CN102305152A - Hybrid exhaust aircraft engine - Google Patents
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- CN102305152A CN102305152A CN201110131487A CN201110131487A CN102305152A CN 102305152 A CN102305152 A CN 102305152A CN 201110131487 A CN201110131487 A CN 201110131487A CN 201110131487 A CN201110131487 A CN 201110131487A CN 102305152 A CN102305152 A CN 102305152A
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- 238000002485 combustion reaction Methods 0.000 claims description 19
- 239000012530 fluid Substances 0.000 claims description 10
- 238000006073 displacement reaction Methods 0.000 claims 1
- 238000010304 firing Methods 0.000 description 29
- 238000000034 method Methods 0.000 description 9
- 230000008569 process Effects 0.000 description 8
- 239000000203 mixture Substances 0.000 description 5
- 230000008859 change Effects 0.000 description 4
- 230000008676 import Effects 0.000 description 4
- 230000008901 benefit Effects 0.000 description 3
- 238000010168 coupling process Methods 0.000 description 3
- 230000015572 biosynthetic process Effects 0.000 description 2
- 230000008878 coupling Effects 0.000 description 2
- 238000005859 coupling reaction Methods 0.000 description 2
- 239000000446 fuel Substances 0.000 description 2
- 241000883990 Flabellum Species 0.000 description 1
- 230000006835 compression Effects 0.000 description 1
- 238000007906 compression Methods 0.000 description 1
- 238000004590 computer program Methods 0.000 description 1
- 238000001816 cooling Methods 0.000 description 1
- 230000007812 deficiency Effects 0.000 description 1
- 238000009792 diffusion process Methods 0.000 description 1
- 238000010894 electron beam technology Methods 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 239000000295 fuel oil Substances 0.000 description 1
- 238000010438 heat treatment Methods 0.000 description 1
- 239000003350 kerosene Substances 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 239000003921 oil Substances 0.000 description 1
- 230000001105 regulatory effect Effects 0.000 description 1
- 238000007493 shaping process Methods 0.000 description 1
- 230000008646 thermal stress Effects 0.000 description 1
- 230000008719 thickening Effects 0.000 description 1
- 238000003466 welding Methods 0.000 description 1
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Abstract
The invention relates to a hybrid exhaust aircraft engine with a rear-mounted fan. The hybrid exhaust aircraft engine comprises an engine rotor, an inner housing, an outer housing, a main combustion chamber and an interstage combustion chamber which are formed in an inner culvert, and the rear-mounted fan, wherein an outer culvert is formed between the inner housing and the outer housing, and the inner culvert is formed in the inner housing; the engine rotor comprises a front-mounted fan rotor and a high-pressure compressor rotor which are positioned in front of the main combustion chamber, a high-pressure turbine rotor positioned between the main combustion chamber and the interstage combustion chamber, and a low-pressure turbine rotor positioned in the rear of the interstage combustion chamber; the rear-mounted fan is arranged in the rear of the low-pressure turbine rotor in the flowing direction of fluid, and comprises a plurality of fan blades and turbine blades which are arranged circumferentially and an annular margin plate connected with the turbine blades and the fan blades; and the fan blades are positioned in the outer culvert, and the turbine blades are positioned in the inner culvert.
Description
Technical field
The present invention relates to rearmounted fan mixing aeroengine, particularly relate to a kind of rearmounted fan mixing aeroengine with the inter-stage firing chamber.
Background technique
That truly can realize at present supersonic cruise flight in the military aircraft has only the F22 fighter; Though bomber XB-70, attack plane A-12, reconnaissance plane SR-71 can realize that under afterburning state Mach number is up to 3 supersonic flight; But foundation the 4th generation fighter standard, the use of tail pipe burner has greatly reduced the stealthy performance of aircraft.And the raising of the performance of aeroengine is mainly when increasing the gas compressor overall pressure tatio, and improves turbine inlet temperature.At present; F100-PW229 motor overall pressure tatio is near 34; The turbine inlet temperature of F119 can reach 2050K, under the existing technology condition, further increases the stall that overall pressure tatio causes gas compressor easily; The firing chamber temperature rise also receives the restriction of turbine blade high-temperature capability, improves the combustor exit temperature and awaits further developing of materials science.As if in existing overall pressure tatio and temperature rise, aero-engine performance has arrived the limit.
California, USA university in 1997 proposes that the inter-stage firing chamber is set and can effectively improve the motor overall performance between high and low pressure turbine; Patent No. CN101566353A, CN101070961A, CN1858498A have proposed different inter-stage chamber structures, and patent CN101709656A is the calculating coupling process that has proposed inter-stage firing chamber blade cooling effectiveness and combustion efficiency.In fact, behind the increase inter-stage firing chamber, except chamber structure is designed; Also need change the motor overall structure, because after setting up the inter-stage firing chamber, low-pressure turbine acting ability increases; Blow down ratio reduces, and under former engine structure, the main duct stagnation pressure will be higher than the by-pass air duct stagnation pressure; And true all show with theory: only when inside and outside culvert pressure approximately equal; Just it is minimum that inside and outside culvert air-flow mixes the back loss, in order to contain the pressure coupling inside and outside making, the modification scheme that generally believes both at home and abroad is to increase fan pressure ratio or increase bypass ratio.But change the structure of former motor, need expend more financial resources input.Even keep core engine constant, only transform fan, will increase former motor overall pressure tatio, cause the inner aerodynamics problem of gas compressor easily.
The rearmounted turbofan engine of the first in the world desk fan is the CF700 of AM General motor company nineteen fifty-nine development; The also similar motor for the fan postposition of the propfan engine with by-pass air duct of former Soviet Union's development, patent CN1654806A and application number CN101881237A are the rearmounted branch row engine structure of fan.The problem that more than exists in the invention is the same as the CF700 motor; Fan blade is in the atmosphere environment, subzero approximately 50 degree of high aerial ambient temperature, and the intension turbine blade will bear the above high temperature of 500 degree; So on blade radial, produce higher thermal stress; Separately the common bypass ratio of motor of exhaust is bigger simultaneously, so blade is longer, the centrifugal force that produces during high speed rotating can make the turbine blade of lower end can't bear the heavy load; This is that CF700 does not obtain the reason promoted, also is that separately exhaust steam turbine adopts the remarkable deficiency of interposed structure behind the fan.
Summary of the invention
For solving at least one aspect of technical problem of the prior art, the present invention is proposed.The present invention increases a rearmounted fan behind low-pressure turbine, consume the part heat that the inter-stage firing chamber increases, and is used to improve by-pass air duct pressure, makes losses by mixture reach minimum.
According to an aspect of the present invention, proposed a kind of rearmounted fan mixing aeroengine, having comprised: engine rotor; Interior casing and outer casing form by-pass air duct between interior casing and the outer casing, form main duct in the interior casing; Be arranged on main combustion chamber and inter-stage firing chamber in the main duct; Low Pressure Turbine Rotor part after said engine rotor comprises preposition fan propeller and high-pressure compressor rotor part, the part of the High Pressure Turbine Rotor between main combustion chamber and inter-stage firing chamber before main combustion chamber and is positioned at the inter-stage firing chamber; Said motor also comprises: rearmounted fan; Said rearmounted fan places on fluid flow direction after the Low Pressure Turbine Rotor part; Said rearmounted fan comprises: be positioned at a plurality of fan blade of by-pass air duct, circumferential arrangement, be formed on the passage of flaring on the fluid flow direction between the adjacent fans blade; Be positioned at a plurality of turbine blades of said main duct, circumferential arrangement, be formed on the passage of convergent on the fluid flow direction between the adjacent turbine blades; And annular listrium; Have the annular blade joint and the first annular extension part; Said blade joint connects turbine blade and fan blade, and the said first annular extension part extends the end with casing in contiguous said from said blade joint towards the end of interior casing.
Further, said annular listrium also can comprise the second annular extension part, and the said second annular extension part extends away from the end of interior casing from said blade joint.The said first annular extension part can be fir-tree type.
Further, form mechanical seal between the opposed end near the end of interior casing and interior casing of the said first annular extension part.
Said annular listrium can adopt the mode of global formation to be connected with turbine blade and fan blade.Perhaps, said annular listrium can adopt mechanical part to be connected with turbine blade and fan blade.
Alternatively, said rearmounted fan also comprises: turbine blade setting wheel disc above that; The wheel shaft outstanding and that be wholely set with wheel disc from said wheel disc; The bearing that rotatably keeps said wheel shaft; And the support that is used for fixing bearing, said support passes interior casing and is fixedly connected between bearing and the outer casing, and perhaps said support is fixedly connected between bearing and the interior casing.
Said support can be isolated a plurality of fixed strips of equal angles or fixed plate.
Alternatively, the turbine blade of said rearmounted fan directly is fixed on the Low Pressure Turbine Rotor part.
Through hereinafter with reference to accompanying drawing to the description that the present invention did, other purpose of the present invention and advantage will be obvious, and can help that the present invention is had comprehensive understanding.
Description of drawings
Fig. 1 is the structural representation of rearmounted fan mixing aeroengine according to an embodiment of the invention;
Fig. 2 is the schematic representation of the Placement of rearmounted fan according to an embodiment of the invention;
Fig. 3 is the local detailed drawing of the rearmounted fan of three-dimensional according to an embodiment of the invention.
Fig. 4 is the structural representation of rearmounted fan mixing aeroengine according to another embodiment of the invention;
Fig. 5 is according to main duct thermodynamic cycle process P-V figure of the present invention;
Fig. 6 is according to by-pass air duct thermodynamic cycle process P-V figure of the present invention.
Embodiment
Pass through embodiment below, and combine accompanying drawing, do further bright specifically technological scheme of the present invention.In specification, same or analogous drawing reference numeral is indicated same or analogous parts.Following explanation to embodiment of the present invention is intended to present general inventive concept of the present invention is made an explanation with reference to accompanying drawing, and is not to be understood that to a kind of restriction of the present invention.
Like Fig. 1 and 3, shown in 4, mixing aeroengine according to the present invention comprises: engine rotor; Interior casing 55 and outer casing 56 form by-pass air duct 53 between interior casing 55 and the outer casing 56, form main duct 52 in the interior casing; Be arranged on main combustion chamber 58 and inter-stage firing chamber 61 in the main duct 52; Wherein, Said engine rotor be included on the fluid flow direction the preposition fan propeller 51 before the main combustion chamber 58 and high-pressure compressor rotor part 54, between main combustion chamber 58 and inter-stage firing chamber 61 High Pressure Turbine Rotor part 59 and be positioned at inter-stage firing chamber 61 after Low Pressure Turbine Rotor part 62; And wherein: said motor also comprises: rearmounted fan (comprising parts 64,65,66 in the accompanying drawings); Rearmounted fan places on fluid flow direction after the Low Pressure Turbine Rotor part 62; Shown in Fig. 1-3, rearmounted fan comprises a plurality of fan blade 64 that are positioned at by-pass air duct 53, circumferential arrangement, is formed on the passage of flaring on the fluid flow direction between the adjacent fans blade; Be positioned at a plurality of turbine blades 66 of said main duct 52, circumferential arrangement, be formed on the passage of convergent on the fluid flow direction between the adjacent turbine blades; And annular listrium 65; Have the annular blade joint 651 and the first annular extension part 652; Said blade joint 651 connects turbine blade 66 and fan blade 64, and the said first annular extension part 652 extends the end with casing 55 in contiguous said from said blade joint 651 towards the end of interior casing 55.
As shown in Figure 3, the number of the number of fan blade 64 and turbine blade 66 can be identical, also can be inequality.
Conventional engine is followed successively by along each parts of airflow direction: preposition fan, gas compressor, firing chamber, high-pressure turbine, low-pressure turbine, mixer, jet pipe.
Be followed successively by along each parts of airflow direction among the present invention: preposition fan propeller 51, gas compressor (high-pressure compressor rotor part 54), main combustion chamber 58; High-pressure turbine (High Pressure Turbine Rotor part 59), inter-stage firing chamber 61, low-pressure turbine (Low Pressure Turbine Rotor part 62); Rearmounted fan (fan blade 64, annular listrium 65, turbine blade 66); Mixer 71, jet pipe 70.Duoed inter-stage firing chamber 61 and rearmounted fan than conventional engine, here, the purposes of rearmounted fan is not matched by the pressure that the inter-stage firing chamber causes for balance.
The know-why of rearmounted fan is following: under the former engine structure, inside and outside duct air-flow is the stagnation pressure approximately equal at the mixer entrance place, but after having increased the inter-stage firing chamber; Low-pressure turbine import stagnation temperature raises, and does same merit, and the stagnation pressure of outlet will be significantly higher than the outlet of conventional engine low-pressure turbine; This moment, inside and outside duct air-flow with imbalance, increased losses by mixture, among the present invention at the stagnation pressure of mixer entrance; Consume the heat that a part of inter-stage firing chamber produces, promote the turbine blade acting of rearmounted fan, the bottom of rearmounted fan is divided into turbine blade; Top is divided into fan blade, and turbine blade drives fan blade and rotates, the air-flow that by-pass air duct slowed down the back, had uniform temperature for the pressurization through front fan this moment; Because the pressure ratio of preposition fan is lower, so the stagnation pressure of by-pass air duct is less than the air-flow stagnation pressure of main duct, then after the blade diffusion deceleration of by-pass air duct air-flow through rearmounted fan; The by-pass air duct stagnation pressure further raises, and the main duct air-flow is because do work, and the stagnation pressure stagnation temperature all reduces; Inside and outside duct air-flow stagnation pressure gets into mixer mixing under approximately equal, the losses by mixture of this moment is minimum.
Owing to after the inter-stage firing chamber is positioned at high-pressure turbine,, only affect low-pressure turbine to not influence of high-pressure turbine.In addition; Because the bypass ratio of mixing exhaust is very low, fan blade length is shorter, has also reduced the influence of centrifugal force to turbine blade; The annular inner rim plate places inside and outside duct air-flow respectively with turbine blade part and fan blade part simultaneously, avoids mixing too early the loss that causes.Also comprise at annular listrium 65 under the situation of the second annular extension part 653 that extends away from the end of interior casing 55 from said blade joint, can more effectively avoid mixing too early the loss that causes.
Advantageously, the first annular extension part 652 is a fir-tree type.Form mechanical seal between the opposed end near the end of interior casing 55 and interior casing 55 of the first annular extension part 652.
It should be noted that said annular listrium can be to adopt the mode of global formation to be connected with turbine blade and fan blade.The advantage of doing like this is: reduced amount of parts, can produce deviation otherwise number of components is multi-link, and the joint needs thickening, can increase engine weight.But the disadvantage of whole machine shaping is exactly that a blade damages, and whole aft-fan rotatable parts all will be changed.Perhaps, said annular listrium can be connected turbine blade with parts such as bolt or pins with fan blade.
With reference to accompanying drawing 2, said rearmounted fan also comprises: turbine blade setting wheel disc 69 above that; The wheel shaft 68 outstanding and that be wholely set with wheel disc 69 from said wheel disc 69; The bearing 67 that rotatably keeps said wheel shaft 68; And the support 63 that is used for fixing bearing 67, said support 63 passes interior casing 55 and is fixedly connected between bearing 67 and the outer casing 56, and perhaps said support 63 is fixedly connected between bearing 67 and the interior casing 55.Said support 63 can be isolated a plurality of fixed strips of equal angles or fixed plate.Fig. 2 shows a kind of connecting means of rearmounted fan; Rearmounted fan is made up of fan blade 64, turbine blade 66 and fir-tree type listrium 65 in the left hand view; Rearmounted fan is connected through the tenon notch with wheel disc 69, wheel disc 69 and wheel shaft 68 employing electron beam weldings, and through bearing 67 and pass interior casing 55; The support 63 that is fixed on the outer casing 56 connects, and has constituted the rearmounted fan that a cover can freely rotate.Right part of flg is the sectional view of A-A section.Fig. 3 is the three-dimensional local detailed drawing of rearmounted fan.
As shown in fig. 1; The opposite side of a side with said wheel shaft 68 places of said wheel disc 69 is provided with circular cone portion 72; This circular cone portion 72 is a spin axis and away from Low Pressure Turbine Rotor part convergent, to guarantee the flaring passage of turbine meridian plane with the spin axis of engine rotor.
In the technological scheme in Fig. 1 and Fig. 2, can guarantee that rearmounted fan and engine rotor rotate with different speed.
Fig. 3 is the three-dimensional local detailed drawing of rearmounted fan.Because turbine blade 66 is different with fan blade 64 working principles; So its denseness, established angle, curvature also have than big-difference; The power division of compression work that can distribute expansion work and the fan blade of turbine blade through the adjusting vane mounting type; The pressure approximately equal that makes main duct 52 and by-pass air duct 53 interior two strands of air-flows get into mixer 71, minimum to guarantee loss.
Rearmounted fan and the engine rotor scheme of rotation synchronously has been shown among Fig. 4.Particularly, the flabellum of said rearmounted fan directly is fixed on the Low Pressure Turbine Rotor part, for example is connected with former Low Pressure Turbine Rotor wheel disc through the wheel disc of lag bolt with rearmounted fan.
Below in conjunction with the running of accompanying drawing 1,4 descriptions according to motor of the present invention.
Fig. 1 is first exemplary embodiment of the present invention, divides two strands of entering motors after the 51 preliminary superchargings of atmosphere process fan propeller in the environment, and is most of through main duct 52 entering core engine; The by-pass air duct 53 that interior casing 55 of fraction entering and outer casing are 56, the main duct air further compresses through high-pressure compressor rotor 54, gets into main combustion chamber 58; With the aviation kerosine mixed combustion of oil nozzle 57, high-pressure gas promotes High Pressure Turbine Rotor 59 expansion actings, consumes part energy; The back gets into inter-stage firing chamber 61, with the fuel oil mixed combustion of 60 ejections of fuel nozzle in the inter-stage firing chamber, promotes Low Pressure Turbine Rotor 62 actings; Energy remaining continues turbine blade 66 actings to rearmounted fan; Drive 64 pairs of further superchargings of by-pass air duct gas of fan blade of rearmounted fan, inside and outside culvert air-flow mixes in mixer 71, discharges through jet pipe 70.
Fig. 4 is second exemplary embodiment of the present invention, and numbering 0-9 carries out each parts cross section signal that the cycle analysis of motor heating power is divided, and 0 is distant place atmosphere; 2.2 be the front fan outlet; 3 is the main combustion chamber import, and 4 are the main combustion chamber outlet, and 4.5 is the import of inter-stage firing chamber; 4.8 be the inter-stage combustor exit, 5 is that mixer import 6 is the jet pipe outlet for mixer outlet 9; 5 ', 6 ', 9 ' is respectively the engine agitator inlet/outlet of not being with the inter-stage firing chamber, jet pipe outlet parameter.Engine structure before second embodiment's mesolow turbine is identical with first embodiment; Different is rearmounted fan structure; The wheel disc of rearmounted fan is connected through lag bolt with former motor Low Pressure Turbine Rotor wheel disc; Its rotating speed is consistent with Low Pressure Turbine Rotor, and this structure is compared change in rotational speed with first embodiment and is restricted, but the structure of rearmounted fan is simplified.
Fig. 5 is main duct thermodynamic cycle process P-V figure of the present invention, and 0-3-4-5 '-6 '-9 ' is a conventional engine main duct thermodynamic cycle process, and 0-3-4-4.5-4.8-5-6-9 is the firing chamber motor main duct thermodynamic cycle of band inter-stage.Make the main duct temperature reduce owing in mixer, mix, so the isobaric temperature-fall period of 5-6 (5 '-6 ') is arranged among the Ideal Cycle figure.
Fig. 6 is by-pass air duct thermodynamic cycle process P-V figure of the present invention, and 0-2.2-6 '-9 ' is the conventional engine thermodynamic cycle process, and 0-2.2-5-6-9 is the firing chamber motor by-pass air duct thermodynamic cycle of band inter-stage.Make the by-pass air duct temperature raise owing in mixer, mix, increased by-pass air duct circulation merit, so separately the by-pass air duct thermodynamic cycle of exhaust is a curve, and the by-pass air duct thermodynamic cycle of mixing exhaust is an enclosed areas.Lack the thermodynamic cycle that part equals by-pass air duct in the main duct thermodynamic cycle; Can be seen by the addition of inside and outside culvert circulation area: band inter-stage firing chamber varieties of engine thermodynamic cycles process 0-3-4-4.5-4.8-5-6-9 has more a part of area than the conventional engine 0-3-4-5 '-6 '-9 ' that circulates, and is the increment of band inter-stage firing chamber motor than conventional engine circulation merit.
Under the prerequisite of considering component efficiencies; Adopt motor overall performance one dimension computer program, when the Mach number 1.6, the calculating of 11km high cruise flight shows: after increasing the inter-stage firing chamber to the F22 fighter; Compare conventional engines and increase by 7% fuel consumption; Can increase by 21% ratio and push away, adopt after interposed structure is regulated the pressure coupling behind the fan, increase to 30% than pushing away.Fan is rearmounted to be can former engine structure and aeroperformance not to be exerted an influence than the advantage that increases preposition fan pressure ratio or bypass ratio, has reduced lead time and cost.
Although illustrated and described embodiments of the invention; For those of ordinary skill in the art; Be appreciated that under the situation that does not break away from principle of the present invention and spirit and can change that scope of the present invention is limited accompanying claims and equivalent thereof to these embodiments.
Claims (9)
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| CN201110131487A CN102305152A (en) | 2011-05-20 | 2011-05-20 | Hybrid exhaust aircraft engine |
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| CN201110131487A CN102305152A (en) | 2011-05-20 | 2011-05-20 | Hybrid exhaust aircraft engine |
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Cited By (18)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105673088A (en) * | 2016-01-19 | 2016-06-15 | 哈尔滨工业大学 | Oil cooling turbine moving blade |
| CN106194497A (en) * | 2016-06-21 | 2016-12-07 | 杨京生 | Fixing shaft type, tubular gas turbine, the turbofan aero-jet engine of single combustion chamber structure |
| CN106840643A (en) * | 2017-04-06 | 2017-06-13 | 中国科学院工程热物理研究所 | The measurement apparatus of casing thermal deformation under a kind of impingement heat transfer |
| CN107965397A (en) * | 2016-10-20 | 2018-04-27 | 谢爱梅 | The efficient aviation new work engine structure of multistage centrifugal turbofan |
| CN108278165A (en) * | 2017-12-14 | 2018-07-13 | 雷静桃 | A kind of adjustable aero-engine of the direction of motion |
| CN108506111A (en) * | 2018-05-04 | 2018-09-07 | 西北工业大学 | A kind of microminiature fanjet |
| CN108533418A (en) * | 2018-04-28 | 2018-09-14 | 上海啸风航空科技有限公司 | A kind of double duct mixing exhaust fanjets of novel microminiature |
| CN108626026A (en) * | 2018-04-28 | 2018-10-09 | 上海啸风航空科技有限公司 | A kind of novel microminiature fan postposition fanjet |
| CN109441635A (en) * | 2018-12-18 | 2019-03-08 | 王立芳 | Three axis triple channels become duct variable diameter in series and parallel and adaptively recycle to turning jet engine |
| CN109948231A (en) * | 2019-03-14 | 2019-06-28 | 中国航发湖南动力机械研究所 | Engine cycle parameters analysis method and device |
| CN112524641A (en) * | 2020-12-02 | 2021-03-19 | 西北工业大学 | Novel turbine interstage combustion chamber |
| CN113204842A (en) * | 2021-04-28 | 2021-08-03 | 中国航发沈阳发动机研究所 | Engine thermodynamic cycle design method |
| CN113738530A (en) * | 2021-10-15 | 2021-12-03 | 清华大学 | Multi-duct aero-engine casing structure with blade tip fan |
| CN114542510A (en) * | 2022-02-23 | 2022-05-27 | 中国航发沈阳发动机研究所 | Self-adaptive variable-cycle engine fan rotor structure |
| CN114810350A (en) * | 2022-05-06 | 2022-07-29 | 中国科学院工程热物理研究所 | Methane precooling turbine-based combined cycle engine system with interstage combustion chamber |
| CN115183273A (en) * | 2022-07-21 | 2022-10-14 | 中国航发沈阳发动机研究所 | Afterburning engine combustion chamber |
| CN115653790A (en) * | 2022-10-15 | 2023-01-31 | 中国科学院力学研究所 | Dual-mode variable-cycle turbine rocket engine |
| CN116181518A (en) * | 2023-05-04 | 2023-05-30 | 中国航发沈阳发动机研究所 | Interstage duct aeroengine |
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Cited By (24)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105673088A (en) * | 2016-01-19 | 2016-06-15 | 哈尔滨工业大学 | Oil cooling turbine moving blade |
| CN106194497A (en) * | 2016-06-21 | 2016-12-07 | 杨京生 | Fixing shaft type, tubular gas turbine, the turbofan aero-jet engine of single combustion chamber structure |
| CN107965397A (en) * | 2016-10-20 | 2018-04-27 | 谢爱梅 | The efficient aviation new work engine structure of multistage centrifugal turbofan |
| CN106840643B (en) * | 2017-04-06 | 2023-04-18 | 中国科学院工程热物理研究所 | Measuring device for thermal deformation of casing under impact heat exchange |
| CN106840643A (en) * | 2017-04-06 | 2017-06-13 | 中国科学院工程热物理研究所 | The measurement apparatus of casing thermal deformation under a kind of impingement heat transfer |
| CN108278165A (en) * | 2017-12-14 | 2018-07-13 | 雷静桃 | A kind of adjustable aero-engine of the direction of motion |
| CN108533418A (en) * | 2018-04-28 | 2018-09-14 | 上海啸风航空科技有限公司 | A kind of double duct mixing exhaust fanjets of novel microminiature |
| CN108626026A (en) * | 2018-04-28 | 2018-10-09 | 上海啸风航空科技有限公司 | A kind of novel microminiature fan postposition fanjet |
| CN108506111A (en) * | 2018-05-04 | 2018-09-07 | 西北工业大学 | A kind of microminiature fanjet |
| CN108506111B (en) * | 2018-05-04 | 2023-11-17 | 西安觉天动力科技有限责任公司 | Microminiature turbofan engine |
| CN109441635A (en) * | 2018-12-18 | 2019-03-08 | 王立芳 | Three axis triple channels become duct variable diameter in series and parallel and adaptively recycle to turning jet engine |
| CN109948231A (en) * | 2019-03-14 | 2019-06-28 | 中国航发湖南动力机械研究所 | Engine cycle parameters analysis method and device |
| CN112524641A (en) * | 2020-12-02 | 2021-03-19 | 西北工业大学 | Novel turbine interstage combustion chamber |
| CN113204842A (en) * | 2021-04-28 | 2021-08-03 | 中国航发沈阳发动机研究所 | Engine thermodynamic cycle design method |
| CN113204842B (en) * | 2021-04-28 | 2024-05-24 | 中国航发沈阳发动机研究所 | Engine thermodynamic cycle design method |
| CN113738530A (en) * | 2021-10-15 | 2021-12-03 | 清华大学 | Multi-duct aero-engine casing structure with blade tip fan |
| CN114542510A (en) * | 2022-02-23 | 2022-05-27 | 中国航发沈阳发动机研究所 | Self-adaptive variable-cycle engine fan rotor structure |
| CN114810350A (en) * | 2022-05-06 | 2022-07-29 | 中国科学院工程热物理研究所 | Methane precooling turbine-based combined cycle engine system with interstage combustion chamber |
| CN114810350B (en) * | 2022-05-06 | 2023-12-22 | 中国科学院工程热物理研究所 | Methane precooling turbine-based combined cycle engine system with interstage combustion chamber |
| CN115183273A (en) * | 2022-07-21 | 2022-10-14 | 中国航发沈阳发动机研究所 | Afterburning engine combustion chamber |
| CN115653790A (en) * | 2022-10-15 | 2023-01-31 | 中国科学院力学研究所 | Dual-mode variable-cycle turbine rocket engine |
| CN115653790B (en) * | 2022-10-15 | 2025-03-25 | 中国科学院力学研究所 | A dual-mode variable cycle turbo rocket engine |
| CN116181518A (en) * | 2023-05-04 | 2023-05-30 | 中国航发沈阳发动机研究所 | Interstage duct aeroengine |
| CN116181518B (en) * | 2023-05-04 | 2023-12-15 | 中国航发沈阳发动机研究所 | Interstage duct aeroengine |
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Application publication date: 20120104 |