CN1692216A - Integrated bypass turbojet engines for air craft and other vehicles - Google Patents

Abstract

Turbojet engines (10) and aircraft configurations (40, 50, 60, 70, 80, 200, 220, 230, 250, 280, 300) for advantageous use of the turbojet engines ; the turbojet engines utilizing ram air turbine unuts that centrifugally compress air isothermally for use in various combustion configurations designed for stoichiometric combustion, wherein a stream of by-pass ram-air jets is mixed with combustion gas jets for discharge in a common discharge nozzle.

Description

Be used for the integrated bypass turbojet engine on aircraft and other means of transportation

Background of invention

The application requires the right of following patent application: U.S.No.10/383462, on March 6th, 2003 submission day, and U.S.No.10/337032, submitting day to is on January 6th, 2003, and U.S.No.10/292829, submitting day to is on November 12nd, 2002.

The application requires the right of following provisional application: U.S.No.60/372618, on April 15th, 2002 submission day, and U.S.No.60/374737, submitting day to is on April 23rd, 2003, and U.S.No.60/405460, submitting day to is on August 23rd, 2002.

Technical field

The present invention relates to the application on aircraft and other means of transportation of a kind of novel turbojet engine and turbojet engine.This turbojet engine is such one type, and wherein bypass fan blade and compressor fuse, and compressor be connected by the turbine blade that annular combustion chamber surrounded.Turbojet engine of the present invention is the improvement to the described motor of following patent: U. S. Patent 4845941, title are " gas turbine engine working procedure ", announce on July 11st, 1989; U. S. Patent 5003766, title are " gas turbine engine ", announce on April 2nd, 1991; U. S. Patent 5177954, title are " gas turbine engine that has cooling turbine bucket ", announce on January 12nd, 1993; U. S. Patent 5341636, title are " gas turbine engine method of work ", announce on August 30th, 1994.

Background technique

Traditional turbojet engine is made up of independent module, comprises the bypass manifold fan, axial and centrifugal compressor, firing chamber, and combustion gas turbine.When coaxially assembling, these modules of conventional turbine air breathing engine and part make up and have formed a long narrow motor, lack the required compactedness of multiple application as herein described.

The bypass fan is main propulsion die, and other all modules and parts co-operation and finally drive this module.The complexity of modern turbojet engine has reached the limit, and cost has surpassed the limit that most of in the world developers can afford.

The thermal efficiency of conventional turbine air breathing engine is limited in 30% when full load, be limited near 10% when sub load.

Specific power is limit by the maximum temperature of firing chamber.At 25% o'clock of the maximum chemical stoichiometric level, air fuel ratio was 60/1, rather than stoichiometry 15/1, and this makes that the size of all turbojet engines is 4 times of motor that are operated in the stoichiometry level at least, and the result is expensive, inefficient.

For military project application, particularly unmanned aircraft and cruise missile, the high cost that the aircraft that is designed to lose in fight is paid is main spending of of national defence and state budget.

Main purpose of the present invention provides a kind of high efficiency turbojet engine, this turbojet engine carries out isothermal compression to the suction air that a part is used for stoichiometric(al) combustion, in conjunction with bypass air the pressurized air of delivering to the firing chamber is cooled off simultaneously.

As this application (continuation-in-part application) that continues part, the additional structure that has comprised the turbojet engine of type described in patent application 10/292829 (submitting day on November 12nd, 2002 to), the theme embodiment of the turbojet engine in the application's book is designed in the high-altitude commercialization and military spacecraft of atmosphere and space flight.Two embodiments' of this high-altitude turbojet engine common ground has been to use the rocket promotion, wherein in atmospheric flight, uses the oxygen in the atmosphere, and in high-altitude and space flight, uses to concentrate liquid oxygen.

These hybrid systems provide a kind of general propulsion system, can be applicable in a series of military project and commercial the application, and have reduced implementation cost largely.

Hindering gas turbine engine and turbojet engine, to have given play to high performance biggest obstacle be the restriction of combustion gas maximum temperature and the restriction of the pressure ratio of air compression in engine cycles.The metallic character of gas-turbine blade and to the limited cooling scheme of gas-turbine blade has seriously limited the intake temperature of turbo machine.And for intake temperature is remained in the scope of modern turbine blade designing institute acceptance, needing high air fuel ratio be 50/1 to 60/1.Embodiment with this turbojet engine of rocket ability solves these problems, and can provide maximum absolute thermodynamic property for the aircraft of atmosphere and space flight.

Summary of the invention

The present invention relates to turbojet engine, the present invention is integrated into a combination bypass fan single, that pipeline links to each other with all functions of traditional turbojet engine module, this fan has common fan blade and inner centrifugal pressing chamber, links to each other with peripheral turbine blade and is surrounded by an annular combustion chamber.The common jet stream of the circulation combustion gas of bypass gas flow and annular combustion chamber is finished driving jointly.

The integrated overall dimensions that greatly reduces complexity, weight, cost and pushing unit of this module.These reductions make and can work in the stoichiometry level and maximize the absolute heat energy ability and the thermal efficiency.This ability is by isothermal ground cooled compressed gas and in a particular embodiment by fuel being injected in the turbine blade and the result of inner cooling turbine bucket.

Therein among embodiment, an axial motor compressor and a centrifugal electronic wind eddies wheel Duct-Burning Turbofan that is attached thereto keep constant compression and combustion pressure with the identical Placement of energy under all states, make to produce a constant maximum thermal efficiency.

In certain embodiments, main turbogenerator of the present invention act as starter motor with one or angle of rake conventional turbine air breathing engine combines.

In other embodiments, the rotatable parts of this motor not with contacts fuel gas, thereby make turbogenerator keep cooling.

Among embodiment, turbogenerator is designed to the navigation propulsion system therein, and this system can append on the naval ship, as the main or auxiliary propelling energy, by air jet being produced propelling force forward in the water.

Fu Jia embodiment is designed to allow the turbojet engine of described type to be used for high-altitude and space flight in this article.

Although once proposed to be useful on the commerce application of high-altitude or space flight, in rocket system, use the superelevation cost of liquid oxygen to make this rocket flight not accepted by the commerce application.Even having in the atmospheric flight of a large amount of available oxygen around, traditional rocket system also uses liquid oxygen.Be adapted to the normally noticeable inadequately and shortage imagination of practical commercial plan of space flight.For example, the high speed forwarding plan is only imagined in ionosphere and is stopped blink, and most flight then has in the atmospheric layer or upper atmosphere of a large amount of oxygen around.

Two kinds of solutions that proposed herein, its common characteristic is: have such ability, in atmospheric flight, use the rocket propulsion of aerial oxygen, and in high-altitude or space flight, will concentrate the liquid oxygen conversion, make these two kinds of solutions become the general propulsion system that can be applicable to all application.So for the first time the propelling of commercial aviation is become a reality.

Second common feature of these two embodiments is when motor uses the atmospheric air burning, compare 100-200/1 by in engine cycles, carrying out the air compression that isothermal compression obtains a superelevation, make absolute thermodynamic efficiency obtain maximization, reach 80% approximately.

Two embodiments' the 3rd common feature is the ability that can burn in the stoichiometry level, thereby the system that makes has maximized absolute thermodynamic power propulsion capability.These and other characteristics will seem clearer in following DETAILED DESCRIPTION OF THE PREFERRED:

Versatility of the present invention illustrates in the example that comprises aircraft in the present invention.

Description of drawings

Fig. 1 is a kind of schematic diagram of integrated bypass centrifugal turbine air breathing engine.

Fig. 2 is the schematic diagram of the axial centrifugal turbine air breathing engine of a kind of integrated bypass.

Fig. 3 is the schematic diagram of the axial electronic centrifugal turbine air breathing engine of a kind of integrated bypass.

Fig. 4 is the schematic diagram of the jet cruise missile of the motor-driven ahead turbine of a kind of integral body.

Fig. 5 is a kind of schematic diagram of whole motor traction turbojet cruise missile.

Fig. 6 is the schematic diagram of the motor-driven dual turbojet engine of a kind of integral body.

Fig. 7 is the schematic diagram of the motor-driven wing formula of a kind of integral body turbojet engine.

Fig. 8 is the schematic diagram of the motor-driven VTOL of a kind of integral body (vertical takeoff and landing) wing formula turbojet engine.

Fig. 9 is a kind of schematic diagram of two-stage turbine Duct-Burning Turbofan.

Figure 10 is the T-S chart of the pressure ratio of corresponding two kinds of levels.

Figure 11 is a kind of two-stage turbine circuit schematic diagram.

Figure 12 is a kind of two-stage turbine circuit schematic diagram that has the backward rotation compressor that is driven by preposition no air turbine.

But Figure 13 is a kind of schematic diagram that has the dual bypass turbofan of transition loop.

But Figure 14 is a kind of schematic diagram that has transition loop, subsonic speed to ultrasonic triple bypass turbofans.

Figure 15 A is the schematic diagram that common flight is converted to the VTOL ability.

Figure 15 B is the schematic diagram that the conversion among Figure 15 A is converted to the VTOL ability.

Figure 16 A is the principle plan view of a rectangle universal machine moving platform.

Figure 16 B is the principle side view of the universal machine moving platform shown in Figure 16 A.

Figure 17 A is the principle plan view that has the general maneuvering-vehicle of propulsion die on the fuselage.

Figure 17 B is the principle rear view of the general maneuvering-vehicle shown in Figure 17 A.

Figure 18 A is the principle plan view that has the general maneuvering-vehicle of propulsion die on the wing.

Figure 18 B is the principle rear view of the general maneuvering-vehicle shown in Figure 18 A.

Figure 19 is a kind of schematic diagram of comprehensive VTOL aircraft.

Figure 20 A is a principle plan view that has the general motor-driven lengthening means of transportation of a plurality of vector jet plane.

Figure 20 B is the principle side view of the general motor-driven lengthening means of transportation shown in Figure 20 A.

Figure 21 is the perspective view of the turbojet engine of a kind of fundamental type as shown in Figure 1.

Figure 22 is a kind of schematic diagram that is used for the turbojet engine of navigation propulsion system.

Figure 23 is the enlarged view that is used for the turbojet engine of navigation system as shown in figure 22.

Figure 24 is the schematic diagram that has the turbine bypass turbofan engine of the plurality of oxygen injectors that is used for space propultion.

Figure 25 is the schematic diagram that is in three grades of turbojet engines in the turbine circulation.

Figure 26 is the turbojet engine that is positioned at the turbine the gondola a kind of and combination of conventional turbine air breathing engine.

Figure 27 is as shown in figure 26 turbojet engine is eliminated hot turbine blade through revising a schematic diagram.

Figure 28 is turbojet engine and schematic diagram that common starter motor makes up as shown in figure 27.

Figure 29 is the turbojet engine shown in Figure 27 and 28 shows the split rotor structure through amplifying a schematic diagram.

Figure 30 is a schematic diagram of eliminating the improvement turbojet engine of hot turbine blade.

Figure 31 is the hyperpressure ratio that has peripheral blade-tip turbine and parallel burning, the cross-sectional schematic diagram of punching press gas turbine rocket motor.

Figure 32 has central authorities' burning and eliminates the hyperpressure ratio of hot rotatable parts, the cross-sectional schematic diagram of punching press gas turbine rocket motor.

Figure 33 is the hyperpressure ratio that has convertible turbine/rocket-propelled additional central combustion system, the cross-sectional schematic diagram of punching press gas turbine rocket motor.

Figure 34 is the cross-sectional enlarged view of the part of motor as shown in figure 33.

Figure 35 is a kind of convertible turbine/rocket-propelled hyperpressure ratio, punching press gas turbine rocket motor that substitutes additional central combustion system that have.

Figure 36 is the cross-sectional enlarged view of the part of motor as shown in figure 35.

Embodiment

Turbojet engine 10 of the present invention describes in a plurality of embodiments with reference to accompanying drawing.Should be understood that, principles illustrated among the figure generally is half a engine structure of the motor that shows axial symmetry of the mode with sectional elevation, and the fan blade of being mentioned, turbine blade and compressor blade are represented with an individual blade or a part with blade of multiple function usually.

A kind of integrated bypass centrifugal turbine air breathing engine has been described in Fig. 1, by label 10 expressions, a rotor disk unit 11 is wherein arranged, this unit comprises the whole fan unit 10.a of a combination, this comprises as holistic unit: a fan 11.a, centrifugal pressing chamber 11.b, and the turbine blade 11.c that links together with outer cylinder 12, described firing chamber 12 has a fuel injector 13, the inside shell block 15 of 14, one porous of an outside shell block, nozzle 16, and an airspace 17, this ventilation system 17 is communicated with centrifugal chamber 11.b by hole 18.

The inwall of turbine blade 11.c is cooled off by the fuel that sparger 19 is ejected into inner passage 11.d, is also cooled off by the centrifugal pressurized gas by chamber 11.b in addition.

Fuel is from sparger 13 and inner fuel passage 20 and provide at first, and described fuel channel comes from central shaft 21 and fuel connecting line 22.

The rotor disk unit 11 of combination is supported by bearing 23, and bearing 23 is connected with center body 24, and the center body is by inner prop 25,26,27 supports, and inner prop 25,26,27 are connected on the shell block structure 28 with the form of turbo jet engine gondola.

Air enters into fan blade 11.d by opening 29, and in air chamber 11.b by centrifugal compression, be cooled by contacting then with rotor disk unit 11.

Main shaft by rotor disk unit 11 arrives isothermal level, consumption of energy minimum to the air-flow concentrated area with the radially centrifugal air cooling of compression process.Pressurized gas partly is to arrive in the porous air follow board space 17 by through hole 18 to be directed in the firing chamber 12 again, part then is radially to be directed to the firing chamber 12 from the inside of turbine blade 11.c, wherein the fuel mix among the pressurized gas in turbine blade 11.c and the passage 11.d that is injected into blade 11.c.

The mixture that is rich in pressurized gas and evaporative cooling fuel in turbine blade 11.c is from the tip of blade and be directly injected to the firing chamber 12, after burning, combustion gas is discharged by nozzle 16 at high speed, thereby drive turbine blade 11.c and fan, finally from outside propulsion jet nozzle 8, join in the main jet gas, be used for mixing at common jet nozzles 9 by bypass fan 11.a.

Novel turbojet engine of the present invention is the simplest, effective, powerful and motor cheaply, mainly is designed to in-flight.

Described same integrated bypass turbojet engine 10 in Fig. 2, this motor adds and has an axial compressor 30.Identical among every other parts and Fig. 1, and have same function.

In Fig. 3, special structure is a novel axial compressor 31, and this compressor is driven by a motor 32.The whole fan unit 10.a that has the combination of a rotation fan 11a and centrifugal compressor and turbine 33 and be combined into one, the structural similarity of its structure and above-mentioned rotor disk unit 11, and be used to drive motor generator set 34.Motor generator set 34 is connected electrically on the control box 35, and is connected on the motor 32, is used for to motor 32 power supplies, and drives backward rotation axial compressor 31.

Pressure transducer 36 is indicated the pressure in the firing chamber in real time.If pressure descends when fan compression turbine 33 low speed, control box 35 will be given an order, and improves the rotating speed of axial compressor 31, thereby is effective constant level with pressure recovery.Surpass predetermined level if pressure rises to, control box 35 will be given an order, and reduces the rotating speed of axial compressor 31.

Final result is the thermal efficiency that can both keep maximum under all loads and state.

First application of bypass turbojet engine 10 with the fan unit 11.a that is combined into one has been described in Fig. 4, by with turbojet engine 10 as turbojet engine 41 and with the mode of action of " advance A ", append on the fuselage 43 of cruise missile 40 by hinged universal ball joint 42, be used to drive cruise missile 40.This cruise missile 40 has a cross 44 that has rudder 45.The angle of turbojet engine 41 connects and deviation, together produces the full maneuverability that has comprehensive instant performance with cross rudder 45.This aircraft can be by the emission of tubulose pad.

Second application that is used to drive cruise missile 50 described in Fig. 5, by turbojet engine 10 is appended on the fuselage 53 of cruise missile by a hinged universal ball joint 52 with the mode of action of " traction B " as turbojet engine 51, this cruise missile has a cross 54 that has rudder 55, is used to produce full maneuverability.This embodiment can start on any pad and launch.

The 3rd and general maneuvering-vehicle application have been described in Fig. 6, wherein turbojet engine 10 is as two turbo jet engines 61 and 62, these two jet plane are hinged on the both sides of means of transportation 63 rotationally, make that means of transportation 63 can be in any direction motion of A and B, comprise vertical, level and rotation, and the combination in any of above-mentioned various directions.This is a kind of absolute motor-driven aircraft.

Described four application of turbojet engine 10 as turbo jet engine 71 and 72 in Fig. 7, wherein these two jet plane are connected on the fuselage 74 of the aircraft 70 with wing 73 by universal joint.This aircraft 70 also can have the all-purpose maneuverability, and this is the most important change of contemporary aircraft.

In Fig. 8, described the 5th application that turbojet engine 10 is used for a kind of VTOL wing formula means of transportation 80, wherein be positioned at the place ahead of VTOL wing formula means of transportation 80 in the integrated turbine bypass jet plane 81 described in first embodiment of the invention described above.This VTOL (vertical takeoff and landing) wing formula means of transportation 80 has a gas-entered passageway 82, and this gas-entered passageway is coated on the front side of fuselage fully, to avoid the detection of radar.The air fuel gas propelling force is by variable deflection mouth 86a and 86b and be deflected right passage 84 and the left passage 85 from central passage 83.The fuel gas flow that horizontal vertical openings 87a and 87b shift selectivity imports to every side of wing, is used for relative aerofoil surface and exhaust up and down.The variable guiding centre outlet 88 that is positioned at the aircraft afterbody axially guides main fuel gas flow and jet vector on principal direction, perhaps make it perpendicular to aerofoil surface in one of them upper and lower direction.Auxiliary wing plate 89a, 89b, 90a, 90b, 91a and 91b provide the conventional control surface that is used to control means of transportation flight.Opening 87a, the combination of the effect of 87b and variable guiding centre empennage 88 makes means of transportation to make this means of transportation become the absolute change of a general motor-driven VTOL wing formula aircraft and flight from any position, to the landing vertically of any direction.For the application of navigation, the vertical position on the means of transportation afterbody can allow airplane carrier to carry maximum VTOL wing formula aircraft, and they can take off simultaneously, thereby have promoted attacking ability the biglyyest.

Traditional wing formula aircraft is to leave on level runway and the platform, and can only sequentially take off, thereby has limited attacking ability.

Turbojet engine 10 in two-stage turbine Duct-Burning Turbofan gondola embodiment has been described in Fig. 9, described gondola is represented by label 90 usually, wherein the single moving element of the centrifugal compressed gas turbo machine of this integrated fan is a rotor unit 11, has an additional front fan 91, axial compressor blade 92a with the connection, 92b and 92c, they are by axial compressor stator blade 93a, and 93b and 93c separate.Fan attachment blade 94 comprises the rotor blade of this bypass turbofan.Other inner body and Fig. 1 are described, and those are identical.

It in Figure 10 the comparison diagram of the thermal cycle of thermal cycle of conventional gas turbine machine and turbo jet engine of the present invention.Have limited pressure ratio 30-40 and changeable compression (1-2), have turbine-entry temperature (3), be subjected to the modern combustion gas turbine of the reality of structural condition restriction, its traditional hot circulates in the T-S chart shown in circulation point 1-2-3-4-5-1.This circuit useful work is by shown in the following circulation point: 1-2-3-4-1.When identical circulation is extended to stoichiometry maximum temperature (3LPS), the turbine power that can obtain to increase, represented by following circulation point: 1-2-3LPS-4S-5S-1.Cycle through isothermal compression (1-2) expansion when identical, turbo machine deducts minimum isothermal compression power in the peak output of stoichiometry level and equals a circuit maximum effective power, has limited pressure ratio 30-40.

More high pressure circulation for turbo jet engine of the present invention, pressure ratio is 50-100, isothermal compression, (1-2HP), have maximum chemical metering temperature (3HPS), maximum absolute thermal cycle is represented by following circulation point: 1-2HP-3HP-3HPS-4AS-5AS-1.When partial load, this circulation is restricted to as following circulation point represented: 1-2HP-3HP-1.For this limited partial load circulation, the thermal efficiency is actual to be 100%, and this is because all entropy that circulation produced are negative.

Described among Figure 11 and had two-stage turbine circuit turbojet engine 10.This turbojet engine 10 is arranged in a turbine gondola, this turbine gondola is represented with label 100 usually, turbojet engine 10 comprises a dual fan compression turbine rotor unit 110, this rotor unit has a centrifugal passage 111 in inside, this centrifugal passage passes two hollow turbine blade 112a and 112b, and the pressurized air of a part is directed in the air pressure airspace 114 by by-pass hole 113, this air pressure airspace surrounds refluence firing chamber 115.Firing chamber 115 has one or more fuel injectors 116 and nozzle wheel blade 117 and 118, and described wheel blade is separated two levels of turbo machine.

Rotor unit 110 has additional fuel injector 119 and 120, is used to inject fuel in hollow turbine vane 112a and the 112b.Chamber wall 121 in that ambient air follow board space, firing chamber has a porous has internal hanger 122 and 123 on this wall.

Rotor unit 110 has first order fan blade 124, this blade and stator fan blade 125 operate together.Rotor unit 110 also has axial compressor rotor blade 126a, and 126b and 126c, they and stationary axle be to compressor blade 127a, 127b and 127c operate together.Rotor unit 110 centers on a stationary axle 128, and is supported by bearing 129 and 130.Fuel is transported in hollow turbine vane 112a and the 112b by fuel line 131.

The aerodynamics fuselage of turbo jet engine gondola 100 is made up of outer shell base 132, in this outer shell base, and front hanger 133 and empennage 135 after after gallows 134 is connected, and preceding taper 136 surrounds stationary axle 128.

This axial and centrifugal air compression is by the compression heat of scattering and disappearing of the fan part 110a with the heat exchange of first order fan blade 124 and rotor unit 110.

The heat of discharging from isothermal compression is passed in the bypass air, and by regeneration effect actuating air is heated.The inside cooling energy of turbine blade 112a and 112b turns back to the circulation by the regeneration effect of the preheated air and the fuel of discharging in 115 from the firing chamber.

Air from ram intake 137 is divided into two air-flow paths, is respectively bypass passageways 138 and compressed air channel 139.By the high by-pass ratio 12-20 of a uniqueness, air in the bypass passageways through fan nozzle 140 and with the expansion combustion gas mixing of discharging by gas of combustion nozzle 141.This mixed airflow has formed total propelling media in combined exhaust gas nozzle 142.Final result has developed to have the most effective turbine fan jet machine ultralow, infrared characteristic.

Described in Figure 12 and have two-stage turbine circuit turbojet engine 10, backward rotation compression that is provided by an additional preposition freewheel air turbine is provided for it.This birotary burbine jet plane gondola is generally represented with label 146.Main two-stage combination fan is connected with the preposition freewheel air turbine of a backward rotation rotor unit 149 with turbine rotor unit 148.This freewheel air turbine rotor unit 149 has turbine blade 150, is used to drive rotor unit 149 and axial compressor blade 152a, and 152b and 152c, these blades are positioned on the extension wheel hub 151 of rotor unit 149.

This two-stage fan and turbine rotor unit 148 have axial compressor blade 153a, 153b and 153c, these blades are positioned at and extend on the cover 154, the axial compressor blade 152a of their driving direction and freewheel air turbine rotor unit 149, and 152b and 152c's is opposite.Depend on flying speed, when the angle geometry control that is subjected to guiding blade 155, the effect of power ram-air can be converted to the power rotate effect.Air that is produced and turbine blade 150 operate together, and drive backward rotation, freewheel air turbine rotor unit 149.

The speed change of the backward rotation axial compressor 167 that the combiner by two rotor units 148 and 149 forms enters radially or before the centrifugal compressor 166 of combination fan and turbine rotor unit 148 at air, and air is carried out initial compression.This preparation compression makes and can both keep a final combination constant pressure ratio under any practical flight speed.

The rotation on stationary axle 160 upper bearing (metal)s 156 and 157 of this combination fan and turbine rotor unit 148.In an identical manner, freewheel air turbine rotor unit 149 is in the rotation on bearing 158 and 159 on the stationary axle 160, but sense of rotation and combination fan and turbine rotor unit 148 is opposite.Birotary burbine jet plane gondola 146 has an outer shell base 161, wherein have a suspension bracket 162 supporting preceding wind sleeve 163, and suspension bracket 165 is supporting empennage 164.Identical among miscellaneous part and as shown in figure 11 the embodiment.

Described in Figure 13 and have the turbojet engine 10 of the general structure of turbo jet engine gondola 146 as shown in figure 12, this turbojet engine has an improved firing chamber, but has a double bypass turbofan that has transition loop.The convertible turbo jet engine gondola 180 of among Figure 13 this has a firing chamber 181, this firing chamber has a hinged variable geometry bypass mechanism 182, and this bypass mechanism is used for firing chamber 181 is transformed into a firing chamber with second bypass discharge leg mouth 183 from a firing chamber with single annular floss hole 141.This variable geometry bypass mechanism 182 has nozzle Lens element 184, can control the size of discharge nozzle 185 according to operating conditions.

When the aircraft high-speed flight, dynamically the punching press compression is fairly obvious.This punching press effect is that compressor stage improves compression ratio, to keep required compression ratio.And this can also reduce required power of turbine stage.Variable geometry bypass mechanism 182 is directly with the parts discharging of firing chamber combustion gas, is discharged to the air that comes out from fan nozzle 140 with from the turbine combustion gas that main combustion chamber emission nozzle 141 comes out by a variable discharge nozzle 185.

Described a turbojet engine 10 with three bypass turbofan gondolas 190 in Figure 14, it has a convertible circulation, is applicable to the speed of subsonic speed and ultrasonic means of transportation.Turbofan engine gondola 190 have as shown in figure 13 a turbofan engine is integrated into structure in a variable geometry air inlet and the exit casing 197.This air inlet and exit casing 197 have a preceding variable geometry suction port 192 that has traverse 198, a bypass suction port 193 that has traverse 199, and a variable geometry exhaust nozzle 195 that has lens wing flap (lens flag) 196.In the exhaust jet stream 142 of combination, an afterburner 194 is arranged.

When subsonic flight, 198 withdrawals of nose air intake 192 standard-sized sheets and traverse, suction port is through suitable adjustment, makes that aircraft can be that the work of turbofan engine keeps optimal conditions near velocity of sound and ultrasonic speed the time.When higher supersonic speed, bypass suction port 193 is opened gradually, and afterburner 194 is lighted, thereby circulation is transformed into the turbine punching press jet work of combination.In order farthest to avoid the detection of radar, suction port 192 and 193 is disposed in the top of turbofan engine gondola 190.

In Figure 15 A and 15B, described a common aircraft and be transformed into aircraft, shown in label 200 with VTOL ability.This common aircraft 200 is with a common aircraft 201 and propulsion die 202a and 203b combination, and the type of described propulsion die 202a and 203b is shown in this specification previous drawings.The direction of propulsion die 202a and 203b is propulsive jet downward vertically, with the lift that produces aircraft and take off vertically.

Shown same aircraft 201a has identical propulsion die 202a and 203b, and their direction is arranged horizontally, to produce horizontal propelling force.After propulsion die 202a and 203b returned to original vertical fluid work, aircraft can vertically land.This simple especially common VTOL technology can be changed for conventional aircraft produces a flight.

A universal machine moving platform means of transportation has been described, shown in label 220 in Figure 16 A and 16B.This means of transportation has a common platform, for example, the quadrilateral structure 221 of a rectangle has comprehensive propulsion die 222 and 224, such as described in the present invention, these propulsion dies are hinged on each of four angles of platform structure 221 and 225.These comprehensive propulsion dies 222 and 224 can rotate with vertical axis 223V, also can rotate with horizontal axis 223H.

Propulsion die 222 and 224 combination are orientated aircraft 220 comprehensively and provide and comprising the aloft universal machine kinetic force of VTOL.In the time of on the ground, aircraft has a mixed propulsion system described in patent formerly.

Described a general maneuvering-vehicle in Figure 17 A and 17B, this aircraft has the propulsion die that is fixed on the fuselage, and aircraft is shown in label 230.The aircraft 230 of this combination has a fuselage 231, and this fuselage has comprehensive propulsion die 232 and 233, and they are fixed on suspension bracket 234 and 235.

By with comprehensive propulsion die 232 and 233 vertical orientations, just this aircraft 230 has the ability of VTOL, and with described propulsion die horizontal orientation, this aircraft just can horizontal flight as conventional aircraft.

Described a general maneuvering-vehicle in Figure 18 A and 18B, this aircraft has the comprehensive module that is fixed on the wing tip, and this aircraft is shown in label 250.The aircraft 250 of this combination has a fuselage 251 and wing 252.Comprehensive propulsion die 153 is connected on the wing tip 255.

On empennage 254, comprehensive module 253 also can be connected on the wing tip 256.This layout of comprehensive module 253 can produce the VTOL ability with unique universal machine dynamic response.

A comprehensive VTOL aircraft has been described, shown in label 280 in Figure 19.The aircraft of this combination has the cross direction wheel blade 284 of a fuselage 281 and a symmetry, has the wing 282 and the comprehensive module 283 that links to each other of symmetry on its middle machine body 281.

This symplex structure of flight structure, make this aircraft move along axis X-X and Y-Y, at direction D1, D2 symmetrical flight with when vertical axis C rotates, all have stable especially performance.

Additional VTOL ability makes this aircraft become the unique in history general maneuvering-vehicle of flight.

Described a general motor-driven extended type aircraft in Figure 20, shown in label 300, this aircraft has a plurality of vector jet plane.The aircraft 300 of this combination has a fuselage 301, has on the fuselage to mix wing 302, to form a comprehensive wing 305 of flight.The comprehensive propulsion die 303 that is hinged on along on the cursor 304 of axis X-X is installed on this flight wing 305, has the arched position of acquisition A, B, the ability of C, thus can make aircraft along direction A, B, C, D, the E motion comprises gyratory directions R.All these variablees have produced total VTOL ability.

In Figure 21, turbojet engine 10 of the present invention shows in the mode of perspective, comes the three-dimensional feature of description scheme parts with cut-away section.

Described the turbojet engine 10 of turbo jet engine 320 types that are preferably as shown in figure 12 in Figure 22, this turbojet engine is applied in a general navigation and advances in the gondola 322, is used for naval's means of transportation 324 is advanced.This general navigation advances gondola 322 that a suction tude 326 is incorporated on naval's means of transportation 324, is used for air is drawn into turbo jet engine 320.Common exhaust and bypass air jet pipe 328 are with the firing chamber combustion gas and bypass air mixes and by being positioned at a plurality of motion gas ejectors 330,332 and 334 ejections under the water line 335.

What describe in Figure 23 is as shown in figure 22 turbo jet engine 320, and it has turbojet engine 10 as shown in figure 12, and is applied as a general navigation propulsion system 322 through revising.Turbo jet engine 320 shown in Figure 23 has through the shell of revising 336 (display unit), and suction tude 326 and common exhaust and bypass air jet pipe 328 are arranged on this shell.Should be noted in the discussion above that the nuclear means of transportation for naval, the gas turbine part 330 of this turbo jet engine 320 is modified to a steamturbine part, or is replaced by a motor fan unit, is the identical air-spray of water injection generation of motion gas.

A turbojet engine 10 of type has as shown in figure 12 been described in Figure 24.This turbojet engine is the form of three bypass turbofan gondolas 190, as shown in figure 14, is modified to a transcontinental aviation propulsion die 191 with oxygen supply battery of a liquid oxygen sparger 197.In the air along with the minimizing of inlet air amount, the supply of liquid oxygen increases, to keep burning at height.For stratosphere and space flight, the oxygen of being supplied is total support of fuel combustion, and when nose air intake 192 and top air bypass 193 were closed, this turbojet engine 10 was worked with regard to being converted into as a rocket.

In all working of motor 10, can use any liquid fuel, particularly including liquid hydrogen (LH2) and liquified natural gas (LNG).Use for commerce, liquified natural gas low-cost, high energy content is preferred.When using low temp fuel, the low emission of fuel or zero-emission in conjunction with obtain adiabatic compression and do not have hot feature advantage, make low temp fuel be specially adapted to military project and aerospace applications.

One type as shown in figure 12, have freewheel air turbine rotor unit 149, the turbojet engine 10 in a turbo jet engine gondola 480 have been described in Figure 25.Turbo jet engine gondola 480 among Figure 25 has an additional freewheel rotor unit 482, and this unit has the turbine 484 and the fan 486 of an integral body.Turbine 484 has turbine blade 488, and these blades are disposed in the turbine blade 112a of main rotor unit 148 of turbojet engine 10 and the downstream of 112b.These three grades of turbines and additional fan 486 combinations and produce high compression and high the expansion are for the application of the bigger propelling force of needs provides high air bypass pressure ratio.

Turbojet engine 10 at turbo jet engine gondola 500 has been described in Figure 26, it has an additional conventional axial gas turbine jet machine 502, this jet plane 502 act as a starter motor and advancing means, with the power amplification with main turbojet engine 10.Main turbojet engine 10 in turbo jet engine gondola 500 and axially conventional turbine jet plane 502 by will mixing from the bypass air of passage 504 and the waste gas of discharging from the waste gas and the centre gangway from axial turbo jet engine 502 jet blowers 508 507 of passage 506, thereby the generation propelling force.

This turbojet engine 10 is modifications of the motor in as shown in Figure 9 the turbo jet engine gondola 90, and comprising an annular combustion chamber 510 that has variable geometry discharge nozzle 512, this variable geometry discharge nozzle 512 has an inner annular dividing plate 513 and outside hinged member 514 and 516.The waste gas of discharging from nozzle 512 produces a jet and advances, and this jet advances the fuel that is subjected to the variable geometry discharging mechanism and is assigned in the fuel injector 517 of firing chamber 510 to control.

Turbojet engine 10 as shown in figure 11, a freewheel fan 520 is parts of backward rotation rotor 522, this backward rotation rotor 522 comprises an additional inside stage compressor 524, is used for the air that imports to the radial compressor fan propeller 526 of freewheel fan 520 switched in opposite is further compressed.Adjustable stator blade 528 provides spacing control for the deflection ram-air, with the freewheel fan 520 on the optimization co-rotor will 522 and the rotation of inner axial compressor 524.

Radial compressor fan propeller 526 is centrifugal by inner passage 530 with the axial compression air, and its bypass air that is passed across the fan inside 532 of compressor fan rotor 526 is cooled off.The cooling that turbine blade 519 is radially added injects fuel into inner vanes tip channel 538 by inner fuel service duct 536 and internal spray device 537 in addition, to produce rich oil air mixture.Part radial compression air is diverted by bypass opening 533 and arrives in the porous air follow board space 540 around firing chamber 510, and part is the most advanced and sophisticated nozzle 539 by turbine blade 519 then.This cooling bypass air provides isothermal compression for the air in the inner passage 530.In the embodiment of Figure 26, be preferably most air and be directed in the air pressure airspace 540, there enter firing chamber 510 and with fuel mix from fuel injector 517, be used for the burning of phase I sparger 517.The rich oil air mixture that ejects by vane tip nozzle 539, the Turbulence Mixed compound that forms owing to effect at a high speed make the stoichiometric(al) combustion in the firing chamber 510 to produce to flow through the combustion gas of the annular jet nozzle 512 that formed by dividing plate 513 and hinged member 514 and 516.

This conventional axial gas turbine jet machine 502 has a suction port 546, and this suction port leads to an axial compressor 548 that pressurized gas is provided for firing chamber 550.High-pressure turbine 552 of gas driven and a low-pressure turbine 556, wherein this high-pressure turbine passes through spool 554 and Driven Compressor 548, and low-pressure turbine drives a centrifugal spool 557 that is connected to radial compressor fan propeller 526.When starting, the axially rotation of the main turbojet engine 10 of gas turbine jet machine 502 startings is to produce the main thrust of turbo machine gondola 500.

For the rocket propulsion of aircraft, the speed of aircraft produces a ram-air at the suction port 558 of bird shell 559, with Driven Compressor fan propeller 522 and backward rotation radial compressor fan propeller 526.The work done process of system is amplified with the proportional ram-air of aircraft speed.The treble cut propulsive jet is by the combustion gas of bypass fan air, ejection from controlled jet nozzle 512, and the gas of ejection from injector nozzle 508 provides.In the combining unit of Figure 26, the thrust that turbojet engine 10 is provided is 10 times to 20 times of auxiliary conventional turbine jet plane 502.

In Figure 27, described have additional conventional axial gas turbine jet machine 502, at the turbojet engine 10 of a turbo jet engine gondola 560, wherein Fu Jia conventional axial gas turbine jet machine 502 act as the starter motor and the advancing means of main turbojet engine 10.This turbojet engine 10 is modifications that have the motor that is arranged in turbo jet engine gondola 146 of freewheel air turbine rotor unit 149 among Figure 12, and comprise annular combustion chamber 510 with variable geometry discharge nozzle 512, wherein this variable geometry discharge nozzle 512 has as shown in figure 26 inner annular dividing plate 513 and outside hinged member 514 and 516.In the embodiment of Figure 27,, pass across axially and the bypass air of the fan 561 of radial compression machine rotor 563 is provided for isothermal compression through two most advanced and sophisticated nozzles 562 and 564 and be injected in the firing chamber 510 and and before being injected in the porous follow board space 538 at pressurized air through bypass opening 533.A two-stage combustion process is started by an additional fuel injector 568.In the embodiment of Figure 27, two most advanced and sophisticated nozzles 562 and 564 do not provide the additional peripheral axially turbine shown in Figure 11 embodiment.Therefore, the embodiment of Figure 27 is a cooled rotor system that is arranged in turbo jet engine 10 on the whole, has minimum work isothermal compression and stoichiometric(al) combustion.This conception of species provides a strong propulsion system and an air line that does not have the high heat effect of high-temperature turbine blade.

Turbojet engine 10 in the turbo jet engine gondola 570 has been described in Figure 28, wherein the structure of this turbo jet engine gondola and working principle are identical with turbo jet engine gondola among Figure 27, but do not have combination conventional axial gas turbine jet machine 502.For the rotor 565 that starts freewheel fan and implements spatial scalable compression machine and the backward rotation of radial compressor fan propeller 563, turbojet engine 10 includes a traditional starter system 571, and wherein starter system 571 has first a common starter motor 572 that engages with freewheel rotor 565 and the common starter motor 574 that engages with radial compressor fan propeller 563.These common starter motors can be any types, comprise that pressurized air type, explosive emission type, motor and other are applicable to the system of this application.In certain was used, during with transmitted at high speed, ram-air had enough started backward rotation to the turbo jet engine gondola from a pipeline, and therefore starter system 571 can omit in this application.

Turbo jet engine gondola 570 among Figure 28 is full punching press gas-powered turbines, has isothermal cooled compressed machine rotor and zero calory moving element.This gondola comprises a concentric gun hose 576,578,580 and 582 battery, wherein said concentric gun hose mainly concentrates on the nozzle effect of gas jet, and described gas jet part is by element 514 and 516 guiding of the variable geometry discharge nozzle 512 of firing chamber 510.The finally ejection from the common waste gas nozzle 584 of gondola shell 560 of the partially mixed combustion gas of gun hose 576-582 and bypass air of flowing through.

The enlarged portion view of in Figure 29, describing that is included in the radial compressor fan propeller 563 in the turbojet engine 10 in Figure 25 and 27.This radial compressor fan propeller 563 is made of two-part 586 and 588, and has an O-ring seals 590 to surround each outstanding nozzle 562 of rotor 563 and 564 bifurcated inner passage 530.

What describe in Figure 30 is a turbojet engine 10 in the turbo jet engine gondola 600, this turbojet engine has a backward rotation fan and compressor assembly 602, this assembly has first fan and compressor drum 604, and second backward rotation fan and compressor drum 606.The rotor 604 of this backward rotation fan and compressor assembly 602 and 606 is worked as axial backward rotation air turbine, and is driven by ram-air, and wherein the parts of Zhuan Donging separate with the combustion gas of heat, thereby first cryogenic turbo air breathing engine is provided.Jet-flow jet advance in the air-flow mixes with the bypass fan air and finally from common waste gas nozzle 612, sprays before, the high temperature of turbojet engine 10 by thermal insulation in firing chamber 608 and variable geometry nozzle 610.

Outside ram-air turbine rotor 606 is by being connected to pillar 618 on the gondola shell 622 and bearing 614 and 616 supports of 620.A division center 623 is additionally by pillar 624 and 626 supports.Adjustable stator blade 628 at first is directed to ram-air in the fan 630 of rotor 604, and then is directed in the backward rotation fan 632 of rotor 606.Air turbine drives classification backward rotation axial compressor 634 and 636 effectively, thereby the ram-air that a part that is fed in the firing chamber 608 is sucked compresses, and sparger 638 is to described firing chamber burner oil simultaneously.The hinged member 640 of variable geometry nozzle 610 and the core gas stream of 642 pairs of combustion gas that eject from reaction nozzle 644 are controlled. Gun hose 646 and 648 leads the bypass air air-flow and is mixed in the fuel gas flow, and they mix in common waste gas nozzle 612 afterwards.

In order to start the work of ram-air turbine, the rotor 604 that has the blade 650 of first fan 630 is rotated by an axle 652 that is operably connected on the common starter motor 654.In case rotate beginning, force the blade 656 that air will backward rotation second fan 632.Then, backward rotation axial compressor 634 and the 636 pairs of a part of ram-airs that sucked that integrate with fan propeller 604 and 606 compress respectively, with the fuel in the burning firing chamber 608.Then, as mentioned above, combustion gas is discharged by variable geometry nozzle 610 and common waste gas nozzle.

Contrast Figure 31, first embodiment of high-altitude turbojet engine uses label 700 represented usually.This turbojet engine 700 comprises a fuselage 701, has a suction port 702, pillar 703, and variable geometry air guide device 704.This air guide device 704 is positioned near the preceding rotor 705, and described preceding rotor 705 comprises a ram-air turbine 706, and this ram-air turbine 706 has hollow blade 707, thereby forms a centrifugal isothermal air compressor.The cold compression air is supplied by Hollow Pillar 708, described Hollow Pillar 708 is supplied to by bearing 713 at the air that is compressed and cool off, 714 and 716 that supported, have a grading plant 710, before 711 and 712 the backward rotation compressor, act as cold machine 709 in the auxiliary air.

Second rotor centrifugal compressor 717 is a kind of air bypass fans 718, and it has hollow peripheral gas-turbine blade 719 and 720.771 pairs of compressions of starter motor and turbine process start.

External concentric firing chamber 727 has two combustion zones, wherein first zone 722 is used for primary combustion, and second zone 723 is used for total burning, wherein said total burning is activated under the rocket mode of operation, this moment, the firing chamber was opened, be applicable to that the bypass rocket Gas Jet by variable geometry discharge nozzle 724 and 725 advances, thereby produce pure rocket propulsion jet 726.

Exhausting air in the propulsive jet 727 of gas turbine is mixed with gas and bypass air jet 728 in the pure rocket propulsion jet 726, thereby forms final total propulsive jet 729.

Isothermal compression in rotor 707 and 717, and in the hollow gas-turbine blade 719,720 that is cooled in the cold machine 709 the coldest pressurized air is provided.These gas-turbine blades have the inner fuel sparger 770 as describing in detail among the above-mentioned embodiment.Compressor and in the cooling of cold machine and inner fuel combinations of injections get up to produce and be applicable to the strongest cooling of gas-turbine blade, and burning can be carried out on the maximum chemical stoichiometric level.On very high altitude, when airborne oxygen content reduces, liquid oxygen sparger 730 will be activated and eject oxygen enrichment pressurized air, to keep maximum burning capacity.

Embodiment among contrast Figure 32 is the evolution patterns of the embodiment 700 among Figure 31 with label 750 represented high-altitude turbojet engines generally.This turbojet engine has a fuselage 751, comprise ram-air forepieces all among Figure 31, also comprise second ram-air turbine 752 in addition, has hollow blade, act as by Hollow Pillar 754 and supply compressed-air actuated final stage isothermal centrifugal compressor 753, with the cooling air guide in the central combustion chamber 755 that is surrounded by an air pressure airspace 756, described firing chamber 755 is separated by a firing chamber liner 758 that is preferably stupalith.

Jet rocket gas after the burning is subjected to the control of variable geometry discharge nozzle 759,760, for the burning optimization under all speed goes out a constant compression force ratio, thereby keeps the maximum thermal efficiency under all flight conditions.

Rocket gas jet 761 finally mixes with bypass air jet 762, forms final combination propulsive jet 763.

A key property of this air-rocket-bypass propulsion system is the moving element without any a heat, thereby makes cost lower, and has maximum absolute thermokinetics performance.

Contrast Figure 33, the 3rd high-altitude turbojet engine is represented by label 800.This turbojet engine 800 has the forepiece identical with motor 750, but has different interior rear wall parts, and these parts form one and are applicable to convertible turbine/rocket-propelled central burner system 802.In Figure 33 and 34, ram-air turbine 752 and final stage isothermal centrifugal compressor 753 are supplied pressurized air by first air passageways 803 in the Hollow Pillar 754 and the bypass channel 805 around pillar 754.

Shown in more details among Figure 34, air passageways 803 in the pillar 754 is divided into three passages 806,807 and 808, respectively by a concentric outer air pressure airspace 810, a concentric inner air pressure airspace 811, and one form with the concentric middle porous annular combustion chamber 812 of external air pressure airspace 810 and internal air pressure airspace 811.Fuel injector 814 sprays into annular combustion chamber 812 with fuel, with the generation combustion gas, thereby produces the motion gas that drives hollow blade turbine 816.Hollow fuel nozzle 815 laterally internally air pressure airspace 811 pass annular combustion chamber 812 and reach external air pressure airspace 810, and have internal spray device 817, be used for fuel gas mixture is transported to the outer cylinder 819 of concentric outer air pressure airspace 810 ends.Burner oil in pressurized air and the nozzle 815 cooled off the combustion gas in the annular combustion chamber 812 before the hollow turbine vane 809 of the center of driving turbine 816.Hollow turbine vane 809 obtains pressurized air in the air pressure airspace 811 internally, and with it fuel mix with inner fuel nozzle 818 ejections, and then centrifugal atomization is in outer cylinder 819.At this, the air in mixture and the external air pressure airspace and mix from the oil laden air of hollow fuel nozzle 815 ejections.This spent mixture is ejected by a variable geometry discharge nozzle 820, becomes elementary propelling source, and wherein said variable geometry discharge nozzle 820 has hinged member 821 and 822, is used to control flowing of rocket jet 824.

The ram-air of rocket jet 824 and bypass channel 805 and from the turbine drives combustion gas mixing of discharging jet 824 ejections internally of annular combustion chamber 812.

A turbine shaft 825 that is supported by bearing 826 provides a pto for a driving mechanism such as gear-box driver 828, and to be connected on the centrifugal compressor 829, described centrifugal compressor 829 produces the basic air compression that is used to start burning.Gear-box driver 828 also is connected on the motor-driven starter 771 (not drawing among the figure), shown in Figure 32 principle.

In the turbojet engine 800 of Figure 33 and 34, isothermal compression as described in Figure 32 is relevant with air velocity, and when compressing at a high speed, it is higher that air concentration also becomes.When high-speed flight, the ram-air compression that ram-air turbine carried out reduces back gas turbine 816 required power.When speed-raising work, the parallel rocket propulsion of outer cylinder 819 produces main thrust.By continuous control variable geometry discharge nozzle 820, can obtain a constant compression force ratio, thereby make thermal efficiency maximization, and from partial load in fully loaded all working scope and all keep this maximized thermal efficiency under all practical flight speed.

Contrast Figure 35, the 4th embodiment of high-altitude turbojet engine represents with label 850.This turbojet engine 850 has the forepiece identical with motor 750, but has different interior rear wall parts, and these parts form one and are applicable to convertible turbine/rocket-propelled central burner system 852.In Figure 35 and 36, ram-air turbine 752 and final stage isothermal centrifugal compressor 753 are supplied pressurized air by first air passageways 853 in the Hollow Pillar and the bypass channel 854 around pillar 754.

Shown in more details among Figure 36, air passageways 853 in the pillar 754 is divided into three passages 855,856 and 857, respectively by a concentric outer air pressure airspace 860, a concentric inner air pressure airspace 861, and one form with concentric intermediate arrangement, adverse current, the porous annular combustion chamber 862 of external air pressure airspace 810 and internal air pressure airspace 811.

Annular combustion chamber 862 has primary combustion region 858 and secondary combustion zone 859.Primary combustion region 858 has a fuel injector 863 and an adverse current conduit 865, described adverse current conduit be used for combustion gas from primary combustion region 858 import to the horizontal air fuel nozzle 866 that has inner fuel sparger 870 and center two-stage gas turbine 867 in.Described two-stage gas turbine 867 has hollow gas-turbine blade 868a and 868b and inner fuel sparger 869.Air enters into the side canal 880 of guide blades 868a from concentric inner air pressure airspace 860, and with public hollow area 881 in fuel mix, then from vane tip 882a and 882b with ejection at a high speed, as shown in figure 36.

This air-fuel mixture cools off horizontal air/fuel nozzle and turbine blade 868a and 868b.The centrifugal air/fuel mixture that comes out from turbine blade 868a and 868b provides sufficient mixing and effectively burning in adverse current annular combustion chamber 862, wherein said adverse current annular combustion chamber 862 extends in primary and secondary combustion zone 863 and 864 always.

To the mode similar with the embodiment shown in 34 as Figure 33,883 pairs of rocket jets 873 of a variable geometry discharge nozzle with hinged member 871 and 872 are controlled, and wherein said rocket jet 873 mixes with the exhausting air of turbo jet engine 874 and the ram-air bypass jet 875 in the final waste gas.The control of this variable geometry discharge nozzle produces a constant compression force ratio, all makes thermal efficiency maximization under all loads and the flying speed thereby make.

Figure 35 and embodiment in 36 have utilized as identical live axle 825 in Figure 33 and 34 illustrated embodiments, bearing 826 and 827, and the gear drive 828 that is connected with starter motor 771 (not having among the figure to show).

Described two embodiments are designed to allow engine operation to obtain to produce the stoichiometric(al) combustion of absolute peak output, make crucial parts obtain suitable cooling simultaneously, guarantee to work effectively in the thermal limit scope.For the isothermal compression of above-mentioned acquisition pressure ratio 100-200, can obtain maximum and be about 80% the thermal efficiency.

Turbojet engine 800 and 850 is designed to being operated in effectively in subsonic speed, the integrated turbofan engine of supersonic speed, turbo jet engine and the rocket engine as a change notion.

Although in the above description embodiments of the invention are had been described in detail, purpose is for full disclosure content of the present invention, but it is apparent that, for the those of skill in the art in those present technique fields, can carry out multiple modification within spirit of the present invention and the principle not departing from.

Claims (30)

1. turbojet engine, comprise a turbo jet engine structure, it has a suction port, a rotor disk unit, this rotor disk unit has a fan unit, this fan unit has centrifugal pressing chamber, turbine blade, and outer cylinder, this firing chamber has at least one fuel injector and nozzle, be used for the turbine blade of gas emission to the rotor disk unit, the bypass air fan unit cooling air flow of flowing through and in centrifugal pressing chamber, compressing wherein, this bypass air stream is sprayed into the firing chamber, and wherein the combustion gas from turbine blade mixes with the bypass air of the fan unit of flowing through in a common discharge nozzle.

2. turbojet engine as claimed in claim 1, wherein the rotor disk unit has fuel channel, and wherein fuel is injected in the centrifugal pressing chamber at this fuel channel, is used for the pressurized air of cooled compressed chamber, isothermal ground before entering the firing chamber.

3. turbojet engine as claimed in claim 2, wherein this rotor disk unit comprises an axial compressor, this compressor compresses the centrifugal chamber that enters the rotor disk unit and the air in the turbine blade.

4. turbojet engine as claimed in claim 2, wherein this turbo jet engine structure comprises a backward rotation axial compressor, this backward rotation axial compressor has a motor, is used for driving the axial compressor that the air to the centrifugal chamber that enters the rotor disk unit and turbine blade compresses.

5. turbojet engine as claimed in claim 4, wherein this turbo jet engine structure comprises a motor generator set that is connected on the rotor disk unit, this generator is used for motor is powered.

6. turbojet engine as claimed in claim 5, it comprises a controller, is used to control the speed of axial compressor.

7. turbojet engine as claimed in claim 2, it has an additional front fan, and this fan has the axial compressor blade that is connected to the rotor disk unit and is connected to structural stator fan blade of turbo jet engine and stator compressor blade.

8. turbojet engine as claimed in claim 2, wherein the rotor disk unit comprises a fan compressor turbine rotor unit, this fan compressor turbine rotor unit has and is used for the two hollow turbine vanes of a two-stage turbine circuit, and separates this pair hollow turbine vane nozzle vane in two stages.

9. turbojet engine as claimed in claim 8, wherein this fan compressor turbine rotor unit has an additional front fan, and this fan has the axial compressor blade that is connected on the rotor unit and is connected to structural stator fan blade of turbo jet engine and stator compressor blade.

10. turbojet engine as claimed in claim 8, it has a preceding freewheel air turbine, this freewheel air turbine has a backward rotation freewheel air turbine rotor unit that has the air turbine blade, described air turbine vane drive air turbine rotor unit and axial compressor blade, wherein this rotor disk unit has the axial compressor blade opposite with the axial compressor blade rotation direction of freewheel air turbine unit, is used for the air that enters this fan compressor drum unit is carried out precompression.

11. turbojet engine as claimed in claim 10, wherein this firing chamber has a variable geometry bypass discharge leg nozzle that is used for convertible circulation.

12. turbojet engine as claimed in claim 1, with an axial gas turbine jet machine combination, wherein these axial gas turbine jet facility have a turbine that is pivotally connected on the rotor disk unit, are used for the starting turbine air breathing engine and increase the power of this group shot.

13. turbojet engine as claimed in claim 12, wherein heart is arranged in this turbo jet engine structure in this axial gas turbine jet machine, and have a nozzle, be used for the common nozzle of combustion gas from turbojet engine is injected to bypass air and the fuel gas flow.

14. turbojet engine as claimed in claim 13, wherein outer cylinder comprises a variable geometry discharge nozzle, and wherein the part combustion gas sprays from this variable geometry discharge nozzle and mixes with the bypass air of the turbine blade that does not drive the rotor disk unit.

15. turbojet engine as claimed in claim 14, it has a preceding freewheel air turbine, this freewheel air turbine has a backward rotation freewheel air turbine rotor unit that has the air turbine blade, described air turbine blade is used to drive air turbine rotor unit and axial compressor blade, wherein this rotor disk unit has the axial compressor blade opposite with the axial compressor blade rotation direction of freewheel air turbine unit, is used for the air that enters this fan compressor drum unit is carried out precompression.

16. a turbojet engine comprises a turbo jet engine structure, it has a suction port; Public combustion gas and air nozzle; A rotor disk unit, this rotor disk unit has an air fan, and this air fan has inner compressor passage and radial discharge nozzle; An outer cylinder, this outer cylinder has a porous air follow board space, and wherein this air fan has and is arranged near the side opening radial discharge nozzle, that be used for pressurized air is fed to from the compressor passage porous air follow board space; Inject fuel into the fuel injector in the outer cylinder, wherein flow into the air separated into two parts in the air inlet, a part is the bypass air by air fan, another part is the pressurized air that enters into the compressor passage, described air is discharged in the firing chamber by the radial discharge nozzle with by described hole and air pressure airspace, wherein this firing chamber has a variable geometry discharge nozzle, be used for gas emission being used for spraying from common nozzle together with bypass air to bypass air.

17. turbojet engine as claimed in claim 16, it has a preceding freewheel air turbine, this freewheel air turbine has a backward rotation freewheel air turbine rotor unit that has the air turbine blade, described air turbine vane drive air turbine rotor unit and axial compressor blade, wherein this rotor disk unit has the opposite axial compressor blade of axial compressor blade rotation direction with the freewheel air of the inner compressor passage that enters the rotor disk unit.

18. turbojet engine as claimed in claim 17, it has the device of the rotation that is used to start the rotor disk unit.

19. turbojet engine as claimed in claim 18, the device that wherein is used to start the rotation of rotor disk unit comprises a motor.

20. turbojet engine as claimed in claim 19 is combined with an axial turbine unit, this axial turbine unit has a turbine that is connected on the rotor disk unit, and has the device of the rotation that is used to start the rotor disk unit.

21. the turbojet engine in turbo jet engine, this turbojet engine comprises a turbo jet engine structure, this structure has a suction port, public air and gas nozzle, backward rotation air fan rotor, each rotor has an air fan, an axial compressor that has by the backward rotation compressor blade of backward rotation air fan rotor driven, a firing chamber that has fuel injector, wherein in this firing chamber from the pressurized air in the compressor and fuel mix and produce combustion gas, and combustion gas spray nozzle, wherein this turbo jet engine structure has from the bypass air air-flow of air inlet by the air fan of air fan rotor, and the compressed air by compressor, the combustion gas that sprays from the combustion gas spray nozzle mixes with the bypass air air-flow, and sprays from public air and combustion gas spray nozzle.

22. turbojet engine as claimed in claim 1, with an aircraft combination, wherein this turbo jet engine structure is a gondola, and this gondola comprises turbojet engine, and the gondola gimbal structure is connected to aircraft with gondola.

23. turbojet engine as claimed in claim 22, wherein this combination comprises a plurality of turbojet engines, and each turbojet engine is comprised in the gondola, and this gondola has gondola is connected to carry-on gimbal structure.

24. turbojet engine as claimed in claim 1, with the combination of navigation naval vessels, wherein this turbo jet engine structure is connected on these naval vessels, and wherein public spray nozzle is disposed in gas is sprayed on the position in the water.

25. a turbojet engine comprises a fuselage, it has a suction port and public combustion gas and air spray nozzle, and has betwixt: be positioned at the front strut on the air inlet; A preceding rotor unit, this preceding rotor unit and near the variable geometry air guide device cooperating that is positioned at it, this rotor unit comprises a ram-air turbine, this ram-air turbine has the hollow blade of centrifugal compressor form; Hollow Pillar; An axial compressor that has the backward rotation stage, the centrifugal pressurized air that wherein comes out from centrifugal compressor is supplied in the Hollow Pillar, and is fed in the axial compressor by pillar; Centrifugal compressor and have the bypass fan of hollow gas-turbine blade; And concentric firing chamber, wherein this gas-turbine blade has pressurized air is discharged into end in the firing chamber by hollow gas-turbine blade from axial compressor, then before gas emission is arrived public combustion gas and air spray nozzle, this firing chamber is led combustion gas and is got back to gas-turbine blade, wherein this firing chamber comprises a variable geometry nozzle, be used for combustion gas directly is discharged into public combustion gas and air spray nozzle, to produce rocket propulsion.

26. a turbojet engine comprises a fuselage, it has a suction port and public combustion gas and air spray nozzle, and has betwixt: be positioned at the front strut on the air inlet; A preceding rotor unit, this preceding rotor unit be positioned near variable geometry air guide device cooperating, this rotor unit comprises a ram-air turbine, this ram-air turbine has the hollow blade of centrifugal compressor form; Hollow Pillar; An axial compressor that has the backward rotation stage, the centrifugal pressurized air that wherein comes out from centrifugal compressor is supplied in the Hollow Pillar, and is fed in the axial compressor by pillar; A ram-air turbine, this ram-air turbine has the hollow blade of second centrifugal compressor form; Second Hollow Pillar; And central combustion chamber, wherein the pressurized air from axial compressor is supplied in second centrifugal compressor, and be fed to central combustion chamber by pillar, this firing chamber has a variable geometry discharge nozzle, is used for the rocket gas jet is discharged into public combustion gas and air spray nozzle.

27. turbojet engine as claimed in claim 26, wherein central combustion chamber comprises an annular porous burner chamber, this porous burner chamber has a concentric outer air pressure airspace and a concentric inner air pressure airspace, and this annular combustion chamber then places between these two air pressure airspaces.

28. turbojet engine as claimed in claim 27, it also has a center gas turbine that has hollow blade, wherein this gas turbine of the gas driven in the annular combustion chamber.

29. turbojet engine as claimed in claim 28, wherein said hollow blade has the inner fuel sparger, and from concentric inner air pressure airspace, obtain air, at this, air/fuel mixture is discharged into the external air pressure airspace, with the discharging of control by the variable geometry discharge nozzle.

30. turbojet engine as claimed in claim 28, should annular porous burner chamber be an adverse current annular combustion chamber wherein, it has primary combustion region and secondary combustion zone, wherein the gas flow in primary combustion region leads in the blade of center gas turbine and drives this turbine, and wherein the combustion gas in the secondary combustion zone is discharged by the variable geometry discharge nozzle.

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