32results about "Composite engine plants" patented technology

An air-breathing turbojet engine (301) for a hypersonic vehicle is shown. The engine comprises a pump for pumping a cryogenic fuel, an inlet (302) configured to compress inlet air by one or more shocks, a cooler (303) to cool the compressed inlet air using the cryogenic fuel, and a turbo-compressor (304) to compress the air further. A precooler (311) between the inlet (302) and the cooler (303) cools compressed inlet air using compressed air from the turbo-compressor (304). A combustor (305) receives compressed cooled air from the turbo-compressor and a first portion of the cryogenic fuel for combustion. A first turbine (306) expands and is driven by combustion products, and a second turbine (307) expands and is driven by a second portion of the cryogenic fuel. The first turbine and the second turbine drive the turbo-compressor via a shaft. An afterburner (309) receives combustion products from the first turbine and the second portion of the cryogenic fuel from the second turbine for combustion therein.

The application discloses a combined engine and a control method thereof. The combined engine comprises a control assembly, an air inlet channel, a turbine core engine, a nuclear reactor system, a propellant supply system and a nozzle assembly which are connected through pipelines. The control assembly is used for controlling the air inlet end of the nuclear reactor system to be selectively connected with the air inlet channel, the turbine core engine or the propellant supply system through the pipelines. The air outlet end of the nuclear reactor system is connected with the nozzle assembly. The air inlet end of the nuclear reactor system is connected with the air inlet channel, the turbine core engine and the propellant supply system through the pipelines respectively by the control assembly, so that the combined engine enters the nuclear ram mode, the nuclear turbine mode and the non-suction mode respectively. The combined engine has the characteristics of high fuel economy, working in the atmosphere-free environment, realizing different speed domain flights by using different engine modes and meeting the needs of space transportation of the super-speed aircraft.

The invention discloses a double-flow-channel combustion chamber and a rocket-based combined power engine. The problem that a combustion chamber cannot stably work in a wide range in a high-performance mode is solved. The first flow channel is provided with a first isolation section and a first combustion chamber, and a first flame stabilizer is arranged on the side, close to the first isolation section, of the first combustion chamber; the second flow channel is provided with a second isolation section and a second combustion chamber, a second flame stabilizer is arranged on the side, close to the second isolation section, of the second combustion chamber, the first flame stabilizer and the second flame stabilizer are both used for injecting fuel, the second flow channel is parallel to the flowing direction of the first flow channel, and the expansion ratio of the second flow channel is larger than or equal to that of the first flow channel; the first combustion chamber and the second combustion chamber communicate with the second-stage combustion chamber, and the expansion ratio of the first flow channel and the expansion ratio of the second flow channel are both smaller than the expansion ratio of the second combustion chamber. The power device communicates with the second-stage combustion chamber, and the jetting direction of the power device inclines towards the flowing direction of the second flow channel and is used for jetting high-temperature wake flow into the second-stage combustion chamber.

A system and method for ram air intake for pulse combustion systems is disclosed that improves the ability of pulse combustions to ingest air into the inlet pipe when the pulse combustion system is moving in a direction opposite the direction the open end of the inlet pipe is facing and the system and method includes the ability to increase the thrust output from the pulse combustion system.

The present disclosure relates to the technical field of engines, and particularly provides a wide-speed-range ramjet engine and a combined power device. The ramjet engine comprises a ram channel, a front-section strut fuel injector, a front-section cavity flame stabilizer, a rear-section strut fuel injector, a rear-section cavity flame stabilizer and a high-temperature jet flame stabilizer combination. The ram channel comprises, in sequence, an air inlet channel, a separation section, a front section of a scramjet combustion chamber, a rear section of the scramjet combustion chamber and a divergent nozzle. The front-section strut fuel injector and the front-section cavity flame stabilizer are sequentially arranged on the inner wall of the front section of the scramjet combustion chamber along the front-rear direction. The rear-section strut fuel injector and the rear-section cavity flame stabilizer are sequentially arranged on the inner wall of the rear section of the scramjet combustion chamber along the front-rear direction. The high-temperature jet flame stabilizer combination is arranged on the inner wall of the divergent nozzle close to the rear section of the scramjet combustion chamber. The present disclosure can reduce the comprehensive technical difficulty of combining a high-speed turbine engine with a scramjet engine.

Disclosed is an aircraft having a rocket-based combined cycle (RBCC) propulsion system including a rotating detonation rocket engine (RDRE) and an airbreathing jet engine. The RDRE is located within a throat area of the jet engine and is configured to be cooled by air bypassed from upstream of the jet engine.

The application relates to an RBCC engine propellant supply system and a control method thereof, wherein the RBCC engine propellant supply system comprises a liquid nitrogen storage tank, a liquid oxygen storage tank communicated with the liquid nitrogen storage tank through a first pipeline, a first flow valve, a first booster pump and a first gasifier arranged on the first pipeline, and the first gasifier is close to the liquid oxygen storage tank; the liquid oxygen storage tank is communicated with a combustion chamber of a rocket engine through a second pipeline, and a second flow valve and a second booster pump are arranged on the second pipeline; a kerosene storage tank is communicated with the liquid nitrogen storage tank through a third pipeline, a third flow valve, a third booster pump and a third gasifier are arranged on the third pipeline, and the third gasifier is close to the liquid oxygen storage tank; the kerosene storage tank is communicated with the combustion chamber of the rocket engine through a fourth pipeline, and a fourth flow valve and a fourth booster pump are arranged on the fourth pipeline; and the kerosene storage tank is communicated with a combustion chamber of a ramjet engine through a fifth pipeline, and a fifth flow valve and a fifth booster pump are arranged on the fifth pipeline.

The application discloses a kind of helium cooling inner channel air multi-modal combined power cycle system and method, system can carry out the conversion of 4 kinds of working modes by adjustable geometry, especially in turbine ram switching mode, helium is as closed quasi-brayton cycle working substance, by cooling and heating to the air of inner channel air compressor front air, significantly reduce the power consumption of turbine mode compressor, improve the thrust of mode conversion process;At the same time, helium is as intermediate medium and transmits the heat of inner channel air to outer channel air, improves the thrust of outer channel air.The turbine ram combined engine of the present application compared with the conventional jet precooling improves the thrust of mode conversion process, will not affect the oxygen content of compressor inlet flow, the total pressure loss of flow is small, can improve the pressure ratio of compressor under turbine ram switching mode, improve the working environment of engine, solve the "thrust gap" problem when working mode switches, realize the working requirement of wide speed range.

The invention relates to an aircraft and a parallel rotary detonation turbine ramjet combined engine thereof. The parallel type rotary detonation turbine and ramjet combined engine comprises a turbine engine, a ramjet engine and a mode switching mechanism. The turbine engine comprises a main combustion chamber, a gas compressor and a turbine. In the working process of the parallel rotary detonation turbine ramjet combined engine, when the speed of an aircraft is lower than the Mach number 2.5, the air inlet of the rotary detonation combustion chamber is closed by the mode switching mechanism, and the air inlet of the main combustion chamber is opened. When the speed Mach number of the aircraft ranges from 2.5 to 3.0, the mode switching mechanism opens air inlets of the rotary detonation combustion chamber and the main combustion chamber. When the speed of the aircraft is higher than the Mach number 3.0, the mode switching mechanism opens the rotary detonation combustion chamber and closes the air inlet of the main combustion chamber, so that the working service performance of the multi-stage afterburner under the high-temperature condition is improved, it is guaranteed that the speed mode switching of the aircraft is stable, and the flight stability of the aircraft is guaranteed.

The embodiment of the application provides an engine and an aircraft. The engine comprises an air intake system, a pre-cooling supercharging system, a turbocharging system, a ram combustion system and an exhaust system; the pre-cooling supercharging system and the ram combustion system comprise a high-pressure compressor, the turbocharging system comprises a low-pressure compressor, in a first working mode, airflow enters the turbocharging system for supercharging, is combusted into gas in the turbocharging system and the ram combustion system, and finally expands in the exhaust system to generate thrust; in a second working mode, airflow is pre-cooled and supercharged by the pre-cooling supercharging system, then enters the turbocharging system for supercharging, is combusted into gas in the turbocharging system and the ram combustion system, and finally expands in the exhaust system to generate thrust; in a third working mode, airflow is combusted into gas by the ram combustion system, and finally expands in the exhaust system to generate thrust. The engine solves the problems of zero-speed starting and crossing "thrust trap", and realizes continuous and efficient operation.

This invention discloses a combined power cycle system and method for cooling incoming air using ammonia-water, relating to the aerospace field. The system can switch between turbine-ramjet operating modes through an adjustable structure. Particularly during turbine-ramjet switching, ammonia-water is used to cool the high-temperature incoming air in the intake duct. The absorbed ammonia-water is then distilled to separate high-temperature ammonia gas and water. This invention combines heat exchanger precooling and jet precooling, effectively increasing the pressure ratio of the fan and compressor through multiple air cooling processes. This improves the compressor outlet gas temperature, thereby reducing compressor power consumption and providing greater thrust. Ammonia, as a supplementary fuel, has a higher specific impulse, addressing the thrust gap during engine mode transitions and effectively widening the operating speed range of the turbine engine.

The application relates to the field of aerospace power design and manufacturing, and discloses a multi-combustion mode combined hydrogen fuel aviation power system for a super-speed aircraft, which is based on a single-rotor turbojet engine and is additionally provided with an outer channel detonation combustion chamber and a scramjet engine after the turbine, is compact and simple in structure, is in series combination of various parts, can minimize the volume and weight of the power system, is more suitable for small and medium-sized aircrafts for super-speed, can meet the wide-speed-domain flight of the aircraft, can effectively utilize the high-speed and high-pressure airflow of the outer channel by adjusting the opening and closing mode of the nozzle with the suction hole to adapt to different working modes of the combined engine, and can meet the technical requirements of low carbon, high energy efficiency, high speed and long-time flight when the turbine engine works independently, the turbine engine and the detonation combustion chamber work together, the scramjet engine works independently and the scramjet and the turbine engine work together.

Disclosed is an aircraft having a rocket-based combined cycle (RBCC) propulsion system including a rotating detonation rocket engine (RDRE) and an airbreathing jet engine. The RDRE is located within a throat area of the jet engine and is configured to be cooled by air bypassed from upstream of the jet engine.

The invention provides an inner cavity air entraining type axisymmetric rocket-based combined engine. The inner cavity air entraining type axisymmetric rocket-based combined engine comprises a cone, a rocket engine, a shell and an air entraining component. The cone is divided into a cone part and a containing part in the axis direction, and the end, away from the cone part, of the containing cavity is an open end. A bent part is arranged at the joint of the cone part and the accommodating part; the rocket engine is provided with an injection part and a body part, and a flow guide gap is formed between the body part and the containing part at the opening end. One end of the shell sleeves the accommodating part, and an air inlet channel is formed between the shell and the accommodating part; the air entraining component is arranged in the containing part, one end of the air entraining component communicates with the outside at the position of the bent part, and the other end communicates with the shell through the flow guide gap. The additional air flow is introduced near the outlet of the spray pipe, so that the speed and pressure distribution of the convergent flow beam are improved, the shock wave intensity is weakened, and the total pressure loss is further reduced. The interference between shock waves and a boundary layer is reduced through the suction effect of the air entraining guide pipe on the boundary layer of the air inlet cone, so that the flow stability of the air inlet channel is improved.

Methods and apparatus for power generation and / or thermal management on a high speed flight vehicle include circulating a power generation loop working fluid through a power generation loop, which absorbs heat associated with the flight vehicle. A generator operably coupled to the power generation loop generates electrical power. Additionally, a heat transport loop working fluid may circulate through a heat transport loop to provide thermal management through heat transfer.

This invention discloses a nozzle control oil supply device for a parallel turbo-ramjet combined engine, comprising a turbine nozzle control component, a ramjet nozzle control component, a turbopump, a ramjet pump, and an oil circuit switching device. The oil circuit switching device includes a first reversing valve and a second reversing valve. The first reversing valve has a first inlet, a second inlet, a first outlet, and a second outlet on its valve body. The second reversing valve has a third inlet, a fourth inlet, a third outlet, and a fourth outlet on its valve body. The first and third inlets communicate with the outlet of the ramjet pump, the second and fourth inlets communicate with the outlet of the turbopump, the second outlet communicates with the turbine nozzle control component, and the third outlet communicates with the ramjet nozzle control component. The purpose of this invention is to provide a nozzle control oil supply device for a parallel turbo-ramjet combined engine capable of maintaining control oil in both the turbine duct nozzle and the ramjet duct nozzle for an extended period.

An engine in which thrust is achieved by converting electrical energy into high temperature plasma discharges that, in turn, apply thermal, pressure, and / or kinetic energy to a stream of passing air. The engine comprises a plasma region that includes a pair of gapped electrodes, such that the plasma discharges occur in the electrode gap. An energy storage device generates voltage pulses between the electrodes that electrically break down the air as the operating medium within the electrode gap and create plasma discharges.

The invention relates to an aircraft and a tandem rotary detonation turbine ramjet combined engine thereof. The invention discloses a tandem type rotary detonation turbine ramjet combined engine which comprises a main combustion chamber, a rotary detonation afterburner, a gas compressor, a ramjet duct and a mode switching mechanism. In the operation process of the tandem type rotary detonation turbine ramjet combined engine, the mode switching mechanism is arranged at an air inlet of the ramjet duct and an air inlet of the turbine engine and used for opening or closing the ramjet duct and the turbine engine, when the speed is lower than the Mach number 2.5, the mode switching mechanism closes the ramjet duct, and when the speed is lower than the Mach number 2.5, the mode switching mechanism closes the ramjet duct; and an air inlet of the turbine engine is opened, and the first state is switched. And when the speed Mach number ranges from 2.5 to 3.0, the mode switching mechanism opens the stamping duct and an air inlet of the turbine engine to be switched to the second state, and smooth switching of the speed mode is ensured. And when the speed is higher than the Mach number 3.0, the mode switching mechanism opens the stamping duct, the air inlet of the turbine engine is closed, and the third state is switched.

A high-speed, air-breathing propulsion engine includes a rotating detonation combustor configured to burn a fuel-air mixture and at least one ram scoop. The at least one ram scoop is configured to separate incoming air into a core flow stream and a pilot flow stream. The at least one ramp scoop is configured to cause the separation of the core flow stream and the pilot flow stream at a location disposed within a range of locations. The engine is configured to combine the pilot flow stream and the core flow stream in the rotating detonation combustor to enable a rotating detonation wave in the rotating detonation combustor.

Scramjet engine (1) for a terrestrial vehicle (1) comprising a case (3) having a front part (4) and a rear part (5), an air inlet portion (6) at the front part (4) of the case (3) facing air flow for receiving and compressing incoming air, a combustion portion (7) and a fuel injection module (8) coupled to the combustion portion (7) located downstream the air inlet portion (6) for heating and expanding air, a nozzle portion (9) downstream the combustion portion (7) at the rear part (5) of the case (3) for expelling air and generating a thrust propelling the vehicle (2), and a control unit (10) connected at least to the air inlet portion (6) to modulate the air flow inside the engine, wherein the air inlet portion (6) comprises a multi-phase inlet and compression module (11) connected to the control unit (10), the multi-phase inlet and compression module (11) having an adjustable compression action for progressively compressing air and modulating the speed of the incoming air to a maximum speed value that is lower than the sound speed.

The application discloses a time sequence control rotating detonation engine system combined with a turbojet engine, which comprises a turbojet engine, a shell sleeved on the periphery of the tail nozzle of the turbojet engine, a cylindrical base coaxially arranged on the inner side of the shell, and a fan-shaped detonation tube. A cover driven by an electromagnetic push rod is arranged at the top end of each detonation tube. A sequential ignition control device is arranged, and a fuel supply unit of the sequential ignition control device comprises an oil pump, a fuel coil pipe coiled in the cylindrical base, a common rail ring outside the shell, and a plurality of fuel injectors. A control unit adopts a mechanical time sequence controller, and through rotating a moving contact on an outer rotor, the mechanical time sequence controller sequentially connects static contacts on an electrode distributor, synchronously triggers electromagnetic push rods and fuel injectors of corresponding detonation tubes, and realizes multi-tube sequential cycle detonation. The application recovers exhaust heat energy to preheat fuel through a coaxial integrated structure, supplies air by using an ejecting effect, replaces a complex electronic system with a simple and reliable mechanical control, and realizes efficient, compact, wide working condition adaptation and thrust enhancement.

An air breathing engine system has an engine inlet configured to receive a flow of air. A combustor is in fluid communication with the engine inlet. A rocket ejector is positioned between the engine inlet and combustor. The rocket ejector is connected to the combustor with a mixing tube. The rocket ejector adjustably outputs high-temperature gas into air in the mixing tube. The high-temperature gas changes a pressure and thermal energy of the air within the mixing tube, which can be used to optimize combustor operation, such as based on a trajectory, flight condition, or speed of an aircraft propelled by the air breathing engine system.

A jet engine for propelling aircraft, capable of providing thrust from rest to high speeds. Characterised in that it employs a liquid oxidant and two liquid fuels of different molecular weights. The engine has an axial compressor (16) or several axial compressors located on the same plane, but differs from a conventional gas turbine in that the turbine is driven by a gas generator. At the outlet of the turbine there is a gasification chamber (23) into which more fuel is injected. Combustion of the gases from the gasification chamber is performed in two combustion chambers (18) with a rectangular cross-section, separated by a central body (10). The exhaust of the gases is performed in nozzles with a square convergent / divergent cross-section (19) and (21). The cross-section of the throats (26) can be adjusted by means of two mobile elements (20). The final section of the central body (10) forms a wedge-shape (27), enabling the continued expansion of the exhaust gases.

The application discloses a kind of earth Mars round trip powder metal fuel rotating detonation rocket ram combined cycle engine, belong to aerospace engine technical field, combined engine includes: rocket rotating detonation engine and ram rotating detonation engine, whole engine system uses powder metal as main fuel, by rocket rotating detonation engine and ram rotating detonation engine respectively organize combustion and jointly generate thrust.The application can give full play to rocket engine, ram engine and rotating detonation combustion mode respective advantage, utilize the advantages that powder metal fuel stability is good, energy density is high and can occur combustion reaction with multiple oxidizing agents, can satisfy future reusable aerospace vehicle, earth Mars round trip interplanetary spacecraft etc combined power propulsion demand.

A rocket-based combined cycle propulsion system comprising an air inlet and a centerbody extending along a center axis. The centerbody at least partially defines the air inlet. The system further includes a first propulsion system (e.g., a rotating detonation rocket engine, RDRE) and a second propulsion system (e.g., a scramjet) coupled with the first propulsion system. The system further includes an air bleed assembly including an inlet aperture formed in the centerbody, an outlet aperture positioned in the first propulsion system, and a duct fluidly connecting the inlet aperture and the outlet aperture.

The application discloses a shaft symmetrical three-channel adjustable turbine ram combined power nozzle, wherein an inner channel and an outer channel are turbine engine nozzle and ram engine nozzle respectively, the turbine engine nozzle channel is divided into an inner duct flow channel and an outer duct flow channel, an inner plug cone and an inner duct wall form the flow channel of the inner duct, the inner plug cone can be simply moved in translation, an outer plug cone is in the outer duct and forms the flow channel of the outer duct together with the outer duct wall, and the inner and outer plug cones form physical throat channels, the outer flow channel is a ram channel, the inlet is supersonic airflow and has no contraction section, the opening and closing of the inner and outer ducts are controlled by adjusting the positions of the inner and outer plug cones, the shape of the joint is inclined by a certain angle, the size of the throat channel and the expansion ratio can be adjusted, the movements of the inner and outer plug cones are independent, and the independent and linkage conditions can be realized to achieve different flow rates and working states, and the combined power engine can meet the demand under complex working conditions.

The disclosure discloses a fan variable flow path series turbine-based combined engine, which has two channels of a turbine engine and a ramjet engine. The turbine engine has two internal and external ducts, and contains components: inter-stage bleed fan, high-pressure compressor, main combustion chamber, high-pressure turbine, low-pressure turbine; the ramjet engine includes components: ramjet duct; after the turbine engine and the ramjet engine are connected in series to form a combined engine, the components used in common include: mixing chamber, afterburner / ramjet combustion chamber, integrated two-element nozzle of the ramjet duct. The fan variable flow path series turbine-based combined engine is provided with a front duct splitter ring, a fan inter-stage splitter ring, a duct conversion valve and a rear duct ejector structure. The engine adopts a framework design of double rotors, three ducts, mixed exhaust, main combustion chamber and afterburner / ramjet combustion chamber arranged in series, and the area ratio of the ramjet duct to the turbine duct is set to be small, and the overall cross-sectional area of the combined engine is small.