228results about "Ram jet engines" patented technology

A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

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 relates to the field of scramjet engines, in particular to a method for inhibiting combustion asymmetry of a scramjet engine, which comprises the following steps of: releasing high-temperature airflow at the upstream of shock wave string head shock wave in an isolation section of the engine, or releasing and igniting auxiliary fuel in the isolation section of the engine, therefore, high-temperature airflow is formed at the upstream of shock waves of the shock wave string head. According to the embodiment of the invention, the flow field in the isolation section is reconstructed through combustion of high-temperature airflow or auxiliary fuel, so that the quality of the flow field at the inlet of the combustion chamber is improved, and the purpose of inhibiting combustion asymmetry of the combustion chamber is finally achieved.

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 aero-engines, and discloses an efficient thermal protection hybrid thermodynamic cycle combined engine and a control method thereof.The combined engine composed of an external rotary detonation engine and an internal turbine engine is adopted, and modal switching of the combined engine is achieved through an inlet adjusting valve and an adjustable exhaust nozzle; the working upper limit of a turbine engine can be effectively joined, power relay can be completed, so that a thrust gap is spanned, the structural design complexity is relatively low, the system stability is relatively high, and in a stress application rotating detonation mode, the reliability is high. Combustible mixtures sprayed by the rotary detonation engine are guided into the afterburner of the turbine engine through the afterburner diverter valve for detonation combustion, an original cooling device of the afterburner is used for thermal protection, and an outer casing is prevented from being damaged due to high temperature.

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 a three-dimensional curved surface isolation section structure with a turning section and a design method, the isolation section comprises an isolation section part, a turning section part and a connected transition section, and the height, position and midpoint parameters of a two-dimensional molded line of a symmetry plane of an isolation section inlet section and a combustion chamber inlet section are given; an isolation section outlet section having a certain offset distance and expansion ratio with an isolation section inlet molded line is constructed as a transition section, an isolation section is a molded surface section which is relatively straight from an isolation section inlet to the transition section, and a turning section is a molded surface turning part extending from the transition section to a combustion chamber inlet; smooth transition from an isolation section inlet to a combustion chamber inlet can be achieved by adjusting the position and the rotation angle of the transition section, compared with other configurations, the configuration can flexibly adjust the contraction degree of the side wall of the isolation section so as to meet the overall layout requirement of an engine, and an aircraft can be more compact in structure, higher in volume ratio and higher in engineering application value.

A detachable strut structure in a combustion chamber of a scramjet-oblique detonation engine includes two strut components and two hydraulic telescopic mechanisms. The strut components are disposed in the combustion chamber, the combustion chamber defines an inlet and an outlet, and a connection between the inlet and the outlet is a centerline, and the two strut components are mirrored and arranged on two sides of the centerline respectively. Each strut component includes a first surface, a second surface, a third surface, the first surface is arranged parallel to the centerline, the second surface is arranged on a side of the first surface close to the inlet. An angle α is defined between the second and first surfaces. The third surface is arranged on a side of the first surface close to the outlet, an angle β is defined between the third surface and an extension line of the first surface.

The invention provides an ammonia fuel cell power generation dual-fuel scramjet engine capable of achieving energy gradient utilization, belongs to the technical field of aircraft thermal protection and power generation, and solves the problems that a traditional airborne power generation system is difficult to use due to the high incoming flow Mach number in hypersonic flight, and coking is caused in a regeneration cooling channel of a fuel decomposition gas turbine. Ammonia and aviation kerosene generate high-temperature fuel gas in a combustion chamber of the scramjet engine to push the engine to generate thrust; high-temperature waste heat of the combustion chamber is recycled through the regeneration cooling channel, fuel is heated, a part of ammonia is decomposed, and decomposed gas drives a turbine generator to generate electric energy. According to the power generation system for gradient utilization of energy, factors such as actual use temperature and pressure are fully considered, the energy utilization efficiency is effectively improved, energy waste is reduced, and reasonable adaptation of energy is achieved, so that the power generation efficiency of the whole system is improved, and the electric energy requirement of a hypersonic aircraft is greatly met.

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.

This invention belongs to the field of ramjet technology and provides a wide-speed range ramjet fuel supply system and method driven by a pyrolysis oil-gas turbine. The system comprises a fuel tank, a centrifugal pump, a motor, an oil-gas turbine, a battery, an electromagnetic control module, fuel lines, a pyrolysis fuel vapor line, a combustion chamber, and a regenerative cooling channel on the outer wall of the combustion chamber. Through the connection between the centrifugal pump, the regenerative cooling channel, and the oil-gas turbine, the system utilizes the cracking of hydrocarbon fuel into small-molecule pyrolysis kerosene vapor to drive the turbine rotor to produce work. The motor then assists in starting the ramjet, thereby resolving the power generation challenges of the ramjet mode and alleviating thermal protection issues.

The invention provides a high-Mach-number boron powder ramjet engine based on air fluidization. The high-Mach-number boron powder ramjet engine comprises a boron powder conveying module and a kerosene conveying module. The boron powder conveying module can convey boron powder to the combustion part; the kerosene conveying module can convey kerosene to the combustion part; and the combustion part is provided with a stable combustion support plate. The problems of boron powder conveying and ignition stable combustion of the high-Mach-number boron powder ramjet engine based on air fluidization are solved, and incoming flow air is directly used as driving gas and fluidizing gas to fluidize boron powder fuel; kerosene is combusted, and injected fluidized air generates a high-temperature low-speed oxygen-enriched area behind the stable combustion support plate, so that removal of an oxide layer on the surface of the boron powder fuel is accelerated, and the combustion effect of the boron powder fuel is improved. The problem that in the prior art, a boron powder fuel conveying system is too complex to be integrated with an engine system is solved, the defects that reliable ignition and stable combustion are difficult to achieve due to the poor ignition combustion characteristic of boron powder fuel are overcome, and high integration of the boron powder fuel conveying system and the engine and ignition stable combustion of the boron powder fuel are achieved.

The application discloses a strut plate embedded plasma igniter suitable for a scramjet engine and relates to the technical field of the scramjet engine.The strut plate embedded plasma igniter comprises a strut plate main body, and fuel main flow channels, oxygen main flow channels and an embedded channel are arranged on the strut plate main body; a plurality of fuel injection holes are arranged on the side wall of the fuel main flow channel; the oxygen main flow channel is in communication with the embedded channel, and an anode core is arranged in the embedded channel; a detachable head is arranged at one end of the embedded channel facing a combustion chamber, and a plurality of communication holes are arranged on the detachable head. A circular axis-symmetrical hypersonic combustion chamber is connected to the outside of the strut plate main body. The strut plate embedded plasma igniter suitable for the scramjet engine has the advantages of a stable combustion strut plate and a plasma igniter, has the characteristics of good insulation and partial stable combustion, and can ensure stable ignition of the axis-symmetrical supersonic combustion chamber.

The invention discloses a two-stage supporting plate injector applied to a wide-working-condition scramjet engine combustion chamber, belongs to the technical field of scramjet engines, solves the problem that an existing alternating supporting plate is low in fuel and air mixing efficiency under the conditions of a high Mach number and a high equivalence ratio, and comprises a supporting plate body. One end of the supporting plate body is a wedge-shaped tail edge formed by arranging a plurality of wedge-shaped slopes in a staggered mode, and first-stage fuel oil spraying holes are formed in the staggered points of the wedge-shaped slopes in the wedge-shaped tail edge. The two side faces of the supporting plate body are symmetrically provided with a plurality of second-stage fuel oil spraying holes inclined in the wedge-shaped tail edge direction. The first-stage fuel injection hole and the second-stage fuel injection hole form a two-stage fuel injection system, the fuel input mode is changed fundamentally, and the fuel flow ratio of the two-stage fuel injection system is dynamically adjusted according to different inlet Mach numbers and equivalence ratios, so that the two-stage support plate injector can be used in various working conditions, and the service life of the two-stage support plate injector is prolonged. And the optimal mixing efficiency can be realized through the optimal fuel distribution strategy.

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 invention discloses a distributed discharge-bypass pressurization coupled wide-area ramjet engine and a working method thereof, and belongs to the field of wide-area ramjet engines, the distributed discharge-bypass pressurization coupled wide-area ramjet engine comprises a bypass pressurization flow channel and a dual-mode ramjet combustion chamber flow channel, and the dual-mode ramjet combustion chamber flow channel comprises a fixed geometric air inlet channel; the bypass supercharging flow channel comprises a distributed drainage hole communicating with the compression face side of the fixed geometric air inlet channel, and a drainage cavity, a first Laval nozzle, a bypass supercharging chamber, a second Laval nozzle and a bypass exhaust nozzle which communicate with the distributed drainage hole in sequence. According to the distributed discharge-bypass supercharging coupled wide-area ramjet engine and the working method thereof, by arranging the bypass supercharging flow channel, the problem of structure and control complexity and difficulty caused by adjustable parts in integrated adjustment of the engine is solved, meanwhile, the starting capacity and thrust gain under the low Mach number are increased, and the starting efficiency is improved. And the engine has the wide-speed-range working capacity.

The invention provides a solid powder pre-combustion assembly and a heater assembly, comprising: a fluid interconnection device having a first interface for receiving fluidized powder fuel, a second interface for receiving fluidized powder fuel, and a third interface for receiving fluidized powder fuel; the first fuel gas generator is used for receiving auxiliary gas and generating high-temperature fuel gas, and a gas outlet of the fuel gas generator is communicated with the second connector of the fluid interconnection device; and an inlet of the pre-combustion chamber communicates with the third connector, and the pre-combustion chamber is used for receiving high-temperature fuel gas and powder fuel, conducting pre-combustion and removing a surface oxidation layer of the powder fuel. Compared with the prior art, the solid powder pre-combustion assembly and the heater assembly are simple in structure, and the solid powder combustion completeness and combustion efficiency can be improved.

The application discloses a combustion chamber thrust control method in the acceleration process of a scramjet engine, and adopts a gradually expanding series connection cavity combustion chamber, fuel injection holes are arranged on the upstream of two cavities, the first cavity is a pre-combustion zone, and the second cavity is a main combustion zone; the thrust control method is as follows: when the combustion chamber inlet Mach number is low, the combustion mode of the pre-combustion zone is controlled to be kept in a weak subsonic combustion mode, and the combustion mode of the main combustion zone is controlled to be kept in a strong subsonic combustion mode; when the combustion chamber inlet Mach number is high, the combustion mode of the pre-combustion zone is controlled to be kept in a supersonic combustion mode, and the combustion mode of the main combustion zone is controlled to be kept in a strong subsonic combustion mode. The application is applied to the field of scramjet engines, can guarantee that the combustion chamber can generate higher thrust in a wide flight Mach number range, simultaneously avoids thermal choking, and the thrust mutation caused by combustion mode conversion is smaller, thrust control is more stable, and smooth transition of thrust during combustion mode conversion can be realized.

This invention relates to an inlet structure and determination method for a subsonic and scramjet engine. The outer compression section of the inlet consists of an outer compression section horizontal plane, a first panel, a second panel, and a third panel. An angled lip is provided at the inlet of the lower end face of the inner compression section. When determining the inlet structure parameters, the design of the subsonic and scramjet engine inlet is combined. Only the inlet Mach number needs to be determined to automatically identify and select the appropriate ramjet engine inlet geometry determination process, ultimately outputting the optimal inlet geometry that meets the design requirements. This invention enables rapid two-dimensional inlet structure design for subsonic and scramjet engines and provides optimized shape and performance parameters, laying the foundation for subsequent overall propulsion performance analysis of the ramjet engine.

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.

An intentional fluid manipulation apparatus (IFMA) assembly that includes an upstream intentional momentum shedding apparatus (IMSA) configured to impart a first induced velocity to a local free stream flow during a nominal operation requirement. The upstream IMSA creates a streamtube. The IFMA includes a downstream IMSA, with some or all of the downstream IMSA being located in a downstream portion of the streamtube. The downstream IMSA imparts a second induced velocity to the local free stream flow within the streamtube. The second induced velocity at the location of the downstream IMSA has a component in a direction opposite to the direction of the first induced velocity at the location of the downstream IMSA.

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 central water injection type water ramjet engine which comprises a front end socket, a combustion chamber, a rear end socket, an afterburning chamber and a spray pipe which are connected in sequence. The front explosive blocking ring is arranged at the position of the front sealing head, the rear explosive blocking ring is assembled on a boss of the rear sealing head, the combustion chamber is filled with the propellant grain, the jetting assembly is arranged on the central axis in the combustion chamber, and the ignition assembly is arranged at the position of an ignition connector of the rear sealing head. The engine is integrally assembled in the underwater navigation body to serve as a propelling system. The central injection assembly is adopted, the structure is simple, installation is easy, and the manufacturing cost is reduced; water is directly injected into the center of the combustion chamber through central water injection, so that the water and fuel can be rapidly mixed, the reaction is more sufficient, and the combustion efficiency can be greatly improved.

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 invention discloses an end face combustion solid-liquid fuel gas generator which comprises a combustion chamber and a solid grain arranged in the combustion chamber, an exhaust port communicated with a combustion cavity is formed in one end of the combustion chamber, and the end, facing the exhaust port, of the solid grain is a combustion end. A space formed by the end face of the combustion end and the cavity wall of the end, provided with the exhaust port, of the combustion cavity is a combustion space, the spring is arranged in the combustion cavity, and the cavity wall, located between the end of the non-combustion end of the solid grain and the cavity wall of the combustion cavity, of the combustion cavity is provided with an annular groove capable of abutting against the combustion end. The movement of the solid grain towards the exhaust port is limited; the annular groove enables the two ends of the spring to abut against the solid grain and the combustion cavity respectively and to be in a compressed state, so that elastic force for moving towards the exhaust port is continuously applied to the solid grain, and the combustion space is kept unchanged in the combustion process of the solid grain.

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 invention belongs to the field of ramjet engines, and particularly relates to a combustion oscillation restraining structure and method and a ramjet engine, and the combustion oscillation restraining structure comprises a combustion oscillation forward sensing assembly, a pressure potential energy flow guiding device and a controller; the combustion oscillation forward sensing assembly comprises two pressure sensors arranged between an inlet of the combustion chamber and the injector; the pressure potential energy flow guide device comprises a flow guide pipe and a valve arranged in the flow guide pipe, an inlet of the flow guide pipe communicates with the inner wall of the combustion chamber and is located between the two pressure sensors, and an outlet of the flow guide pipe communicates with the spray pipe; the controller is used for controlling the opening and closing or the opening degree of the valve according to the combustion oscillation condition obtained by the combustion oscillation forward sensing assembly. According to the combustion oscillation suppression structure, combustion oscillation suppression is conducted on the basis of the internal mechanism of combustion oscillation of the ramjet, the pressure potential energy flow guiding device fully utilizes the pressure difference existing in the reaction flow field to adjust the flow field in a self-adaptive mode, and the thermodynamic cycle efficiency cannot be reduced.