CN105953266A - Oblique flow combustion chamber structure - Google Patents

Oblique flow combustion chamber structure Download PDF

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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
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CN
China
Prior art keywords
described
inner liner
burner inner
head
combustion chamber
Prior art date
Application number
CN201610288486.2A
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Chinese (zh)
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CN105953266B (en
Inventor
胡斌
赵庆军
赵巍
徐建中
Original Assignee
中国科学院工程热物理研究所
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Priority to CN201610288486.2A priority Critical patent/CN105953266B/en
Publication of CN105953266A publication Critical patent/CN105953266A/en
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Publication of CN105953266B publication Critical patent/CN105953266B/en

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Classifications

    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04Air inlet arrangements
    • F23R3/06Arrangement of apertures along the flame tube
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42Continuous 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

A kind of oblique flow chamber structure

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:

α = a r c c o s ( w · c o s β c )

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.

α = a r c c o s ( w · c o s β c )

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:
α = a r c c o s ( w · c o s β c )
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.
CN201610288486.2A 2016-05-04 2016-05-04 A kind of oblique flow chamber structure CN105953266B (en)

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Cited By (1)

* Cited by examiner, † Cited by third party
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

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
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

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
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
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

Cited By (1)

* Cited by examiner, † Cited by third party
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

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