38 results about "Scramjet" patented technology

The invention discloses a modified scramjet combustion chamber and a design method of a swirler thereof. The swirler capable of causing air flow to generate helical motion is arranged in a combustion chamber and comprises a rotational flow generation upper wallboard, a rotational flow generation lower wallboard, a back end surface wallboard and a sidewall mounting surface, wherein the rotational flow generation upper wallboard is used for making the air flow entering the combustion chamber generate oblique shock waves, and the rotational flow generation lower wallboard is used for making the air flow entering the combustion chamber generate expansion waves. The sidewall mounting surface is connected with the combustion chamber, and the front edge of the swirler faces to an inlet of the combustion chamber; the back end surface wallboard is arranged opposite to an outlet of the combustion chamber; the interior of the swirler is provided with a fuel conveying embedded pipe; an inlet of the fuel conveying embedded pipe is arranged on the sidewall mounting surface; and an injection opening of the fuel conveying embedded pipe is arranged at the back end surface wallboard. Therefore, the invention has the advantages of improving the mixing efficiency and the combustion efficiency of the fuel of a scramjet, shortening the lengths of an engine isolating section and the combustion chamber, further optimizing the structure of the engine, increasing the thrust-weight ratio of the engine and improving the whole performances of the engine.

Composite combustion catalyst particles are described and disclosed. A metal core of a combustible metal can be coated with a metal oxide coating. Additionally, a catalyst coating can at least partially surround the metal oxide coating to form a composite catalyst particle. The composite catalyst particles can be dispersed in a variety of fuels such as propulsion fuels and the like to form an enhanced fuel. During initial stages of combustion, the catalyst coating acts to increase combustion of the fuel. As combustion proceeds, the metal core heats sufficiently to disturb the metal oxide coating. The metal core then combusts in highly exothermic reactions with an oxidizer and the catalyst coating to provide improved energy densities to the enhanced fuel. Enhanced fuels using these composite combustion catalyst particles can exhibit decreased ignition delay times and increased energy densities sufficient for use in high performance propulsion applications such as pulse detonation engines, scramjets, and ramjets.

A design method of a multi-channel parallel three-power combined engine belongs to the field of combined engines, and comprises the following steps of 1) making overall performance requirements according to a flight task, determining the capture area of an air inlet channel, and obtaining a three-dimensional inward rotation air inlet channel through streamline tracking; 2) calculating inlet and outlet parameters of the scramjet combustion chamber, and designing a scramjet channel and the scramjet combustion chamber; 3) designing an upper molded surface and a lower adjusting plate of the exhaust nozzle; 4) arranging turbine channels, turbine engines and turbine channel splitter plates at openings in the two side walls of the three-dimensional internal rotation air inlet channel according tothe flow requirements of the engines at the Ma0-2 stages; 5) calculating an injection rocket-sub-combustion channel maximum flow demand based on the thrust demands of the working state of the Ma2-3 rocket engine and the working state of the Ma3-4.5 sub-combustion chamber, so that an inlet of the injection rocket-sub-combustion channel, the rocket engine with corresponding size and the injection rocket-sub-combustion channel splitter plate are arranged. According to the invention, the higher low-speed climbing performance and high-speed cruising performance can be provided while a thrust gap can be effectively spanned.

The invention relates to a novel aerospace plane which is conceived on the basis of thought on defense of the galaxy, and belongs to the technical field of aircraft engines. A flying disk engine disclosed by the invention integrates top three of the existing engine technologies of human beings into one, and is an epoch-making aerospace plane. A jet engine and a scramjet direct jet engine have a common feature, i.e., a detonation problem, and however, the flying disk engine disclosed by the invention solves the detonation problem tactfully and can fly and hover at very high speed. The engine is originated from a kite; all aircrafts invented by human beings are based on the imitation of birds which run hard towards the wind with two large wings; from an original propeller to a ramjet engine, all aircrafts made by human beings have wings. The kite is invented in China a long time ago, and flies based on the principle of being blown by the wind, which is completely different from the birds. The basic flying principle of a flying disk is specifically as follows: imagining that a person carrying two wings stands still, a fan is installed in front of the person, and the person is blown by the wind provided by the fan to fly, after the person leaves the ground, no relative motion is generated, another person carrying two wings stands in opposition to the person and is blown by the wind provided by another fan, the two persons are joined together hand in hand, and if people are joined together to form a circle, wind blows towards all directions from the centre, so that the flying disk can fly.

A design method of a turbofan ramjet combined engine with an external culvert and a built-in rocket belongs to the field of combined engines, and comprises the following steps: firstly, formulating overall performance requirements of the engine, designing a basic flow field of an air inlet channel, and then carrying out streamline tracking to obtain a three-dimensional internal rotation air inletchannel; Inlet and outlet parameters of the scramjet combustion chamber are obtained according to the overall performance of the engine, and the scramjet combustion chamber is designed at an outlet ofthe three-dimensional internal rotation air inlet channel; Designing an exhaust nozzle with an adjustable geometric area at an outlet of the scramjet chamber on the basis of parameters of the scramjet chamber and the thrust requirement of the engine at the Ma4.5-6 + stage; Selecting a turbofan engine on the basis of engine performance requirements at stages Ma0-2, and arranging turbofans at openings in the upper wall surfaces of the three-dimensional inward rotating air inlet channels; A sub-combustion combination channel and a turbofan engine; A rocket engine is arranged in an outer culvertchannel of the turbofan engine; based on the thrust requirement of an engine in the Ma0-5 state, and in combination with the performance parameters of a turbofan engine in the Ma0-2 state and a rocketengine in the Ma2-3 state, a sub-combustion chamber capable of switching working modes is designed behind the turbofan engine.

The invention relates to a scramjets and associated aircraft and methods. The scramjet includes a converging inlet, a combustor configured to introduce a fuel stream into an air stream in a combustionchamber and to combust the fuel air mixture stream to create an exhaust stream, and a diverging exit nozzle configured to accelerate the exhaust stream. The combustor includes a fuel injection systemincluding at least one arcjet. A method of creating thrust for an aircraft includes compressing a supersonic incoming air stream in a converging inlet, injecting a fuel stream into the air stream ina combustion chamber to create a fuel air mixture stream, igniting the fuel air mixture stream to create an exhaust stream, and exhausting the exhaust stream from a diverging exit nozzle. The injecting the fuel stream into the air stream includes injecting the fuel stream at a fuel speed sufficient to create shear between the fuel stream and the air stream.

The invention discloses an axisymmetric adjustable air inlet channel for a Ma0-6+ wide-range precooling and stamping combined engine. Turbojet and dual-mode stamping dual channels mainly consist of afixed cowl profile, a front-and-back translation center cone and a flow splitting plate part, and are in an open state all the time within the Mach number range of Ma0-6+. As the incoming Mach numberrises, the center cone moves forwards to guarantee a wide-range internal shock-on-lip, and the flow coefficient of the air inlet channel is 1.0; the flow splitting plate is linked with the center coneto guarantee that the shrinkage ratios of the turbojet and stamping channels are both increased along with rise of the Mach number, and flow distribution of the two channels is adjusted according todemands. Under a high Mach number, the stamping channel can be transformed into a scramjet mode from a ramjet mode.

The invention discloses a gas compressor multi-rotor impeller and turbine multi-rotor full contra-rotating aero-engine. According to the aero-engine of the new concept, a central hollow shaft is arranged as a longitudinal framework of the aero-engine, and a contra-rotating power shaft in the hollow shaft penetrates through the aero-engine front and back, so that the aero-engine has the advantages that power can be conveniently output front and back, the aero-engine has full coverage of the functions of a turbojet aero-engine, a turbofan aero-engine, a turboshaft aero-engine, a turboprop aero-engine and a gas turbine, and the front end of a gas compressor is provided with a front-pointed-cone-shaped network-shaped high-strength protective cover to prevent impact of birds, ice, snow and sundries. A two-stage folding type conical air duct is arranged on the periphery of the air compressor in a sleeving mode and used for adjusting the windward area, stamping air inlet adjustment is achieved, and the aero-engine has the function of a scramjet new conceptual aero-engine. Dynamic structures of work conversion, power transmission and the like of the aero-engine with the new concept are all suitable for upgrading of gas turbines, steam turbines, air compressors and water turbines.

The invention discloses a multi-stage rotary detonation rocket ramjet combined engine, and belongs to the field of combined engines. The combined engine comprises a ramjet engine, and a multi-stage rotary detonation rocket engine embedded between an air inlet channel and a combustion chamber of the ramjet engine, wherein the multi-stage rotary detonation rocket engine comprises a plurality of single-stage rotary detonation rocket engines distributed in a stepped mode, and thrust is provided in a rotary detonation combustion mode; and according to the whole engine system, the ramjet engine and the multi-stage rotary detonation rocket engine jointly generate thrust. Compared with a single-stage rotary detonation structure, the multi-stage rotary detonation structure can enable multi-stage rotary detonation to jointly generate an ejection and suction combination in a rocket ejection mode with a low Mach number, and the low-speed ejection pressurization efficiency is improved; in a subsonic combustion and scramjet mode, the positions of induced flames and a flame stable area can be adjusted, the working stability of the engine is improved, meanwhile, a larger thrust range can be provided, thrust adjustment is more flexible, and the thrust performance is improved.

The invention discloses a rocket based combined cycle engine rocket ejector T-type layout comprising a central supporting plate rocket ejector and a plurality of side-wall rocket ejectors. The side-wall rocket ejectors are located above the central supporting plate rocket ejector. The front portions of the side-wall rocket ejector are connected to rocket combustion chamber pipelines, the central supporting plate rocket ejector is connected to a combustion chamber through upper-side pipelines, and the section of two sets of rockets is the interface of an insulating segment and a mixing segment. By means of limitation on T-type rocket ejector layout size, and a layout manner with smaller duty ratio than a full throughout central supporting plate layout manner and larger wetted perimeter than a side-wall layout manner is provided; congestion effect on the insulating segment of a full throughout supporting plate type rocket ejector and flight drag of scramjet mode are reduced, and cutting area of a primary jet flow and secondary jet flow are increased; the scalability of the rocket ejector T-type layout is high, and the rocket ejector layout is designed according to different mixing segment inlet section sizes so as to acquire excellent engine performance.

The invention relates to a scramjet pulse engine. The explosive power is controlled and converted into kinetic energy capable of driving missiles and aircrafts to move at a high speed, the thrust generated by incident shock waves of an impact energy cylinder cup and the reverse thrust generated by reflected shock waves are applied to push the missiles and aircrafts to move and fly at the 23 times hypersonic speed and 28,260 km/h, accordingly, any missile can not catch and any radar can not detect the missiles and aircrafts. The power of detonation wave front of explosives is 1010 watts per square cm, and the thrust generated by explosion of 1 kg TNT is 800,000 kg. A cover type scramjet pulse engine catapults the aircrafts to take off, and a buffer spring extends to catapult the aircrafts again. An energy storage catapult aircraft device uses the explosion of explosives to compress a high pressure spring to retract to store all energy of explosion, and then releases tens of meters of the compressed spring to extend and continuously force to catapult the aircrafts about more than 100 meters to take off. The scramjet pulse engine is superior to the electromagnetic catapult.

The invention relates to a support plate ejection rocket-based combined cycle power engine based on a unit propellant. The support plate ejection rocket-based combined cycle power engine comprises an air inlet channel, a central support plate single-component rocket, a powder fuel supply system, a nitric acid hydroxylamine-based propellant supply system, a high-pressure air source, a combustion chamber and a spray pipe. A nitric acid hydroxylamine single-component propellant which is good in normal-temperature storage performance and can be rapidly catalyzed and started at normal temperature is used as a propellant of a support plate rocket in an engine, the structure of a support plate rocket system is simplified, and the reliability of the engine is improved; the aluminum powder is adopted to replace traditional hydrocarbon fuel, the density specific impulse of the engine is improved, the fuel dissociation phenomenon in a scramjet mode is overcome, and the engine performance is maintained. According to the engine, mode conversion in a flight process is achieved by adjusting the flow of a nitric acid hydroxylamine-based propellant.

The invention discloses an electromagnetic hypersonic thrust vector jet engine. The engine comprises a shell, an adjustable air inlet, an electromagnetic turbofan, a force bearing frame, an outer duct, an inner duct, an air compressor, an oxygen supply and oil supply pipe, a thrust augmentation nozzle, a combustion chamber, a generator, a double-turbine engine, a thrust vectoring nozzle, an adjustable nozzle, a main shaft, a bearing, a frequency converter, a transmission, a sparking plug, an oil sprayer, a lithium battery, a charger, a computer, a cable, an instrument and an electronic switch. Compared with an existing turbojet engine, the electromagnetic hypersonic thrust vector jet engine is more advanced, simple in structure, easy to machine, manufacture and maintain, low in cost and capable of controlling airflow balance of an air inlet and an air outlet and regulating and controlling the temperature of fuel gas in a combustion chamber and the supercharge ratio of an air compressor, so that the fuel gas achieves subsonic combustion or scramjet, the heat efficiency and the propelling efficiency of the turbojet engine are improved, the fuel gas is utilized to generate power for an electromagnetic turbofan, the lithium battery and the computer, and the electromagnetic hypersonic thrust vector jet engine which is low in noise, low in oil consumption, high in efficiency, high in reliability and good in running performance can be achieved.

The invention discloses a layout method of igniters in a large-size scramjet and aims to solve the problems that ignition can only be conducted close to the wall faces of scramjets through existing ignition methods, and the existing methods cannot be applied to combustion of large-size scramjets. The method comprises the steps that the effective ignition diameters of the igniters are tested; and grid layout is conducted, specifically, the multiple igniters are arranged in a combustion chamber of the scramjet, the positions of the igniters are set with the congestion ratio being lower than 15%,on the cross section of the scramjet, the center-to-center distance between every two adjacent igniters is (0.7-1.3)*D, and every three adjacent igniters are arranged in a regular triangle shape. According to the provided layout method of the igniters in the combustion chamber of the large-size scramjet, by reasonably designing the positions of the igniters, the ignition and combustion efficiencyin the large-size scramjet is improved, and a main flow region is effectively used for air intake. The layout method is applied to the field of scramjets.

The invention discloses an aerodynamic layout structure of a reusable vehicle and a design method, and relates to the technical field of aerodynamic layout of vehicles, the aerodynamic layout structure comprises a vehicle body including a fuselage and wings, the wings are arranged at two sides of the fuselage, and the wings and the fuselage form a wing body fusion body layout; a carrying plane which is arranged on the upper surface of the machine body; a scramjet stamping power system which is arranged on the lower side of the machine body; and a vertical fin which is arranged at the tail part of the machine body; the wings and the fuselage of the aerodynamic layout structure form a wing body fusion body layout, the performance requirement of a wide space/speed domain is met, the aerodynamic layout structure is provided with a carrying plane, the carrying capacity is improved, the aerodynamic layout structure is provided with empennages, and the control performance is improved.