CN109404165B - The continuous pinking rocket base engine and aircraft of thruster vector control - Google Patents
The continuous pinking rocket base engine and aircraft of thruster vector control Download PDFInfo
- Publication number
- CN109404165B CN109404165B CN201811597713.5A CN201811597713A CN109404165B CN 109404165 B CN109404165 B CN 109404165B CN 201811597713 A CN201811597713 A CN 201811597713A CN 109404165 B CN109404165 B CN 109404165B
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- Prior art keywords
- fuel
- chamber
- oxidant
- air collecting
- end cap
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Classifications
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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
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous 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
-
- 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
-
- 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
- F02K9/80—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by thrust or thrust vector control
Abstract
The continuous pinking rocket base engine and aircraft of a kind of thruster vector control provided by the invention, are related to technical field of aerospace, comprising: inner core-body；Outer barrel have axially through hollow cavity, for being set in the outside of inner core-body；The first gap is formed between the outer wall of inner core-body and the inner wall of outer barrel；At least one preknock pipe structure is provided on outer barrel；The head end of inner core-body and outer barrel is arranged in air collecting chamber end cap, and connect with inner core-body and outer barrel；The side wall of air collecting chamber end cap is provided with fuel inlet, and the inside of air collecting chamber end cap has a fuel cavity of annular, and annular fuel cavity it is circumferentially spaced be at least two fuel unit chambers；The end face of air collecting chamber end cap is provided with oxidant inlet, and the inside of air collecting chamber end cap has oxidant cavity, and it is at least two oxidants unit chamber corresponding with fuel unit chamber that oxidant cavity is circumferentially spaced.
Description
Technical field
The present invention relates to technical field of aerospace, start more particularly, to a kind of continuous pinking rocket base of thruster vector control
Machine and aircraft.
Background technique
Carrier rocket (rocket launcher) is used for artificial earth satellite, manned spaceship, spaceport or interplanetary spy
It surveys device etc. and is sent into planned orbit.Final stage has instrument room, built-in guidance and control system, telemetry system and launching site security system.
In the design and manufacture of carrier rocket, rationally adjust to motor power is to realize carrier rocket flight environment
The necessary means of the active control abilities such as control, trajectory optimization, not only liquid engine will have thrust regulating power, Gu
Body engine will also realize thrust control by reasonably designing.
But in the prior art in the carry-on all kinds of engines such as carrier rocket, propulsive performance to be poor, structure is multiple
Miscellaneous, to meet thrust adjustment, the increased weight of institute is big, changes cooling caused by nozzle structure and sealing difficulty increases.
Summary of the invention
The purpose of the present invention is to provide the continuous pinking rocket base engine and aircraft of a kind of thruster vector control,
The technical problems such as propulsive performance to solve engine existing in the prior art is poor, structure is complicated.
A kind of continuous pinking rocket base engine of thruster vector control provided by the invention, comprising:
Inner core-body；
Outer barrel, the outer barrel have axially through hollow cavity, for being set in the outside of the inner core-body；
The first gap is formed between the outer wall of the inner core-body and the inner wall of the outer barrel, which is used for structure
Circularize combustion chamber；At least one preknock pipe structure is provided on the outer barrel, for lighting a fire to the toroidal combustion chamber；
Air collecting chamber end cap, the air collecting chamber end cap are arranged in the head end of the inner core-body and the outer barrel, and with it is described
Inner core-body is connected with the outer barrel；
The side wall of the air collecting chamber end cap is provided with fuel inlet, and the inside of the air collecting chamber end cap has the fuel of annular
Chamber, and the annular fuel cavity it is circumferentially spaced be at least two fuel unit chambers；The fuel inlet passes through the fuel
Unit chamber is connected to the toroidal combustion chamber；
The end face of the air collecting chamber end cap is provided with oxidant inlet, and the inside of the air collecting chamber end cap has oxidant
Chamber, and it is at least two oxidants unit chamber corresponding with the fuel unit chamber that the oxidant cavity is circumferentially spaced；It is described
Oxidant inlet is connected to by the oxidant unit chamber with the toroidal combustion chamber；
Fuel tank and at least two first adjustable Venturi tubes, the fuel tank pass through the described first adjustable Wen
Pipe is connect with each fuel unit chamber, for being adjusted into the fuel quantity of each fuel unit chamber；With or,
Oxidant storage tank and at least two second adjustable Venturi tubes, the oxidant storage tank are adjustable by described second
Venturi tube is connect with each oxidant unit chamber, for being adjusted into the oxidant content of each oxidant unit chamber.
Further, in an embodiment of the present invention, the inner core-body two end faces opposite with the air collecting chamber end cap
Between be formed with circumferential weld, the fuel cavity is connected to by the circumferential weld with the toroidal combustion chamber.
Further, in an embodiment of the present invention, it is arranged on the inner core-body close to the end face of the air collecting chamber end cap
There is annular groove, is formed with annular chamber between the annular groove and the end face of the air collecting chamber end cap.
Further, in an embodiment of the present invention, the oxidant cavity is connected to by the annular chamber with the circumferential weld,
And it is connected to by the circumferential weld with toroidal combustion chamber.
Further, in an embodiment of the present invention, it is logical that several fuel are provided between the fuel cavity and the circumferential weld
Road；
Matrix is distributed in the inside of the air collecting chamber end cap to several fuel channels in a ring.
Further, in an embodiment of the present invention, several oxygen are provided between the oxidant cavity and the annular chamber
Agent channel；
Matrix is distributed in the inside of the air collecting chamber end cap to several oxidant channels in a ring.
Further, in an embodiment of the present invention, the continuous pinking rocket base engine of the thruster vector control is also
Including cone；
The cone is connected to the tail end of the inner core-body；From the head end of the inner core-body to the direction of tail end, the cone
The diameter of body is gradually reduced；
The outer sheath of the cone is equipped with convergence cylinder, and the head end of the convergence cylinder and the tail end of the outer barrel connect
It connects；From the head end of the inner core-body to the direction of tail end, diameter the subtracting with the cone diameter of the convergence cylinder inboard wall
It is small and reduce；
It is described convergence cylinder inner wall and the cone outer wall between be formed with the second gap, second gap with it is described
Toroidal combustion chamber is connected to and constitutes a part of the toroidal combustion chamber.
Further, in an embodiment of the present invention, the continuous pinking rocket base engine of the thruster vector control is also
Including expanding cylinder；
The expansion cylinder is set in the outside of the cone, and the head end of the expansion cylinder and the convergence cylinder
Tail end connection；
From the head end of the inner core-body to the direction of tail end, the diameter of the expansion cylinder inboard wall is with the cone diameter
Reduction and reduce, and the inner wall of the expansion cylinder and the distance between the outer wall of the cone are gradually increased；
It is described expansion cylinder inner wall and the cone outer wall between be formed with third space, the third space with it is described
Second gap is connected to and constitutes a part of the toroidal combustion chamber.
Further, in an embodiment of the present invention, circumferential annular of the fuel unit chamber along the air collecting chamber end cap
Matrix distribution；The oxidant unit chamber is distributed along the circumferential ring quasi array of the air collecting chamber end cap.
The present invention also provides a kind of aircraft, and the continuous pinking rocket base including the thruster vector control starts
Machine.
In the above-mentioned technical solutions, relative to conventional rocket engine, this continuous pinking rocket base engine has more
High efficiency of combustion, simpler engine structure and bigger thrust ratio.In conjunction with continuous pinking rocket base engine itself
Controllability, the fuel of injection engine in the unit time, oxidation can be made by increasing the flow of its fuel, oxidant
Dosage generates difference, to realize the vector controlled to thrust.
In conclusion the features such as high using the above-mentioned continuous pinking rocket base engine thermal efficiency, structure is simple, can assign
The more efficient propulsive performance of thruster vector control technology and simpler, reliable system structure.
Detailed description of the invention
It, below will be to specific in order to illustrate more clearly of the specific embodiment of the invention or technical solution in the prior art
Embodiment or attached drawing needed to be used in the description of the prior art be briefly described, it should be apparent that, it is described below
Attached drawing is some embodiments of the present invention, for those of ordinary skill in the art, before not making the creative labor
It puts, is also possible to obtain other drawings based on these drawings.
Fig. 1 is the section view of the continuous pinking rocket base engine of thruster vector control provided by one embodiment of the present invention
Figure；
Fig. 2 is the fuel cavity of the continuous pinking rocket base engine of thruster vector control provided by one embodiment of the present invention
Separation structure schematic diagram；
Fig. 3 is the oxidant of the continuous pinking rocket base engine of thruster vector control provided by one embodiment of the present invention
The separation structure schematic diagram of chamber；
Fig. 4 is the air collecting chamber of the continuous pinking rocket base engine of thruster vector control provided by one embodiment of the present invention
The perspective view of end cap；
Fig. 5 is the explosion of the continuous pinking rocket base engine of thruster vector control provided by one embodiment of the present invention
Figure；
Fig. 6 is the perspective view of inner core-body provided by one embodiment of the present invention；
Fig. 7 is the cross-sectional view of inner core-body provided by one embodiment of the present invention；
Fig. 8 is the perspective view of outer barrel provided by one embodiment of the present invention；
Fig. 9 is the cross-sectional view of outer barrel provided by one embodiment of the present invention；
Figure 10 is the perspective view of convergence cylinder provided by one embodiment of the present invention；
Figure 11 is the cross-sectional view of convergence cylinder provided by one embodiment of the present invention；
Figure 12 is the perspective view of cone provided by one embodiment of the present invention；
Figure 13 is the cross-sectional view of expansion cylinder provided by one embodiment of the present invention.
Appended drawing reference:
1- inner core-body；2- outer barrel；3- air collecting chamber end cap；
4- fuel inlet；5- oxidant inlet；6- cone；
7- restrains cylinder；8- expands cylinder；
21- toroidal combustion chamber；
31- fuel cavity；32- circumferential weld；
33- oxidant cavity；34- annular chamber；
35- fuel channel；36- oxidant channel；
The second gap 71-；81- third space.
Specific embodiment
Technical solution of the present invention is clearly and completely described below in conjunction with attached drawing, it is clear that described implementation
Example is a part of the embodiment of the present invention, instead of all the embodiments.Based on the embodiments of the present invention, ordinary skill
Personnel's every other embodiment obtained without making creative work, shall fall within the protection scope of the present invention.
In the description of the present invention, it should be noted that term " center ", "upper", "lower", "left", "right", "vertical",
The orientation or positional relationship of the instructions such as "horizontal", "inner", "outside" be based on the orientation or positional relationship shown in the drawings, merely to
Convenient for description the present invention and simplify description, rather than the device or element of indication or suggestion meaning must have a particular orientation,
It is constructed and operated in a specific orientation, therefore is not considered as limiting the invention.In addition, term " first ", " second ",
" third " is used for descriptive purposes only and cannot be understood as indicating or suggesting relative importance.
In the description of the present invention, it should be noted that unless otherwise clearly defined and limited, term " installation ", " phase
Even ", " connection " shall be understood in a broad sense, for example, it may be being fixedly connected, may be a detachable connection, or be integrally connected；It can
To be mechanical connection, it is also possible to be electrically connected；It can be directly connected, can also can be indirectly connected through an intermediary
Connection inside two elements.For the ordinary skill in the art, above-mentioned term can be understood at this with concrete condition
Concrete meaning in invention.
Fig. 1 is the section view of the continuous pinking rocket base engine of thruster vector control provided by one embodiment of the present invention
Figure；Fig. 2 is the fuel cavity 31 of the continuous pinking rocket base engine of thruster vector control provided by one embodiment of the present invention
Separation structure schematic diagram；Fig. 3 is the continuous pinking rocket base engine of thruster vector control provided by one embodiment of the present invention
Oxidant cavity 33 separation structure schematic diagram；Fig. 4 is the continuous quick-fried of thruster vector control provided by one embodiment of the present invention
Shake the perspective view of the air collecting chamber end cap 3 of rocket base engine；Fig. 5 is thruster vector control provided by one embodiment of the present invention
The explosive view of continuous pinking rocket base engine；Fig. 6 is the perspective view of inner core-body 1 provided by one embodiment of the present invention；Fig. 7 is
The cross-sectional view of inner core-body 1 provided by one embodiment of the present invention；Fig. 8 is the vertical of outer barrel 2 provided by one embodiment of the present invention
Body figure；Fig. 9 is the cross-sectional view of outer barrel 2 provided by one embodiment of the present invention；As shown in figs 1-9, provided in this embodiment one
The continuous pinking rocket base engine of kind thruster vector control, comprising:
Inner core-body 1；
Outer barrel 2, the outer barrel 2 have axially through hollow cavity, for being set in the outside of the inner core-body 1；
It is formed with the first gap between the outer wall of the inner core-body 1 and the inner wall of the outer barrel 2, which is used for
Constitute toroidal combustion chamber 21；At least one preknock pipe structure is provided on the outer barrel 2, for the toroidal combustion chamber
21 igniting；
Air collecting chamber end cap 3, the air collecting chamber end cap 3 are arranged in the head end of the inner core-body 1 and the outer barrel 2, and with
The inner core-body 1 and the outer barrel 2 connect；
The side wall of the air collecting chamber end cap 3 is provided with fuel inlet 4, and the inside of the air collecting chamber end cap 3 has annular
Fuel cavity 31, and the annular fuel cavity 31 it is circumferentially spaced be at least two fuel unit chambers；The fuel inlet 4 passes through
The fuel unit chamber is connected to the toroidal combustion chamber 21；
The end face of the air collecting chamber end cap 3 is provided with oxidant inlet 5, and the inside of the air collecting chamber end cap 3 has oxidation
Agent chamber 33, and it is at least two oxidants unit corresponding with the fuel unit chamber that the oxidant cavity 33 is circumferentially spaced
Chamber；The oxidant inlet 5 is connected to by the oxidant unit chamber with the toroidal combustion chamber 21；
Fuel tank and at least two first adjustable Venturi tubes, the fuel tank pass through the described first adjustable Wen
Pipe is connect with each fuel unit chamber, for being adjusted into the fuel quantity of each fuel unit chamber；With or,
Oxidant storage tank and at least two second adjustable Venturi tubes, the oxidant storage tank are adjustable by described second
Venturi tube is connect with each oxidant unit chamber, for being adjusted into the oxidant content of each oxidant unit chamber.
According to above structure it is found that being formed between first between the outer wall of the inner core-body 1 and the inner wall of the outer barrel 2
Gap, the toroidal combustion chamber 21 which is constituted can provide combustion space, start for inputting into the continuous pinking rocket base
Oxidant and fuel in machine burn.Originally this continuous pinking rocket base engine need to only detonate once can be continuous
Propagation go down, during the work time, propellant is along axial spray, and detonation wave is circumferentially propagated, and the two direction is vertical.It is quick-fried
The height and intensity of seismic wave are influenced by parameters such as combustion chamber configuration, mixture activity, mass flows simultaneously.It is sent out by experiment
It is existing, if detonation engine is divided into several regions, by adjusting different region promotion agent mass flows, difference can be made
Region generates different thrust, has itself adjustable characteristic.This based on continuous pinking rocket base engine itself can
Control characteristic, the application are circumferentially spaced for different fuel lists by the fuel cavity 31 of the annular of continuous pinking rocket base engine
First chamber, and it is corresponding that the oxidant cavity 33 is circumferentially spaced for different oxidant units corresponding from the fuel unit chamber
Chamber is based on this structure, in conjunction with the use of the first adjustable Venturi tube and the second adjustable Venturi tube, to being input to the annular
The different different intracavitary fuel and oxidation of oxidant unit of fuel unit chamber and the oxidant cavity 33 of fuel cavity 31
The amount of agent is adjusted, to achieve the purpose that continuous pinking rocket base engine thruster vector control.
The first adjustable Venturi tube and the second adjustable venturi tube structure are identical, are all using Wen's tube body and adjusting
Needle cone is constituted, and metering needle bores the throat center for being plugged on Wen's tube body, and by motor control, makes metering needle cone in Wen's tube body
Throat center disengaging, and then control fuel and oxidant flow.Those skilled in the art can be according to the actual situation to institute
It states the first adjustable Venturi tube and structure, model, the type of the second adjustable Venturi tube etc. to be adjusted, with reasonable control combustion
The flow of material and oxidant, herein just without limitation.
It carries on the continuous pinking rocket base engine of the thruster vector control of above-mentioned offer for providing fuel and oxygen
The fuel tank and oxidant storage tank of agent.Fuel in fuel tank mainly includes the combustible gases such as gaseous fuel, such as hydrogen
Body；Oxidant in oxidant storage tank mainly includes oxygen or oxygen-containing gas.Those skilled in the art can adjust according to demand
The type of whole fuel and oxidant, herein just without limitation.
In the course of work, preknock pipe structure can light a fire to the toroidal combustion chamber 21, by once lighting a fire just
Engine continuous can be made to work, it during this period, can be by corresponding controller to the described first adjustable Venturi tube and
Motor in two adjustable Venturi tubes is controlled (prior art can be used in controller), and then controls the first adjustable text
The aperture of family name pipe and the second adjustable Venturi tube.Wherein, the fuel in the fuel tank can pass through the described first adjustable text
The fuel inlet 4 of the air collecting chamber end cap 3 is entered after the regulation of family name's pipe with flow appropriate, and passes through different fuel unit chambers
Enter the corresponding position of the toroidal combustion chamber 21；Similarly, the oxidant in the oxidant storage tank will also pass through described
The oxidant inlet 5 of the air collecting chamber end cap 3 is entered after two adjustable Venturi tube regulations with flow appropriate, and by different
Oxidant unit chamber enter the corresponding position of the toroidal combustion chamber 21, be mixed and burned with the fuel, described
The different location of toroidal combustion chamber 21 forms different thrust, realizes the vector controlled to thrust.
So can be carried out to the aperture of the described first adjustable Venturi tube and the second adjustable Venturi tube using controller
Adjustment in real time, so as to the amount and the different intracavitary oxygen of oxidant unit for being input to the intracavitary fuel of different fuel units
The amount of agent carries out real-time control, and by adjusting injection fuel, the flow of oxidant, fuel, oxidant blend can be changed
Amount and its mixing proportion can adjust the thrust of 21 different location of toroidal combustion chamber by changing the amount and ratio of blend
It is whole.So when the amount of the amount and the different intracavitary oxidants of oxidant unit that are input to the intracavitary fuel of different fuel units
When changing, it will be able to carry out vector adjustment to the thrust of continuous pinking rocket base engine, realize continuous pinking rocket base
The vectored thrust of engine controls.
Relative to conventional rocket engine, this continuous pinking rocket base engine has higher efficiency of combustion, simpler
Single engine structure and bigger thrust ratio.In conjunction with the controllability of continuous pinking rocket base engine itself, Ke Yitong
It crosses in the intracavitary flow for increasing its fuel, oxidant of different fuel unit chamber and oxidant unit, makes injection hair in the unit time
Fuel, the oxidant content of motivation different location generate difference, to realize the vector controlled to thrust.
In conclusion the features such as high using the above-mentioned continuous pinking rocket base engine thermal efficiency, structure is simple, can assign
The more efficient propulsive performance of thruster vector control technology and simpler, reliable system structure.
Also, works as fuel to enter after the fuel inlet 4, the fuel cavity 31 of annular can be first into, lead to
The ring structure for crossing fuel cavity 31 forms the effect of buffering, and disperses fuel uniformly simultaneously, then using fuel cavity 31
Different fuel unit chamber enter in toroidal combustion chamber 21.It so can be so that fuel be input to the toroidal combustion chamber 21
During it is more stable and uniform, guarantee the stability of pinking, adjusted in the thrust to continuous pinking rocket base engine
It still is able to make continuous pinking rocket base engine smooth flight when whole.
Similarly, after oxidant enters the oxidant inlet 5, it can be first into oxidant cavity 33,
The effect of buffering is formed, and disperses oxidant uniformly simultaneously, then using the different oxidant units of oxidant cavity 33
Chamber enters in toroidal combustion chamber 21.So can so that oxidant during being input to toroidal combustion chamber 21 more
It is stable and uniform, stable mixing is formed between fuel, guarantees the stability of pinking, to continuous pinking rocket base engine
Thrust still be able to make continuous pinking rocket base engine smooth flight when being adjusted.
It continues to refer to figure 1, in an embodiment of the present invention, two opposite with the air collecting chamber end cap 3 of the inner core-body 1
Circumferential weld 32 is formed between end face, the fuel cavity 31 is connected to by the circumferential weld 32 with the toroidal combustion chamber 21.
During fuel enters the toroidal combustion chamber 21 by the fuel cavity 31, additionally it is possible to first pass around institute
Circumferential weld 32 is stated, which has the function of current limliting, has more accurate control effect to the input of fuel.So passing through
More accurate control also can directly carry out thrust to the accurate control of fuel.
It continues to refer to figure 1, in an embodiment of the present invention, close to the end face of the air collecting chamber end cap 3 on the inner core-body 1
It is provided with annular groove, is formed with annular chamber 34 between the annular groove and the end face of the air collecting chamber end cap 3.
During oxidant enters the toroidal combustion chamber 21 by the oxidant cavity 33, additionally it is possible to pass through first
It crosses in the annular chamber 34 by being formed between the annular groove and the end face of the air collecting chamber end cap 3, so oxidant enters oxygen
After agent chamber 33 buffers, it can be entered in toroidal combustion chamber 21 along the annular chamber 34 is circumferentially uniform, and and fuel
It is mixed, the input of oxidant can be made more uniform in this way, also directly enhance the stabilization of continuous pinking rocket base engine
Property.
It continues to refer to figure 1, in an embodiment of the present invention, the oxidant cavity 33 passes through the annular chamber 34 and the ring
32 connection of seam, and be connected to by the circumferential weld 32 with toroidal combustion chamber 21.The oxidant cavity 33 by the annular chamber 34 with
The circumferential weld 32 is connected to, so fuel and oxidant respectively can be first in circumferential welds 32 after fuel cavity 31 and oxidant cavity 33
Contact and mix, then enter back into toroidal combustion chamber 21, this structure can the amount to fuel and oxidant carry out essence simultaneously
Quasi- control, guarantee carry out more accurate control to thrust.
It continues to refer to figure 1, in an embodiment of the present invention, is provided between the fuel cavity 31 and the circumferential weld 32 several
Fuel channel 35；
Matrix is distributed in the inside of the air collecting chamber end cap 3 to several fuel channels 35 in a ring.
So by several equally distributed fuel channels 35, it can be so that fuel is circumferentially uniform to be input to the ring
In seam 32, uniformly mixed with oxidant.Wherein, ring quasi array is to indicate that fuel channel 35 is one week structure, fuel
Channel 35 can be the seam of one week connection, be also possible to discrete pipeline.
It continues to refer to figure 1, in an embodiment of the present invention, is provided between the oxidant cavity 33 and the annular chamber 34
Several oxidant channels 36；
Matrix is distributed in the inside of the air collecting chamber end cap 3 to several oxidant channels 36 in a ring.
It similarly, can be so that oxidant is circumferentially uniform to be input to institute by several equally distributed oxidant channels 36
It states in circumferential weld 32, is uniformly mixed with fuel.Wherein, ring quasi array is to indicate that oxidant channel 36 is one week structure,
Oxidant channel 36 can be the seam of one week connection, be also possible to discrete pipeline.
Figure 10 is the perspective view of convergence cylinder 7 provided by one embodiment of the present invention；Figure 11 is one embodiment of the invention
The cross-sectional view of the convergence cylinder 7 of offer；Figure 12 is the perspective view of cone 6 provided by one embodiment of the present invention；As Figure 10-12 institute
Show, and continue to refer to figure 1, in an embodiment of the present invention, the continuous pinking rocket base engine of the thruster vector control is also
Including cone 6；
The cone 6 is connected to the tail end of the inner core-body 1；It is described from the head end of the inner core-body 1 to the direction of tail end
The diameter of cone 6 is gradually reduced；
The outer sheath of the cone 6 is equipped with convergence cylinder 7, the head end of the convergence cylinder 7 and the tail end of the outer barrel 2
Connection；From the head end of the inner core-body 1 to the direction of tail end, the diameter of convergence 7 inner wall of cylinder is with 6 diameter of cone
Reduction and reduce；
It is formed with the second gap 71 between the inner wall of the convergence cylinder 7 and the outer wall of the cone 6, second gap 71
A part of the toroidal combustion chamber 21 is connected to and constituted with the toroidal combustion chamber 21.
So by accessing cone 6 and the convergence cylinder 7 that matches in its tail end, can be improved engine thrust and
Specific impulse.
Figure 13 is the cross-sectional view of expansion cylinder 8 provided by one embodiment of the present invention；As shown in figure 13, and with continued reference to figure
1, in an embodiment of the present invention, the continuous pinking rocket base engine of the thruster vector control further includes expansion cylinder 8；
The expansion cylinder 8 is set in the outside of the cone 6, and the head end and the convergence cylinder of the expansion cylinder 8
The tail end of body 7 connects；
From the head end of the inner core-body 1 to the direction of tail end, the diameter of expansion 8 inner wall of cylinder is with the cone 6
The reduction of diameter and reduce, and the inner wall of the expansion cylinder 8 and the distance between the outer wall of the cone 6 are gradually increased；
It is formed with third space 81 between the inner wall of the expansion cylinder 8 and the outer wall of the cone 6, the third space 81
A part of the toroidal combustion chamber 21 is connected to and constituted with second gap 71.
So by accessing the expansion cylinder 8 matched with cone 6 in the tail end of convergence cylinder 7, it can be further
Improve the thrust and specific impulse of engine.
It continues to refer to figure 1, in an embodiment of the present invention, circumferential direction of the fuel unit chamber along the air collecting chamber end cap 3
Ring quasi array distribution；The oxidant unit chamber is distributed along the circumferential ring quasi array of the air collecting chamber end cap 3.So by right
The fuel unit chamber and the oxidant unit chamber carry out ring quasi array distribution, can push away to continuous pinking rocket base engine
Power forms balanced vector controlled, keeps the flight of continuous pinking rocket base engine more steady.Wherein, ring quasi array distribution can
With reference to record above.
The present invention also provides a kind of aircraft, and the continuous pinking rocket base including the thruster vector control starts
Machine.
Finally, it should be noted that the above embodiments are only used to illustrate the technical solution of the present invention., rather than its limitations；To the greatest extent
Pipe present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that: its according to
So be possible to modify the technical solutions described in the foregoing embodiments, or to some or all of the technical features into
Row equivalent replacement；And these are modified or replaceed, various embodiments of the present invention technology that it does not separate the essence of the corresponding technical solution
The range of scheme.
Claims (10)
1. a kind of continuous pinking rocket base engine of thruster vector control characterized by comprising
Inner core-body；
Outer barrel, the outer barrel have axially through hollow cavity, for being set in the outside of the inner core-body；
The first gap is formed between the outer wall of the inner core-body and the inner wall of the outer barrel, first gap is for constituting ring
Shape combustion chamber；At least one preknock pipe structure is provided on the outer barrel, for lighting a fire to the toroidal combustion chamber；
Air collecting chamber end cap, the air collecting chamber end cap are arranged in the head end of the inner core-body and the outer barrel, and with the inner core
Body is connected with the outer barrel；
The side wall of the air collecting chamber end cap is provided with fuel inlet, and the inside of the air collecting chamber end cap has the fuel cavity of annular,
And the annular fuel cavity it is circumferentially spaced be at least two fuel unit chambers；The fuel inlet passes through the fuel unit
Chamber is connected to the toroidal combustion chamber；
The end face of the air collecting chamber end cap is provided with oxidant inlet, and the inside of the air collecting chamber end cap has oxidant cavity, and
The circumferentially spaced oxidant cavity is at least two oxidants unit chamber corresponding with the fuel unit chamber；The oxidant
Entrance is connected to by the oxidant unit chamber with the toroidal combustion chamber；
Fuel tank and at least two first adjustable Venturi tubes, the fuel tank by the described first adjustable Venturi tube with
Each fuel unit chamber connection, for being adjusted into the fuel quantity of each fuel unit chamber；With or,
Oxidant storage tank and at least two second adjustable Venturi tubes, the oxidant storage tank pass through the described second adjustable Wen
Pipe is connect with each oxidant unit chamber, for being adjusted into the oxidant content of each oxidant unit chamber.
2. the continuous pinking rocket base engine of thruster vector control according to claim 1, which is characterized in that in described
Be formed with circumferential weld between core two end faces opposite with the air collecting chamber end cap, the fuel cavity by the circumferential weld with it is described
Toroidal combustion chamber connection.
3. the continuous pinking rocket base engine of thruster vector control according to claim 2, which is characterized in that in described
End face on core close to the air collecting chamber end cap is provided with annular groove, between the annular groove and the end face of the air collecting chamber end cap
It is formed with annular chamber.
4. the continuous pinking rocket base engine of thruster vector control according to claim 3, which is characterized in that the oxygen
Agent chamber is connected to by the annular chamber with the circumferential weld, and is connected to by the circumferential weld with toroidal combustion chamber.
5. the continuous pinking rocket base engine of thruster vector control according to claim 4, which is characterized in that the combustion
Several fuel channels are provided between material chamber and the circumferential weld；
Matrix is distributed in the inside of the air collecting chamber end cap to several fuel channels in a ring.
6. the continuous pinking rocket base engine of thruster vector control according to claim 4, which is characterized in that the oxygen
Several oxidant channels are provided between agent chamber and the annular chamber；
Matrix is distributed in the inside of the air collecting chamber end cap to several oxidant channels in a ring.
7. the continuous pinking rocket base engine of thruster vector control according to claim 1 to 6, feature
It is, further includes cone；
The cone is connected to the tail end of the inner core-body；From the head end of the inner core-body to the direction of tail end, the cone
Diameter is gradually reduced；
The outer sheath of the cone is equipped with convergence cylinder, and the head end of the convergence cylinder is connect with the tail end of the outer barrel；From
The head end of the inner core-body to the direction of tail end, the diameter of the convergence cylinder inboard wall subtracts with the reduction of the cone diameter
It is small；
It is formed with the second gap between the inner wall of the convergence cylinder and the outer wall of the cone, second gap and the annular
Combustion chamber and a part for constituting the toroidal combustion chamber, second gap subtract with the reduction of institute's cone diameter
It is small.
8. the continuous pinking rocket base engine of thruster vector control according to claim 7, which is characterized in that further include
Expand cylinder；
The expansion cylinder is set in the outside of the cone, and the tail end of the head end of the expansion cylinder and the convergence cylinder
Connection；
From the head end of the inner core-body to the direction of tail end, diameter the subtracting with the cone diameter of the expansion cylinder inboard wall
It is small and reduce, and the distance between the inner wall of the expansion cylinder and the outer wall of the cone are gradually increased；
It is formed with third space between the inner wall of the expansion cylinder and the outer wall of the cone, the third space and described second
Gap is connected to and constitutes a part of the toroidal combustion chamber.
9. the continuous pinking rocket base engine of thruster vector control according to claim 1 to 6, feature
It is, the fuel unit chamber is distributed along the circumferential ring quasi array of the air collecting chamber end cap, and the oxidant unit chamber is along institute
State the circumferential ring quasi array distribution of air collecting chamber end cap.
10. a kind of aircraft, which is characterized in that the company including thruster vector control as claimed in any one of claims 1-9 wherein
Continuous pinking rocket base engine.
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| CN112555114B (en) * | 2020-12-01 | 2022-06-17 | 中国人民解放军战略支援部队航天工程大学 | Electromagnetic combined vector acceleration spray pipe for laser ablation propulsion |
| CN115342381A (en) * | 2022-07-26 | 2022-11-15 | 清航空天(北京)科技有限公司 | Detonation combustion chamber module and detonation combustion chamber |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| US3807306A (en) * | 1957-07-17 | 1974-04-30 | Us Army | Follow through device capable of injecting material (liquid) through hole formed by shaped charge |
| DE2829002C2 (en) * | 1978-07-01 | 1985-04-04 | Messerschmitt-Bölkow-Blohm GmbH, 8012 Ottobrunn | Warhead |
| DE19601507C1 (en) * | 1996-01-17 | 1997-06-19 | Rheinmetall Ind Ag | Flame blow back tester for engine component groups |
| US7517215B1 (en) * | 2004-07-09 | 2009-04-14 | Erc Incorporated | Method for distributed ignition of fuels by light sources |
| CN101985904B (en) * | 2010-10-28 | 2013-06-05 | 西北工业大学 | Control method of detonation pipe for high-frequency pulse detonation engine |
| CN104895698A (en) * | 2015-05-19 | 2015-09-09 | 西北工业大学 | Boosting structure of large-pipe-diameter pulse detonation rocket engine and control method thereof |
| CN205190059U (en) * | 2015-11-30 | 2016-04-27 | 清华大学 | Modular rocket engine |
| CN105257429B (en) * | 2015-11-30 | 2017-03-08 | 清华大学 | Composite rocket engine |
| RU2644798C1 (en) * | 2016-03-18 | 2018-02-14 | Владимир Дмитриевич Шкилев | Pulsed detonation rocket engine |
| CN107013369B (en) * | 2017-04-27 | 2019-04-02 | 江苏大学 | A kind of open and close type rotation unidirectional priming device of detonation engine |
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