CN103195612B - Multifunctional turbofan jet engine - Google PatentsMultifunctional turbofan jet engine Download PDF
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- CN103195612B CN103195612B CN201310118431.3A CN201310118431A CN103195612B CN 103195612 B CN103195612 B CN 103195612B CN 201310118431 A CN201310118431 A CN 201310118431A CN 103195612 B CN103195612 B CN 103195612B
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The invention discloses a multifunctional turbofan jet engine which includes a fan, a gas compressor, a second combustor, a turbine and an exhaust nozzle, wherein the fan is used for sucking air, the gas compressor is used for doing work on the sucked air, the second combustor is used for igniting gas on which the gas compressor does work, and the exhaust nozzle is used for exhausting the combusted gas. The multifunctional turbofan jet engine further comprises a first combustor, a first splitter plate and a second splitter plate, wherein the first combustor is arranged in front of the second combustor, the first splitter plate and the second splitter plate are used for splitting gas, the first splitter plate and the first combustor are correspondingly arranged, the second splitter plate and the second combustor are correspondingly arranged, and the first splitter plate and the second splitter plate are respectively provided with a travel switch I and a travel switch II which are used for rotating the corresponding splitter plate. The invention aims to provide the multifunctional turbofan jet engine which has the advantages of inlet air cooling, accurate gas splitting, balanced ignition, uniform-velocity starting, level-changing acceleration, high-efficiency, safety, thrust vectoring, easy operation and versatility.
The present invention relates to Aero-Space technical field of engines, particularly relate to a kind of turbofan adopting multi-cylinder technology.
Present aircarrier aircraft, can only fly to the region of 10 kilometers high; Fighter, generally fly to 20 kilometers of high left and right, then higher position difficult for very much.The aircraft can breaking through 20 kilometers high is considerably less.Satellite can not fly to less than 100 kilometers, because satellite runs around the earth below 100 kilometers do not have power, highly drop to apart from 100 kilometers, ground, satellite will soon fall down.At present, the proximity space in 20-100 kilometer high-altitude is a clear area substantially.How present each state all in research, develops proximity space aircraft.Who can break through and capture this commanding height ， Who and just grasp the control of the air except space.Develop proximity space aircraft, the feature of proximity space must be grasped.The feature of proximity space is: have air, but thin; Centrifugal force is less than low latitude, but larger than space.The feature of aeroengine is High Temperature High Pressure.Do you overcome centrifugal force, by high temperature? obviously be not.Can only by overcoming gravitational reaction force, this reaction force is exactly thrust concerning motor.From mode of operation: the thrust of turbofan is little, speed is slow, but fuel-efficient; The thrust of whirlpool spray is large, and speed is fast, but not fuel-efficient; What thrust was maximum works as number punching press.Thrust is mainly from high pressure, instead of high temperature.High temperature not only dilutes the density of inlet air, reduces the quality of inlet air flow, increases oil consumption, and temperature is too high also can destroy cylinder body and blade, affects the life-span of motor.Exactly because the thermal management capabilities of China's motor is not high, cooling is done good not, causes the life-span of China's motor not as Foreign Engine power.Therefore, necessary control temperature, increases thrust, acquisition speed.I thinks will develop proximity space aircraft, and the turbofan mode of operation only according to motor is not all right.Must be turbofan, whirlpool spray adds punching press, and based on press work pattern, because press work pattern, aircraft is made to obtain at a high speed with high thrust, to make the air of proximity space become no longer thin at a high speed, make to lower the temperature the density of air increase, obtain high-quality air-flow.Therefore, proximity space aircraft only has and main adopts press work pattern, just most possibly breaks through the height of more than 20 kilometers and free flight.
The great advantage of press work pattern is supersonic speed, and disadvantage is can not be automatic.Rocket carries the similar press work pattern of liquid hydrogen liquid oxygen.With general pressed engine unlike before engine charge with air inlet after its air-flow all maintain more than the hypersonic speed of 5 Mach.General pressed engine then needs flow slowing down supercharging.But airspeed is more than the hypersonic speed reaching 5 Mach, and the high pressure-temperature intensity that flow slowing down supercharging brings can exceed engine material and bear the limit.So best solution is exactly with hypersonic speed air-breathing and hypersonic speed ejection at once after burning.The static pressure static temperature be detained in such scramjet engine would not threaten motor normal operation.And one of scramjet engine key technical problems is firing technique, in hypersonic speed, add fuel and light a fire being tantamount in tornado, light a match! Scramjet engine key technology comprises the injection of fuel, blending, igniting.The airspeed flowing through scramjet engine is always ultrasound velocity, air flows through in aircraft body the holdup time usually only having several milliseconds, want within the time short like this, complete compression, supercharging, and complete low loss, high efficiency blending with fuel rapidly, uniformly and stably in supersonic flows state, point is fought, and burn be very difficult, fine or not igniting length and the heat load directly affecting motor of blending of fuel and air.Because incoming flow is uneven, the work of the firing chamber of scramjet engine is very complicated.Therefore, firing chamber design and to test the research of particularly supersonic combustion process extremely important.
According to external research, Russia to begin one's study from the sixties scramjet engine, and target is the civil transport of M number 5 ~ 7, Single Stage To Orbit space shuttle and hypersonic cruise missile.The eighties in 20th century, this research institute and the cooperation of center air fluid dynamic research Suo Deng unit have carried out the Technical Development Program of " cold " hypersonic speed, main development test rectangle and axisymmetric Dual-mode Scramjet.1991 ~ 1998 years, carried out the confirmatory flight test of 5 scramjet engines altogether, flight M number is the highest by 6.5, and motor uses hydrogen fuel.Wherein cooperate with France for the 2nd, 3 time, cooperate with the U.S. for the 4th, 5 time.It is said that second time is the most successful, the data of acquisition are the most complete.At present, this research institute is carrying out the technical research that speed is the hypersonic aircraft scramjet engine of 6 ~ 7 times of velocities of sound, and application target is military-civil hypersonic aircraft.
The U.S. carries out scramjet engine technical research country comparatively early.The sixties begin one's study, to the eighties, one of them important achievement in research is exactly so-called bimodal motor (Dual-mode scramjet), it is that one is applicable to medium flight Mach number (4 ~ 8), not only can carry out subsonic combustion but also the pressed engine of supersonic combustion can have been carried out, widen the application lower limit of scramjet engine.It is a kind of annular inlet structure, comprises subsonic speed and supersonic speed two intake ducts, and under different flight Mach numbers and fuel equivalence ratio situation, the mode that motor realizes sub-combustion and super burn automatically transforms.Current NASA, air and naval force have oneself development plan.On March 27th, 2004, unmanned research aircraft X-43A successfully reaches the speed of M number 7 in the 2nd flight test, become flying speed in the world the fastest take air breathing jet engine as the aircraft of power plant.Expecting 2025, take scramjet engine as power is main hypersonic sky and space plane, will likely come into operation.
Why is air speed fast not as rocket speed? because the heart of rocket (firing chamber) is larger than the heart (firing chamber) of aircraft.Take a broad view of the development history of air breathing engine, all that the heart (firing chamber) is little, health (fan, gas compressor, turbine, nozzle) is large, eight, after the nineties, progressed into the epoch only having a large annular combustion chamber to rule all the land, the first generation turbofan engine JT3D of the use multitube can annular type combustor only having Pu Hui company to produce the sixties is an exception.
Pu Hui company utilizes oneself technological reserve on turbojet engine, have employed the intension core-engine of very ripe J57 as new turbofan engine.JT3D employs the connular combustor of 8 burner inner liners, pushes away motor in 8 tonnes of thrusts.Boeing 707 is dress JT3C turbojet engine originally, compared with JT3C, JT3D takeoff thrust increases 50%, cruise thrust increases 27%, the oil consumption rate that cruises reduces by 13%, and the improvement brought to aircraft is thus: ultimate run increases by 27.6%, and rate of climb improves 110%, maximum cruise improves 8.2%, and takeoff distance reduces 29.4%.The use of JT3D is very wide, and be not only contained on civilian Boeing 707, DC-8 aircraft, the B-52H bomber in later stage and the military aircraft such as C-141A military transportation airplane, E-3A early warning plane also fill the military version TF-33 motor of JT3D.
It is said the large aircraft of the of great reputation fortune 10 of China's the eighties, be subject to inspiring from the research of JT3D and get out, flown more than 100 hour, flown over several big cities.It's a great pity that fortune 10 is got down from horse owing to lacking the many reasons such as leadership structure when doing two bombs and one satellite and policy input, become before China does not produce large aircraft, the pain that Chinese are eternal in the heart.The good news is that we have had fortune 20 now.But had the foreign heart (firing chamber) just not have the Chinese heart (firing chamber), key is reproduced and the strong Chinese heart (firing chamber).
Recently, the development of various countries' aeroengine has had many new developments: CE to announce its adaptive universal engine technology, and the thermal management capabilities announcing it can be managed through the air quantity of engine core part and convert by control flow check the performance of motor.Think more by the flow of core engine, the thrust of motor is larger, and speed is higher; Reduce the flow of core engine, with regard to fuel saving.
According to Britain's the Daily Telegraph newspaper November 28, utilize this new technology that the scientists of air breathing engine company of engineering company of Britain is invented, the air entering motor can be reduced to subzero 150 degrees Celsius from 1000 degrees Celsius within the time of centisecond, but also can not produce any frost.This disruptive technology is the very thin pipe utilizing a row to make " whirlpool " shape, loads onto the heat in concentrated nitrogen draw air, and makes air be cooled to subzero 150 degree before feeding motor.New refrigeration technology will allow air breathing engine with more high power safe handling, and it is overheated to there will not be, and this means that its speed can exceed 2000 miles per hour.These researchers claim, having a reusable and more efficient motor will reduce the cost of space flight greatly.They wish, this motor just can come into operation in 10 years.
Russia's chemical industry the Automation Design office (KBKhA) claims: be successfully made the new engine test that thrust is 7.5 tons in September, 2012, this motor provides power by liquid oxygen and LNG Liquefied natural gas.Russia and Italy are by cooperation research and development liquid oxygen LNG Liquefied natural gas motor.
Recently, China thrust maximum Launch Vehicle Engine of new generation---120 tonnes of oxygen kerosene high pressure afterburning cycle engines are succeeded in developing, this h type engine h using the power system as China New Generation carrier rocket, for the national key special subjects task such as manned space flight, moon exploration provides safeguard.Expert claims, and this h type engine h succeeded in developing is the high pressure afterburning cycle engine that First Chinese type has independent intellectual property right.
No matter be high pressure afterburning, or concentrated nitrogen cooling, no matter be improve thermal management capabilities, still fresh fuel is used, all provide good condition for we innovate further in raising thermal management capabilities, but we are subject to again long-term Traditional Thinking and Xi to be used to the constraint of fixed pattern, and we only make an accurate selection of break-through point, just can capture commanding height simultaneously.
Now, we are necessary to do to analyze further to the firing chamber of JT3D.Should affirm the achievement of JT3D, before the sixth of the twelve Earthly Branches did introduce.But JT3D also has problems, mainly the connular combustor of JT3D is formed a circle by 8 burner inner liners and formed, and between burner inner liner with burner inner liner, having flame-transferring tube to be connected to ensure, the outlet gas pressure of each burner inner liner is large to equal.Even the gaseous-pressure so within each burner inner liner is also or can not be completely equal, but small gaseous-pressure in each burner inner liner also deficiency think trouble.But in the outlet port of each burner inner liner because adjacent two combustion gas that burner inner liner sprays can overlap, so higher than the temperature in other places in the temperature of the outlet adjacent of each burner inner liner.The temperature contrast of the outlet temperature field of burner inner liner brings certain infringement can to the combustion gas guider of turbine front portion, and the part that temperature is high can be accelerated ablated.Such as on the JT3D of connular combustor employing 8 burner inner liners, only has 1/3rd of normal blade in the life-span of its gas diversion blade of burner inner liner wake flame overlapping.So JT3D has got down from horse like this.Proposed here Railway Project: 1, burner inner liner wake flame overlapping temperature high to blade accelerate this problem ablated really can't solve Do? does is 2, air inlet how Fen Do enters the connular combustor of 8 burner inner liners? 3, how different according to thrust needs, allow the connular combustor entering dozens or even hundreds of burner inner liner into Qi Fen Do?
Existing aeroengine, no matter be whirlpool spray or turbofan, no matter be domestic or external, all vat (firing chambers), not only also exist that temperature is high, oil consumption is large, thrust is little, part is many, speed is slow, can not significantly raise speed and the problem of speed change, more it is a risk that there is potential safety hazard, once the only large-scale combustion room (vat) of motor is broken down, tend fatal crass.
Summary of the invention
For above deficiency of the prior art, the object of the present invention is to provide a kind of air inlet to lower the temperature, accurately divide gas, balanced to light a fire, at the uniform velocity start, become level acceleration, highly effective and safe, thrust vectoring, the Multifunctional turbofan jet engine that easy to operate, purposes is various.
For achieving the above object, technological scheme of the present invention is:
A kind of Multifunctional turbofan jet engine, comprise: for suck air fan, to the gas compressor of the air acting sucked, for the jet pipe lighting the second firing chamber of gas after gas compressor acting, turbine and discharged by the gas after burning, also comprise the first firing chamber being arranged at described second combustion chamber charge end, in wherein said first firing chamber and the second firing chamber, be provided with tubular shape deflagrating jar; On the cylindrical jacket of described first firing chamber and the second firing chamber, point Do is provided with the outlet of the import of the cooling pipe of band electrically switchable grating and the cooling pipe of band electrically switchable grating;
For the first flow distribution plate and second flow distribution plate of point gas, the hole be circumferentially provided with for point gas of described first flow distribution plate and the second flow distribution plate, described hole is corresponding with the tubular shape deflagrating jar position on tubular shape deflagrating jar to be arranged and quantity is equal; The diameter in described hole and the equal diameters of tubular shape deflagrating jar;
Described first flow distribution plate and the first firing chamber are combined to form the first firing chamber modules A, and described second flow distribution plate and the second firing chamber are combined to form the second firing chamber module B;
Described first flow distribution plate correspondence is arranged at the inlet end of the first firing chamber, described second flow distribution plate correspondence is arranged at the second combustion chamber charge end, described first flow distribution plate and the second flow distribution plate being also respectively arranged with the position limit switch I for rotating described first flow distribution plate and the second flow distribution plate and position limit switch II, between described position limit switch I and position limit switch II, being also provided with the double-action mechanism that the first flow distribution plate and the second flow distribution plate are not relatively rotated;
Described second firing chamber and the first firing chamber are made up of circumferentially at least three layers of several tubular shape deflagrating jar equally distributed;
Described gas compressor is connected with turbine by central shaft, and wherein said first firing chamber modules A and the second firing chamber module B are arranged with central shaft is concentric.
Described gas compressor comprises low pressure compressor and high-pressure compressor, is provided with the import of the cooling pipe of band electrically switchable grating and is connected with gas-entered passageway after wherein said high-pressure compressor;
Described low pressure compressor is provided with provided with internal duct and external duct air balance switch, high-pressure compressor is also provided with provided with internal duct and external duct air balance switch.
The exhaust pipe of described jet pipe is provided with is with the activity of automatically controlled rotation position switches to be vented flow distribution plate and fixing bypassing exhaust plate.
Further, when circumferentially first layer tubular shape deflagrating jar n1 is 4, second layer tubular shape deflagrating jar n2 is 8, and third layer tubular shape deflagrating jar n3 is 16.
Closer, when described circumferentially third layer tubular shape deflagrating jar n3 is 16, the move angle of the position limit switch on described flow distribution plate is 45 degree.
The quantity of n1 is larger, and the total quantity of tubular shape deflagrating jar is more, and the move angle of the position limit switch on the flow distribution plate of place is less.
Advantage of the present invention and beneficial effect as follows:
1, owing to adopting concentric discs flow distribution plate technology, and position limit switch cooperating, achieve the air inlet cooling of motor, the accurate control of air inflow and uniform distribution, decrease the oil consumption of aircraft, create high Zhi and measure steady air flow, improve thermal management capabilities.
2, the n cylinder n layer owing to have employed concentric pipe divides cylinder technology, adopts multi-cylinder arrangement, is convenient to obtain actual parameter data, eliminates the drawback that motor vat is difficult to obtain actual parameter data and potential safety hazard, improves the safety coefficient of aircraft.
3, owing to have employed the module series winding technology of concentric circle pipe-type chamber, staged combustion cycle, high pressure afterburning, punching is than increasing, motor can be made by different mode of operation, the reinforcing of motor, acceleration and variable speed capability are increased substantially, add adding of Thrust Vectoring Technology, on the exhaust pipe of jet pipe, be provided with and be with the activity of automatically controlled rotation position switches to be vented flow distribution plate and fixing bypassing exhaust plate, point gear effectively controls the direction of exhaust, aircraft can be made to fly get Geng Gao, economize sooner, more, farther, more skilful, better.
4, the variable cycle engine of the present invention's invention, for the deficiency of domestic and international available engine, successfully solves the Five problems of air inlet cooling, precisely air inlet, safe navigation, boosting speed-change, thrust vectoring, cost performance is high, thrust weight ratio is large, rush than high, oil consumption is low, voyage is far away, long service life, safety coefficient is high, has dual-use wide prospect of the application, is rich in high market development and is worth, likely terminate the situation that a vat is monopolized the world, the real arrival in declaration multi-cylinder epoch.
Accompanying drawing explanation
Fig. 1 is the structural representation of a kind of Multifunctional turbofan jet engine of one embodiment of the present invention;
Fig. 2 be in the present invention the first flow distribution plate along the sectional view of A-A line;
Fig. 3 be in the present invention the second flow distribution plate along the sectional view of B-B line;
Fig. 4 is the structural representation of the first firing chamber and tubular shape deflagrating jar in the present invention;
Fig. 5 is the structural representation of the second firing chamber and tubular shape deflagrating jar in the present invention;
Fig. 6 is movable exhaust manifold structure schematic diagram in the present invention;
Fig. 7 fixes bypassing exhaust plate structure schematic diagram in the present invention.
Wherein, 1-fan; 2-1 low pressure compressor; 2-2 high-pressure compressor; 3-1 first flow distribution plate; 3-2 second flow distribution plate; 4-central shaft; 5-turbine; 6-jet pipe; 7-1 first firing chamber; 7-2 second firing chamber; 8-tubular shape deflagrating jar; 9-1 position limit switch I, 9-2 position limit switch II.10-serial verb construction; 11-1 low-pressure machine provided with internal duct and external duct air balance switch; 11-2 high-pressure unit provided with internal duct and external duct air balance switch; The import of 12-1 cooling pipe, the outlet of 12-2 cooling pipe; 13-1 activity exhaust flow distribution plate; 13-2 fixes bypassing exhaust plate; The automatically controlled rotation position switches of 13-3.
The aeroengines such as the existing Taihang of current China and the Kunlun all adopt turbofan engine to be prototype.
The invention will be further elaborated to provide the embodiment of an indefiniteness below in conjunction with accompanying drawing.
With reference to shown in Fig. 1, Fig. 2, Fig. 3, Fig. 4, Fig. 5 and Fig. 6, a kind of Multifunctional turbofan jet engine, comprise: for suck air fan 1, to the gas compressor 2 of the air acting sucked, for the jet pipe 6 lighting the second firing chamber 7-2 of gas after gas compressor acting, turbine 5 and discharged by the gas after burning, also comprise the first firing chamber 7-1 being arranged at described second firing chamber 7-2 inlet end, in wherein said first firing chamber 7-1 and the second firing chamber 7-2, be provided with tubular shape deflagrating jar 8; On the cylindrical jacket of described first firing chamber and the second firing chamber, point Do is provided with the outlet 12-2 of the import 12-1 of the cooling pipe of band electrically switchable grating and the cooling pipe of band electrically switchable grating;
For the first flow distribution plate 3-1 of point gas and the hole be circumferentially provided with for point gas of the second flow distribution plate 3-2, described first flow distribution plate 3-1 and the second flow distribution plate 3-2, described hole is corresponding with the position of tubular shape deflagrating jar 8 to be arranged and quantity is equal; The diameter in described hole and tubular shape deflagrating jar 8 equal diameters;
Described first flow distribution plate 3-1 and the first firing chamber 7-1 is combined to form the first firing chamber modules A, and described second flow distribution plate 3-2 and the second firing chamber 7-2 is combined to form the second firing chamber module B;
Described first flow distribution plate 3-1 correspondence is arranged at the inlet end of the first firing chamber 7-1, described second flow distribution plate 3-2 correspondence is arranged at the second firing chamber 7-2 inlet end, described first flow distribution plate 3-1 and the second flow distribution plate 3-2 being also respectively arranged with position limit switch I 9-1 for rotating described first flow distribution plate 3-1 and the second flow distribution plate 3-2 and position limit switch II 9-2, between described position limit switch I 9-1 and position limit switch II 9-2, being also provided with the double-action mechanism that the first flow distribution plate 3-1 and the second flow distribution plate 3-2 is not relatively rotated;
Described second firing chamber 7-2 and the first firing chamber 7-1 is made up of at least three layers of several tubular shape deflagrating jar 8 equally distributed;
Described gas compressor 2 is connected with turbine 5 by central shaft 4, and wherein said first firing chamber modules A and the second firing chamber module B are arranged with central shaft 4 is concentric;
Described gas compressor 2 comprises low pressure compressor 2-1 and high-pressure compressor 2-2, is wherein provided with the import of the cooling pipe of band electrically switchable grating after wherein said high-pressure compressor 2-2 and is connected with gas-entered passageway (figure does not indicate);
Described low pressure compressor 2-1 is provided with provided with internal duct and external duct air balance switch 11-1, high-pressure compressor 2-2 is also provided with provided with internal duct and external duct air balance switch 11-2;
Described firing chamber 7 is made up of at least three layers of equally distributed some tubular shape deflagrating jar 8;
Wherein the second firing chamber is identical with the first firing chamber size, and tubular shape deflagrating jar quantity is identical with position;
The exhaust pipe of described jet pipe is provided with the activity exhaust flow distribution plate 13-1 and fixing bypassing exhaust plate 13-2 of automatically controlled rotation position switches 13-3.
Preferably, when first layer tubular shape deflagrating jar n1 is 4, second layer tubular shape deflagrating jar n2 is 8, and third layer tubular shape deflagrating jar n3 is 16; In like manner, first layer tubular shape deflagrating jar n1 also can be 3 tubular shape deflagrating jars, then second layer n2 is just 6 tubular shape deflagrating jars, and third layer n3 is just 12 tubular shape deflagrating jars, can also continue to add combustible zone n4, n5
Preferably, when described third layer deflagrating jar n3 is 16 tubular shape deflagrating jars, the move angle of the position limit switch 9 on described flow distribution plate 3 is 45 degree.Be that the deflagrating jar of outermost one deck combustible zone and n3 is connected and burns during original state, when position limit switch is spent to right rotation 45//2=22.5, namely the cylinder of n2 layer and n3 layer is connected and is burnt, it is now turbofan form, when position limit switch is to anticlockwise 45//2=22.5 degree, then when continuing to spend to anticlockwise 45//2=22.5, namely the cylinder of n1, n2 layer and n3 layer is connected and is burnt, now all connect, speed reaches maximum;
Each tubular shape deflagrating jar is all containing hydraulic pressure oil nozzle and spark plug, when air is through intake duct, air is made to be pressed into gas compressor by fan, the flabellum blade rotation of gas compressor is done work to air-flow, the pressure and temperature of air-flow is raised, tubular shape deflagrating jar is entered by flow distribution plate after high pressure draught cooling subsequently, the hydraulic pressure oil nozzle ejection fuel oil of tubular shape deflagrating jar, spark ignitor simultaneously, produce high temperature and high pressure gas, discharge backward, high-temperature high-pressure fuel gas flows through turbine backward, can expand in turbine in part and be converted into mechanical energy, driving turbine rotates, because turbine and gas compressor are contained on same central shaft, the energy that the maximum thrust that aircraft requirements improves motor is extracted from turbine, be converted into shaft work, driving gas compressor rotates, thus the air that Repeated Compression enters, and very large MAF is provided, part high-temperature high-pressure fuel gas is discharged from jet nozzle with high speed from turbine backward through jet pipe.Thus the reaction thrust produced motor, order about aircraft through bypassing exhaust plate flight forward in certain direction.
For reaching technological scheme above-mentioned purpose of the present invention, must design from eight aspects such as cooling, point gas, point cylinder, speed change, igniting, reinforcing, module series winding, thrust vectoring.
Cooling.One is as previously mentioned, and high temperature not only dilutes the density of inlet air, reduces the quality of inlet air flow, increases oil consumption, but also can destroy cylinder body and blade, affects engine life, therefore suitably must lower the temperature to motor.Two is according to ram effect, and when speed increases, momentum drag can make net thrust reduce.On the one hand when inlet pressure increases, current density and flow quality are increased all to some extent, makes the increase of nozzle pressure ratio, and then gross thrust is increased.The rising of another aspect press temperature, will improve the gas flow temperature of inlet-pressure mechanism of qi.When engine intake temperature improves, the density of inlet air reduces, and flow quality reduces, and the energy can added in gas flow is also reducing, and thrust also can reduce, and the rising of ambient temperature makes thrust sharply decline.Notice specific thrust and unit fuel consumption rate have contrary characteristic with the increase of ambient temperature.Therefore, reinforcing be realized, must suitably lower the temperature.Three is that the cooling of concentrated nitrogen provides successful story.The new refrigeration technology of Britain will allow air breathing engine with more high power safe handling, and it is overheated to there will not be, and its speed can exceed 2000 miles per hour, and can have a reusable and more efficient motor will reduce the cost of space flight greatly.Say the necessity of the cooling that is over, retell the possibility of cooling.How to carry out cooling? one be the admission line before module flow distribution plate with firing chamber installs additional band electrically switchable grating conveying concentrate nitrogen cooling manage and be connected lead to (figure does not indicate).Two is can also have C, D at each modules A, B(, C, D are after A, B, figure does not indicate) firing chamber cylindrical jacket on open two apertures 12-1,12-2, the conveying installing band electrically switchable grating additional concentrates the pipe of nitrogen cooling and is communicated with, pore if desired as the concentrated nitrogen cooling of conveying uses, wherein, 12-1 is the import of cooling pipe, and 12-2 is the outlet of cooling pipe.These two kinds of ways when mode of operation is turbofan can, mode of operation be whirlpool spray and punching press time can use.Come with preparation, available also can, with need during preemergency.Three is that C, D module also can be used as the use of conveying concentrated nitrogen cooling pipeline.Four is because the pressure ratio of the best has equal stagnation pressure at low-pressure turbine exit with by-pass air duct air-flow, at this moment the provided with internal duct and external duct balance cock 11-2 opening low pressure compressor 11-1 and high-pressure compressor is needed, both bypass ratio was increased, be conducive to realizing low-pressure turbine exit and have equal stagnation pressure with by-pass air duct air-flow, suitably reduce engine intake temperature again.
Divide gas.Because flow distribution plate and firing chamber and central shaft are that unique concentric disjunctor of disk (hollow) construct, can by the automatically controlled rotation position switches manipulation of flow distribution plate, precisely control air inlet area, the tubular shape deflagrating jar (firing chamber) high quality air-flow being entered specify.Point pneumoelectric control rotation position switches achieves from all layering tubular shape deflagrating jar area of contact that minimum (absolute value is 0, and this time-division pneumoelectric control rotation position switches is in closed condition.), the accurate control moved in circles between maximum to all layering tubular shape deflagrating jar area of contact, air inlet area is larger, and high pressure ratio is less, and air inlet area is less, and high pressure ratio is larger.Start with from minimizing suction port area unlike, this case with west thinking, realize accurately controlling air inlet, just improve high pressure ratio, and economical.Increasing the most effective method of thrust is exactly mass flow rate by increasing motor.It is emphasized that, the effect of the flow distribution plate controlled by automatically controlled rotation position switches is outstanding: one is under the effect of flow distribution plate, inlet total pressure distortion can not be there is, inlet swirl flow dis tortion can not occur, the even situation of inlet total temperature Bu Zhuo can not occur, high Zhi will be produced and measure steady air flow, improve thermal management capabilities.Two is that the time-division stream plate air inlet area that starts to light a fire is less, and tubular shape deflagrating jar is lighted a fire higher into shop rate.Improve owing to entering shop rate, solve the ignition problem of tubular shape deflagrating jar.Three is solve JT3D Problems existing; divide owing to taking same layer tubular shape deflagrating jar interval function; carry out with interlayer every one semi-circle tubular deflagrating jar burning; with interlayer every second half tubular shape deflagrating jar to misfire burning; in the outlet port of each burner inner liner; efficiently avoid the combustion gas that two adjacent burner inner liners spray to overlap, effectively protect turbine blade, extend the working life of turbine blade.Four is because module flow distribution plate has had accurate control to main duct air inlet, there is again the provided with internal duct and external duct air balance switch on low compressor and high pressure mechanism of qi controlling, can precisely control provided with internal duct and external duct air balance, can make motor under different environmental conditionss, according to the difference of executing the task, normally work by different mode of operations (turbofan, whirlpool spray and punching press) of starting, add the manipulation of the multifactor cloud computing of computer and digital control panel, thus aircraft is possessed adaptivity function, become a variable cycle engine.
Divide cylinder.From west thinking, reforming key point is placed on outside firing chamber different, I is placed on reforming key point on firing chamber.Divide how many tubular shape deflagrating jars suitable actually? does is the diameter of tubular shape deflagrating jar how many? does is the length of tubular shape deflagrating jar how many? same tubular shape deflagrating jar or size tubular shape deflagrating jar mixing arrangement? contact or and connect? the how oiling of tubular shape deflagrating jar? how to light a fire? the present invention proposes the Design Conception of disk concentric disjunctor variable cycle engine: on hollow concentric disjunctor disk, equally distributed several tubular shape deflagrating jars (Small Combustion room) of being kept off control by point n cylinder n floor 4 become radial arrangement to combine.Due to the circular hole of first, second flow distribution plate described and diameter, the structure identical with position (except position limit switch) of first, second firing chamber tubular shape deflagrating jar; Described first flow distribution plate and the first firing chamber form a modules A; Described second flow distribution plate and the second firing chamber form a module B.When we are when solving point gas problem, meanwhile, also just solve a point cylinder problem.The principle of point cylinder, layering and point gear: must be with one heart+equidistant+and connect+outer tubular shape deflagrating jar quantity multiplication (or quantity is equal)+4 gear.In technological process, high temperature high voltage resistant wanted by the endless tube of tubular shape deflagrating jar, endless tube inwall adopts THE BARCHAN DUNE VORTEX technology to increase contact surface area, first reserve the position of oil nozzle and spark plug, after burn-on, the installation of tubular shape deflagrating jar is carried out one by one successively by the order of n1, n2, n3, n4, the surface of tubular shape deflagrating jar of finally burn-oning again and cylindrical jacket.Because first, second flow distribution plate has point gas position limit switch carrying out interlock, can realize from tubular shape deflagrating jar area of contact that minimum (absolute value is 0, and this time-division gas position limit switch is in closed condition.) maximum to tubular shape deflagrating jar area of contact between the accurate control that moves in circles, like this, a point cylinder just solves oil consumption problem, and this is one.When n1 is 8, n2 is 16, n3 is 32, n4 is just 64 ... in numerous tubular shape deflagrating jars (little cylinder), just in case the tubular shape deflagrating jar of Ge Do (little cylinder) breaks down, also there will not be fatal crass's tragedy to occur, this is two.Meanwhile, owing to there being numerous tubular shape deflagrating jar (little cylinder), just established good basis for we solve speed change problem, this has been three.
Speed change.For the problem that former aeroengine can not significantly raise speed, how significantly or exponentially to raise speed? mentality of designing of the present invention is by position limit switch as required, divides the air inlet of all tubular shape deflagrating jars in 4 gear management and control gears, igniting and discharge.The position limit switch of 1 level can only be opened at every turn.Such as: open the 1st gear, the tubular shape deflagrating jar of outermost one deck is started working; Open the 2nd gear, the tubular shape deflagrating jar of outermost two layers is started working; Open the 3rd gear, the tubular shape deflagrating jar coming back to outermost one deck is started working; Open the 4th gear, the tubular shape deflagrating jar of structure at all levels is all started working; Then oppositely return.Like this, just successfully speed change is achieved.
Igniting.For anti-turbulent flow occurs, igniting must be carried out by symmetrical equilibrium.Because tubular shape deflagrating jar is too much, igniting-throttle interlock technology can be adopted to each tubular shape deflagrating jar, being concatenated between the igniting-throttle of each tubular shape deflagrating jar with a wire, when occurring lighting a fire, will automatically closing tubular shape deflagrating jar throttle because circuit is obstructed.Disk increases by 1 spark plug for subsequent use separately all again at all tubular shape deflagrating jar afterbodys, carries out dual ignition (punching press pattern).The igniting of tubular shape deflagrating jar afterbody, according to flight environment and task needs, can light a fire, also can misfire.The benefit done like this is: even if Ge Do tubular shape deflagrating jar breaks down, and affects aircraft also not too largely and runs, more there will not be the tragedy of fatal crass to occur.
Afterburning.First, due to the One's name is legion of tubular shape deflagrating jar, the quantity using position limit switch to increase tubular shape deflagrating jar just achieves reinforcing.Open whole tubular shape deflagrating jar switch just more to power.Secondly, use and become cylinder technology, increase and increase length and the capacity of tubular shape deflagrating jar cylinder body, also can realize reinforcing.Again, the mode of operation (turbofan, whirlpool spray and punching press) more reengined, can realize reinforcing better.
Module is contacted.When similar tubular shape deflagrating jar modules A, B, C, D contact, the concept of the first and second firing chambers is relative, can transform.Time normal, A module is that the first flow distribution plate adds the first firing chamber 3-1+7-1, and B module is that the second flow distribution plate adds the second firing chamber 3-2+7-2, and C module can be used as tail pipe burner, and D module can become breather line.If A module is injured in fight, B, C module can be used as the first and second firing chambers, and D module becomes breather line.When A, B module be the first flow distribution plate and the first firing chamber 3-1+7-1 and the second flow distribution plate and the second firing chamber 3-2+7-2 time, C, D module also can be made air-flow path and use, or the pipeline of lowering the temperature as the concentrated nitrogen of conveying is for subsequent use.My theory is that each tubular shape deflagrating jar of each module will be accomplished: complete (fuel conducting tube, cooling conducting tube, oil nozzle and igniter etc.), comes with preparation, available also can, with need during preemergency.The benefit of module series winding is a lot: one is power and energy.Each tubular shape deflagrating jar can work as firing chamber use, also can only when air-flow path use, and modules and each tubular shape deflagrating jar, can realize taking turns to operate, and increases working life.Seem waste like this, but from the aircraft of protection costliness, from the more expensive pilot of protection, be completely necessary.Meanwhile, because C, D module adds the power and energy of motor, necessary especially.Two is that oil consumption is successively decreased, and thrust increases progressively.If the oil consumption of A module is the oil consumption of 1, B module be likely just the oil consumption of 0.5, C module be likely just the oil consumption of 0.2, D module is likely just 0.1; And thrust is between A, B, C, D module, significantly promote successively as tail pipe burner.Three is due to likely by pilot and computer to the flight environment of outside (highly, brightness, wind-force, wind direction), enemy's situation (type, distance, speed, angle), task (is destroyed, forced landing) and the situation (performance of self motor, means) etc. multifactorly carry out cloud computing automatic control, due to several similar tubular shape deflagrating jar modules of likely contacting, staged combustion cycle can be carried out as rocket, high pressure afterburning, increase punching ratio, thrust (in theory) can be made can to realize from 0 to infinitely-great acceleration, the deceleration from infinity to 0 can be realized again, (in theory) absolute liBerty circulation achieving from 0 to infinity, this is concerning space shuttle, even more important.
Thrust vectoring.Thrust vectoring is a comprehensive very strong technology, and it comprises thrust steering nozzle technology and airframe/propelling/control system integrated technique.Do not adopt the aircraft of Thrust Vectoring Technology, the jet flow of motor is all that also along axis forward, the thrust of motor is just for the resistance overcome suffered by aircraft and the power provided needed for aircraft in this case for the thrust of generation with the dead in line of aircraft.And adopt the aircraft of Thrust Vectoring Technology, be then deflected by nozzle, the thrust utilizing motor to produce obtains unnecessary control moment, realizes the gesture stability of aircraft.Its outstanding feature is that control moment and motor are closely related, and not by the impact of the attitude of aircraft own, can ensure that aircraft is at low speed, At High Angle of Attack maneuvering flight and a few near failure of primary control surface time utilize thrust vectoring to provide additional manipulation moment to control air maneuver.Vectored thrust technological merit is a lot, it can provide the maneuverability under post stall maneuver ability and At High Angle of Attack, F/A-22, the outstanding opportunity of combat such as Su-37 is exactly use the masterpiece of thrust vectoring, emulation air battle was once done with X-31 validator and F/A-18 by the U.S., result military success is 32:1, its advantage has some idea of, and use the fighter of thrust vectoring all to have excellent STOL ability, Thrust Vectoring Technology improves the control efficiency of aircraft in addition, the pneumatic control of aircraft can be made to look like vertical fin and vertical tail reduces greatly, aircraft weight can be alleviated, improve the stealth capabilities of aircraft.How to accomplish thrust vectoring? the bypassing exhaust plate being with automatically controlled rotation position switches is installed in this case design additional in jet pipe exhaust pipe, and regulation and control exhaust airstream flows to, with jet pipe moving vane close fit, and conversion and control discharge directions (see Fig. 6 and Fig. 7).Bypassing exhaust plate forms with movable exhaust flow distribution plate 13-1 with one heart by fixing bypassing exhaust plate 13-2,13-2 with 13-1 diameter is identical, and the distance of concentric is equal.Wherein, fix bypassing exhaust plate 13-2 by the balanced equidistant arrangement of several tap holes on bypassing exhaust plate, and be fixed on jet pipe exhaust pipe; Movable exhaust flow distribution plate 13-1 by the tap hole be less than on fixing bypassing exhaust plate 13-2 and with it balanced equidistant arrangement on bypassing exhaust plate, the tap hole of movable exhaust flow distribution plate 13-1 is with some tap hole equal diameters (except automatically controlled rotation position switches) of fixing bypassing exhaust plate 13-2, running orbit is identical, control rotating distance with automatically controlled rotation position switches 13-3,360 degree of rotations can be carried out.The invention will be further elaborated to provide the embodiment of an indefiniteness below in conjunction with accompanying drawing.When fixing bypassing exhaust plate 13-2 by the balanced equidistant arrangement of 8 tap holes on bypassing exhaust plate, and be fixed on jet pipe exhaust pipe, movable exhaust flow distribution plate 13-1 is by 3 adjacent equal diameters, the identical tap hole composition of running orbit, 360 degree of rotating distances of flow distribution plate 13-1 are vented by automatically controlled rotation position switches 13-3 control activity, the tap hole often crossed on fixing bypassing exhaust plate 13-2 can be used as a gear, first grade for three holes shown in Fig. 7 be downwards original state, if counterclockwise rotate to the right, if now initial position first grade must be got back to left, and then be rotated clockwise to predetermined gear, be divided into 8 gear conversion and control discharge directions, along with 3 tap holes of activity exhaust flow distribution plate 13-1 angle clockwise or inhour rotate change, just define certain clockwise or counterclockwise vectored thrust, aircraft just can be followed one's bent up and down to shift gears and be converted different flight paths, this is conducive to the flexibility and the Stealth Fighter that improve aircraft.
In the present invention, the multi-cylinder design concept of motor is equally applicable to turbojet engine and pressed engine, can carry out corresponding change, comprise in the present invention equally as the prototype of motor.
These embodiments are interpreted as only being not used in for illustration of the present invention limiting the scope of the invention.After the content of reading record of the present invention, technician can make various changes or modifications the present invention, and these equivalence changes and modification fall into the scope of the claims in the present invention equally.
1. a Multifunctional turbofan jet engine, comprise: for sucking the fan (1) of air, to the gas compressor (2) of the air acting sucked, for lighting second firing chamber (7-2) of the gas after gas compressor acting, turbine (5) and the jet pipe (6) that the gas after burning is discharged, it is characterized in that: also comprise the first firing chamber (7-1) being arranged at described second firing chamber (7-2) inlet end, tubular shape deflagrating jar (8) is provided with in wherein said first firing chamber (7-1) and the second firing chamber (7-2), on the cylindrical jacket of described first firing chamber and the second firing chamber, point Do is provided with the outlet (12-2) of the import (12-1) of the cooling pipe of band electrically switchable grating and the cooling pipe of band electrically switchable grating, the exhaust pipe of described jet pipe (6) is provided with activity exhaust flow distribution plate (13-1) and the fixing bypassing exhaust plate (13-2) of the automatically controlled rotation position switches of band (13-3),
For the first flow distribution plate (3-1) and second flow distribution plate (3-2) of point gas, described first flow distribution plate (3-1) and the second flow distribution plate (3-2) are provided with the hole for point gas, and described hole is corresponding with the position of tubular shape deflagrating jar (8) to be arranged and quantity is equal; The diameter in described hole and tubular shape deflagrating jar (8) equal diameters;
Described first flow distribution plate (3-1) correspondence is arranged at the inlet end of the first firing chamber (7-1), described second flow distribution plate (3-2) correspondence is arranged at the second firing chamber (7-2) inlet end, described first flow distribution plate (3-1) and the second flow distribution plate (3-2) are also respectively arranged with the position limit switch I (9-1) for rotating described first flow distribution plate (3-1) and the second flow distribution plate (3-2) and position limit switch II (9-2), the double-action mechanism (10) that the first flow distribution plate (3-1) and the second flow distribution plate (3-2) are not relatively rotated also is provided with between described position limit switch I (9-1) and position limit switch II (9-2),
Described second firing chamber (7-2) and the first firing chamber (7-1) are made up of circumferentially at least three layers of several tubular shape deflagrating jar (8) equally distributed;
Described first flow distribution plate and the first firing chamber are combined to form the first firing chamber modules A, and described second flow distribution plate and the second firing chamber are combined to form the second firing chamber module B;
On the cylindrical jacket of described first firing chamber modules A and the second firing chamber module B, point Do is provided with the outlet (12-2) of the import (12-1) of the cooling pipe of band electrically switchable grating and the cooling pipe of band electrically switchable grating;
Described gas compressor (2) is connected with turbine (5) by central shaft (4), and wherein said first firing chamber modules A and the second firing chamber module B are arranged with central shaft (4) is concentric.
2. Multifunctional turbofan jet engine according to claim 1, it is characterized in that: described gas compressor (2) comprises low pressure compressor (2-1) and high-pressure compressor (2-2), be provided with the import of the cooling pipe of band electrically switchable grating after wherein said high-pressure compressor (2-2) and be connected with gas-entered passageway.
3. Multifunctional turbofan jet engine according to claim 2, it is characterized in that: described low pressure compressor (2-1) is provided with provided with internal duct and external duct air balance switch (11-1), high-pressure compressor (2-2) is also provided with provided with internal duct and external duct air balance switch (11-2).
4. Multifunctional turbofan jet engine according to claim 1, is characterized in that: when circumferentially first layer tubular shape deflagrating jar n1 is 4, and second layer tubular shape deflagrating jar n2 is 8, and third layer tubular shape deflagrating jar n3 is 16.
5. Multifunctional turbofan jet engine according to claim 4, it is characterized in that: when described circumferentially third layer tubular shape deflagrating jar n3 is 16, the move angle of the position limit switch (9) on described flow distribution plate (3) is 45 degree.
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|Application Number||Priority Date||Filing Date||Title|
|CN201310118431.3A CN103195612B (en)||2013-04-08||2013-04-08||Multifunctional turbofan jet engine|
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|CN201310118431.3A CN103195612B (en)||2013-04-08||2013-04-08||Multifunctional turbofan jet engine|
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|CN104373205A (en) *||2013-08-16||2015-02-25||袁丽君||Novel engine|
|CN105484898A (en) *||2015-12-25||2016-04-13||中国航空工业集团公司沈阳发动机设计研究所||Mode switching device of variable cycle engine|
|US9777633B1 (en) *||2016-03-30||2017-10-03||General Electric Company||Secondary airflow passage for adjusting airflow distortion in gas turbine engine|
|CN106368851A (en) *||2016-09-13||2017-02-01||中国民用航空飞行学院||Multi-fan propelling device|
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|GB1436796A (en) *||1972-08-22||1976-05-26||Mtu Muenchen Gmbh||Gas turbine ducted fan engines of multi-shaft and multi-flow construction|
|US20060086078A1 (en) *||2004-10-21||2006-04-27||Paul Marius A||Universal Carnot propulsion systems for turbo rocketry|
|US7752836B2 (en) *||2005-10-19||2010-07-13||General Electric Company||Gas turbine assembly and methods of assembling same|
|US7730714B2 (en) *||2005-11-29||2010-06-08||General Electric Company||Turbofan gas turbine engine with variable fan outlet guide vanes|
|US20090067993A1 (en) *||2007-03-22||2009-03-12||Roberge Gary D||Coated variable area fan nozzle|
|CN203214192U (en) *||2013-04-08||2013-09-25||魏汉章||Multi-functional turbofan engine|
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