53results about "Composite engine plants" patented technology

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.

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.

A combined power intake duct with a translational linkage flow regulation geometry device and its operating method are disclosed. This invention features a wide-range intake flow regulation range. Based on the principles of changing the capture flow and increasing the throat flow capacity for stable intake operation, this invention proposes a linkage flow regulation geometry device involving the translational lip and throat. The geometry device consists of a translational lip, a translational throat, and a four-bar linkage. It can achieve multi-dimensional optimization of intake performance by coordinating and controlling the actual capture flow and throat compression through single-degree-of-freedom adjustment. The geometry device operates in two states during movement: 1) When the four-bar linkage moves backward, the actual capture flow decreases, the ram throat height increases, and the turbine throat height decreases; 2) When the four-bar linkage reaches a certain position and continues to move backward, the actual capture flow decreases, the ram throat height decreases, and the turbine throat height increases.

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.

The application provides a rotary ram shock wave supercharged gas turbine engine and belongs to the technical field of turbine engines, which comprises an outer casing and an inner casing, an outer channel is formed between the inner casing and the outer casing, a guide vane is installed in the outer casing and is used for guiding airflow, a first mechanism is installed in the inner casing and is used for supercharging and heating the airflow and doing work, a second mechanism is arranged on one side of the inner casing and is used for supercharging and heating the airflow for the second time. The application solves the problems of large windward area of the centrifugal compressor, many parts of the multistage axial flow compressor and low efficiency of small flow by using the rotary ram compressor, higher flow and pressure ratio are obtained in compact size, greater thrust is obtained under the condition of smaller outer size of the original engine, the outer size is reduced by about 32%, and the thrust is increased by 17.3% under the condition of the same fuel consumption rate when the engine is at rest.

The invention relates to the technical field of rocket engines, in particular to a solid-liquid mixed combined type ramjet engine which comprises a rocket body, supporting parts are fixedly installed at the front end of the rocket body and distributed annularly, and a spray pipe system is fixedly installed at the front end of each supporting part and connected with the rocket body. The supporting part comprises a positioning device, a transmission device and a displacement device, the displacement device is slidably inserted into the inner center of the positioning device, and the transmission device is fixedly installed in the positioning device and located above the displacement device; the positioning device comprises a positioning cavity, a combustion chamber, a heat insulation sleeve layer, a solid fuel column, a guide sleeve, a connecting base frame, a driving motor, a transmission worm, a movable hole and a limiting groove, the combustion chamber is fixedly installed at the front end of the positioning cavity, the heat insulation sleeve layer is fixedly installed on the inner wall of the combustion chamber, and the solid fuel column is fixedly installed on the inner wall of the heat insulation sleeve layer. And through the arrangement of the supporting component, the purpose that an oxidizing agent can be injected into different areas in the combustion column is achieved.

The invention discloses a water-air dual-purpose engine. The engine takes air or water as a combustion medium; the engine includes: a medium introduction part for introducing a combustion medium into the engine; the trapped vortex combustion part is used for accommodating the fuel and the combustion medium under the condition that the combustion medium is air, so that the fuel is combusted; the metal powder combustion part is used for accommodating fuel and a combustion medium under the condition that the combustion medium is water, so that the fuel is combusted; and the substance spraying part is used for spraying substances generated by combustion out of the engine so as to generate traction force. According to the water-air dual-purpose engine, the two different propelling modes of turbojet and metal powder water stamping are integrated, the two propelling modes are compatible at the same time according to different environments and working conditions, and the working states are synchronously switched during environment switching so as to adapt to combustion media corresponding to the different environments; therefore, the engine can provide power for the aircraft in different propulsion modes in the air environment and the water environment respectively, and the stability of fuel combustion is kept.

The present application relates to a ram air turbine emergency power drive device, comprising: a first set of turbine blades and a second set of turbine blades coaxially arranged in an axial direction, wherein the second set of turbine blades can be rotated relative to the first set of turbine blades around the axial direction from a first position to a second position and kept in the second position, wherein in the first position, the first set of turbine blades is superposed with the second set of turbine blades, and in the second position, the first set of turbine blades is at a predetermined non-zero angle with the second set of turbine blades. In this way, the number of blades is increased without affecting the space in the ram air turbine cabin, so that the turbine starting speed is faster, and the RAT access to the power grid time is reduced. In addition, increasing the number of blades can also improve the efficiency of the turbine in converting wind energy, thereby improving the efficiency of the emergency power generation. The present application also relates to an aircraft comprising the ram air turbine emergency power drive device.

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 invention belongs to the technical field of aero-engine control, and discloses an anti-interference switching control method for a rocket ramjet combined engine. The working mode and the control requirement of the rocket ramjet combined engine are analyzed, and the control loop switching rate based on the mode and protection is designed. For different working modes and flight conditions, feedforward adjustment is introduced on the basis of a linear active disturbance rejection algorithm, and the rapidity and disturbance rejection performance of adjustment of the rocket ramjet in all states are guaranteed. Furthermore, the invention provides a mixed residence time method depending on a system state, so that the conservative property of switching limitation is reduced, and the effects of ensuring the action sensitivity when the engine is close to the switching demand and ensuring the switching stability of the engine are achieved. The smooth switching strategy is introduced at the switching moment, a control system is built, and finally the self-adaptive switching control function of the rocket ramjet combined engine is achieved.

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.

The invention discloses a binary strong pre-cooling air inlet system with a flow channel switching function and a power device, and belongs to the field of hypersonic flight vehicle power devices. The device comprises a binary air inlet channel, a strong pre-cooling section with a flow channel switching function, an inlet transition section and a power device. A pre-cooling channel is closed under the low Mach number through a two-dimensional flow channel switching function based on rotation adjustment, so that air flow bypasses a strong pre-cooler and directly enters a turbine engine, and the total air inlet pressure loss is reduced; and the pre-cooling channel is opened under the high Mach number, so that the airflow flows through the strong pre-cooler to be pre-cooled, the service life of engine parts is prolonged, the performance of the engine is improved, and the low Mach number performance and the high Mach number performance of the strong pre-cooling turbine engine are both considered.

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 provides a super-combustion ramjet engine spanwise pulse injection device, which comprises a spanwise injection cylinder and at least two pulse injection assemblies in parallel and communicated with the spanwise injection cylinder, the pulse injection assembly comprises a first valve body, a second valve body, a movable valve core arranged at an outlet end of the second valve body and a nozzle seat which are sequentially arranged along a fuel flow direction; at least two buffer cavities which are symmetrical along a second valve body axis are arranged at a connection position of a third valve seat and a second valve seat, a buffer elastic member is arranged in the buffer cavities, and a pressure relief pipeline which is communicated with the second cavities is arranged at one end of the buffer spring close to the buffer cavities. Advantageous effects: through the design of four buffer cavities, a pressure relief pipeline and a buffer elastic member, when the pressure in the pipeline changes, the buffer elastic member is compressed or rebounds, the water hammer effect is effectively reduced, the rate of fuel entering the second cavities from the pipeline is accelerated, the injection rate is further accelerated, the pulse lag phenomenon is reduced, and the pulse frequency is improved.

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.

The application discloses a combined engine system, an aerospace vehicle and a flight control method and device, and relates to the field of aerospace vehicles.The combined engine system comprises a detonation ramjet engine device, a detonation rocket engine device and a turbojet engine device, wherein the detonation rocket engine device is arranged inside the detonation ramjet engine device, and the turbojet engine device is arranged outside the detonation ramjet engine device; the detonation rocket engine device and the turbojet engine device are arranged to be capable of working independently or working in combination to form a combined mode; and the working modes of the combined engine system include a ramjet mode, a rocket mode, a turbojet mode and a combined mode.The combined engine system is suitable for more application scenarios and can reduce the design of an oxidant cabin.

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.

The invention discloses a water ramjet engine with a bluff body cavity powder inlet structure. The water ramjet engine comprises a combustion chamber; the fuel gas generator is used for providing high-temperature fuel gas for the combustion chamber, and the fuel gas generator is fixedly connected to the front end of the combustion chamber; the powder conveying pipe is fixedly assembled in the fuel gas generator, the front end of the powder conveying pipe is connected with the powder supply device, and the rear end of the powder conveying pipe extends into the combustion chamber from a gas inlet in the front end of the combustion chamber; the exhaust nozzle is fixedly communicated with the rear end of the combustion chamber; and the bluff body is fixedly communicated with the rear end of the powder conveying pipe, and an opening of the bluff body faces the exhaust nozzle. The device has a good mixing combustion effect, and meanwhile deposition on the wall face of the combustion chamber and the surface of the bluff body can be reduced.

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.

The application discloses a kind of aircraft energy and heat comprehensive management system, belong to control, adjusting technical field.The system integrates turbine-based combined cycle engine subsystem, fuel heat management subsystem, environmental control subsystem, aircraft accessory subsystem, through from TBCC engine subsystem carry out air bleeding and energy extraction, coordinate the joint work between each subsystem, realize whole machine energy distribution and heat management regulation and control, to respond different engine working state and energy efficiency demand, system has stage regulation and mode switching capability, can be according to the multiple flight conditions in flight envelope including take-off, climb, ram cruise, return carry out power distribution and heat management regulation and control, greatly enrich the selection of effective heat sink and improve heat sink utilization efficiency, realize the temperature of optimal regulation and control different position heat source, effectively reduce weight compensation, improve the maneuverability and long endurance capability of aircraft and reduce flight cost.

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.

A device for improving the jet cooling effect of a TBCC combined engine provides a jet cooling technical route of an engine air inlet channel in which a cooling medium is mixed with carbon powder, the device comprises a carbon powder supply mechanism and a mixed jet mechanism, and the carbon powder enters the mixed jet mechanism from the carbon powder supply mechanism and is mixed with the cooling medium to form carbon-containing liquid drop particles to realize jet. The method comprises the following steps: before the carbon powder is mixed with a cooling medium, pretreating the carbon powder with a surfactant to improve the wettability of the surface of carbon powder particles, then mixing the carbon powder with the cooling medium, and wrapping the carbon powder with the cooling medium to form carbon-containing liquid drop particles with larger particle sizes. Then the carbon-containing liquid drop particles are injected into a high-temperature air inlet channel of the engine at a high speed, the outer-layer cooling medium in the carbon-containing liquid drop particles is used for achieving the effect of enhancing heat exchange through phase change, and the remaining carbon powder particles in the carbon-containing liquid drop particles can enter a combustion chamber of the engine to serve as fuel for auxiliary combustion; and finally, the jet cooling effect of the TBCC combined engine is improved.