CN110566371A - detachable rocket engine for test - Google Patents
detachable rocket engine for test Download PDFInfo
- Publication number
- CN110566371A CN110566371A CN201910955006.7A CN201910955006A CN110566371A CN 110566371 A CN110566371 A CN 110566371A CN 201910955006 A CN201910955006 A CN 201910955006A CN 110566371 A CN110566371 A CN 110566371A
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- China
- Prior art keywords
- outer sleeve
- nozzle
- head
- fuel
- thrust chamber
- Prior art date
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- 238000012360 testing method Methods 0.000 title abstract description 17
- 239000000446 fuel Substances 0.000 claims description 53
- 239000007800 oxidant agent Substances 0.000 claims description 50
- 230000001590 oxidative effect Effects 0.000 claims description 43
- 238000002485 combustion reaction Methods 0.000 claims description 30
- 239000007788 liquid Substances 0.000 claims description 22
- 238000002347 injection Methods 0.000 claims description 14
- 239000007924 injection Substances 0.000 claims description 14
- 238000001816 cooling Methods 0.000 claims description 11
- 238000011144 upstream manufacturing Methods 0.000 claims description 8
- 230000007704 transition Effects 0.000 claims description 7
- 230000013011 mating Effects 0.000 claims description 4
- 239000007921 spray Substances 0.000 abstract description 4
- 239000000463 material Substances 0.000 description 10
- RYGMFSIKBFXOCR-UHFFFAOYSA-N Copper Chemical compound [Cu] RYGMFSIKBFXOCR-UHFFFAOYSA-N 0.000 description 9
- 229910052802 copper Inorganic materials 0.000 description 9
- 239000010949 copper Substances 0.000 description 9
- 239000007789 gas Substances 0.000 description 8
- 238000000034 method Methods 0.000 description 8
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 description 6
- 239000003350 kerosene Substances 0.000 description 6
- 239000001301 oxygen Substances 0.000 description 6
- 229910052760 oxygen Inorganic materials 0.000 description 6
- 238000005507 spraying Methods 0.000 description 5
- SBYXRAKIOMOBFF-UHFFFAOYSA-N copper tungsten Chemical compound [Cu].[W] SBYXRAKIOMOBFF-UHFFFAOYSA-N 0.000 description 4
- 230000008595 infiltration Effects 0.000 description 4
- 238000001764 infiltration Methods 0.000 description 4
- 230000008569 process Effects 0.000 description 4
- 230000009286 beneficial effect Effects 0.000 description 3
- 229910000851 Alloy steel Inorganic materials 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 235000013290 Sagittaria latifolia Nutrition 0.000 description 1
- 229910000831 Steel Inorganic materials 0.000 description 1
- 238000002679 ablation Methods 0.000 description 1
- 230000006978 adaptation Effects 0.000 description 1
- 238000005452 bending Methods 0.000 description 1
- 230000000903 blocking effect Effects 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 235000015246 common arrowhead Nutrition 0.000 description 1
- 239000012141 concentrate Substances 0.000 description 1
- 238000005034 decoration Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000013461 design Methods 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 230000003628 erosive effect Effects 0.000 description 1
- 239000002737 fuel gas Substances 0.000 description 1
- 230000014509 gene expression Effects 0.000 description 1
- 238000009434 installation Methods 0.000 description 1
- 238000011900 installation process Methods 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 210000002445 nipple Anatomy 0.000 description 1
- 238000011160 research Methods 0.000 description 1
- 238000007789 sealing Methods 0.000 description 1
- 238000004088 simulation Methods 0.000 description 1
- 239000007787 solid Substances 0.000 description 1
- 239000000243 solution Substances 0.000 description 1
- 239000010935 stainless steel Substances 0.000 description 1
- 229910001220 stainless steel Inorganic materials 0.000 description 1
- 239000010959 steel Substances 0.000 description 1
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 description 1
Classifications
-
- 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
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Testing Of Engines (AREA)
Abstract
The invention provides a detachable rocket engine for a test, which comprises a bracket and an ejector rocket arranged on the bracket, wherein the ejector rocket comprises a head component, a thrust chamber component and a tail nozzle; the front end of the thrust chamber component is connected with the head component, and the rear end of the thrust chamber component is connected with the tail spray pipe. The invention provides a detachable rocket engine for a test, which is convenient for the test.
Description
Technical Field
the invention belongs to the field of rocket engines, and relates to a small-size detachable rocket engine. The rocket engine of the invention is mainly applied to a small rocket engine using oxygen/kerosene.
background
Rocket engines have become the main power of various spacecrafts, and are mainly divided into solid rocket engines and liquid rocket engines at present. The combustion chamber wall is easily burnt in the liquid rocket engine due to the high temperature of combustion products. Therefore, active thermal protection and passive thermal protection methods are proposed for the wall surface of the combustion chamber, wherein the active thermal protection is usually implemented by cooling the wall surface by using fuel or water, and the defect is that the structure is too complex; the passive thermal protection mainly adopts high-temperature resistant materials such as copper and the like as the wall surface of the thrust chamber, and the passive thermal protection becomes a research hotspot due to simple structure.
At present, the following problems exist by adopting passive thermal protection:
(1) The high-temperature resistant material is heavy in copper ratio, and the wall surface of the combustion chamber is completely made of copper, so that the combustion chamber is not suitable for the test requirement of a small engine.
(2) the high-temperature resistant material copper material is soft and has insufficient strength, so that the pressure measuring hole is not suitable for being arranged on the wall surface of the combustion chamber, and the strength of the engine is not suitable for being ensured.
the design of the thrust chamber wall surface is restricted by the problems of the thrust chamber wall surface of the liquid rocket engine, and how to organically combine the strength of common high-temperature steel and the high-temperature resistance of copper is particularly important for designing an engine scheme with an excellent structure.
Disclosure of Invention
The invention aims to provide a detachable rocket engine for a test, which is convenient for the test.
the technical scheme adopted by the invention is as follows:
The invention provides a detachable rocket engine for a test, which comprises a bracket and an ejector rocket arranged on the bracket, wherein the ejector rocket comprises a head component, a thrust chamber component and a tail nozzle; the front end of the thrust chamber component is connected with the head component, and the rear end of the thrust chamber component is connected with the tail spray pipe;
The head assembly comprises an igniter, and a fitting part matched with the igniter is arranged on the outer side of the igniter;
a head outer sleeve is arranged on the outer side of the matching piece, the front end of the head outer sleeve and the matching piece are mutually enclosed to form a fuel collecting cavity, a gap between the head outer sleeve and the matching piece forms a fuel nozzle, and the fuel nozzle is communicated with the fuel collecting cavity;
the rear end of the head jacket along the air inflow direction is provided with a jetting piece, the head jacket and the jetting piece are mutually enclosed to form an oxidant liquid collecting cavity, an oxidant nozzle is arranged on the jetting piece, and the oxidant nozzle is communicated with the oxidant liquid collecting cavity;
The bracket comprises a limiting block for preventing the ejection rocket from moving in the direction opposite to the air inflow direction.
optionally, the support includes a front support and a rear support, the front support is disposed at a connection position of the head assembly and the thrust chamber assembly, and the rear support is disposed at a connection position of the thrust chamber assembly and the tail nozzle;
The limiting block is arranged on the rear support.
Optionally, the thrust chamber assembly includes an equal straight section, a throat portion and a thrust chamber outer sleeve, the equal straight section is arranged at the front end of the throat portion, the thrust chamber outer sleeve is arranged on the outer sides of the equal straight section and the throat portion in a sleeved mode, and the thrust chamber outer sleeve extends to the front end of the equal straight section.
Optionally, the thrust chamber outer sleeve sequentially comprises an outer sleeve front end, an outer sleeve body and an outer sleeve tail end from front to back; the outer sleeve body is arranged on the periphery of the equal straight section and the throat part, and the front end of the outer sleeve, the injection piece and the rear end of the head outer sleeve as well as the tail end of the outer sleeve and the tail nozzle are connected through connecting pieces.
optionally, a combustion chamber is formed in the equal straight section;
The front end of the outer sleeve is provided with a pressure measuring hole for installing a pressure measuring pipeline, and the pressure measuring pipeline is communicated with the inside of the front end of the outer sleeve; the inner wall of the front end of the outer sleeve is also provided with a positioning groove.
optionally, the fitting comprises a thick portion, a thin portion and a fitting end portion, a transition portion is arranged between the thick portion and the thin portion, and the fitting end portion is arranged at the front end of the thick portion;
the cross sections of the end part, the thick part, the transition part and the thin part of the fitting part are gradually reduced;
the thin part is internally provided with a threaded hole and is connected with the igniter in a matching way through the threaded hole.
Optionally, an oxidant port and a fuel port are arranged on the head outer sleeve, the oxidant port is used for installing an oxidant nozzle, and the oxidant nozzle is communicated with the oxidant liquid collecting cavity;
The fuel port is used for mounting a fuel nozzle which is communicated with the fuel liquid collecting cavity;
the front end of the head outer sleeve is connected with the end part of the fitting piece through a connecting piece.
Optionally, a plurality of circumferential seam ribs are arranged on the inner wall of the injection piece to form a plurality of air film cooling circumferential seams; and a positioning key matched with the positioning groove is arranged at the rear end of the spraying piece.
Optionally, the fuel nozzle includes an upstream portion, a downstream portion, and a bend disposed between the upstream portion and the downstream portion.
The invention has the beneficial effects that:
1. The invention provides a detachable rocket engine for a test, which is convenient for the test. The head assembly and the thrust chamber assembly of the engine and the tail nozzle and the thrust chamber assembly are detachably connected, so that the experimental part is changed to test according to different experimental working conditions.
2. The straight section of the engine of the embodiment of the invention adopts red copper, and the throat part adopts tungsten copper infiltration as a heat-proof material; the thrust chamber outer sleeve and the tail nozzle are made of high-temperature alloy steel materials. Not only ensures that the inner wall surfaces of the combustion chamber and the throat part can bear the erosion of high-temperature heat flow, but also ensures that the engine has enough strength, and is also beneficial to arranging the pressure measuring holes. The simulation result shows that under the working condition of large flow, the throat temperature is 1950K, which is far less than the temperature resistance of the tungsten copper infiltration (more than 3000K), and the tungsten copper infiltration device can normally work for 20 s. The alloy steel is mainly positioned at the front end of an oxidant impact point, and the combustion flame is ensured to be positioned at the red copper material part.
3. the embodiment of the invention has compact structure and small size, and is suitable for small-sized rocket engines using oxygen/kerosene. According to the embodiment of the invention, the injection piece is designed to be a direct current self-impact type, the spark plug igniter is arranged in the center of the head component, and compared with a centrifugal nozzle, the scheme can reduce the outer diameter of the fire injection arrow head component, and is more beneficial to designing a small-size rocket engine.
4. In the embodiment of the invention, the wall surface of the combustion chamber is cooled by adopting a gas film cooling mode, wherein the gas film seams are three gas film cooling circular seams, so that compared with a common gas film hole, the gas film cooling circular seam reduces a gas film blind area and also reduces the ablation effect of high-temperature heat flow on the wall surface of the nozzle; such a circumferential seam is also more advantageous to manufacture than an inclined nozzle.
5. in the embodiment of the invention, all parts are detachably connected, such as the front end and the fitting piece of the head outer sleeve, the rear end of the head outer sleeve, the injection piece, the thrust chamber component and the tail nozzle, which are detachably connected by screws.
Drawings
FIG. 1 is a schematic perspective view of a detachable rocket engine for testing according to an embodiment of the present invention;
FIG. 2 is a schematic diagram of a quarter section of a detachable rocket engine for testing according to an embodiment of the present invention;
FIG. 3 is an enlarged view of a portion of FIG. 2 at A;
FIG. 4 is a schematic view of a head cover according to an embodiment of the present invention;
FIG. 5 is a schematic view of a fitting configuration in an embodiment of the present invention;
FIG. 6 is a schematic view of a structure of a spray member in an embodiment of the present invention;
FIG. 7 is a schematic view of a thrust chamber outer sleeve according to an embodiment of the present invention;
FIG. 8 is a simulated graph of the change of the monitored throat temperature with time according to the embodiment of the invention.
1, a bracket; 10 a front bracket; 11, a rear support; 110 a limiting block; 4, a gasket; 5, a tail spray pipe; 6, connecting pieces;
2 a head assembly; 20 head cover; 200 an oxidant liquid collection chamber; 201 oxidant port; 202 a fuel port; 21 a fitting member; 210 a fuel collection plenum; 211 a fuel nozzle; 2110 upstream part; 2111 downstream portion; 2112 bending part; 212 fitting ends; 213 thick portion; 214 details; 215 a transition portion; 22 an injector; 221 an oxidant nozzle; 222 film cooling the annular seam; 223 circumferential seam ribs; 224 an orientation key; 23 an igniter; 24 an igniter sleeve; 25 oxidant nozzle; 26 a fuel nipple;
3, a thrust chamber component; 30 equal straight sections; 300 a combustion chamber; 31 a throat portion; 33 pressure measuring pipelines; 34 a thrust chamber housing; 340 a jacket front end; 3400 pressure measuring hole; 3401 locating slot; 341 a jacket body; 342 at the end of the jacket.
Detailed Description
the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the present invention, and not all embodiments. All other embodiments, which can be obtained by a person skilled in the art without any inventive step based on the embodiments of the present invention, belong to the scope of the present invention. It should be noted that the directional terms such as "front" and "rear" related to the embodiments of the present invention are determined by referring to the incoming flow direction of air, and are not intended to limit the scope of the present invention.
The embodiment of the invention provides a detachable rocket engine for a test, which comprises a bracket 1 and an ejector rocket arranged on the bracket 1, wherein the ejector rocket comprises a head component 2, a thrust chamber component 3 and a tail nozzle 5; the front end of the thrust chamber component 3 is connected with the head component 2, and the rear end of the thrust chamber component 3 is connected with the tail nozzle 5;
the head assembly 2 comprises an igniter 23, a head outer sleeve 20, a fitting piece 21 and a spraying piece 22, wherein the fitting piece 21 matched with the igniter 23 is arranged on the outer side of the igniter 23;
The electrical wires of the igniter 23 of this embodiment are directly connected to an external power source at the front end of the head assembly.
A head outer sleeve 20 is arranged outside the fitting 21, the front end of the head outer sleeve 20 in the air inflow direction and the fitting 714 mutually enclose to form a fuel collecting cavity 210, a gap between the head outer sleeve 20 and the fitting 21 forms a fuel nozzle 211, and the fuel nozzle 211 is communicated with the fuel collecting cavity 210;
The rear end of the head jacket 20 in the air inflow direction is provided with a spraying piece 22, the head jacket 20 and the spraying piece 22 enclose each other to form an oxidant liquid collecting cavity 200, the spraying piece 22 is provided with an oxidant nozzle 221, and the oxidant nozzle 221 is communicated with the oxidant liquid collecting cavity 200; the oxidant nozzle 221 of this embodiment may be of a chamfered type, and may concentrate the injected oxidant to a point in the combustion chamber 300, so as to facilitate combustion.
If the oxidant liquid collecting cavity is sleeved outside the fuel liquid collecting cavity, or the fuel liquid collecting cavity is sleeved outside the oxidant liquid collecting cavity, the diameter of the injection rocket is inevitably increased. In the embodiment of the invention, the fuel liquid collecting cavity 210 and the oxidant liquid collecting cavity 200 are arranged in parallel, the structure is compact, the diameter of the rocket ejector can be smaller, and the influence of air inflow can be reduced.
The bracket 1 comprises a limiting block 110 for blocking the thrust chamber component 3 of the ejector rocket from moving in the direction opposite to the air inflow direction. During test operation, air has a backward force along the direction of the incoming air flow, and the ejector rocket has a force in the direction opposite to the direction of the incoming air flow, so that the ejector rocket can possibly move in the direction opposite to the direction of the incoming air flow, and therefore the limiting block 110 is adopted in the embodiment to play a limiting role, and the ejector rocket is prevented from moving in the direction opposite to the direction of the incoming air flow.
further, the support 1 comprises a front support 10 and a rear support 11, the front support 10 is arranged at a joint of the head assembly 2 and the thrust chamber assembly 3, and the rear support 11 is arranged at a joint of the thrust chamber assembly 3 and the tail nozzle 5 and is fixed through a connecting piece 6. The connecting member 6 of this embodiment is a screw.
the limiting block 110 is arranged on the rear bracket 11.
Further, the thrust chamber assembly 3 includes an equal straight section 30, a throat portion 31 and a thrust chamber outer sleeve 34, the equal straight section 30 is disposed at the front end of the throat portion 31 along the air inflow direction, the thrust chamber outer sleeve 34 is sleeved outside the equal straight section 30 and the throat portion 31, and the thrust chamber outer sleeve extends to the front portion of the equal straight section 30 towards the front end of the air inflow direction.
Further, the thrust chamber housing 34 includes a housing front end 340, a housing body 341, and a housing end 342 in sequence from front to back;
Referring to fig. 1, the outer sleeve end 342 has a larger cross-sectional area than the outer sleeve body 341, the stopper 110 is disposed at the front of the outer sleeve end 342, and the stopper 110 is disposed below the outer sleeve body 341 and spaced apart from the outer sleeve body 341.
The outer sleeve body 341 is arranged on the periphery of the equal straight section 30 and the throat 31, the outer sleeve front end 340, the injection piece 22 and the rear end of the head outer sleeve 20 are connected through a connecting piece 6, and the outer sleeve tail end 342 and the tail nozzle 5 are also connected through the connecting piece 6. The connecting piece 6 of the embodiment of the invention can adopt a screw or a bolt, and the invention does not limit the structural form of the connecting piece 6 as long as the detachable connection can be realized.
Further, a combustion chamber 300 cavity is formed in the equal straight section 30;
the outer sleeve front end 340 is provided with a pressure measuring hole 3400, the pressure measuring hole is located the front side of the combustion chamber 300 and is used for installing a pressure measuring pipeline 33, and the pressure measuring pipeline 33 is communicated with the inside of the outer sleeve front end 340 and is used for realizing the real-time control of the pressure in the combustion chamber 300. The inner wall of the front end 340 of the outer sleeve is also provided with a positioning groove 3401.
further, the fitting 21 includes a thick portion 213, a thin portion 214, and a fitting end 212, a transition portion 215 is disposed between the thick portion 213 and the thin portion 214, and the fitting end 212 is disposed at a front end of the thick portion 213;
The cross-section of the fitting end 212, the thick portion 213, the transition portion 215 and the thin portion 214 is gradually reduced;
The thin part 214 is internally provided with a threaded hole and is matched and connected with the igniter 23 through the threaded hole.
Further, an oxidant port 201 and a fuel port 202 are arranged on the head outer sleeve 20, the oxidant port 201 is used for installing an oxidant nozzle 25, and the oxidant nozzle 25 is communicated with the oxidant liquid collecting cavity 200;
The fuel port 202 is used for installing a fuel nozzle 26, and the fuel nozzle 26 is communicated with a fuel collecting cavity 210;
in this example, the oxidant used was oxygen and the fuel used was kerosene. External oxygen enters the oxidizer manifold 200 through the oxidizer nozzle 25 and external kerosene enters the fuel manifold 210 through the fuel nozzle 26.
the front end of the head housing 20 and the fitting end 212 are connected by a connecting member 6.
Further, a plurality of annular seam ribs 223 are arranged on the inner wall of the injection piece to form a plurality of air film cooling annular seams 22; the rear end of the injection member 22 is provided with a positioning key 224 which is matched with the positioning groove 3401. An embodiment of the present invention has three film cooling circumferential seams 222. A part of the oxidant in the oxidant liquid collecting cavity 200 passes through the gas film cooling annular seam 222 and enters the combustion chamber along the inner wall of the combustion chamber, so that the inner wall of the combustion chamber can be prevented from being burnt, and the inner wall of the combustion chamber is protected.
Further, the fuel nozzle 211 includes an upstream portion 2110, a downstream portion 2111, and a bent portion 2112 provided between the upstream portion 2110 and the downstream portion 2111. The bent portion 2112 is inclined to facilitate the flow of fuel, so that the fuel can sufficiently flow from the upstream portion 2110 to the downstream portion 2111. The downstream portion 2111 of the fuel nozzle 211 is spaced approximately one slot at 0.4-0.6mm intervals to vaporize the liquid fuel and thereby inject the gaseous fuel into the combustion chamber 300.
The specific installation process of the embodiment of the invention is as follows:
The head assembly 2: the front end of the head housing 20 and the fitting member 21 are fixed by screws, and the igniter 23 is connected to the fitting member 21 by screw threads. While the screw fixation is assisted by the igniter sleeve 24. The igniter sleeve 24 of the present embodiment is conventional and functions as a wrench to screw in its igniter 23 and engage the mating member 21. When the igniter 23 is mounted, the igniter sleeve 24 is removed and the igniter sleeve 24 is pulled out.
the thrust chamber assembly 3: the throat 31 is firstly placed in the thrust chamber outer sleeve 34, the equal straight section 30 is placed in the thrust chamber outer sleeve 34, and then the thrust chamber outer sleeve 34 and the tail nozzle 5 are fixed through screws. In the embodiment of the invention, the equal straight section 30 is made of red copper material, the throat part 31 is made of tungsten copper infiltration material, and the thrust chamber outer sleeve 34 is made of stainless steel material.
finally, the injector 22, the rear end of the head jacket 20, and the front end 340 of the jacket are fixed to each other by screws with reference to the injector 22. Thereby completing the installation.
In consideration of the connection sealing property, gaskets 4 are provided on the end surfaces of the injection member 22 engaged with the rear end of the head outer jacket 20, the injection member 22 engaged with the front end 340 of the outer jacket, the front end of the head outer jacket 20 engaged with the engaging member 21, and the engaging member 21 engaged with the igniter 23. The gaskets 4 are all red copper gaskets.
the working process of the embodiment of the invention is as follows:
Oxidant (oxygen is used as oxidant in the examples of the present invention):
The oxidant enters the oxidant liquid collecting cavity 200 through the oxidant nozzle 25, and a part of the oxidant is obliquely sprayed into the combustion chamber 300 through the oxidant nozzle 221 and is mixed through self-impact; another portion of the oxidant is injected into the combustion chamber 300 through the film cooling annulus 222 and adjacent to the walls of the combustion chamber 300, thereby protecting the walls of the combustion chamber 300 from direct contact with the high temperature gas stream and burning.
fuel (kerosene was used as fuel in the examples of the present invention):
the fuel enters the fuel collecting cavity 210 through the fuel nozzle 26, the fuel is atomized through the fuel nozzle 211 and enters the combustion chamber 300, when the igniter 23 (spark plug) is electrified, the spark plug generates high temperature by breaking down air at the rear end of the spark plug, so that the fuel is ignited, and combustion products are sprayed out through the tail nozzle 5 to generate thrust.
in the test ignition process, the pressure gauge is connected to the pressure measuring pipeline 33, and high-temperature fuel gas generated after oxygen and kerosene in the combustion chamber 300 are combusted is transmitted to the pressure gauge through the pressure measuring pipeline 33, so that real-time control of the pressure in the combustion chamber 300 is realized.
It should be noted that, throughout the specification, the terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.
The principles and embodiments of the present invention are explained herein using specific examples, which are presented only to assist in understanding the method and its core concepts of the present invention. It should be noted that there are no more than infinite trial-and-error modes objectively due to the limited character expressions, and it will be apparent to those skilled in the art that various modifications, decorations, or changes may be made without departing from the spirit of the invention or the technical features described above may be combined in a suitable manner; such modifications, variations, combinations, or adaptations of the invention using its spirit and scope, as defined by the claims, may be directed to other uses and embodiments.
Claims (9)
1. The detachable rocket engine is characterized by comprising a bracket (1) and an ejector rocket arranged on the bracket, wherein the ejector rocket comprises a head component (2), a thrust chamber component (3) and a tail nozzle (5); the front end of the thrust chamber component (3) is connected with the head component (2), and the rear end of the thrust chamber component (3) is connected with the tail nozzle (5);
the head assembly (2) comprises an igniter (23), and a fitting piece (21) matched with the igniter (23) is arranged on the outer side of the igniter (23);
A head outer sleeve (20) is arranged on the outer side of the fitting piece (21), the front end of the head outer sleeve (20) and the fitting piece (714) are mutually enclosed to form a fuel collecting cavity (210), a gap between the head outer sleeve (20) and the fitting piece (21) forms a fuel nozzle (211), and the fuel nozzle (211) is communicated with the fuel collecting cavity (210);
The rear end of the head outer sleeve (20) in the air inflow direction is provided with a jetting piece (22), the head outer sleeve (20) and the jetting piece (22) are mutually enclosed to form an oxidant liquid collecting cavity (200), an oxidant nozzle (221) is arranged on the jetting piece (22), and the oxidant nozzle (221) is communicated with the oxidant liquid collecting cavity (200);
The bracket (1) comprises a limiting block (110) for preventing the ejection rocket from moving in the direction opposite to the air inflow direction.
2. A detachable rocket engine according to claim 1, wherein said frame (1) comprises a front frame (10) and a rear frame (11), said front frame (10) being located at the connection between said head assembly (2) and said thrust chamber assembly (3), said rear frame (11) being located at the connection between said thrust chamber assembly (3) and said tail nozzle (5);
The limiting block (110) is arranged on the rear support (11).
3. A collapsible rocket engine according to claim 1 wherein the thrust chamber assembly (3) comprises an equal straight section (30), a throat (31) and a thrust chamber housing (34), the equal straight section (30) being provided at the front end of the throat (31), the thrust chamber housing (34) being provided outside the equal straight section (30) and the throat (31) and extending towards the front end to the front of the equal straight section (30).
4. A collapsible rocket engine according to claim 3 wherein said thrust chamber housing (34) comprises, in order from front to back, a housing front end (340), a housing body (341) and a housing end (342); the outer sleeve body (341) is arranged on the periphery of the equal straight section (30) and the throat (31), and the front end (340) of the outer sleeve, the injection piece (22) and the rear end of the head outer sleeve (20) and the tail end (342) of the outer sleeve are connected with the tail nozzle (5) through connecting pieces (6).
5. a detachable rocket engine according to claim 4, wherein said equal straight section (30) forms a combustion chamber (300);
The outer sleeve front end (340) is provided with a pressure measuring hole (3400) for installing a pressure measuring pipeline (33), and the pressure measuring pipeline (33) is communicated with the inside of the outer sleeve front end (340); the inner wall of the front end (340) of the outer sleeve is also provided with a positioning groove (3401).
6. A collapsible rocket engine according to claim 1, wherein said mating member (21) comprises a thick portion (213), a thin portion (214), and a mating member end portion (212), a transition portion (215) being provided between said thick portion (213) and said thin portion (214), said mating member end portion (212) being provided at a front end of said thick portion (213);
the cross section of the fitting end (212), the thick part (213), the transition part (215) and the thin part (214) is gradually reduced;
the thin part (214) is internally provided with a threaded hole and is matched and connected with the igniter (23) through the threaded hole.
7. A detachable rocket engine according to claim 6 wherein said head casing (20) is provided with an oxidizer port (201) and a fuel port (202), said oxidizer port (201) is used for installing an oxidizer nozzle (25), said oxidizer nozzle (25) is communicated with an oxidizer collecting cavity (200);
the fuel port (202) is used for installing a fuel nozzle (26), and the fuel nozzle (26) is communicated with a fuel collecting cavity (210);
the front end of the head outer sleeve (20) is connected with the end part (212) of the fitting piece through a connecting piece (6).
8. A collapsible rocket engine according to claim 5 wherein said injection member (22) is provided with a plurality of circumferential seam ribs (223) on its inner wall forming a plurality of film cooling circumferential seams (22); and a positioning key (224) matched with the positioning groove (3401) is arranged at the rear end of the injection piece (22).
9. The removable rocket engine of claim 1, wherein the fuel nozzle (211) comprises an upstream portion (2110), a downstream portion (2111), and a bend (2112) provided between the upstream portion (2110) and the downstream portion (2111).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910955006.7A CN110566371A (en) | 2019-10-09 | 2019-10-09 | detachable rocket engine for test |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910955006.7A CN110566371A (en) | 2019-10-09 | 2019-10-09 | detachable rocket engine for test |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN110566371A true CN110566371A (en) | 2019-12-13 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201910955006.7A Pending CN110566371A (en) | 2019-10-09 | 2019-10-09 | detachable rocket engine for test |
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| CN (1) | CN110566371A (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
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| CN112485012A (en) * | 2020-11-13 | 2021-03-12 | 东北大学 | Solid rocket engine experiment table and stress testing method |
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| CN112485012A (en) * | 2020-11-13 | 2021-03-12 | 东北大学 | Solid rocket engine experiment table and stress testing method |
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