CN105953266A - Oblique flow combustion chamber structure - Google Patents
Oblique flow combustion chamber structure Download PDFInfo
- Publication number
- CN105953266A CN105953266A CN201610288486.2A CN201610288486A CN105953266A CN 105953266 A CN105953266 A CN 105953266A CN 201610288486 A CN201610288486 A CN 201610288486A CN 105953266 A CN105953266 A CN 105953266A
- Authority
- CN
- China
- Prior art keywords
- inner liner
- burner inner
- head
- combustion chamber
- combustor
- 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.)
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Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/06—Arrangement of apertures along the flame tube
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
Abstract
The invention provides an oblique flow combustion chamber structure, and relates to combustion chamber structures. Non-axial direction air inlet of a combustion chamber can be achieved. The combustion chamber is of an oblique flow type flame tube structure, that is, the head of each flame tube deflects by a certain angle relative to the axis of the combustion chamber. The heads of the different flame tubes are connected through head side plates; cooling holes are formed in the head side plates of the flame tubes, and the head side plates of the flame tubes are perpendicular to a head panel. Secondary air holes are formed in the inner walls and the outer walls of the flame tubes; the secondary air holes are divided into a plurality of groups according to the number of the heads of the flame tubes; and the arraying direction of the secondary air holes in each group is parallel to the outer edge of the head panel of the corresponding flame tube. According to the oblique flow combustion chamber structure, outlet guide vanes of an axial flow compressor on the upper stream can be canceled, the length of an entire aircraft engine can be reduced, the mass of an entire machine can be lightened, and the thrust-weight ratio of the entire machine can be improved. In addition, the oblique flow combustion chamber further can reduce the folding angle of inlet guide vanes of a downstream turbine of the combustion chamber, the design difficulty of the inlet guide vanes of the turbine can be lowered, and the turbine efficiency is improved.
Description
Technical field
The present invention relates to aeroengine combustor buring room design field, be that one can realize aero-engine combustion
Burn the chamber structure of the non axial air inlet in room, a kind of combustor with diagonal flow type burner inner liner
Structure.
Background technology
In aero-engine, first air is compressed by compressibility, and the high pressure gas after compression flows to
Entering combustor, fuel injection system is oil spout in high pressure draught, the most substantially effectively fires
Burning, form high temperature, high-pressure gas after burning, this gas driven turbine provides the merit needed for compressibility.
For being directed to gas turbine complete machine, shorter compressor, combustor and turbine length mean higher
Thrust-weight ratio.Ensureing that Capability of Compressor shortens the main of compressor and axial turbine length on the premise of constant
Approach is to improve level load, but great level load can cause leaf design difficulty to strengthen, and blade strength drops
Low.For combustor, combustor needs enough length to ensure that fuel oil has enough at combustor
The time of staying, and then keep higher efficiency of combustion.The length to height ratio of current combustion room is by 6 development in early days
To about 2.If on the premise of ensureing that burner inner liner height of head is certain, shortening chamber length further
Then need to introduce more advanced combustion technology, otherwise can seriously undermine the performance of combustor.Therefore, work
Journey technical staff needs to have innovated in the design of aero-engine structure, improves aero-engine further
Complete machine thrust-weight ratio.
It is known that the airintake direction of conventional combustion room is axially, the shadow rotated by compressor last stage movable vane
Ring, last stage movable vane exit flow in a certain angle with the axis, in order to movable vane exit flow direction is adjusted to
Axially, and then meet the air inlet requirement of combustor, need to lead in compressor last stage movable vane added downstream one-level
Leaf adjusts airflow direction.Therefore, if blower outlet stator can be cancelled by certain technological means,
Realize the seamless connection of compressor last stage movable vane and combustor, then can effectively alleviate electromotor overall mass,
Shorten complete machine axial length, improve thrust-weight ratio.
Summary of the invention
The present invention relates to aeroengine combustor buring room design field, be that one can realize aero-engine combustion
Burn the chamber structure of the non axial air inlet in room.Upstream axle can be cancelled by the diagonal flow type combustor of the present invention
Flow air compressor exit guide blade, effectively shortens aero-engine whole length, alleviates overall mass;Additionally,
The diagonal flow type combustor of the present invention can also reduce combustor downstream turbine inlet guide vane and turn back angle, weakens
Turbine inlet stator design difficulty, promotes turbine efficiency.
For realizing above-mentioned technical purpose, the oblique flow chamber structure of the present invention is real by the following technical programs
Existing:
A kind of chamber structure, in order to realize the non axial air inlet in aeroengine combustor buring room;Described aviation is sent out
Motivation includes compressor, combustor and turbine;Described compressor is axial-flow type;Described combustor is following current
Formula, including diffuser, burner inner liner and inside and outside casing, is positioned at described axial flow compressor downstream;Described fire
Flame cylinder is oblique flow structure, it is characterised in that: described oblique flow burner inner liner be annular, comprise burner inner liner head and
The inside and outside wall of burner inner liner;Described burner inner liner head is multiple, uniform along combustor circumference, including eddy flow
Device, head panel, oil baffle disc and head side plate, each burner inner liner head relative combustion chamber axis deflection one
Determine angle;Described head panel has Cooling Holes;Described cyclone is installed on described head panel, institute
State cyclone central axis upright in head panel;Described oil baffle disc is positioned at described head surface plate high temperature side,
And keep at a certain distance away with described head panel;Use described head side plate even between described burner inner liner head
Connect;Described head side plate has Cooling Holes;The inside and outside wall of described burner inner liner passes through described burner inner liner head
Panel and head side plate connect, in described burner inner liner inside and outside wall leading edge and described burner inner liner head panel,
Outer rim and the inside and outside edge of burner inner liner head side plate are concordant;The inside and outside wall of described burner inner liner has auxiliary air
Hole;Described auxiliary air hole is divided into some groups according to burner inner liner head number, often the arrangement of group auxiliary air hole
The corresponding burner inner liner head panel outer rim in direction is parallel.
Preferably, described burner inner liner head with the angle α of combustor axis is:
Wherein, β is that compressor last stage movable vane exports geometry angle, and w is exit flow relative velocity, and c is for going out
Implication stream absolute velocity.
The oblique flow chamber structure of the present invention has significant technique effect: Neng Goushi compared to prior art
The existing non axial air inlet in aeroengine combustor buring room, effectively shortens aero-engine whole length, alleviates complete machine
Quality, improves complete machine thrust-weight ratio.
Accompanying drawing explanation
Fig. 1 is the combustor (3/4 sectional view) with oblique flow flame tube structure, and the burner inner liner in figure is
Diagonal flow type burner inner liner.
Fig. 2 is that the diagonal flow type burner inner liner of the present invention launches signal with certain profile shaft flow air compressor last stage movable vane circumference
Figure.
Fig. 3 is the diagonal flow type burner inner liner of the present invention, 45 ° of views of (A) diagonally forward, (B) front view,
(C) 45 ° of views of rear ramp.
In figure, symbol description is as follows:
1, cyclone;2, head panel;3, head side plate;4, oil baffle disc;5, burner inner liner inwall
Face;6, burner inner liner outside wall surface;7, auxiliary air hole;8, diffuser;9, burner inner liner;10, interior machine
Casket;11, outer casing;12, axial flow compressor last stage movable vane.
Detailed description of the invention
For making the purpose of the present invention, technical scheme and advantage clearer, develop simultaneously referring to the drawings
Embodiment, the present invention is described in more detail.
Fig. 1 is the toroidal combustion chamber (3/4 sectional view) with oblique flow flame tube structure, comprise diffuser 8,
Burner inner liner 9 and inside and outside casing 10,11, described burner inner liner 9 is diagonal flow type burner inner liner, burner inner liner head
Relative combustion chamber axis deflection certain angle.Fig. 2 is diagonal flow type burner inner liner 9 and certain profile shaft stream of the present invention
Compressor last stage movable vane 12 circumference launches schematic diagram.Wherein, movable vane outlet geometry angle beta is 55 °, outlet
Air-flow relative velocity w is 210m/s, and exit flow absolute velocity c is 150m/s, draws fire according to following formula
Flame cylinder head is 36.58 ° with the angle α of combustor axis.
Fig. 3 is the diagonal flow type burner inner liner of the present invention.Wherein, described diagonal flow type burner inner liner comprises twin-stage radially
Cyclone 1, head panel 2, head side plate 3, oil baffle disc 4 and the inside and outside wall of burner inner liner 5,6;
Described diagonal flow type burner inner liner comprises 12 heads, uniform along combustor circumference;Described twin-stage radial swirler
1 central shaft is vertical with described head panel 2;Described head panel 2 is vertical with described head side plate 3,
Cooling Holes, Cooling Holes diameter 1.5mm is had on it;It is high that described oil baffle disc 4 is positioned at described head panel 2
Temperature side, and and described head panel 2 be spaced 2mm;The inside and outside wall of described burner inner liner 5,6 passes through institute
State burner inner liner head panel 2 and head side plate 3 connects;The inside and outside wall of described burner inner liner 5,6 leading edge with
The inside and outside edge of described burner inner liner head panel 2 and the inside and outside edge of burner inner liner head side plate 3 are concordant;Described fire
Flame drum outer wall face 6 has auxiliary air hole, and wherein primary holes aperture 10mm, circumferentially distributed along combustor
12 groups, often group 4;Blending hole aperture 8mm, circumferentially distributed 12 groups along combustor, often group 4;
Wall cooling hole aperture 1.5mm, circumferentially distributed 12 groups along combustor, often group arranges 3 along air flow direction
Row, often row 24;Described burner inner liner internal face 5 has primary holes, and aperture 11mm, along combustor week
To being distributed 12 groups, often group 1.The all auxiliary air holes of the inside and outside wall of described burner inner liner 5,6 arrange
The corresponding burner inner liner head panel outer rim in direction is parallel.
Air is after axial flow compressor last stage movable vane 12 flows out, with combustor axis to be the angle of 36.58 °
Being directly entered combustor, air is sufficiently mixed with fuel in burner inner liner 9, burns, and combustion gas is from combustor
Export and flow out into turbine acting with angle identical with axis.Owing to eliminating blower outlet stator,
Shortening complete machine axial length, alleviate overall mass, complete machine thrust-weight ratio is improved;Additionally, due to
Combustor exit air-flow, with necessarily prewhirling, reduces turbine inlet stator to a certain extent and turns back angle
(even can cancel inlet guide vane), is conducive to promoting turbine efficiency, increases complete machine thrust further.
The foregoing is only presently preferred embodiments of the present invention, not in order to limit the present invention, all at this
Within the spirit of invention and principle, any modification, equivalent substitution and improvement etc. done, should be included in
Within the scope of the present invention.
Claims (2)
1. a chamber structure, in order to realize the non axial air inlet in aeroengine combustor buring room.Described boat
Empty electromotor includes compressor, combustor and turbine;Described compressor is axial-flow type;Described combustor is
Downflow type, including diffuser, burner inner liner and inside and outside casing, is positioned at described axial flow compressor downstream;Institute
Stating burner inner liner is oblique flow structure, it is characterised in that: described oblique flow burner inner liner is annular, comprises burner inner liner head
Portion and the inside and outside wall of burner inner liner;Described burner inner liner head is multiple, uniform along combustor circumference, including
Cyclone, head panel, oil baffle disc and head side plate, each burner inner liner head relative combustion chamber axis is inclined
Turn certain angle;Described head panel has Cooling Holes;Described cyclone is installed on described head panel,
Described cyclone central axis upright is in head panel;Described oil baffle disc is positioned at described head surface plate high temperature side,
And keep at a certain distance away with described head panel;Use described head side plate even between described burner inner liner head
Connect;Described head side plate has Cooling Holes;The inside and outside wall of described burner inner liner passes through described burner inner liner head
Panel and head side plate connect, in described burner inner liner inside and outside wall leading edge and described burner inner liner head panel,
Outer rim and the inside and outside edge of burner inner liner head side plate are concordant;The inside and outside wall of described burner inner liner has auxiliary air
Hole;Described auxiliary air hole is divided into some groups according to burner inner liner head number, often the arrangement of group auxiliary air hole
The corresponding burner inner liner head panel outer rim in direction is parallel.The diagonal flow type combustor of the present invention can be cancelled
Upstream axial flow compressor exit guide blade, effectively shortens aero-engine whole length, alleviates overall mass,
Improve thrust-weight ratio;Additionally, the diagonal flow type combustor of the present invention can also reduce combustor downstream turbine entrance
Stator is turned back angle, weakens turbine inlet stator design difficulty, promotes turbine efficiency.
2. oblique flow flame tube structure as claimed in claim 1, is characterized in that: burner inner liner head and combustion
Burning chamber axis angle α is:
Wherein, β is that compressor last stage movable vane exports geometry angle, and w is exit flow relative velocity, and c is for going out
Implication stream absolute velocity.
Priority Applications (1)
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CN201610288486.2A CN105953266B (en) | 2016-05-04 | 2016-05-04 | A kind of oblique flow chamber structure |
Applications Claiming Priority (1)
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CN201610288486.2A CN105953266B (en) | 2016-05-04 | 2016-05-04 | A kind of oblique flow chamber structure |
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CN105953266A true CN105953266A (en) | 2016-09-21 |
CN105953266B CN105953266B (en) | 2018-08-10 |
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Cited By (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN108826357A (en) * | 2018-07-27 | 2018-11-16 | 清华大学 | The toroidal combustion chamber of engine |
CN109404968A (en) * | 2017-08-16 | 2019-03-01 | 中国航发商用航空发动机有限责任公司 | A kind of combustion chamber of aero-engine |
CN112577069A (en) * | 2020-12-17 | 2021-03-30 | 中国科学院工程热物理研究所 | Oblique flow combustion chamber side wall surface structure suitable for small head inclination angle |
CN112902230A (en) * | 2021-03-11 | 2021-06-04 | 西北工业大学 | Inclined inlet double-head two-stage swirler combustion chamber |
CN114777161A (en) * | 2022-04-11 | 2022-07-22 | 南京航空航天大学 | Inclined combustion chamber scheme for integrated design of coupling compressor and turbine |
CN115076719A (en) * | 2022-05-11 | 2022-09-20 | 南京航空航天大学 | Brand-new folding fan inclined swirl combustion chamber |
CN115095885A (en) * | 2022-06-06 | 2022-09-23 | 中国船舶集团有限公司系统工程研究院 | Combined multi-point LDI oblique-feeding combustion chamber |
CN115355528A (en) * | 2022-09-02 | 2022-11-18 | 中航通飞华南飞机工业有限公司 | Oblique flow type combustion chamber of turboprop engine |
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GB619232A (en) * | 1946-06-24 | 1949-03-07 | Adrian Albert Lombard | Improvements in or relating to gas turbine plants |
US2567079A (en) * | 1945-06-21 | 1951-09-04 | Bristol Aeroplane Co Ltd | Gas turbine power plant |
US2809493A (en) * | 1951-03-19 | 1957-10-15 | American Mach & Foundry | Centrifugal flow compressor and gas turbine power plant with a centrifugal flow compressor, toroidal combustion chamber, and centripetal flow turbine |
US5946902A (en) * | 1997-10-01 | 1999-09-07 | Siemens Aktiengesellschaft | Gas turbine engine with tilted burners |
CN101799174A (en) * | 2010-01-15 | 2010-08-11 | 北京航空航天大学 | Main combustible stage tangential oil supply premix and pre-evaporation combustion chamber |
US20100313570A1 (en) * | 2006-10-20 | 2010-12-16 | Ihi Corporation | Gas turbine combustor |
US20110209482A1 (en) * | 2009-05-25 | 2011-09-01 | Majed Toqan | Tangential combustor with vaneless turbine for use on gas turbine engines |
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2016
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US2567079A (en) * | 1945-06-21 | 1951-09-04 | Bristol Aeroplane Co Ltd | Gas turbine power plant |
GB619232A (en) * | 1946-06-24 | 1949-03-07 | Adrian Albert Lombard | Improvements in or relating to gas turbine plants |
US2809493A (en) * | 1951-03-19 | 1957-10-15 | American Mach & Foundry | Centrifugal flow compressor and gas turbine power plant with a centrifugal flow compressor, toroidal combustion chamber, and centripetal flow turbine |
US5946902A (en) * | 1997-10-01 | 1999-09-07 | Siemens Aktiengesellschaft | Gas turbine engine with tilted burners |
US20100313570A1 (en) * | 2006-10-20 | 2010-12-16 | Ihi Corporation | Gas turbine combustor |
US20110209482A1 (en) * | 2009-05-25 | 2011-09-01 | Majed Toqan | Tangential combustor with vaneless turbine for use on gas turbine engines |
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Cited By (10)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN109404968A (en) * | 2017-08-16 | 2019-03-01 | 中国航发商用航空发动机有限责任公司 | A kind of combustion chamber of aero-engine |
CN109404968B (en) * | 2017-08-16 | 2020-08-07 | 中国航发商用航空发动机有限责任公司 | Combustion chamber of aircraft engine |
CN108826357A (en) * | 2018-07-27 | 2018-11-16 | 清华大学 | The toroidal combustion chamber of engine |
CN112577069A (en) * | 2020-12-17 | 2021-03-30 | 中国科学院工程热物理研究所 | Oblique flow combustion chamber side wall surface structure suitable for small head inclination angle |
CN112902230A (en) * | 2021-03-11 | 2021-06-04 | 西北工业大学 | Inclined inlet double-head two-stage swirler combustion chamber |
CN114777161A (en) * | 2022-04-11 | 2022-07-22 | 南京航空航天大学 | Inclined combustion chamber scheme for integrated design of coupling compressor and turbine |
CN115076719A (en) * | 2022-05-11 | 2022-09-20 | 南京航空航天大学 | Brand-new folding fan inclined swirl combustion chamber |
CN115095885A (en) * | 2022-06-06 | 2022-09-23 | 中国船舶集团有限公司系统工程研究院 | Combined multi-point LDI oblique-feeding combustion chamber |
CN115095885B (en) * | 2022-06-06 | 2024-02-09 | 中国船舶集团有限公司系统工程研究院 | Combined multi-point LDI inclined combustion chamber |
CN115355528A (en) * | 2022-09-02 | 2022-11-18 | 中航通飞华南飞机工业有限公司 | Oblique flow type combustion chamber of turboprop engine |
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