91results about "Composite engine plants" patented technology

The invention discloses an integrated closed circulation system for an RBCC engine, and solves the problems that a constant-pressure extrusion mode of an existing RBCC engine propellant supply system is not suitable for variable working condition adjustment of an RBCC rocket thrust chamber, and a high-pressure storage tank is huge in structure, large in weight and not beneficial to application and expansion of the system on the day. According to the system, high-temperature oxygen is generated through catalytic decomposition of hydrogen peroxide and acts on the turbine, then the hydrogen peroxide pump and the coal oil pump are driven to work, integrated pressurization, supply and adjustment of propellants can be carried out on the multiple rocket thrust chambers and the ramjet combustion chamber at the same time, and therefore wide-range fuel oil adjustment is met, and a widely-applied electric pump system can be replaced; and the structural envelope and weight are greatly reduced, so that the power requirement of a longer voyage is met.

The invention discloses a nitrogen pressurization system and method of an RBCC engine. The technical problem that an existing nitrogen pressurization system is poor in safety and testability is solved. The pressurization system comprises a nitrogen gas cylinder, a stamping fuel oil storage tank, an oxidizing agent storage tank, a fuel storage tank and a power controller. The nitrogen gas cylinder is connected with the inlet end of the stamping fuel oil storage tank, and a self-locking valve, a first-stage pressure reducing valve and a second-stage pressure reducing valve are sequentially arranged on a connecting pipeline. The power controller is electrically connected with the self-locking valve; the outlet end of the stamping fuel storage tank is connected with the stamping combustion chamber; the outlet ends of the oxidant storage tank and the fuel storage tank are respectively connected with a rocket thrust chamber, and a first gas explosion valve and a second gas explosion valve are respectively mounted on connecting pipelines; a first branch is arranged between the self-locking valve and the first-stage pressure reducing valve and connected with the first gas explosion valve and the second gas explosion valve. A second branch is arranged between the first-stage pressure reducing valve and the second-stage pressure reducing valve and connected with the inlet end of the oxidizing agent storage box and the inlet end of the rocket fuel storage box.

The invention particularly relates to a hypersonic aerospace engine propelled by electric energy and chemical energy cooperatively. The hypersonic aerospace engine comprises an electric energy propelling assembly and a chemical energy propelling assembly. The electric energy propulsion assembly comprises a first medium-voltage magnetic suspension motor and a second medium-voltage magnetic suspension motor which are of an outer rotor motor structure and an inner rotor motor structure respectively. An air pressurization channel is arranged between the first motor rotor and the second motor rotor; the chemical energy propelling assembly comprises a combustion chamber, a fuel nozzle and an igniter; the air ionization electrification device is used for electrifying the air after ionizing the air; the combustion chamber is connected to the rear side of the electric energy propelling assembly and communicates with the air pressurization channel. The fuel nozzle and the igniter are both arranged at an outlet of the air pressurization channel. The electric energy and chemical energy collaborative propulsion mode has better controllability and operation stability; the zero-speed starting can be directly realized after electrification, an additional starter is not needed, the structural complexity is reduced, the flying distance is higher, and the voyage is farther.

A propulsion system includes a turbine engine and a ramjet engine. A turbine engine inlet section restriction device permits an inlet airflow to a turbine engine inlet in a first (low speed) operating state of the propulsion system, and restricts the inlet airflow to the turbine engine inlet in a second (high speed) operating state of the propulsion system. A combustor inlet flow control device, in the first operating state, provides an airflow from a compressor to the turbine engine combustor, and restricts an airflow from a ramjet bleed air passage, and, in the second operating state, provides a ramjet inlet bleed airflow to the turbine engine combustor, while restricting airflow from the compressor. In the second operating state, the turbine engine drives at least one accessory gearbox to drive at least one accessory component.

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.

The invention discloses a pressure-adjustable fuel gas pressurization system and an RBCC engine. The technical problem that an existing pressurization system cannot meet the variable working condition adjustment requirement of a thrust chamber of the RBCC engine is solved. The system comprises a pressurizing storage tank, a gunpowder starter, a generator, an oxidizing agent storage tank, a fuel storage tank and a controller, the inlet end of the pressurizing storage tank is connected with the gunpowder starter, and the outlet end is connected with the inlet end of the generator through the first electric explosion valve, the electric pump and the first one-way valve; a pressure sensor is installed at the outlet end of the generator and connected with the controller. The outlet end of the generator is connected with the inlet end of the pressurizing storage box and is connected with the inlet ends of the oxidant storage box and the fuel storage box; the outlet ends of the oxidant storage tank and the fuel storage tank are respectively connected with the thrust chamber; a pressure release valve is mounted between the outlet end of the electric pump and the outlet end of the oxidant valve or the fuel valve; the controller is connected with the gunpowder starter, the electric pump and all the valves and used for sending out corresponding control instructions.

The invention relates to the technical field of engines, and discloses a turbine-based combined cycle engine operation method and a turbine-based combined cycle engine, and the turbine-based combined cycle engine operation method comprises the steps that a pulse detonation combustion mode is adopted in a pure turbine mode; a rotary detonation combustion mode is adopted in a pure stamping mode; a pulse detonation combustion mode is adopted in the transition mode, and meanwhile, a rotary detonation combustion mode is adopted. By utilizing the advantages of self-pressurization and high unit thrust of pulse detonation combustion, the high-speed performance is improved, and the working Mach number upper limit of a pure turbine mode is widened; by utilizing the advantages of high combustion speed, high combustion efficiency, simple structure and short length of rotary detonation combustion, the lower limit of the working Mach number of a pure stamping mode is widened. And in the transition mode, the characteristics of a pure turbine mode and a pure stamping mode are achieved at the same time, and enough thrust is provided for the aircraft.

The invention discloses an intercooler integrated device for an engine intermediate case area, which comprises an intermediate case outer case and an intermediate case inner case, a plurality of support plates are arranged between the intermediate case outer case and the intermediate case inner case, and intercooler mounting modules are fixed on the support plates. And the intercooler mounting module is used for mounting the shunting ring-intercooler integrated module between the intermediate case outer case and the intermediate case inner case. The defects in the prior art can be overcome, and the design and installation work of the newly-added intercooler can be completed on the premise that the original structure is reserved to the maximum degree and a new complex structure is not introduced.

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 present invention relates to a rocket-based combined cycle (RBCC) engine with a parallel compressor flow path, which includes a compressor flow path and an RBCC engine flow path. The compressor flow path and the RBCC engine flow path are arranged in parallel, and the two flow paths share an inlet duct. After the inlet duct flow path, a rocket thrust chamber, a primary combustion chamber, a secondary combustion chamber and a nozzle are provided; the airflow in the compressor flow path converges into the secondary combustion chamber and burns in the secondary combustion chamber.

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 present invention belongs to the technical field of tactical weapon system layout and engine structure design, and specifically relates to a solid fuel ramjet rotating detonation combined power system, comprising: an air inlet, an isolation section, a main charge, a pre-detonator tube, and a rotating detonation combustion chamber. To address the current difficulty in organizing mixed combustion in the afterburner, the present invention proposes using a rotating detonation combustion chamber instead of a traditional isobaric combustion chamber, thereby improving the engine cycle thermal efficiency and making the engine structure more compact. Theoretical calculations show that this invention can increase the engine specific impulse by more than 5%. Furthermore, the present invention provides a systematic design solution in conjunction with a solid fuel ramjet rotating detonation combined engine.

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.

The invention discloses a method and device for reducing resistance of an exhaust nozzle in an injection mode in front of the exhaust nozzle. The device is located between a combustion chamber and a single-side expansion nozzle. The front end of the tail spraying embedded plate is hinged to the upper side wall of the rear end of the combustion chamber, and the rear end of the tail spraying embedded plate is close to the expansion face of an inlet of the single-side expansion spraying pipe; the front end of the lower wall face adjustable plate is hinged to the lower side wall of the rear end of the combustion chamber, and the rear end of the lower wall face adjustable plate is close to the lower wall face of the single-side expansion spray pipe inlet. The driving mechanism is used for driving the tail spraying embedding plate and the lower wall surface adjustable plate to rotate by taking the corresponding hinge point as the center, so that the tail spraying embedding plate and the lower wall surface adjustable plate are matched with each other and adapt to the incoming flow speed; flow separation around the single-side expansion spray pipe can be reduced, so that the total resistance is reduced, and when the Mach number is increased, the tail spray embedded plate and the lower wall surface adjustable plate can be regulated and controlled by the servo motor to continuously expand so as to adapt to expansion ratios of different modes.

The invention discloses a multi-angle coordinated impact mixing enhancement device for gas in a solid rocket ramjet engine, which belongs to the technical field of solid rocket ramjet engines. The invention adjusts the uniformity and stability of the gas and air distribution in the flow field by adjusting the angle of mutual impact between multiple gas streams and the deflection angle of the gas toward the inlet side, thereby achieving enhanced mixing of the gas and air. The invention also controls the distance between the multi-angle coordinated impact nozzle and the inlet of the afterburner chamber and the axial, radial, and circumferential orientations of the installation of the multi-angle coordinated impact nozzle by cooperating with the combustion chamber docking section and the nozzle base, thereby adjusting the flow field. The nozzle does not intrude into the afterburner chamber, thereby minimizing the total pressure loss and reducing the difficulty of thermal protection. The invention also controls the multi-angle gas nozzle hole, the axial, radial, and circumferential orientations of the installation of the multi-angle coordinated impact nozzle, and the distance between the multi-angle coordinated impact nozzle and the inlet of the afterburner chamber, thereby achieving multi-dimensional adjustment of the flow field.

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 invention discloses a device for online collection of primary fuel gas of a solid rocket ramjet, and belongs to the field of rocket ramjet. The device comprises a gas storage tank, a flange connecting section, a sealing structure, a piston, an electric push rod and a remote control part. The piston is used as a movable part in the cylinder body, and the piston is of a two-sealing-ring structure. The working condition that negative pressure appears on the expansion section of the spray pipe in the low-pressure environment is adopted, real-time collection of primary fuel gas of the solid ramjet engine is achieved, and the accuracy of primary fuel gas component collection and analysis is improved; the shell, the pipeline and various pipeline joints adopt high-temperature-resistant and high-pressure-resistant design, meanwhile, the piston is additionally arranged in the storage tank, and gas in the storage tank can be completely exhausted by pushing the piston after a single test is finished. The electromagnetic valve and other devices are additionally arranged on the pipeline, opening of the electromagnetic valve or the electric ball valve is controlled through the remote controller or the pressure control starting unit, so that the remote control effect is achieved, and the safety of operators is guaranteed.

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 present application belongs to the technical field of turboramjet combination engine design, and specifically relates to a turboramjet combination engine and a control method thereof. The design separates the complex outer turbine channel from the smooth ramjet channel, and the design shares the space between the afterburner and the ramjet combustion chamber. The design of the central ramjet channel has multiple functions such as the turbine engine stator load-bearing frame, the central bleed channel, the subsonic ramjet rectifier channel, and the scramjet channel, thereby realizing adaptive switching of the four working modes of the turbine engine conventional working mode, the turbine engine central bleed working mode, the subsonic ramjet working mode, and the scramjet working mode. The space of the central ramjet channel is utilized to the greatest extent, so that the turbine mode working speed range is widened and the cross-sectional area is greatly reduced, thereby effectively reducing the flight resistance and realizing a wide speed range operation of Mach number 0-7.

The present invention belongs to the field of combined propulsion systems for hypersonic aircraft, specifically relating to a combined nozzle and its adjustment method. This invention can simultaneously adjust the turbine nozzle throat area while adjusting the diverter plate between the turbine and ramjet channels. This improves the matching between the nozzle's pressure drop ratio and area expansion ratio, alleviates the overexpansion of the ramjet channel during low-Mach-number flight and the underexpansion of the turbine channel during high-Mach-number flight, and enhances the thrust performance of the combined propulsion nozzle across a wide speed range, thereby effectively improving the flight performance of the hypersonic aircraft.