CN103868099A - Aerial engine combustion chamber and aerial engine comprising same - Google Patents
Aerial engine combustion chamber and aerial engine comprising same Download PDFInfo
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- CN103868099A CN103868099A CN201210541591.4A CN201210541591A CN103868099A CN 103868099 A CN103868099 A CN 103868099A CN 201210541591 A CN201210541591 A CN 201210541591A CN 103868099 A CN103868099 A CN 103868099A
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Abstract
The invention provides an aerial engine combustion chamber and an aerial engine comprising the same. The aerial engine combustion chamber comprises an internal casing, an external casing, a flame tube, a fuel spray nozzle and a multi-channel inlet diffuser. Air compressed by an air compressor enters the combustion chamber through the multi-channel inlet diffuser and divided into outer flow, middle flow and inner flow, wherein the outer flow flows into the outer ring cavity between the external casing and an outer flame tube to provide cooling gas and mixed gas penetrating an outer mixing hole array for the outer flame tube, the middle flow flows into the head portion of the flame tube to provide cooling gas and combustion air for the head portion of the flame tube, and the inner flow flows into the inner ring cavity between the internal casing and an inner flame tube to provide cooling gas and mixed gas penetrating an inner mixing hole array for the inner flame tube. The aerial engine combustion chamber can control pollution discharge, prolong the service life of the flame tube at high temperature, improve the outlet temperature distribution and adjustment capabilities of the combustion chamber and guarantee the reliable operation of the combustion chamber.
Description
Technical field
The present invention relates to technical field of engines, particularly relate to a kind of aero-engine that realizes low pollution, long-life aeroengine combustor buring chamber and comprise this combustion chamber.
Background technology
Combustion chamber is one of requisite parts in modern aeroengine, and the chemical energy here containing in fuel reacts 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, increasingly strict to the restriction requirement of aero-engine pollutant emission, even arrive harsh stage.According to current civil aviation engine research, development trend and the market demand, to the year two thousand twenty left and right, the discharge capacity of nitrogen oxide (NOx) also will reduce by 50% left and right than the regulation CAEP6 of existing International Civil Aviation Organization (ICAO) standard.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., safeguard and the angle of operation costs from reducing, be also the problem of paying special attention to the service life of combustion chamber meanwhile.
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 disposal of pollutants, this need to increase substantially the air allocation proportion for participating in burning, bring other difficulty to thus the research and development of combustion chamber, the cooling tolerance of for example burner inner liner reduces, and needs to improve the durability of burner inner liner; Blending tolerance reduces, and needs to guarantee outlet temperature distribution quality.
The combustion chamber flame drum life-span is the combustion chamber topmost decisive factor in service life.The thermal stress that how to reduce burner inner liner wall becomes a key technology of current combustion chamber research and development.
Summary of the invention
In order to address the above problem, the invention provides a kind of aeroengine combustor buring chamber and aero-engine thereof, can realize low pollution emission and long life simultaneously.
According to a first aspect of the invention, provide a kind of aeroengine combustor buring chamber, comprising:
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, interior burner inner liner and the outer burner inner liner around described interior burner inner liner, wherein, described interior burner inner liner 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, provide cold gas and the mixed gas through described outer blending hole battle array to described outer burner inner liner, described intermediate flow flows into described burner inner liner head, provide cold gas and combustion air to described burner inner liner head, and described interior effluent flows into the interior ring cavity between described interior casing and described interior burner inner liner, provide cold gas and the mixed gas through described interior blending hole battle array to described interior burner inner liner.
Wherein, in described multichannel inlet diffuser, be provided with multiple spacer rings, for being divided into outer effluent, intermediate flow and interior effluent through the air of compressor compression.
Wherein, described outer blending hole battle array and described interior blending hole battle array relatively arrange, and form the blending hole battle array liquidating.
Wherein, described interior blending hole battle array is uniformly distributed on the circumferencial direction of described interior burner inner liner.
Wherein, described outer blending hole battle array is uniformly distributed on the circumferencial direction of 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 interior burner inner liner and described outer burner inner liner, the flow area of described middle ring cavity shrinks along airflow direction.
Wherein, described interior burner inner liner and/or described outer burner inner liner comprise abutment wall and floating wall, are fixed with multiple described floating wall on the inner side of described abutment wall, and have 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 chamber.
Length can be shortened in combustion chamber of the present invention, and disposal of pollutants is controlled in weight reduction, extend burner inner liner service life at high temperature, promote combustor exit Temperature Distribution regulating power, guarantee the reliability service of combustion chamber, control disposal of pollutants simultaneously, realized low pollution and long-life.
Accompanying drawing explanation
Structure of the present invention and mode of operation and further object and advantage will be 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 of the present invention chamber;
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
3 interior casing 4 fuel nozzles
The outer burner inner liner of 5 burner inner liner head 6
The outer blending hole battle array of 7 interior burner inner liner 8
9 interior blending hole battle array 10 spacer rings
The outer effluent of 11 air 12
The interior effluent of 13 intermediate flow 14
15 outer ring cavity 16 combustion airs
17 interior ring cavity 18 cold gas
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
The interior side direction of 31 lateral direction 32
The specific embodiment
Discuss enforcement and the use of embodiment below in detail.But, should be appreciated that discussed specific embodiment only exemplarily illustrates and implements 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 chamber of nonrestrictive embodiment of the present invention: adopt multichannel inlet diffuser, the oil-poor head of center classification, burner inner liner moulding and the blending hole battle array that liquidates that ceramic matric composite floating wall, burner inner liner volume reducing back segment shrink fast.
Referring to Fig. 1, aeroengine combustor buring chamber 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 interior burner inner liner 7.Outer burner inner liner 6 is around interior burner inner liner 7, and interior burner inner liner 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.
Burner inner liner head 5 adopts fractional combustion mode, is mainly made up of pre-combustion grade assembly and main combustion stage assembly.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 of existing combustion chamber 23 large (Fig. 3).
As shown in Figure 2, outer burner inner liner 6 and interior burner inner liner 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 interior side direction 32.Cooling-air, by the hole on outer wall, forms jet 29 and impacts cooling floating wall 28, and forms next section of floating wall 28 of air film 30 protections.Preferably, floating wall 28 adopts ceramic matric composite to make, and is connected in abutment wall 27 by connector, between adjacent floating wall 28, has gap.
In the present embodiment, the flow area of the middle ring cavity between outer burner inner liner 6 and interior burner inner liner 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.
On outer burner inner liner 6 and interior burner inner liner 7, there is no primary holes, have 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 relatively arrange.Interior blending hole battle array 9 is uniformly distributed on the circumferencial direction of interior burner inner liner 7, and outer blending hole battle array 8 is uniformly distributed on the circumferencial direction of burner inner liner 6 outside, so that regulate outlet Temperature Distribution.Compared with the blending hole depth of section 25 of conventional combustion chamber, blending hole battle array depth of section 26 of the present invention reduces.
The feature of combustion chamber of the present invention is: burner inner liner moulding and the blending hole battle array that liquidates that oil-poor head, ceramic matric composite floating wall, burner inner liner volume reducing the back segment circulation area of multichannel diffuser, center fractional combustion shunk fast.Air-flow enters combustion chamber by multichannel diffuser, punish into the rational three strands of air-flows of configuration at diffuser exit, wherein, most air-flow participates in burning by adopting the oil-poor head of center grading combustion technology to enter burner inner liner, and inside and outside two strands of air-flows enter inside and outside ring cavity and distribute as the outlet temperature of cold gas and mixed gas cooling flame cylinder and regulating gas respectively.
Beneficial effect of the present invention is as described below:
1. adopt multichannel inlet diffuser, multichannel design can meet the requirement that diffusion is slowed down in shorter length, shortened diffuser length, thereby obtained 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 in oligosaprobic burning temperature range, center grading combustion technology guarantees under different capacity, combustion chamber regional area burns with best fuel-air match pattern, solved the contradiction of CO, UHC discharge under the discharge of NOx under large state and small-power, thereby the income of disposal of pollutants is controlled and is reduced in acquisition.
3. ceramic matric composite floating wall, the expansion clearance of floating wall tile surrounding, solve the thermal stress issues under high-temperature condition, the characteristic 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, heat-resisting ability and the resistance to corrosion of burner inner liner are promoted, reduce the demand of burner inner liner to cooling tolerance, thereby obtain under limited cooling tolerance, extend the burner inner liner income in service life at high temperature.
4. the burner inner liner moulding that burner inner liner volume reducing back segment circulation area are shunk fast, burner inner liner back segment channel height reduces, dwindle on the one hand the volume of burner inner liner, reduce combustion gas residence time, can obtain the income that reduces NOx discharge, blending hole is positioned at burner inner liner back segment on the other hand, the reducing of blending hole place depth of section, be conducive to mixing of blending jet and high-temperature fuel gas, can obtain and under the prerequisite that reduces dilution air demand, obtain the income that is satisfied with combustor exit Temperature Distribution.
5. the blending hole battle array liquidating, the perforate jet of bilateral makes blending jet more complete with mixing of high-temperature fuel gas, effective, can have and more can select regulation scheme can obtain the income that promotes combustor exit Temperature Distribution 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.Be that all equivalences of doing according to the content of the present patent application the scope of the claims change and modify, all belong to protection scope of the present invention.
Claims (10)
1. an aeroengine combustor buring chamber, 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), interior burner inner liner (7) and the outer burner inner liner (6) around described interior burner inner liner (7), wherein, described interior burner inner liner (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), provide cold gas (18) and the mixed gas (20) through described outer blending hole battle array (8) to described outer burner inner liner (6), described intermediate flow (13) flows into described burner inner liner head (5), provide cold gas and combustion air (16) 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 interior burner inner liner (7), provide cold gas (19) and the mixed gas (21) through described interior blending hole battle array (9) to described interior burner inner liner (7).
2. aeroengine combustor buring according to claim 1 chamber, it is characterized in that, in described multichannel inlet diffuser, be provided with multiple spacer rings (10), for being divided into outer effluent (12), intermediate flow (13) and interior effluent (14) through the air of compressor compression.
3. aeroengine combustor buring according to claim 1 and 2 chamber, is characterized in that, described outer blending hole battle array (8) and described interior blending hole battle array (9) relatively arrange.
4. aeroengine combustor buring according to claim 1 and 2 chamber, is characterized in that, described interior blending hole battle array (9) is uniformly distributed on the circumferencial direction of described interior burner inner liner (7).
5. aeroengine combustor buring according to claim 1 and 2 chamber, is characterized in that, described outer blending hole battle array (8) is uniformly distributed on the circumferencial direction of described outer burner inner liner (6).
6. aeroengine combustor buring according to claim 1 and 2 chamber, 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 according to claim 1 and 2 chamber, it is characterized in that, also comprise the middle ring cavity being positioned between described burner inner liner (7) and described outer burner inner liner (6), the flow area of described middle ring cavity shrinks along airflow direction.
8. aeroengine combustor buring according to claim 1 and 2 chamber, it is characterized in that, described interior burner inner liner (7) and/or described outer burner inner liner (6) comprise abutment wall (27) and floating wall (28), on the inner side of described abutment wall (27), be fixed with multiple described floating wall (28), and have gap between adjacent floating wall (28).
9. aeroengine combustor buring according to claim 8 chamber, is characterized in that, described floating wall (28) adopts ceramic matric composite to make.
10. an aero-engine, it comprises according to the aeroengine combustor buring chamber described in any one in claim 1 to 9.
Priority Applications (1)
| 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 |
|---|---|---|---|
| CN201210541591.4A CN103868099B (en) | 2012-12-13 | 2012-12-13 | Aeroengine combustor buring room and aero-engine thereof |
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| CN103868099A true CN103868099A (en) | 2014-06-18 |
| CN103868099B CN103868099B (en) | 2016-02-10 |
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Cited By (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105333456A (en) * | 2014-07-31 | 2016-02-17 | 中航商用航空发动机有限责任公司 | Floating wall tile used for flame tube and floating wall of flame tube |
| 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 |
| CN108626751A (en) * | 2017-03-15 | 2018-10-09 | 中国航发商用航空发动机有限责任公司 | Burner inner liner |
| CN110542119A (en) * | 2018-05-28 | 2019-12-06 | 赛峰航空器发动机 | Combustion module for a gas turbine engine with a combustion chamber bottom stop |
| CN110726562A (en) * | 2019-08-30 | 2020-01-24 | 浙江大学 | Diffuser and flame tube optimization matching experimental research device |
| CN110998189A (en) * | 2017-08-21 | 2020-04-10 | 赛峰飞机发动机公司 | Combustor module for an aircraft turbine engine including markings to aid identification during endoscopy of the combustor |
| CN111396926A (en) * | 2020-04-02 | 2020-07-10 | 西北工业大学 | Combustion chamber with integrated gas discharge type diffuser and flame tube |
| CN112196838A (en) * | 2020-12-09 | 2021-01-08 | 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 |
| CN112577750A (en) * | 2020-12-08 | 2021-03-30 | 中国航发沈阳发动机研究所 | Air inlet heating device for aircraft engine complete machine test |
| CN112696710A (en) * | 2020-12-29 | 2021-04-23 | 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 |
| CN114165814A (en) * | 2021-10-29 | 2022-03-11 | 南京航空航天大学 | Multi-point array synergistic direct-injection lean oil classification cyclone combustion chamber |
| CN115507391A (en) * | 2022-09-16 | 2022-12-23 | 中国航发湖南动力机械研究所 | Ceramic-based flame tube |
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Cited By (17)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105333456A (en) * | 2014-07-31 | 2016-02-17 | 中航商用航空发动机有限责任公司 | Floating wall tile used for flame tube and floating wall of flame tube |
| 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 |
| CN108626751A (en) * | 2017-03-15 | 2018-10-09 | 中国航发商用航空发动机有限责任公司 | Burner inner liner |
| CN110998189B (en) * | 2017-08-21 | 2021-02-26 | 赛峰飞机发动机公司 | Combustor module for an aircraft turbine engine including markings to aid identification during endoscopy of the combustor |
| CN110998189A (en) * | 2017-08-21 | 2020-04-10 | 赛峰飞机发动机公司 | Combustor module for an aircraft turbine engine including markings to aid identification during endoscopy of the combustor |
| CN110542119A (en) * | 2018-05-28 | 2019-12-06 | 赛峰航空器发动机 | Combustion module for a gas turbine engine with a combustion chamber bottom stop |
| CN110726562A (en) * | 2019-08-30 | 2020-01-24 | 浙江大学 | Diffuser and flame tube optimization matching experimental research device |
| CN110726562B (en) * | 2019-08-30 | 2020-10-23 | 浙江大学 | Diffuser and flame tube optimization matching experimental research device |
| CN111396926A (en) * | 2020-04-02 | 2020-07-10 | 西北工业大学 | Combustion chamber with integrated gas discharge type diffuser and flame tube |
| CN112577750A (en) * | 2020-12-08 | 2021-03-30 | 中国航发沈阳发动机研究所 | Air inlet heating device for aircraft engine complete machine test |
| CN112196838A (en) * | 2020-12-09 | 2021-01-08 | 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 |
| 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 |
| CN112696710A (en) * | 2020-12-29 | 2021-04-23 | 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 |
| 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 |
| CN114165814A (en) * | 2021-10-29 | 2022-03-11 | 南京航空航天大学 | Multi-point array synergistic direct-injection lean oil classification cyclone combustion chamber |
| CN115507391A (en) * | 2022-09-16 | 2022-12-23 | 中国航发湖南动力机械研究所 | Ceramic-based flame tube |
| 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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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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