CN106523188B - A kind of distributed air intake duct Exit Cone of Solid Rocket Nozzle aftercombustion device - Google Patents
A kind of distributed air intake duct Exit Cone of Solid Rocket Nozzle aftercombustion device Download PDFInfo
- Publication number
- CN106523188B CN106523188B CN201610881144.1A CN201610881144A CN106523188B CN 106523188 B CN106523188 B CN 106523188B CN 201610881144 A CN201610881144 A CN 201610881144A CN 106523188 B CN106523188 B CN 106523188B
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- CN
- China
- Prior art keywords
- air intake
- intake duct
- air
- aftercombustion
- entrance
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired - Fee Related
Links
- 239000007787 solid Substances 0.000 title claims abstract description 6
- 239000000567 combustion gas Substances 0.000 abstract description 9
- 239000000446 fuel Substances 0.000 abstract description 4
- 239000000203 mixture Substances 0.000 abstract description 3
- 238000002485 combustion reaction Methods 0.000 description 8
- 239000004449 solid propellant Substances 0.000 description 5
- 238000007906 compression Methods 0.000 description 3
- 238000005516 engineering process Methods 0.000 description 3
- 239000007789 gas Substances 0.000 description 3
- 238000002347 injection Methods 0.000 description 2
- 239000007924 injection Substances 0.000 description 2
- 238000011084 recovery Methods 0.000 description 2
- 230000035939 shock Effects 0.000 description 2
- 230000003068 static effect Effects 0.000 description 2
- 239000013589 supplement Substances 0.000 description 2
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000006243 chemical reaction Methods 0.000 description 1
- 230000006835 compression Effects 0.000 description 1
- 238000010276 construction Methods 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 239000002737 fuel gas Substances 0.000 description 1
- 238000002955 isolation Methods 0.000 description 1
- 239000001301 oxygen Substances 0.000 description 1
- 229910052760 oxygen Inorganic materials 0.000 description 1
- 239000003380 propellant Substances 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/08—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
Landscapes
- 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 present invention provides a kind of distributed air intake duct Exit Cone of Solid Rocket Nozzle aftercombustion device, four air intake ducts are installed in the interlude axial symmetry of body, boundary layer diverter is each provided between air intake duct and body, each air intake duct is by entrance, distance piece, additional expansion segment composition, missile wing is provided with above each air intake duct, is internally provided with air incidence mouth, the end set of each air incidence mouth has blanking cover, and each entrance is provided with moveable clapboard.The present invention is introduced at nozzle divergence cone after outside air is compressed by rationally setting air intake duct, fuel not sufficiently combusted in combustion gas is burnt again, to improve thrust and specific impulse.By actuating system, when jet pipe add expansion segment deflection, so as to realize thruster vector control.
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, fuel combustion is insufficient,
Contain a large amount of CO, H2Deng fuel gas.At present, the aftercombustion for combustion gas is mainly in combustion chamber.Existing supplement combustion
It is to introduce air into combustion chamber using air intake duct to burn device, coordinates fuel rich propellant, combustion gas is mixed with air in combustion chamber
Burn again, realize the target for improving 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, is pushed away so as to realize
Force vector controls.But the technology does not carry out second-time burning in nozzle divergence cone, engine boosting power can slightly 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 utilize the O in environment2Solid propellant is fully burnt and a kind of distribution is provided
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, Mei Gejin are installed in the interlude axial symmetry of body
Boundary layer diverter is provided between air flue and body, each air intake duct forms by entrance, distance piece, additional expansion segment, often
Missile wing is provided with above individual air intake duct, is internally provided with air incidence mouth, the end set of each air incidence mouth has blanking cover,
Each entrance is provided with moveable clapboard.
Present invention additionally comprises some such architectural features：
1. the moveable clapboard is arranged on the slide rail set on boundary layer diverter.
Compared with prior art, the beneficial effects of the invention are as follows：The present invention is by rationally setting air intake duct, by outside air
It is introduced to after compression at nozzle divergence cone, fuel not sufficiently combusted in combustion gas is burnt again, discharge heat, makes air and combustion
The mixture expansion acting of gas, improves capacity usage ratio, and then improve thrust and specific impulse.Control clapboard enters various inlet road
Tolerance is different, so as to realize thruster vector control.Namely the present invention utilizes the O in environment2Solid propellant is set fully to burn, energy
It is high to measure utilization rate, thrust and specific impulse improve.Using external-compression type supersonic inlet, simple in construction easily-controllable, operation possibility is strong.
Brief 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, 7- add expansion segment.
Embodiment
The present invention is described in further detail with 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 is compressed to High Mach number incoming, incoming is slowed down and is pressurized.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, guided missile stabilitization is flown.When missile flight speed reaches certain Mach number, blanking cover 6 is opened, and air incidence mouth is easy to add
Expansion segment communicates, and then starts aftercombustion, and air enters additional expansion segment 7 by air incidence mouth 5 after overcompression and mended
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 as to 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, which slows down, to be pressurized.Can be with suitable control air mass flow by the clapboard in air intake duct.By diffuser and isolation
Duan Hou, air enter nozzle divergence cone with velocity of sound, reacted with combustion gas, make combustion gas and the mixing of air using the heat of release
Thing expansion work.Nozzle divergence cone is divided into two parts, is fixed part before air incidence mouth, is appendix after air incidence mouth
Point, the flow in control various inlet road, make the air mass flow for entering additional expansion segment uneven, and then side force is produced, realization pushes away
Force vector controls.When missile flight speed reaches a certain Mach number, the blanking cover at expansion segment is opened, and has just started supplement combustion
Burn.After expansion segment blanking cover is opened, air slows down by air intake duct to be pressurized, and static pressure is more than the static pressure at nozzle divergence cone, makes
Air can enter expansion segment and carry out aftercombustion, the air of injection causes flow increase in jet pipe, caused by aftercombustion
Heat makes air and the mixture expansion of combustion gas do work, so as to add 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
Caused shock wave wave system and flow distribution are uneven, so as to produce side force, so as to realize thruster vector control.Inlet mouth
Set boundary-layer to isolate road and clapboard, different Mach number is adapted to adjust air intake duct.A determining deviation be present 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, 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 for one using the oxygen in air to solid propellant rocket combustion gas in expansion segment
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 incoming air is after clapboard compresses, supercharging of slowing down, after distance piece is stable
Additional expansion segment is entered by air incidence mouth, exothermic heat of reaction, expansion work, so as to improve thrust are carried out with combustion gas in expansion segment
And specific impulse.The flow in clapboard regulation various inlet road is controlled, so as to produce side force, so as to realize thruster vector control.
Claims (2)
- A kind of 1. 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, boundary layer diverter, each air inlet are provided between each air intake duct and body Road forms by entrance, distance piece, additional expansion segment, is provided with missile wing above each air intake duct, is internally provided with air Entrance port, the end set of each air incidence mouth have blanking cover, and each entrance is provided with moveable clapboard.
- A kind of 2. distributed air intake duct Exit Cone of Solid Rocket Nozzle aftercombustion dress according to claim 1 Put, it is characterised in that：The moveable clapboard is arranged on the slide rail set 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 CN106523188A (en) | 2017-03-22 |
| CN106523188B true CN106523188B (en) | 2018-01-19 |
Family
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201610881144.1A Expired - Fee Related 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) |
Families Citing this family (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112412662B (en) * | 2020-11-17 | 2021-11-09 | 哈尔滨工程大学 | Combined thrust vectoring nozzle system and projectile body with same |
| CN118188218B (en) * | 2024-04-23 | 2024-10-11 | 南昌航空大学 | Design method of two-side inlet solid rocket engine spray pipe supplementary combustion device |
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 not_active Expired - Fee Related
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 |
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| Publication number | Publication date |
|---|---|
| CN106523188A (en) | 2017-03-22 |
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