CN103868099B - Aeroengine combustor buring room and aero-engine thereof - Google Patents
Aeroengine combustor buring room and aero-engine thereof Download PDFInfo
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- CN103868099B CN103868099B CN201210541591.4A CN201210541591A CN103868099B CN 103868099 B CN103868099 B CN 103868099B CN 201210541591 A CN201210541591 A CN 201210541591A CN 103868099 B CN103868099 B CN 103868099B
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Abstract
The invention provides a kind of aeroengine combustor buring room and aero-engine thereof, described aeroengine combustor buring room comprises: interior casing, outer casing, burner inner liner, fuel nozzle and multichannel inlet diffuser, air wherein through compressor compression enters in described combustion chamber by described multichannel inlet diffuser, and be divided into outer effluent, intermediate flow and interior effluent, described outer effluent flows into the outer ring cavity outside between casing and outer burner inner liner, outside burner inner liner provides cold gas and the mixed gas through outer blending hole battle array, described intermediate flow flows into burner inner liner head, cold gas and combustion air is provided to burner inner liner head, and described interior effluent flows into ring cavity between interior casing and flame cylinder, inside burner inner liner provides cold gas and the mixed gas through interior blending hole battle array.The present invention can Control pollution discharge, extends burner inner liner service life at high temperature, promotes combustor exit temperature profile adjustment ability, ensure the reliability service of combustion chamber.
Description
Technical field
The present invention relates to technical field of engines, particularly relate to a kind of aero-engine realizing low stain, long-life aeroengine combustor buring room and comprise this combustion chamber.
Background technology
Combustion chamber is one of requisite parts in modern aeroengine, and the chemical energy here contained in fuel is reacted by combustion chemistry, is transformed into heat energy, forms high-temperature combustion product, drives turbine acting.The combination property of modern aeroengine combustion chamber and topology layout have reached quite high level, but still face a large amount of difficult problems and challenge.Its Built to Last of sustainable development application guarantee of new material, new technology, new construction, new ideas.
Along with atmospheric environment problem worldwide receives increasing concern, the restriction of aero-engine pollutant emission is required increasingly strict, even arrived harsh stage.According to current civil engine Development Trend and the market demand, to about the year two thousand twenty, than existing International Civil Aviation Organization (ICAO), the discharge capacity of nitrogen oxide (NOx) specifies that CAEP6 standard also will reduce about 50%.In order to meet the aero-engine emission standard of increasingly stringent, low pollution emission has become an important performance requirement of modern aeroengine combustion chamber.Meanwhile, from reducing the angle safeguarded with operation costs, the service life of combustion chamber is also the problem paid special attention to.
According to the oligosaprobic development trend in modern aeroengine combustion chamber, poor oil firing's pattern of employing center fractional combustion is one of means of most potential Control pollution discharge, this needs the air allocation proportion increased substantially for participating in burning, bring other difficulty to thus the research and development of combustion chamber, such as burner inner liner cooling tolerance reduces, and need improve the durability of burner inner liner; Blending tolerance reduces, and need ensure Exit temperature distribution quality.
The combustion chamber flame drum life-span is the topmost decisive factor in combustion chamber service life.The thermal stress how reducing burner inner liner wall becomes a key technology of current combustion room research and development.
Summary of the invention
In order to solve the problem, the invention provides a kind of aeroengine combustor buring room and aero-engine thereof, low pollution emission and long life can be realized simultaneously.
According to a first aspect of the invention, a kind of aeroengine combustor buring room is provided, comprises:
Interior casing;
Outer casing, it is around described interior casing;
Burner inner liner, it is between described outer casing and described interior casing, and described burner inner liner comprises burner inner liner head, flame cylinder and the outer burner inner liner around described flame cylinder, wherein, described flame cylinder is provided with interior blending hole battle array, and described outer burner inner liner is provided with outer blending hole battle array;
Fuel nozzle, fuel oil enters described burner inner liner head by described fuel nozzle;
Multichannel inlet diffuser, air through compressor compression enters in described combustion chamber by described multichannel inlet diffuser, and be divided into outer effluent, intermediate flow and interior effluent, described outer effluent flows into the outer ring cavity between described outer casing and described outer burner inner liner, cold gas and the mixed gas through described outer blending hole battle array is provided to described outer burner inner liner, described intermediate flow flows into described burner inner liner head, cold gas and combustion air is provided to described burner inner liner head, and described interior effluent flows into ring cavity between described interior casing and described flame cylinder, cold gas and the mixed gas through described interior blending hole battle array is provided to described flame cylinder.
Wherein, in described multichannel inlet diffuser, be provided with multiple spacer ring, for the air compressed through compressor is divided into outer effluent, intermediate flow and interior effluent.
Wherein, described outer blending hole battle array and described interior blending hole battle array are relatively arranged, and form the blending hole battle array liquidated.
Wherein, described interior blending hole battle array being circumferentially uniformly distributed at described flame cylinder.
Wherein, described outer blending hole battle array being circumferentially uniformly distributed at described outer burner inner liner.
Wherein, described burner inner liner head adopts fractional combustion mode, is made up of pre-combustion grade assembly and main combustion stage assembly.
Wherein, also comprise the middle ring cavity between described flame cylinder and described outer burner inner liner, the flow area of described middle ring cavity shrinks along airflow direction.
Wherein, described flame cylinder and/or described outer burner inner liner comprise abutment wall and floating wall, the inner side of described abutment wall are fixed with multiple described floating wall, and there is gap between adjacent floating wall.
Wherein, described floating wall adopts ceramic matric composite to make.
According to a second aspect of the invention, provide a kind of aero-engine, it comprises according to foregoing aeroengine combustor buring room.
Length can be shortened in combustion chamber of the present invention, weight reduction, and Control pollution discharges, extend burner inner liner service life at high temperature, promote combustor exit temperature profile adjustment ability, ensure the reliability service of combustion chamber, Control pollution discharge simultaneously, achieves low stain and long-life.
Accompanying drawing explanation
Structure of the present invention and mode of operation and further object and advantage are better understood by the description below in conjunction with accompanying drawing, wherein:
Fig. 1 is the structural representation of the preferred embodiment of aeroengine combustor buring room of the present invention;
Fig. 2 is the double walled structural representation of burner inner liner of the present invention; And
Fig. 3 is the structural representation of conventional combustion chamber.
Description of reference numerals
The outer casing of 1 multichannel inlet diffuser 2
Casing 4 fuel nozzle in 3
The outer burner inner liner of 5 burner inner liner head 6
The outer blending hole battle array of 7 flame cylinder 8
Blending hole battle array 10 spacer ring in 9
The outer effluent of 11 air 12
Effluent in 13 intermediate flows 14
15 outer ring cavity 16 combustion airs
Ring cavity 18 cold gas in 17
19 cold gas 20 mixed gases
21 mixed gas 22 diffusers
23 conventional burner inner liner head 24 burner inner liners
25 blending hole depth of section 26 blending hole battle array depth of sections
27 abutment wall 28 floating wall
29 jet 30 air films
Direction inside 31 lateral direction 32
Detailed description of the invention
Discuss enforcement and the use of embodiment below in detail.But, should be appreciated that discussed specific embodiment only exemplarily illustrates and implement and use ad hoc fashion of the present invention, but not limit the scope of the invention.
Fig. 1 shows according to an exemplary but aeroengine combustor buring room for nonrestrictive embodiment of the present invention: adopt multichannel inlet diffuser, the oil-poor head of center classification, ceramic matric composite floating wall, burner inner liner volume reducing the burner inner liner moulding of back segment rapid desufflation and the blending hole battle array that liquidates.
See Fig. 1, aeroengine combustor buring room is made up of casing 3, fuel nozzle 4, burner inner liner in multichannel inlet diffuser 1, outer combustion case 2, combustion chamber.Fuel oil enters burner inner liner by fuel nozzle 4, and burner inner liner, between outer casing 2 and interior casing 3, comprises burner inner liner head 5, outer burner inner liner 6 and flame cylinder 7.Outer burner inner liner 6 is around flame cylinder 7, and flame cylinder 7 is provided with interior blending hole battle array 9, and outer burner inner liner 6 is provided with outer blending hole battle array 8.
Multichannel inlet diffuser 1 adopts multi-channel structure, length shorter than the diffuser 22 of conventional combustion room (Fig. 3).Air 11 through compressor compression is divided into multiply by spacer ring 10 and is decelerated to applicable state by diffusion in multichannel diffuser.At diffuser exit, air-flow adapts to the distribution of air flow of combustion chamber, be divided into outer effluent 12, intermediate flow 13 and interior effluent 14, outer effluent 12 flows into the outer ring cavity 15 outside between casing 2 and outer burner inner liner 6, outside burner inner liner 6 provides cold gas 18 and the mixed gas 20 through outer blending hole battle array 8, intermediate flow 13 flows into burner inner liner head 5, cold gas and combustion air 16 is provided to burner inner liner head, interior effluent 14 flows into ring cavity 17 between interior casing 3 and flame cylinder 7, and inside burner inner liner 7 provides cold gas 19 and the mixed gas 21 through interior blending hole battle array 9.
Burner inner liner head 5 adopts fractional combustion mode, primarily of pre-combustion grade assembly and main combustion stage assembly composition.The rotational flow air generation low speed recirculating zone that utilization enters combustion chamber by pre-combustion grade stabilizes the flame, main combustion stage starts after classification point, adopt premix and pre-evaporation combustion method, because this combustion system head air inflow is large, burner inner liner head 5 is than the conventional burner inner liner head 23 large (Fig. 3) of existing combustion chamber.
As shown in Figure 2, outer burner inner liner 6 and flame cylinder 7 adopt double-wall structure, and outer wall is abutment wall 27, and inner layer wall is floating wall 28, and cooling-air flows along lateral direction 31, and combustion gas is flowed along direction, inner side 32.Cooling-air, by hole on outer wall, forms jet 29 impinging cooling floating wall 28, and forms air film 30 and protect next section of floating wall 28.Preferably, floating wall 28 adopts ceramic matric composite to make, and is connected in abutment wall 27 by connector, there is gap between adjacent floating wall 28.
In the present embodiment, the flow area of the middle ring cavity between outer burner inner liner 6 and flame cylinder 7 shrinks along airflow direction, back segment circulation area is shunk fast, compare conventional burner inner liner 24 (Fig. 3), increase the degree of shrinkage of circulation area, under identical head sizes and burner inner liner length, burner inner liner volume reduces.
Outer burner inner liner 6 and flame cylinder 7 do not have primary holes, has outer blending hole battle array 8 and interior blending hole battle array 9.Preferably, outer blending hole battle array 8 and interior blending hole battle array 9 are relatively arranged.Interior blending hole battle array 9 being circumferentially uniformly distributed at flame cylinder 7, outer blending hole battle array 8 being circumferentially uniformly distributed of burner inner liner 6 outside, so that regulate Exit temperature distribution.Compared with the blending hole depth of section 25 of the combustion chamber of routine, blending hole battle array depth of section 26 of the present invention reduces.
The feature of combustion chamber of the present invention is: the oil-poor head of multichannel diffuser, center fractional combustion, ceramic matric composite floating wall, burner inner liner volume reducing the burner inner liner moulding of back segment circulation area rapid desufflation and the blending hole battle array that liquidates.Air-flow enters combustion chamber by multichannel diffuser, the rational three strands of air-flows of configuration are punished at diffuser exit, wherein, most air-flow enters burner inner liner by the oil-poor head of employing center grading combustion technology and participates in burning, and inside and outside two strands of air-flows enter the Exit temperature distribution of inside and outside ring cavity respectively as cold gas and mixed gas cooling flame cylinder and regulating gas.
Beneficial effect of the present invention is as described below:
1. adopt multichannel inlet diffuser, multichannel design can meet the requirement of slowing down to diffusion in shorter length, shorten diffuser length, thus obtain the income of chamber length and weight minimizing, and reasonable disposition enters the flow proportional of burner inner liner head and inside and outside ring cavity.
2. the oil-poor head of center classification, poor oil firing makes fuel be in oligosaprobic burning temperature range, under center grading combustion technology ensures different capacity, combustion chamber regional area burns with best fuel-air match pattern, solve the contradiction of CO, UHC discharge under NOx emission and small-power under large state, thus obtain and control and reduce the income of disposal of pollutants.
3. ceramic matric composite floating wall, the expansion clearance of floating wall tile surrounding, solve the thermal stress issues under high-temperature condition, the characteristics such as ceramic matric composite has heat-resisting ability, intrinsic non-oxidizability, abrasion resistance/aggressivity, lower than the density of high temperature alloy, thermal coefficient of expansion is little, resisting temperature gradient ability is strong, be applied in floating wall, improve heat-resisting ability and the resistance to corrosion of burner inner liner, decrease the demand of burner inner liner to cooling tolerance, thus obtain under limited cooling tolerance, extend the income in burner inner liner service life at high temperature.
4. burner inner liner volume reducing the burner inner liner moulding of back segment circulation area rapid desufflation, burner inner liner back segment channel height reduces, reduce the volume of burner inner liner on the one hand, reduce combustion gas residence time, can obtain the income reducing NOx emission, blending hole is positioned at burner inner liner back segment, the reduction of blending hole place depth of section on the other hand, be conducive to mixing of blending jet and high-temperature fuel gas, the income obtaining under the prerequisite reducing dilution air demand and be satisfied with combustor exit temperature distribution can be obtained.
5. the blending hole battle array liquidated, the perforate jet of bilateral makes blending jet more complete with mixing of high-temperature fuel gas, effective, can have and regulation scheme more can be selected can to obtain the income promoting combustor exit temperature profiling-quality.
Described in the present invention, concrete case study on implementation is only better case study on implementation of the present invention, is not used for limiting practical range of the present invention.Namely all equivalences done according to the content of the present patent application the scope of the claims change and modify, and all belong to protection scope of the present invention.
Claims (8)
1. an aeroengine combustor buring room, is characterized in that, comprising:
Interior casing (3);
Outer casing (2), it is around described interior casing (3);
Burner inner liner, it is positioned between described outer casing (2) and described casing (3), described burner inner liner comprises burner inner liner head (5), flame cylinder (7) and the outer burner inner liner (6) around described flame cylinder (7), wherein, described flame cylinder (7) is provided with interior blending hole battle array (9), and described outer burner inner liner (6) is provided with outer blending hole battle array (8);
Fuel nozzle (4), fuel oil enters described burner inner liner head (5) by described fuel nozzle;
Multichannel inlet diffuser (1), air (11) through compressor compression enters in described combustion chamber by described multichannel inlet diffuser (1), and be divided into outer effluent (12), intermediate flow (13) and interior effluent (14), described outer effluent (12) flows into the outer ring cavity (15) between described outer casing (2) and described outer burner inner liner (6), cold gas (18) and the mixed gas (20) through described outer blending hole battle array (8) is provided to described outer burner inner liner (6), described intermediate flow (13) flows into described burner inner liner head (5), cold gas and combustion air (16) is provided to described burner inner liner head (5), and described interior effluent (14) flows into the interior ring cavity (17) between described interior casing (3) and described flame cylinder (7), cold gas (19) and the mixed gas (21) through described interior blending hole battle array (9) is provided to described flame cylinder (7),
Wherein, described flame cylinder (7) and/or described outer burner inner liner (6) comprise abutment wall (27) and floating wall (28), the inner side of described abutment wall (27) is fixed with multiple described floating wall (28), and there is gap between adjacent floating wall (28);
Wherein, described floating wall (28) adopts ceramic matric composite to make.
2. aeroengine combustor buring room according to claim 1, it is characterized in that, multiple spacer ring (10) is provided with, for the air compressed through compressor being divided into outer effluent (12), intermediate flow (13) and interior effluent (14) in described multichannel inlet diffuser.
3. aeroengine combustor buring room according to claim 1 and 2, is characterized in that, described outer blending hole battle array (8) and described interior blending hole battle array (9) are relatively arranged.
4. aeroengine combustor buring room according to claim 1 and 2, is characterized in that, described interior blending hole battle array (9) being circumferentially uniformly distributed in described flame cylinder (7).
5. aeroengine combustor buring room according to claim 1 and 2, is characterized in that, described outer blending hole battle array (8) being circumferentially uniformly distributed in described outer burner inner liner (6).
6. aeroengine combustor buring room according to claim 1 and 2, is characterized in that, described burner inner liner head (5) adopts fractional combustion mode, is made up of pre-combustion grade assembly and main combustion stage assembly.
7. aeroengine combustor buring room according to claim 1 and 2, it is characterized in that, also comprise the middle ring cavity be positioned between described flame cylinder (7) and described outer burner inner liner (6), the flow area of described middle ring cavity shrinks along airflow direction.
8. an aero-engine, it comprises the aeroengine combustor buring room according to any one of claim 1 to 7.
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| Application Number | Priority Date | Filing Date | Title |
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| CN201210541591.4A CN103868099B (en) | 2012-12-13 | 2012-12-13 | Aeroengine combustor buring room and aero-engine thereof |
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| Application Number | Priority Date | Filing Date | Title |
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| CN201210541591.4A CN103868099B (en) | 2012-12-13 | 2012-12-13 | Aeroengine combustor buring room and aero-engine thereof |
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| Publication Number | Publication Date |
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| CN103868099A CN103868099A (en) | 2014-06-18 |
| CN103868099B true CN103868099B (en) | 2016-02-10 |
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| CN105333456B (en) * | 2014-07-31 | 2017-11-10 | 中国航发商用航空发动机有限责任公司 | Floating wall tile fragment and its floating wall for burner inner liner |
| CN105465831A (en) * | 2016-01-12 | 2016-04-06 | 西北工业大学 | Gas turbine combustion chamber provided with double flame tubes and flame holder |
| CN108626751B (en) * | 2017-03-15 | 2020-05-05 | 中国航发商用航空发动机有限责任公司 | Flame tube |
| FR3070198B1 (en) * | 2017-08-21 | 2019-09-13 | Safran Aircraft Engines | AIRCRAFT TURBOMACHINE COMBUSTION CHAMBER MODULE COMPRISING BRANDS FACILITATING REPERAGE DURING ENDOSCOPIC INSPECTION OF THE COMBUSTION CHAMBER |
| FR3081494B1 (en) * | 2018-05-28 | 2020-12-25 | Safran Aircraft Engines | GAS TURBOMACHINE COMBUSTION MODULE WITH CHAMBER BOTTOM STOP |
| CN110726562B (en) * | 2019-08-30 | 2020-10-23 | 浙江大学 | Diffuser and flame tube optimization matching experimental research device |
| CN111396926B (en) * | 2020-04-02 | 2021-12-07 | 西北工业大学 | Combustion chamber with integrated gas discharge type diffuser and flame tube |
| CN112577750B (en) * | 2020-12-08 | 2022-12-20 | 中国航发沈阳发动机研究所 | Air inlet heating device for aircraft engine complete machine test |
| CN112196838B (en) * | 2020-12-09 | 2021-02-19 | 中国航发上海商用航空发动机制造有限责任公司 | Aeroengine impeller machine and aeroengine |
| CN112696710B (en) * | 2020-12-29 | 2022-11-22 | 中国航发沈阳发动机研究所 | Method and system for determining size of mixing hole of flame tube with funnel |
| CN114165814B (en) * | 2021-10-29 | 2023-02-10 | 南京航空航天大学 | Multi-point array synergistic direct-injection lean oil classification cyclone combustion chamber |
| CN115507391B (en) * | 2022-09-16 | 2023-10-20 | 中国航发湖南动力机械研究所 | Ceramic-based flame tube |
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| EP1074792A1 (en) * | 1999-07-31 | 2001-02-07 | Rolls-Royce Plc | Turbine combustor arrangement |
| EP1688588A1 (en) * | 2005-01-06 | 2006-08-09 | Snecma | Diffusor for an annular combustor, as well as combustor and turboprop with such a diffusor |
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Address after: 200241 Minhang District Lianhua Road, Shanghai, No. 3998 Patentee after: China Hangfa commercial aviation engine limited liability company Address before: 201109 Shanghai city Minhang District Hongmei Road No. 5696 Room 101 Patentee before: AVIC Commercial Aircraft Engine Co.,Ltd. |
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