CN110594037A - Integrated injection rocket engine assembly and engine thereof - Google Patents
Integrated injection rocket engine assembly and engine thereof Download PDFInfo
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- CN110594037A CN110594037A CN201910955017.5A CN201910955017A CN110594037A CN 110594037 A CN110594037 A CN 110594037A CN 201910955017 A CN201910955017 A CN 201910955017A CN 110594037 A CN110594037 A CN 110594037A
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- oxidant
- fuel
- support plate
- wall
- nozzle
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- 238000002347 injection Methods 0.000 title claims description 12
- 239000007924 injection Substances 0.000 title claims description 12
- 239000000446 fuel Substances 0.000 claims abstract description 111
- 239000007800 oxidant agent Substances 0.000 claims abstract description 111
- 230000001590 oxidative effect Effects 0.000 claims abstract description 101
- 239000007788 liquid Substances 0.000 claims description 12
- 230000013011 mating Effects 0.000 claims description 10
- 238000001816 cooling Methods 0.000 claims description 9
- 230000010354 integration Effects 0.000 claims 1
- 238000002485 combustion reaction Methods 0.000 description 22
- 239000007789 gas Substances 0.000 description 10
- 238000009434 installation Methods 0.000 description 10
- 238000000034 method Methods 0.000 description 10
- 238000007789 sealing Methods 0.000 description 10
- 238000003466 welding Methods 0.000 description 9
- 238000009530 blood pressure measurement Methods 0.000 description 8
- 238000012360 testing method Methods 0.000 description 8
- RYGMFSIKBFXOCR-UHFFFAOYSA-N Copper Chemical compound [Cu] RYGMFSIKBFXOCR-UHFFFAOYSA-N 0.000 description 7
- 229910052802 copper Inorganic materials 0.000 description 7
- 239000010949 copper Substances 0.000 description 7
- 230000008569 process Effects 0.000 description 6
- 238000011144 upstream manufacturing Methods 0.000 description 5
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 description 4
- 239000000463 material Substances 0.000 description 4
- 239000001301 oxygen Substances 0.000 description 4
- 229910052760 oxygen Inorganic materials 0.000 description 4
- 230000009286 beneficial effect Effects 0.000 description 3
- SBYXRAKIOMOBFF-UHFFFAOYSA-N copper tungsten Chemical compound [Cu].[W] SBYXRAKIOMOBFF-UHFFFAOYSA-N 0.000 description 3
- 238000013461 design Methods 0.000 description 3
- 230000008595 infiltration Effects 0.000 description 3
- 238000001764 infiltration Methods 0.000 description 3
- 230000000903 blocking effect Effects 0.000 description 2
- 238000011900 installation process Methods 0.000 description 2
- 238000002955 isolation Methods 0.000 description 2
- 239000003350 kerosene Substances 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 238000004080 punching Methods 0.000 description 2
- 230000007704 transition Effects 0.000 description 2
- 229910000851 Alloy steel Inorganic materials 0.000 description 1
- 238000002679 ablation Methods 0.000 description 1
- 230000006978 adaptation Effects 0.000 description 1
- 238000004891 communication Methods 0.000 description 1
- 239000012141 concentrate Substances 0.000 description 1
- 238000005034 decoration Methods 0.000 description 1
- 230000007123 defense Effects 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 239000002737 fuel gas Substances 0.000 description 1
- 230000014509 gene expression Effects 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 238000002156 mixing Methods 0.000 description 1
- 238000012856 packing Methods 0.000 description 1
- 239000011148 porous material Substances 0.000 description 1
- 238000004088 simulation Methods 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
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
- F02K7/10—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines
- F02K7/18—Composite ram-jet/rocket engines
-
- 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
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
-
- 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
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/52—Injectors
-
- 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
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/60—Constructional parts; Details not otherwise provided for
- F02K9/62—Combustion or thrust chambers
- F02K9/64—Combustion or thrust chambers having cooling arrangements
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 an integrated ejection rocket engine assembly, which comprises a stamping air inlet, three support plates and an ejection rocket, wherein the three support plates are respectively an oxidant support plate, a fuel support plate and a pressure measuring circuit board; the three support plates suspend the ejection rocket in the stamping air inlet. The invention also provides a rocket-based combined cycle engine. The invention can reduce the structure size, thereby reducing the influence on the incoming flow of air.
Description
Technical Field
The invention belongs to the field of rocket-based combined cycle (RBCC) engines, and relates to a small-size integrated injection rocket engine assembly and an engine thereof. The invention relates to an integrated injection Rocket engine, which is mainly applied to an injection Rocket engine using gas oxygen/fuel in a Rocket Based Combined Cycle (RBCC) engine.
Background
With the continuous exploration of the aerospace field in various countries, a rocket-based combined cycle (RBCC) engine is one of the most possible future reciprocating propulsion systems by virtue of the advantages of reusability, low cost and the like. The RBCC engine mainly comprises a stamping air inlet channel, an injection rocket, a stamping afterburning chamber, a tail nozzle and the like; the ejector rocket is an important component of the rocket-based combined cycle engine and is a main source of engine thrust in the ejector mode.
At present, different RBCC configuration schemes are designed for different task requirements in each country: the Strutjet engine and the ISTAR engine adopt rectangular inner flow channels, wherein the ejector rocket is arranged in the center of the flow channel in the middle of the combustion chamber; the Japanese JAXA-RBCC engine adopts an inner flow passage with a rectangular cross section, wherein the injection rocket engine is arranged on the top wall of an outlet of the isolation section; the national institute of western's safety space flight power proposes a patent of RBCC engine internal flow channel with central layout in front of rocket, the air inlet channel is designed as curved arc; the national defense science and technology university of the people's liberation army of China proposes a central layout ejector rocket applied to a rocket-based combined cycle engine, and the central layout ejector rocket adopts a plug type tail nozzle to improve the bottom pressure of the lower ejector rocket tail nozzle in a stamping mode, reduce bottom resistance and reduce total pressure loss.
In the different RBCC configuration schemes, the scheme that the ejection rocket is arranged on the top wall of the outlet of the isolation section has poor mixing between the primary flow and the secondary flow; the scheme that the ejector rocket is arranged in the center of the combustion chamber does not consider the influence of the ejector rocket arranged in the stamping air inlet channel on the incoming air flow. Meanwhile, the above schemes do not provide module installation from a test level, so that the idea of changing the configuration at any time for testing is facilitated.
Disclosure of Invention
The invention aims to provide an integrated ejector rocket engine assembly, which can reduce the structural size, thereby reducing the influence on the incoming flow of air.
The technical scheme adopted by the invention is as follows:
the invention provides an integrated ejection rocket engine assembly, which comprises a stamping air inlet, three support plates and an ejection rocket, wherein the three support plates are respectively an oxidant support plate, a fuel support plate and a pressure measuring circuit board; the three support plates suspend the ejection rocket in the stamping air inlet channel;
the ejector rocket comprises a head and a thrust chamber, the head comprises an igniter, a fitting piece matched with the igniter is arranged on the outer side of the igniter, a head clamping wall is arranged on the outer side of the fitting piece, the front end of the head clamping wall in the air inflow direction and the fitting piece are mutually enclosed to form a fuel liquid collecting cavity, a fuel nozzle is formed between the head clamping wall and the fitting piece, and the fuel nozzle is communicated with the fuel liquid collecting cavity;
the rear end of the head part clamping wall along the air inflow direction is provided with a nozzle component, the head part clamping wall and the nozzle component mutually enclose to form an oxidant liquid collecting cavity, the nozzle component is provided with an oxidant nozzle, and the oxidant nozzle is communicated with the oxidant liquid collecting cavity; the outer end of the head part clamping wall is provided with a head part outer sleeve, and the matching piece and the nozzle component are arranged in the head part outer sleeve;
the oxidant support plate comprises an oxidant support plate body, and an oxidant pipeline is arranged in the oxidant support plate body;
the fuel rail includes a fuel rail body, the fuel rail body;
the pressure measuring circuit board comprises a pressure measuring circuit board body, and a pressure measuring pipeline and a circuit pipeline are arranged in the pressure measuring circuit board body.
Preferably, the thrust chamber comprises a thrust chamber front wall, an equal straight section, a throat portion and a tail nozzle outer sleeve member, wherein the equal straight section is arranged between the thrust chamber front wall and the throat portion;
the tail nozzle outer sleeve component comprises a tail nozzle and a thrust chamber outer sleeve, the tail nozzle is arranged at the rear end of the throat part, the thrust chamber outer sleeve is arranged at the outer end of the equal straight section, and the tail nozzle and the thrust chamber outer sleeve are integrally formed.
Preferably, the mating member includes a main body portion, an extension portion extending along the main body portion toward the front end, and a threaded portion extending along the main body portion toward the rear end;
the extension part is provided with a matching piece positioning pin, the outer surface of the main body part is provided with a matching piece rib, and the threaded part is internally provided with a threaded hole;
the igniter is matched with the threaded hole.
Preferably, the head part double-wall comprises a double-wall body, a fuel hole and an oxidant hole are arranged on the double-wall body, and the fuel hole and the oxidant hole are respectively communicated with a fuel pipeline and an oxidant pipeline;
the front end of the wall clamping body is provided with a wall clamping positioning hole matched with the fitting piece positioning pin, and the rear end of the wall clamping body is provided with a wall clamping positioning groove.
Preferably, the nozzle component comprises a nozzle component body and a convex part extending towards the rear end along the nozzle component body, a plurality of annular seam ribs are arranged on the inner wall of the nozzle component body to form a plurality of air film cooling annular seams, a nozzle positioning strip matched with the wall clamping positioning groove is arranged on the nozzle component body, and a thrust chamber positioning pin is arranged on the convex part.
Preferably, a fixing piece and a front wall positioning pin are arranged on the front end face of the front wall of the thrust chamber, an installation groove and a pressure measuring hole are formed in the inner wall of the front wall of the thrust chamber, and the installation groove is matched with the convex part;
and a front wall positioning groove matched with the thrust chamber positioning pin is arranged at the mounting groove.
Preferably, the stamping air inlet is provided with an oxidant branch plate hole, a fuel branch plate hole and a pressure measuring circuit plate hole.
The oxidant support plate is provided with oxidant support plate bosses matched with the oxidant support plate holes, the fuel support plate is provided with oxidant support plate bosses matched with the fuel support plate holes, and the pressure measuring circuit board is provided with pressure measuring circuit board bosses matched with the pressure measuring circuit board holes.
Preferably, the oxidant support plate boss is provided with an oxidant joint, and the fuel support plate boss is provided with a fuel joint;
the oxidant joint penetrates through the oxidant support plate boss and is communicated with the oxidant pipeline, and the fuel joint penetrates through the fuel support plate boss and is communicated with the fuel pipeline;
and the pressure measuring pipeline and the circuit pipeline penetrate through the boss of the pressure measuring circuit board.
Preferably, the ram air inlet forms a gap with the oxidant support plate, the fuel support plate, and the pressure sensing circuit board.
One aspect of the invention provides a rocket-based combined cycle engine comprising the engine assembly described above.
The invention has the beneficial effects that:
1. the invention provides an integrated ejector rocket engine assembly which can reduce the structural size, thereby reducing the influence on the incoming flow of air.
2. In the embodiment of the invention, the support plate, the fuel conveying path and the oxidant conveying path are integrally formed, and compared with a separate design method of the support plate and the pipeline, the thickness of the support plate can be reduced, so that the blocking effect of the support plate on the incoming air flow is reduced; parts of the ejection rocket are connected by welding, so that a flange plate in bolt connection is avoided, the outer diameter of the ejection rocket is reduced, and the air blocking effect of the ejection rocket is reduced; the ejection capacity of the ejection rocket is improved on the whole.
3. In the embodiment of the invention, the equal straight section adopts red copper, the throat part adopts a tungsten copper infiltration passive thermal protection mode, the simulation result shows that under the working condition of large flow, the throat part temperature is 1950K, which is far less than the temperature resistance capability (more than 3000K) of the tungsten copper infiltration, and the equal straight section can normally work for 20 s; considering that red copper is not beneficial to welding, the front end (front wall of the thrust chamber) of the thrust chamber is made of high-temperature-resistant alloy steel which is mainly positioned at the front end of an oxidant impact point, and combustion flame is ensured to be positioned at the red copper material part.
4. In the embodiment of the invention, the nozzle component is designed into a direct current self-impact type, the spark plug igniter is arranged in the center of the head, and compared with a centrifugal nozzle, the scheme can reduce the outer diameter of the head of the injection rocket and is more beneficial to designing a small-size rocket engine.
5. 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 circular seams (gas film cooling circular seams), and 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.
6. In the embodiment of the invention, the rib of the matching piece is specially designed to ensure the accurate size of the fuel nozzle in the installation process by considering that the part is not uniformly heated in the welding process to cause the shaking of the part and cause the blockage of the downstream part of the fuel nozzle.
7. In the embodiment of the invention, in order to ensure the accurate and quick installation of each part, the main body part of each part is provided with the positioning pin, so that the accurate positioning and quick installation in the installation process can be ensured.
8. In the embodiment of the invention, the stamping air inlet, the three support plates and the ejection rocket form an integral module through welding to form the integrated ejection rocket engine assembly, and the engine assembly is connected with the front runner and the rear runner (not shown in the attached drawing) by using stamping wall flanges, so that test parts can be conveniently changed for testing according to different test working conditions.
Drawings
FIG. 1 is a schematic perspective view of an integrated rocket motor assembly according to an embodiment of the present invention;
FIG. 2 is a schematic cross-sectional view of FIG. 1 taken along the axial plane of the oxidant plate;
FIG. 3 is an enlarged view of a portion of FIG. 2 at A;
FIG. 4 is a schematic cross-sectional view of FIG. 1 taken along the axial plane of the fuel rail;
FIG. 5 is a schematic cross-sectional view of FIG. 1 taken along the axial plane of the pressure measurement circuit board;
FIG. 6 is a schematic view of a nozzle member according to an embodiment of the present invention;
FIG. 7 is a schematic view of a head clamp wall structure according to an embodiment of the present invention;
FIG. 8 is a schematic view of a fitting configuration in an embodiment of the present invention;
FIG. 9 is a schematic view of a thrust chamber according to an embodiment of the present invention;
FIG. 10 is a schematic view of a ram air scoop configuration according to an embodiment of the present invention;
FIG. 11 is a schematic view of the fitting, head wall, and nozzle member engagement configuration of an embodiment of the present invention;
fig. 12 is a schematic view of a head structure in an embodiment of the present invention.
1-punching an air inlet channel; 11-oxidant manifold holes; 12-fuel manifold holes; 13-pressure measuring circuit board hole; 14, punching a wall surface flange;
2-an oxidant support plate; 20-an oxidant plate body; 21-oxidant line; 22-oxidant support plate bosses;
3-a fuel support plate; 30-a fuel carrier plate body; 31-fuel line; 32-fuel strut bosses;
4, a pressure measuring circuit board; 40-a pressure measuring circuit board body; 41-pressure measuring pipeline; 42-pressure measuring circuit board boss; 43-line circuit;
5-oxidant linker; 6-fuel joint; 8-clearance;
7-ejecting the rocket; 71-head; 711-head coat; 712-igniter tooling; 713-an igniter; 714-mating parts; 7140 — extension; 71401 — fitting dowel pin; 7141-main body part; 71411 — mating member ribs; 7142-threaded part; 71421-threaded hole; 7144-fuel collection chamber; 7145-fuel nozzle; 71450 — upstream portion; 71451 — downstream portion; 71452-transition;
715-head part clamping wall; 7150 — a double-walled body; 7151-positioning holes in the clamping wall; 7152-fuel holes; 7153-oxidant pores; 7154 wall locating slot; 7155-oxidizer liquid collection chamber;
716-a nozzle member; 7160 — oxidizer nozzle; 7161-nozzle member body; 7162-nozzle positioning strip; 7163 cooling the circular seam with gas film; 7164-circumferential seam ribs; 7165-convex part; 7166-thrust chamber locating pin; 717-igniter sealing ring; 718-a housing mating surface;
72-a thrust chamber; 721-thrust chamber front wall; 7211-a fixing member; 7212-front wall locating pin; 7213-front wall positioning groove; 7214-pressure measuring hole; 7215-mounting grooves; 722-equal straight section; 7221-a combustion chamber; 723-throat; 724-jet nozzle jacket component; 7240 — a tail pipe; 7241-thrust chamber jacket; 725-thrust chamber sealing ring.
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.
Example 1
Referring to fig. 1-12, embodiment 1 of the present invention provides an integrated ejector rocket engine assembly comprising: the device comprises a stamping air inlet 1, three support plates and an injection rocket 7, wherein the three support plates are respectively an oxidant support plate 2, a fuel support plate 3 and a pressure measuring circuit board 4; the three support plates suspend the ejection rocket 7 in the stamping air inlet 1; in the embodiment of the invention, the three support plates suspend the ejection rocket 7 at the central position in the stamping air inlet 1. One ends of the oxidant support plate 2, the fuel support plate 3 and the pressure measuring circuit board 4 are connected with the ejection rocket 7, and the other ends of the oxidant support plate 2, the fuel support plate 3 and the pressure measuring circuit board 4 are connected with the stamping air inlet 1;
the embodiment of the invention adopts a central layout mode, and the three support plates hang the ejection rocket 7 on the stamping air inlet 1. The three support plates are designed to be as thin as possible and have the thickness of about 20 mm. The three support plates are arranged, so that the support performance is good, the turbulent flow can be reduced, and optionally, the included angle between every two adjacent support plates of the three support plates is 120 degrees.
According to the embodiment of the invention, the ejection rocket and the three support plates are welded on the wall surface of the stamping air inlet 1 to form a whole, and the stamping air inlet 1 is connected with other parts through the stamping wall surface flange 14, so that test pieces can be conveniently replaced according to different test conditions; the design of the ejection rocket hung at the center of the stamping air inlet 1 also adopts a welding mode to be integrally formed, so that the structural size is reduced, and the influence of the ejection rocket and the support plate on the incoming air flow is reduced.
The jet rocket 7 comprises a head 71 and a thrust chamber 72, the head 71 comprises an igniter 713, a fitting 714 matched with the igniter 713 is arranged outside the igniter 713, a head clamping wall 715 is arranged outside the fitting 714, the front end of the head clamping wall 715 in the air inflow direction and the fitting 714 mutually enclose to form a fuel collecting cavity 7144, a fuel nozzle 7145 is formed between the head clamping wall 715 and the fitting 714, and the fuel nozzle 7145 is communicated with the fuel collecting cavity 7144;
fuel nozzle 7145 includes an upstream portion 71450, a downstream portion 71451, and a transition portion 71452 connected between upstream portion 71450 and downstream portion 71451. the transition portion 71452 is angled to facilitate the flow of fuel from upstream portion 71450 substantially into downstream portion 71451.
The downstream portion 71451 of fuel nozzle 7145 has a circumferential gap of 0.4-0.6mm, approximately one slot, to vaporize liquid fuel and thereby inject gaseous fuel into the combustion chamber 7221.
A nozzle component 716 is arranged at the rear end of the head clamping wall 715 along the air inflow direction, the head clamping wall 715 and the nozzle component 716 mutually enclose to form an oxidant collecting cavity 7155, an oxidant nozzle 7160 is arranged on the nozzle component 716, and the oxidant nozzle 7160 is communicated with the oxidant collecting cavity 7155; the oxidant nozzle 7160 is formed in a chamfered shape, and can concentrate the sprayed oxidant to one point of the combustion chamber 7221 to facilitate combustion.
In the embodiment of the invention, the fuel liquid collecting cavity 7144 and the oxidant liquid collecting cavity 7155 are arranged in parallel, the structure is compact, the diameter of the rocket for ejecting is smaller, the gap 8 between the stamping air inlet 1 and the oxidant support plate 2, between the fuel support plate 3 and between the pressure measuring circuit boards 4 is larger, and more air can enter the combustion of the next stage. And the ejector rocket 7 has small size, and can reduce the influence of incoming air flow.
The outer end of the head clamping wall 715 is provided with a head outer sleeve 711, and the fitting 714 and the nozzle member 716 are arranged in the head outer sleeve 711;
the oxidant support plate 2 comprises an oxidant support plate body 20, and an oxidant pipeline 21 is arranged in the oxidant support plate body 20;
the fuel plate 3 comprises a fuel plate body 30, and a fuel pipeline 31 is arranged in the fuel plate body 30;
the pressure measurement wiring board 4 includes a pressure measurement wiring board body 40, and a pressure measurement line 41 and a line 43 are provided in the pressure measurement wiring board body 40. The line pipe 43 is used for placing an electric wire or the like (electric wire of an igniter) required for the embodiment of the present invention so as to communicate with an external power source through the line pipe 43.
Further, the thrust chamber 72 includes a thrust chamber front wall 721, an equal straight section 722, a throat 723, and a jet nozzle outer sleeve member 724, the equal straight section 722 being disposed between the thrust chamber front wall 721 and the throat 723; a combustion chamber 7221 cavity is formed within the straight section 722.
The jet nozzle outer sleeve member 724 comprises a jet nozzle 7240 and a thrust chamber outer sleeve 7241, the jet nozzle 7240 is arranged at the rear end of the throat portion 723, the thrust chamber outer sleeve 7241 is arranged at the outer end of the equal straight section 722, and the jet nozzle 7240 and the thrust chamber outer sleeve 7241 are integrally formed.
The traditional connection mode is that a flange is added at the position of the tail nozzle 7240, and then the thrust chamber outer sleeve is connected with the tail nozzle 7240 through the flange, and the design mode tends to enlarge the diameter of the ejector rocket. The exhaust nozzle 7240 and the thrust chamber outer sleeve 7241 are integrally formed, so that the diameter increase caused by the traditional flange connection is avoided, and the influence of air inflow is reduced.
Referring to fig. 8, the fitting 714 includes a main body portion 7141, an extension portion 7140 extending along the main body portion 7141 toward the front end, and a threaded portion 7142 extending along the main body portion 7141 toward the rear end;
the extension part 7140 is provided with a fitting piece positioning pin 71401, the outer surface of the main body part is provided with a fitting piece rib 71411, and a threaded part is internally provided with a threaded hole 71421; the embodiment of the present invention is provided with four fitting member positioning pins 71401 and three fitting member ribs 71411.
The igniter 713 is threadedly engaged with the threaded hole 71421.
Referring to fig. 7, the head-part chuck wall 715 includes a chuck wall body 7150, the chuck wall body 7150 is provided with a fuel hole 7152 and an oxidant hole 7153, the fuel hole 7152 and the oxidant hole 7153 are respectively communicated with the fuel pipeline 31 and the oxidant pipeline 21; specifically, the fuel orifice 7152 communicates with the fuel line 31 and the oxidizer orifice 7153 communicates with the oxidizer line 21. The front end of the wall clamping body 7150 is provided with a wall clamping positioning hole 7151 matched with the matching piece positioning pin 71401, and the rear end of the wall clamping body 7150 is provided with a wall clamping positioning groove 7154.
Further, the ram inlet 1 is provided with an oxidant manifold hole 11, a fuel manifold hole 12, and a pressure measurement line plate hole 13.
The oxidant support plate 2 is provided with oxidant support plate bosses 22 which are matched with the oxidant support plate holes 11, the fuel support plate is provided with oxidant support plate bosses 32 which are matched with the fuel support plate holes 12, and the pressure measuring circuit board 4 is provided with pressure measuring circuit board bosses 42 which are matched with the pressure measuring circuit board holes 13.
The oxidant support plate boss 22 is provided with an oxidant joint 5, and the fuel support plate boss 32 is provided with a fuel joint 6;
the oxidant joint 5 penetrates through the oxidant support plate boss 22 and is communicated with the oxidant pipeline 21, and the fuel joint 6 penetrates through the fuel support plate boss 32 and is communicated with the fuel pipeline 31;
the pressure measurement line 41 and the line 43 pass through the pressure measurement board boss 42.
In the embodiment of the invention, external fuel enters the fuel collecting cavity 7144 through the fuel joint 6, the fuel support plate boss 32, the fuel pipeline 31 and the fuel hole 7152;
external oxygen enters the oxidant liquid collecting cavity 7155 through the oxidant connector 5, the oxidant support plate boss 22, the oxidant pipeline 21 and the oxidant hole 7153.
Referring to fig. 6, the nozzle member includes a nozzle member body 7161 and a protrusion 7165 extending along the nozzle member body 7161 toward the rear end, the inner wall of the nozzle member body 7161 is provided with a plurality of circumferential seam ribs 7164, the plurality of circumferential seam ribs 7164 form a plurality of film cooling circumferential seams 7163, and three film cooling circumferential seams 7163 are formed according to the embodiment of the present invention. A part of the oxidant in the oxidant liquid collecting cavity 7155 passes through the gas film cooling circular seam 7163 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.
The nozzle member body 7161 is provided with a nozzle positioning strip 7162 which is matched with the clamping wall positioning groove 7154, and the convex portion 7165 is provided with a thrust chamber positioning pin 7166.
Further, the thrust chamber front wall 721 is provided at the front end face with a fixing member 7211 and a front wall positioning pin 7212, and the head housing 711 is provided with a head positioning groove (not shown in the drawings) that is fitted with the front wall positioning pin 7212.
The inner wall of the front wall 721 of the thrust chamber is provided with a mounting groove 7215 and a pressure measuring hole 7214, and the mounting groove 7215 is matched with the convex portion 7165; the fixing member 7211 may be a screw.
A front wall positioning groove 7213 is provided at the mounting groove 7215 to mate with the thrust chamber positioning pin 7166.
Further, a gap 8 is formed between the ram air inlet 1 and the oxidizer support plate 2, the fuel support plate 3, and the pressure measuring board 4. An air inflow passes through this gap 8. The direction of the incoming air flow is indicated by the arrows in fig. 1.
The embodiment of the invention adopts a modularized installation mode, and the specific installation is as follows:
ejector rocket head 71: the nozzle component 716 is taken as a reference, matched and fixed through the nozzle positioning strip 7162 and the head clamping wall 715, and then welded along a matching surface welding groove; the fitting 714 is fixed with the positioning slot 7151 by the fitting positioning pin 71401, and is welded along the fitting surface. Because the diameter of the fuel nozzle 7145 is small, considering that the structure is unevenly heated in the welding process to generate sloshing to block the fuel nozzle 7145, the matching piece rib 71411 is arranged at the upstream of the fuel nozzle 7145 for positioning, and the fuel nozzle 7145 is prevented from being blocked due to the sloshing. The igniter 713 is screwed to the mating member 714, and an igniter packing 717 is fitted to the mating end surface to seal the igniter in consideration of the sealing property of the connection. Meanwhile, the igniter tool 712 is used for assisting in thread fixing. Igniter tooling 712 of the present embodiment is conventional and functions as a wrench to screw in its igniter 713 and engage mating member 714.
When the igniter 713 is installed, the igniter tool 712 is removed; the assembled components are then inserted entirely into the head housing 711, positioned by housing mating surfaces 718 (the interface between the head housing 711 and mating member 714), and then welded by the housing mating surfaces 718 to complete the assembly of the head.
Ejection rocket thrust chamber 72: with the jet nozzle outer sleeve member 724 as a reference, the throat 723 is firstly put into the jet nozzle outer sleeve member 724; then the equal straight section 722 is put into the tail nozzle outer sleeve member 724 and is positioned by a positioning groove; the thrust chamber seal 725 is then placed into the gap of the jet nozzle outer housing member 724 and finally the thrust chamber front wall 721 is inserted into the jet nozzle outer housing member 724. In the embodiment of the invention, the straight section 722 is made of red copper material, the throat 723 is made of tungsten copper infiltration material, the front wall 721 of the thrust chamber is made of stainless steel material, and the sealing ring 717 of the igniter and the sealing ring 725 of the thrust chamber are both made of red copper sealing rings.
In order to ensure the overall sealing performance of the thrust chamber, a screw-fixed connection is used between the thrust chamber front wall 721 and the jet nozzle outer sleeve member 724, and the soft equal straight section 722 red copper wall surface and the red copper thrust chamber sealing ring 725 are pressed by tightening the screw to realize sealing. And the installation of the integral thrust chamber is completed.
And finally, the whole thrust chamber is matched with the head 71 through a front wall positioning pin 7212, a front wall positioning groove 7213 and then welded from the matching surface, so that the body of the injection rocket 7 is assembled.
Integral installation: placing the body part of the ejection rocket 7 in the middle of the stamping air inlet 1, manually dragging, then respectively embedding the oxidant support plate 2, the fuel support plate 3 and the pressure measuring circuit board 4 into the oxidant support plate hole 11, the fuel support plate hole 12 and the pressure measuring circuit board hole 13 of the stamping air inlet 1, enabling the oxidant support plate, the fuel support plate and the pressure measuring circuit board holes to be fixed in one-to-one correspondence with the matching surfaces of the ejection rocket 7, and then welding along each matching surface to form sealing; and finally, respectively welding the oxidant joint 5 and the fuel joint 6 to the corresponding support plates, thereby completing the installation of the ejection rocket.
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 pipeline 21 through the oxidant joint 5 and enters the head oxidant collecting cavity 7155 through the oxidant support plate pipeline 21, and a part of the oxidant is obliquely sprayed into the combustion chamber 7221 through the oxidant nozzle 7160 and is mixed by self-impact; another portion of the oxidant is injected into the combustion chamber 7221 through the film-cooled annular gap 7163 and is proximate to the walls of the combustion chamber 7221, thereby protecting the walls of the combustion chamber 7221 from being burned by direct contact with the high temperature gas stream.
Fuel (kerosene was used as fuel in the examples of the present invention):
fuel enters the fuel pipeline 31 through the fuel joint 6, enters the head fuel collecting cavity 7144 through the fuel pipeline 31, is atomized through the fuel nozzle 7145 and enters the combustion chamber 7221, when the igniter 713 (a 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 7240 to generate thrust.
In the test ignition process, the pressure gauge is connected to the pressure measuring circuit board 4, high-temperature fuel gas generated after oxygen and kerosene in the combustion chamber 7221 are combusted flows to a small hole (not shown in the figure) of the thrust chamber outer sleeve 7241 through the pressure measuring hole 7214 of the thrust chamber front wall 721, and finally is transmitted to the pressure gauge through the pressure measuring pipeline 41, so that the real-time control of the pressure in the combustion chamber 7221 is realized.
In an embodiment of the present invention, the pressure tap 7214, a small hole (not shown) of the thrust chamber housing 7241, and the pressure line 41 are in communication with each other.
Example 2
Embodiment 2 of the present invention is a rocket-based combined cycle engine including the engine assembly described in embodiment 1, which is connected to a flow passage of a previous stage and a flow passage of a subsequent stage, respectively, by a stamped wall flange 14.
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 (10)
1. Rocket engine subassembly is penetrated in integration injection, its characterized in that includes: the device comprises a stamping air inlet (1), three support plates and an injection rocket (7), wherein the three support plates are respectively an oxidant support plate (2), a fuel support plate (3) and a pressure measuring circuit board (4); the three support plates suspend the ejection rocket (7) in the stamping air inlet (1);
the rocket ejector (7) comprises a head (71) and a thrust chamber (72), the head (71) comprises an igniter (713), a fitting piece (714) matched with the igniter (713) is arranged on the outer side of the igniter (713), a head clamping wall (715) is arranged on the outer side of the fitting piece (714), the front end of the head clamping wall (715) in the air inflow direction and the fitting piece (714) are mutually enclosed to form a fuel collecting cavity (7144), a fuel nozzle (7145) is formed between the head clamping wall (715) and the fitting piece (714), and the fuel nozzle (7145) is communicated with the fuel collecting cavity (7144);
the rear end of the head clamping wall (715) in the air inflow direction is provided with a nozzle component (716), the head clamping wall (715) and the nozzle component (716) mutually enclose to form an oxidant liquid collecting cavity (7155), the nozzle component (716) is provided with an oxidant nozzle (7160), and the oxidant nozzle (7160) is communicated with the oxidant liquid collecting cavity (7155); a head outer sleeve (711) is arranged at the outer end of the head clamping wall (715), and the matching piece (714) and the nozzle component (716) are arranged in the head outer sleeve (711);
the oxidant support plate (2) comprises an oxidant support plate body (20), and an oxidant pipeline (21) is arranged in the oxidant support plate body (20);
the fuel support plate (3) comprises a fuel support plate body (30), and a fuel pipeline (31) is arranged in the fuel support plate body (30);
the pressure measuring circuit board (4) comprises a pressure measuring circuit board body (40), and a pressure measuring pipeline (41) and a circuit pipeline (43) are arranged in the pressure measuring circuit board body (40).
2. The engine assembly of claim 1, characterized in that the thrust chamber (72) includes a thrust chamber front wall (721), a straight section (722), a throat (723), and a jet nozzle outer sleeve member (724), the straight section (722) being disposed between the thrust chamber front wall (721) and the throat (723);
jet nozzle overcoat component (724) includes jet nozzle (7240) and thrust chamber overcoat (7241), throat (723) rear end is located in jet nozzle (7240), thrust chamber overcoat (7241) are located straight section (722) outer end, and jet nozzle (7240) and thrust chamber overcoat (7241) set up for integrated into one piece.
3. The engine assembly of claim 1, characterized in that the mating element (714) comprises a main body portion (7141), an extension portion (7140) extending along a forward end of the main body portion (7141), and a threaded portion (7142) extending along a rearward end of the main body portion (7141);
a fitting piece positioning pin (71401) is arranged on the extension part (7140), a fitting piece rib (71411) is arranged on the outer surface of the main body part, and a threaded hole (71421) is formed in the threaded part;
the igniter (713) is interfitted with the threaded hole (71421).
4. The engine assembly according to claim 3, characterized in that the head part double wall (715) comprises a double wall body (7150), a fuel hole (7152) and an oxidant hole (7153) are arranged on the double wall body (7150), and the fuel hole (7152) and the oxidant hole (7153) are respectively communicated with the fuel pipeline (31) and the oxidant pipeline (21);
the front end of the wall clamping body (7150) is provided with a wall clamping positioning hole (7151) matched with the fitting piece positioning pin (71401), and the rear end of the wall clamping body (7150) is provided with a wall clamping positioning groove (7154).
5. The engine assembly of claim 4, characterized in that the nozzle component comprises a nozzle component body (7161) and a protrusion (7165) extending along the nozzle component body (7161) towards the rear end, the inner wall of the nozzle component body (7161) is provided with a plurality of circumferential seam ribs (7164) forming a plurality of film cooling circumferential seams (7163), the nozzle component body (7161) is provided with a nozzle positioning strip (7162) matching with the chuck wall positioning groove (7154), and the protrusion (7165) is provided with a thrust chamber positioning pin (7166).
6. The engine assembly according to claim 5, characterized in that the front end face of the thrust chamber front wall (721) is provided with a fixing piece (7211) and a front wall positioning pin (7212), the inner wall of the thrust chamber front wall (721) is provided with a mounting groove (7215) and a pressure measuring hole (7214), and the mounting groove (7215) is matched with the convex part (7165);
and a front wall positioning groove (7213) matched with the thrust chamber positioning pin (7166) is arranged at the mounting groove (7215).
7. An engine assembly according to claim 1, characterized in that the ram inlet (1) is provided with an oxidant manifold hole (11), a fuel manifold hole (12) and a pressure rail hole (13);
the oxidant support plate (2) is provided with oxidant support plate bosses (22) matched with the oxidant support plate holes (11), the fuel support plate is provided with oxidant support plate bosses (32) matched with the fuel support plate holes (12), and the pressure measuring circuit board is provided with pressure measuring circuit board bosses (42) matched with pressure measuring circuit board holes (13).
8. The engine assembly of claim 7, characterized in that the oxidant strut boss (22) has an oxidant fitting (5) disposed thereon and the fuel strut boss (32) has a fuel fitting (6) disposed thereon;
the oxidant joint (5) penetrates through the oxidant support plate boss (22) and is communicated with the oxidant pipeline (21), and the fuel joint (6) penetrates through the fuel support plate boss (32) and is communicated with the fuel pipeline (31);
the pressure measuring pipeline (41) and the line pipeline (43) penetrate through the pressure measuring circuit board boss (42).
9. An engine assembly according to claim 1, characterized in that the ram air inlet (1) forms a gap (8) with the oxidant plate (2), the fuel plate (3) and the pressure sensing circuit board (4).
10. A rocket-based combined cycle engine comprising the engine assembly of any of claims 1-9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910955017.5A CN110594037A (en) | 2019-10-09 | 2019-10-09 | Integrated injection rocket engine assembly and engine thereof |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
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| CN201910955017.5A CN110594037A (en) | 2019-10-09 | 2019-10-09 | Integrated injection rocket engine assembly and engine thereof |
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| CN110594037A true CN110594037A (en) | 2019-12-20 |
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Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| RU2181849C1 (en) * | 2001-02-12 | 2002-04-27 | Орловский государственный технический университет | Ramjet-ejector rocket carrier |
| CN104295406A (en) * | 2014-05-26 | 2015-01-21 | 清华大学 | Rocket stamping combination engine with annular injection structure |
| CN108894893A (en) * | 2018-06-26 | 2018-11-27 | 西北工业大学 | Rocket punching press combined engine Film cooling Rocket ejector motor power room |
| CN109139298A (en) * | 2018-08-14 | 2019-01-04 | 北京控制工程研究所 | A kind of space double elements precise tracking based on two-way single injector centrifugation ejector filler |
| CN109812352A (en) * | 2019-03-06 | 2019-05-28 | 中南大学 | Rocket ejector and its thermal protection structure and thermal protection method |
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2019
- 2019-10-09 CN CN201910955017.5A patent/CN110594037A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| RU2181849C1 (en) * | 2001-02-12 | 2002-04-27 | Орловский государственный технический университет | Ramjet-ejector rocket carrier |
| CN104295406A (en) * | 2014-05-26 | 2015-01-21 | 清华大学 | Rocket stamping combination engine with annular injection structure |
| CN108894893A (en) * | 2018-06-26 | 2018-11-27 | 西北工业大学 | Rocket punching press combined engine Film cooling Rocket ejector motor power room |
| CN109139298A (en) * | 2018-08-14 | 2019-01-04 | 北京控制工程研究所 | A kind of space double elements precise tracking based on two-way single injector centrifugation ejector filler |
| CN109812352A (en) * | 2019-03-06 | 2019-05-28 | 中南大学 | Rocket ejector and its thermal protection structure and thermal protection method |
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Application publication date: 20191220 |