is 160-170 degrees. The invention has large thrust variation range and stable grain structure, when the thrust mode of the engine is switched, the combustion surface variation speed of the grain is low, and the pressure variation speed of the combustion chamber can be effectively reduced, so that the thrust variation speed is alleviated.
Description
Technical Field
The invention relates to an aerospace power technology, in particular to a single-chamber double-thrust solid rocket engine and a rocket.
Background
The solid rocket engine has been widely used in various types of spacecraft due to its simple structure, reliable performance, and no need of maintenance, such as Long-Zheng-eleven. Solid rocket engines typically consist of a combustion chamber housing, propellant grains, insulation, a nozzle, and an igniter. In order to realize rapid switching of braking force, the thrust change of a common single-chamber double-thrust solid engine is rapid, the buffering transition of the change of a combustion surface of a grain is small, the pressure change of a combustion chamber is rapid, but when an aircraft flies at hypersonic speed, the rapid thrust change can cause rapid change of dynamic pressure balance, so that the attitude of an arrow body is rapidly changed, and great burden is brought to a control system. And in severe cases, the flight mission can fail.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: overcomes the defects of the prior art, and provides a single-chamber double-thrust solid rocket engine and a rocket which have large thrust variation range and moderate thrust variation speed and are suitable for hypersonic flight.
In a first aspect, the application provides a single-chamber double-thrust solid rocket engine, which comprises a combustion chamber shell, a heat insulating layer attached to the inner surface of the combustion chamber shell, a propellant grain positioned in the combustion chamber shell, an igniter arranged at one end of the combustion chamber shell and a nozzle assembly arranged at the other end of the combustion chamber shell;
the propellant grain is provided with inner holes penetrating through two ends of the propellant grain; the propellant grain is uniformly provided with radially extending front wing grooves at one end of the inner hole close to the igniter, and radially extending tail wing grooves are uniformly arranged at one end of the inner hole close to the spray pipe assembly;
the ratio of the wing groove depth of the front wing groove and the tail wing groove to the maximum thickness of the propellant grain is more than or equal to 0.81;
the side surface of the tail wing groove far away from the spray pipe assembly is a powder column rear wing axial slope, and the gradient range of the powder column rear wing axial slope is 160-170 degrees.
According to the technical scheme provided by the embodiment of the application, chamfers for connecting two side walls are arranged at the roots of the front wing groove and the tail wing groove; the chamfer is an elliptical chamfer with an axial ratio of 2: 1.
According to the technical scheme provided by the embodiment of the application, two ends of the combustion chamber shell are in an ellipsoid shape with the outer profile of 2: 1.
According to the technical scheme provided by the embodiment of the application, the heat insulation layer is formed by multiple layers of nitrile rubber; one end of the heat insulation layer close to the spray pipe assembly is provided with a plurality of carbon felt plates which are radially inserted into the plurality of layers of nitrile rubber.
According to the technical scheme provided by the embodiment of the application, the spray pipe assembly is sequentially provided with a convergence section and an expansion section which are fixedly connected with the combustion chamber shell; a throat liner protrudes from the convergent section into the combustion chamber shell; and a rubber soft blocking cover clamped on the throat liner and the convergence section is arranged in the shell of the combustion chamber shell.
According to the technical scheme provided by the embodiment of the application, the convergence section is made of a carbon fiber and high silica fiber composite molding material; the expansion section adopts a carbon ribbon-high silica cloth tape/phenolic resin composite winding structure, and the throat lining is made of a puncture carbon/carbon composite material.
According to the technical scheme provided by the embodiment of the application, the propellant grain is formed in the combustion chamber shell by mixing, pouring, curing and shaping a composite solid propellant.
According to the technical scheme provided by the embodiment of the application, the igniter is a basket igniter.
In a second aspect, the present application provides a rocket having mounted thereon a single-chamber, dual-thrust solid rocket engine of any of the above.
Compared with the prior art, the invention has the following beneficial effects:
(1) in the invention, the thrust variation range of the engine is improved by simultaneously using a large wing column type mode that the wing groove depth/thickness (the maximum thickness of the propellant grain) ratio exceeds 0.81 at the head part and the tail part of the grain;
(2) according to the invention, the acute degree of the change of the combustion surface of the explosive column when the thrust of the engine is changed is further reduced by increasing the axial gradient of the wing groove of the explosive column from 135 degrees to 160-170 degrees, the pressure change speed of the combustion chamber is reduced, and the thrust change further has a buffer transition process;
according to the technical scheme provided by the embodiment of the application, the stress is released by changing the round chamfer at the root part of the wing groove of the conventional explosive column into the oval chamfer with the axial ratio of 2:1, and the oval is more smooth and mellow than the round chamfer, so that the structural integrity of the explosive column in the working process of the engine is improved;
according to the technical scheme provided by the embodiment of the application, the solid rocket engine reduces the design difficulty of the control subsystem in the overall design process of the hypersonic aircraft rocket, and can reduce the design requirements on response speed and control moment by the control system.
Drawings
Fig. 1 to 2 are schematic structural views of embodiment 1 of the present application.
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.
Example 1
Please refer to fig. 1 to 2, which are schematic structural diagrams of a single-chamber double-thrust solid rocket engine provided in embodiment 1 of the present application, the rocket engine includes a combustion chamber housing 1, a heat insulating layer 2 attached to an inner surface of the combustion chamber housing 1, a propellant grain 3 located in the combustion chamber housing 1, an igniter 5 disposed at one end of the combustion chamber housing 1, and a nozzle assembly 4 disposed at the other end of the combustion chamber housing 1;
the propellant grain 3 is provided with an inner hole penetrating through two ends of the propellant grain; the propellant grain 3 is uniformly provided with a front wing groove 6 extending in the radial direction at one end of the inner hole close to the igniter 5, and a tail wing groove 7 extending in the radial direction at one end of the inner hole close to the spray pipe component 4;
in this embodiment, the combustion chamber housing 1 is a structural frame of the engine, and is formed by spin welding of D406A steel, the outer diameter is 750mm, the skirt interval is 3760mm, the length is 4290mm, the outer profile of the front and rear end enclosures is an ellipsoid with the ratio of the major axis to the minor axis of 2:1, the wall thickness of the cylinder section reaches 3.5mm, and the wall thickness of the end enclosure section is 4 mm.
In the embodiment, the ratio of the wing groove depth of the front wing groove 6 and the tail wing groove 7 to the maximum thickness of the propellant grain 3 is more than or equal to 0.81; as shown in fig. 2, the wing groove depth is H1, the maximum thickness of the propellant grain 3 is H2, then H1/H2 is greater than or equal to 0.81,
as shown in fig. 2, the side of the tail wing groove 7 away from the nozzle assembly 4 is a rear wing axial slope 30 of the explosive column, and the slope range of the rear wing axial slope 30 of the explosive column is 160-170 degrees.
As shown in fig. 2, the root parts of the front wing groove 6 and the tail wing groove 7 are provided with chamfers connecting two side walls; the chamfer is an elliptical chamfer with an axial ratio of 2: 1.
One end of the combustion chamber shell 1 is connected with the spray pipe component 4 through a flange, the other end of the combustion chamber shell 1 is connected with the igniter 5 through a flange, and the joint is locked, extruded and sealed through bolts by using an O-shaped sealing ring.
Wherein, the spray pipe assembly 4 is sequentially provided with a convergent section 40 and an expansion section 41 which are fixedly connected with the combustion chamber shell 1; the convergent section 40 has a throat insert 42 projecting into the combustion chamber housing 1; the casing of the combustion chamber casing 1 is internally provided with a rubber soft blocking cover 43 clamped on the throat liner 42 and the convergence section 40.
The convergence section 40 is made of a composite die-pressing material of carbon fibers and high silica fibers, namely the convergence section 40 is made of the carbon fibers and the high silica fibers through composite die pressing; the expansion section 41 adopts a carbon ribbon-high silica cloth tape/phenolic resin composite winding structure, namely the expansion section 41 is formed by compositely winding a carbon ribbon, a high silica cloth tape and adhesive phenolic resin, and the throat lining 42 is made of a puncture carbon and carbon composite material. The convergent section 40, the divergent section 41 and the throat insert 42 are fixedly connected by gluing.
Wherein, the propellant grain 3 is formed in the combustion chamber shell 1 by mixing, pouring, solidifying and shaping the composite solid propellant.
The igniter 5 is a basket igniter, the igniter is a starting device of the engine, the ignition powder uses boron/potassium nitrate ignition powder, the ignition tube adopts a standard DHQ-3D ignition tube, and the igniter 5 is positioned at the head of the combustion chamber shell 1.
When the solid rocket engine works, the igniter 5 is firstly ignited to work, the propellant grain 3 is ignited, and the propellant grain 3 is combusted in the combustion chamber shell 1 to generate high-temperature fuel gas which is accelerated and discharged through the nozzle component 4 to generate thrust.
When the invention is applied to a rocket flying at hypersonic speed, the thrust variation range of the engine is large, the grain structure is stable, and when the thrust mode of the engine is switched, the combustion surface variation speed of the grain is low, so that the pressure variation speed of a combustion chamber can be effectively reduced, and the thrust variation speed is alleviated. Meanwhile, the design difficulty of the rocket body control subsystem is effectively reduced, and the control performance of the rocket body flight attitude is guaranteed.
Example 2
This embodiment provides a rocket having the single-chamber, dual-thrust solid rocket engine of embodiment 1 mounted thereon.
The embodiments of the present invention are described in detail above with reference to the drawings, but the present invention is not limited to the described embodiments. It will be appreciated by those skilled in the art that various changes, modifications, substitutions and alterations can be made in these embodiments without departing from the principles and spirit of the invention, which is still within the scope of the invention.
Claims (8)
1. The single-chamber double-thrust solid rocket engine is characterized by comprising a combustion chamber shell (1), a heat insulating layer (2) attached to the inner surface of the combustion chamber shell (1), a propellant grain (3) positioned in the combustion chamber shell (1), an igniter (5) arranged at one end of the combustion chamber shell (1) and a nozzle assembly (4) arranged at the other end of the combustion chamber shell (1);
the propellant grain (3) is provided with an inner hole penetrating through two ends of the propellant grain; the propellant grain (3) is uniformly provided with a front wing groove (6) which extends in the radial direction at one end of the inner hole close to the igniter (5), and a tail wing groove (7) which extends in the radial direction at one end of the inner hole close to the spray pipe component (4);
the ratio of the wing groove depth of the front wing groove (6) and the tail wing groove (7) to the maximum thickness of the propellant grain (3) is more than or equal to 0.81;
the side surface of the tail wing groove (7) far away from the spray pipe assembly (4) is provided with a powder column rear wing axial slope (30), and the gradient range of the powder column rear wing axial slope (30) is 160-170 degrees;
chamfers for connecting the two side walls are arranged at the roots of the front wing groove (6) and the tail wing groove (7); the chamfer is an elliptical chamfer with an axial ratio of 2: 1.
2. The single-chamber dual-thrust solid rocket engine according to claim 1 wherein both ends of said combustion chamber housing (1) are ellipsoid with an outer profile of 2: 1.
3. Single-chamber, double-thrust solid rocket engine according to claim 1 wherein said thermal insulation layer (2) is formed by a plurality of layers of nitrile rubber; one end of the heat insulation layer (2) close to the spray pipe component (4) is provided with a plurality of carbon felt plates which are radially inserted into the plurality of layers of nitrile rubber.
4. Single-chamber, double-thrust solid rocket engine according to claim 1, characterized in that said nozzle assembly (4) is provided in sequence with a convergent section (40) and a divergent section (41) fixedly connected to said combustion chamber casing (1); a throat insert (42) protrudes from the convergent section (40) into the combustion chamber housing (1); a rubber soft blocking cover (43) clamped on the throat liner (42) and the convergence section (40) is arranged in the shell of the combustion chamber shell (1).
5. The single-chamber, double-thrust solid rocket engine according to claim 4, wherein said convergent section (40) is made of a composite molding material of carbon fibers and high silica fibers; the expansion section (41) adopts a carbon ribbon-high silica cloth tape/phenolic resin composite winding structure, and the throat lining (42) is made of a puncture carbon/carbon composite material.
6. Single-chamber, double-thrust solid rocket engine according to claim 1, characterized in that said propellant grains (3) are formed in said combustion chamber casing (1) from a composite solid propellant by mixing, pouring, curing, shaping.
7. The single-chamber, dual-thrust solid rocket engine according to claim 1 wherein said igniter (5) is a basket igniter.
8. A rocket characterized in that said rocket is equipped with a single-chamber, double-thrust solid rocket engine according to any one of claims 1 to 7.
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| Application Number | Priority Date | Filing Date | Title |
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| CN202010594698.XA CN111810318B (en) | 2020-06-28 | 2020-06-28 | Single-chamber double-thrust solid rocket engine and rocket |
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| CN202010594698.XA CN111810318B (en) | 2020-06-28 | 2020-06-28 | Single-chamber double-thrust solid rocket engine and rocket |
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| CN111810318A CN111810318A (en) | 2020-10-23 |
| CN111810318B true CN111810318B (en) | 2022-04-22 |
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Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| USH1082H (en) * | 1991-01-03 | 1992-08-04 | Andrew James W | Main-stage solid-propellant rocket motor design |
| US6148610A (en) * | 1997-06-16 | 2000-11-21 | Aerospatiale Societe Nationale Industrielle | Solid propellant charge for a propulsion unit and propulsion unit equipped with such a charge |
| RU2326261C1 (en) * | 2006-10-30 | 2008-06-10 | Федеральное государственное унитарное предприятие "Научно-исследовательский институт полимерных материалов" | Solid-fuel rocket charge |
| RU2461728C2 (en) * | 2010-12-03 | 2012-09-20 | Федеральное государственное унитарное предприятие "Научно-исследовательский институт полимерных материалов" | Solid-propellant rocket engine |
| CN103644046A (en) * | 2013-11-21 | 2014-03-19 | 北京动力机械研究所 | Charging grain structure of solid rocket engine |
| WO2016128804A1 (en) * | 2015-02-12 | 2016-08-18 | Director General, Defence Research & Development Organisation (Drdo) | Propellant grain for a solid rocket motor |
| CN107322246A (en) * | 2017-08-23 | 2017-11-07 | 湖北三江航天江北机械工程有限公司 | The big jet pipe bonding press-fitting technique of solid propellant rocket |
| CN209654136U (en) * | 2018-12-29 | 2019-11-19 | 西安北方惠安化学工业有限公司 | A kind of fourth hydroxyl complex solid rocket engine boost motor of Design of Single-chamber |
| CN210289958U (en) * | 2019-03-29 | 2020-04-10 | 湖北航天化学技术研究所 | Pressed metal particle generating agent grain structure |
Family Cites Families (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3032975A (en) * | 1957-03-25 | 1962-05-08 | Phillips Petroleum Co | Rocket motor |
| NL292576A (en) * | 1962-05-10 | |||
| EP0184014B1 (en) * | 1984-12-06 | 1988-04-27 | Werkzeugmaschinenfabrik Oerlikon-Bührle AG | A propellant configuration for a solid propellant rocket motor |
| JP3150498B2 (en) * | 1993-07-26 | 2001-03-26 | 雅弘 高野 | Propellant grains for solid rocket motors. |
| JP4215322B2 (en) * | 1998-11-25 | 2009-01-28 | ▲高▼野 雅弘 | Rocket motor |
| RU2196915C1 (en) * | 2001-09-20 | 2003-01-20 | Федеральное государственное унитарное предприятие "Научно-исследовательский институт полимерных материалов" | Solid-propellant rocket engine |
| RU2206778C1 (en) * | 2001-10-18 | 2003-06-20 | Федеральное государственное унитарное предприятие "Научно-исследовательский институт полимерных материалов" | Solid-fuel rocket engine charge |
| US20090205313A1 (en) * | 2008-02-14 | 2009-08-20 | Cavalleri Robert J | Fast Response Solid Fuel Rocket Motor |
| CN203847275U (en) * | 2014-02-25 | 2014-09-24 | 陕西中天火箭技术股份有限公司 | Composite propellant engine |
| CN110145411B (en) * | 2019-05-27 | 2020-09-18 | 上海新力动力设备研究所 | Solid rocket engine inner hole tubular charging structure with inner cavity partition plate and method |
| CN110589034B (en) * | 2019-09-25 | 2021-12-07 | 北京凌空天行科技有限责任公司 | Recyclable high-speed flying rocket and recycling method |
| CN110588977B (en) * | 2019-09-25 | 2021-03-23 | 北京凌空天行科技有限责任公司 | Solid rocket aircraft |
-
2020
- 2020-06-28 CN CN202010594698.XA patent/CN111810318B/en active Active
Patent Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| USH1082H (en) * | 1991-01-03 | 1992-08-04 | Andrew James W | Main-stage solid-propellant rocket motor design |
| US6148610A (en) * | 1997-06-16 | 2000-11-21 | Aerospatiale Societe Nationale Industrielle | Solid propellant charge for a propulsion unit and propulsion unit equipped with such a charge |
| RU2326261C1 (en) * | 2006-10-30 | 2008-06-10 | Федеральное государственное унитарное предприятие "Научно-исследовательский институт полимерных материалов" | Solid-fuel rocket charge |
| RU2461728C2 (en) * | 2010-12-03 | 2012-09-20 | Федеральное государственное унитарное предприятие "Научно-исследовательский институт полимерных материалов" | Solid-propellant rocket engine |
| CN103644046A (en) * | 2013-11-21 | 2014-03-19 | 北京动力机械研究所 | Charging grain structure of solid rocket engine |
| WO2016128804A1 (en) * | 2015-02-12 | 2016-08-18 | Director General, Defence Research & Development Organisation (Drdo) | Propellant grain for a solid rocket motor |
| CN107322246A (en) * | 2017-08-23 | 2017-11-07 | 湖北三江航天江北机械工程有限公司 | The big jet pipe bonding press-fitting technique of solid propellant rocket |
| CN209654136U (en) * | 2018-12-29 | 2019-11-19 | 西安北方惠安化学工业有限公司 | A kind of fourth hydroxyl complex solid rocket engine boost motor of Design of Single-chamber |
| CN210289958U (en) * | 2019-03-29 | 2020-04-10 | 湖北航天化学技术研究所 | Pressed metal particle generating agent grain structure |
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| CN111810318A (en) | 2020-10-23 |
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