CN106523188A - Distributed type air inlet channel solid rocket engine jet pipe divergent section afterflaming device - Google Patents
Distributed type air inlet channel solid rocket engine jet pipe divergent section afterflaming device Download PDFInfo
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- CN106523188A CN106523188A CN201610881144.1A CN201610881144A CN106523188A CN 106523188 A CN106523188 A CN 106523188A CN 201610881144 A CN201610881144 A CN 201610881144A CN 106523188 A CN106523188 A CN 106523188A
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- China
- Prior art keywords
- air inlet
- air
- inlet channel
- section
- jet pipe
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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
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/32—Constructional parts; Details not otherwise provided for
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/26—Burning control
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Jet Pumps And Other Pumps (AREA)
Abstract
The invention provides a distributed type air inlet channel solid rocket engine jet pipe divergent section afterflaming device. Four air inlet channels are installed on the middle section of a missile body in an axial symmetry mode, and a boundary layer separation channel is arranged between each air inlet channel and the missile body. Each air inlet channel comprises an inlet section, a separation section and an additional divergent section. A missile wing is arranged on each air inlet channel, and an air in-jet opening is formed in each air inlet channel. A blocking cover is arranged at the end of each air in-jet opening. A movable wedge-shaped plate is arranged on each inlet section. According to the distributed type air inlet channel solid rocket engine jet pipe divergent section afterflaming device, the air inlet channels are arranged reasonably, after compressed, outside air is led into the jet pipe divergent sections, and fuel which is not fully burnt in fuel gas is burnt again so as to improve thrust and specific impulse. The jet pipe additional divergent sections are deflected through an actuation system, and thus thrust vector control is achieved.
Description
Technical field
The present invention relates to a kind of aftercombustion device, particularly relates to a kind of distributed air intake duct solid-rocket and starts
Machine nozzle divergence cone aftercombustion device.
Background technology
Solid fuel is widely used in various strategies, tactical missile.But solid fuel specific impulse is low, and fuel combustion is insufficient,
Containing a large amount of CO, H2Deng fuel gas.At present, it is indoor mainly in burning for the aftercombustion of combustion gas.Existing supplementary combustion
It is to introduce air into combustion chamber using air intake duct to burn device, coordinates fuel rich propellant, combustion gas is mixed in combustion chamber with air
Burn again, realize improving the target of specific impulse.Gas secondary injection is a kind of technology for realizing thruster vector control, but its main machine
Reason is to introduce high-pressure gas from combustion chamber, injects nozzle divergence cone, forms shock wave, localized thrust inequality occurs, so as to realize pushing away
Force vector is controlled.But the technology does not carry out second-time burning in nozzle divergence cone, and engine boosting power slightly can reduce.Cause
This, designs a kind of device and realizes that nozzle divergence cone aftercombustion has important practical value.
The content of the invention
The invention aims to the O in utilizing environment2Make solid propellant fully burn and provide a kind of distributed
Air intake duct Exit Cone of Solid Rocket Nozzle aftercombustion device.
The object of the present invention is achieved like this：Four air intake ducts are installed in the interlude axial symmetry of body, each enters
Boundary layer diverter is provided between air flue and body, and each air intake duct is constituted by entrance, distance piece, additional expansion segment, often
Missile wing is provided with above individual air intake duct, air incidence mouth is internally provided with, and the end of each air incidence mouth is provided with blanking cover,
Each entrance is provided with moveable clapboard.
Present invention additionally comprises such some architectural features：
1. moveable clapboard described in is arranged on the slide rail arrange on boundary layer diverter.
Compared with prior art, the invention has the beneficial effects as follows：The present invention arranges air intake duct by reasonable, by outside air
It is introduced to after compression at nozzle divergence cone, not sufficiently combusted fuel burns again in making combustion gas, discharges heat, make air and combustion
The mixture expansion acting of gas, improves capacity usage ratio, and then improves thrust and specific impulse.Control clapboard enters various inlet road
Tolerance is different, so as to realize thruster vector control.Namely the present invention is using the O in environment2Solid propellant is made fully to burn, energy
Amount utilization rate is high, and thrust and specific impulse are improved.Using external-compression type supersonic inlet, simple structure is easily-controllable, and operation possibility is strong.
Description of the drawings
Fig. 1 is the overall structure diagram of the present invention；
Fig. 2 is the structural representation of side-looking direction of the present invention.
Description of symbols in figure：1- boundary layer diverters, 2- clapboards, 3- distance pieces, 4- missile wings, 5- air incidence mouths, 6- are blocked up
Lid, the additional expansion segments of 7-.
Specific embodiment
The present invention is described in further detail with specific embodiment below in conjunction with the accompanying drawings.
With reference to Fig. 1 and Fig. 2, the present invention installs four air intake ducts in body stage casing axial symmetry, introduces air into nozzle-divergence
Duan Jinhang aftercombustions.Air intake duct is square, and installed in body stage casing, porch has clapboard to be compressed air, by wedge
Shape plate 2 carrys out stream to High Mach number and is compressed, and makes to flow supercharging of slowing down.There is boundary layer diverter 1 to hinder between body and air intake duct
Only air intake duct suction low energy boundary-layer, improves total pressure recovery coefficient, and distance piece 3 reduces flow distortion.Missile wing 4 is arranged on air intake duct
On, it is guided missile stabilitization flight.When missile flight speed reaches certain Mach number, blanking cover 6 is opened, and air incidence mouth is easy to add
Expansion segment is communicated, and then starts aftercombustion, and air is mended into additional expansion segment 7 by air incidence mouth 5 after compression
Burning is filled, expansion work, so as to improve thrust and specific impulse, improves capacity usage ratio.Various inlet road is adjusted by clapboard 2
Air inflow, so that realize thruster vector control.
The present invention operation principle be：The present invention installs external-compression type supersonic inlet on missile airframe, by high Mach
Number incoming air slows down and is pressurized.Can be with suitable control air mass flow by the clapboard in air intake duct.Through diffuser and isolation
Duan Hou, air enter nozzle divergence cone with velocity of sound, react with combustion gas, make the mixing of combustion gas and air using the heat of release
Thing expansion work.Nozzle divergence cone is divided into two parts, is standing part before air incidence sealing, is appendix after air incidence mouth
Point, the flow in various inlet road is controlled, is made the air mass flow into additional expansion segment uneven, and then is produced side force, realize pushing away
Force vector is controlled.When missile flight speed reaches a certain Mach number, the blanking cover at expansion segment is opened, and has just started supplementary combustion
Burn.After expansion segment blanking cover is opened, air slows down through air intake duct and is pressurized, and static pressure is made more than the static pressure at nozzle divergence cone
Air can carry out aftercombustion into expansion segment, the air of injection causes flow in jet pipe to increase, what aftercombustion was produced
Heat makes air and the mixture expansion of combustion gas do work, so as to increased thrust and specific impulse.Control clapboard makes various inlet road
Air inflow is different, so as to realize thruster vector control.The air incidence amount in various inlet road is different so that expansion segment different parts
The shock wave wave system of generation and flow distribution are uneven, so as to produce side force, so as to realize thruster vector control.Inlet mouth
Boundary-layer isolation road and clapboard are set, and different Mach number are adapted to adjust air intake duct.There is a determining deviation with body in air intake duct,
That is boundary layer diverter, boundary layer diverter can avoid air intake duct low energy boundary-layer, improve total pressure recovery coefficient, and clapboard can be horizontal
It is mobile, air intake duct throat opening area is adjusted, to adapt to different free stream Mach numbers.
To sum up, the present invention is mended in expansion segment to solid propellant rocket combustion gas for one using the oxygen in air
Fill the device of burning.The device is by boundary layer diverter, hemicone, distance piece, missile wing, air incidence mouth, blanking cover and additional expansion
Duan Zucheng.Air intake duct of the present invention is square, and, after clapboard compression, supercharging of slowing down, after distance piece is stable for incoming air
Additional expansion segment being entered by air incidence mouth, exothermic heat of reaction being carried out with combustion gas in expansion segment, expansion work, so that improve thrust
And specific impulse.Control clapboard adjusts the flow in various inlet road, so as to produce side force, so as to realize thruster vector control.
Claims (2)
1. a kind of distributed air intake duct Exit Cone of Solid Rocket Nozzle aftercombustion device, it is characterised in that：In body
Interlude axial symmetry four air intake ducts are installed, be provided with boundary layer diverter between each air intake duct and body, each air inlet
Road is provided with missile wing, is internally provided with air by entrance, distance piece, additional expansion segment composition above each air intake duct
Entrance port, the end of each air incidence mouth are provided with blanking cover, and each entrance is provided with moveable clapboard.
2. a kind of distributed air intake duct Exit Cone of Solid Rocket Nozzle aftercombustion according to claim 1 is filled
Put, it is characterised in that：The moveable clapboard is arranged on the slide rail arrange on boundary layer diverter.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201610881144.1A CN106523188B (en) | 2016-10-10 | 2016-10-10 | A kind of distributed air intake duct Exit Cone of Solid Rocket Nozzle aftercombustion device |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201610881144.1A CN106523188B (en) | 2016-10-10 | 2016-10-10 | A kind of distributed air intake duct Exit Cone of Solid Rocket Nozzle aftercombustion device |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN106523188A true CN106523188A (en) | 2017-03-22 |
| CN106523188B CN106523188B (en) | 2018-01-19 |
Family
ID=58331739
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201610881144.1A Active CN106523188B (en) | 2016-10-10 | 2016-10-10 | A kind of distributed air intake duct Exit Cone of Solid Rocket Nozzle aftercombustion device |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN106523188B (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112412662A (en) * | 2020-11-17 | 2021-02-26 | 哈尔滨工程大学 | Combined thrust vectoring nozzle system and projectile body with same |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5706650A (en) * | 1995-08-09 | 1998-01-13 | United Technologies Corporation | Vectoring nozzle using injected high pressure air |
| CN103899432A (en) * | 2014-03-31 | 2014-07-02 | 西北工业大学 | Improved pneumatic vectoring nozzle structure with function of injecting double secondary flow branches |
| CN104295404A (en) * | 2014-08-22 | 2015-01-21 | 南京航空航天大学 | Two-dimensional fluid type thrust-vectoring power device |
| CN105443268A (en) * | 2015-11-26 | 2016-03-30 | 南京航空航天大学 | Bypass type passive double-throat pneumatic vector spraying pipe with flow regulating function and control method |
-
2016
- 2016-10-10 CN CN201610881144.1A patent/CN106523188B/en active Active
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5706650A (en) * | 1995-08-09 | 1998-01-13 | United Technologies Corporation | Vectoring nozzle using injected high pressure air |
| CN103899432A (en) * | 2014-03-31 | 2014-07-02 | 西北工业大学 | Improved pneumatic vectoring nozzle structure with function of injecting double secondary flow branches |
| CN104295404A (en) * | 2014-08-22 | 2015-01-21 | 南京航空航天大学 | Two-dimensional fluid type thrust-vectoring power device |
| CN105443268A (en) * | 2015-11-26 | 2016-03-30 | 南京航空航天大学 | Bypass type passive double-throat pneumatic vector spraying pipe with flow regulating function and control method |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112412662A (en) * | 2020-11-17 | 2021-02-26 | 哈尔滨工程大学 | Combined thrust vectoring nozzle system and projectile body with same |
Also Published As
| Publication number | Publication date |
|---|---|
| CN106523188B (en) | 2018-01-19 |
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