380results about "Ram jet engines" patented technology

The invention discloses a controllable mass center stamping fuel oil system for a small-sized large-maneuvering aircraft, which comprises a pressurizable vacuum fuel oil tank (1), a pressurizing air entraining pipeline (2), an oil supply pipeline (3) and a fixing unit, the pressurizable vacuum fuel oil tank (1) comprises an oil bag (11) and a plurality of air bags (12), each air bag (12) is connected with the oil bag (11) but not communicated with the oil bag (11), and the oil bag (11) is communicated with the pressurizing air entraining pipeline (2). The pressurizing air entraining pipeline (2) is communicated with each air bag (12), the oil bag (11) is connected with the engine (4) through the oil supply pipeline (3), and the fixing unit is used for fixing the pressurizing vacuum fuel tank (1) and enabling the gravity center of the pressurizing vacuum fuel tank (1) not to deviate. The pressurizable vacuum fuel tank is made of flexible materials, can be made into different shapes according to the structure of a cabin and is composed of the air bag and the oil bag, and in the using process, the air bag pressurizes and directionally extrudes the oil bag to form oil supply pressure.

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 relates to the technical field of ramjet engines, in particular to a composite ramjet engine and a combined power device. According to the composite ramjet engine, a scramjet engine runner serves as a basic runner, and the runner can be switched into an equivalent subsonic combustion ramjet engine runner through profile adjustment. The basic flow channel comprises an air inlet channel, an isolation section, a combustion chamber front section, a combustion chamber rear section and an expanding exhaust nozzle. A flow channel molded surface adjusting mechanism is arranged on the inner wall of the expanding exhaust nozzle; the mechanism has two states: in the first state, the adjusting mechanism is attached to the inner wall of the expansion type exhaust nozzle, and the molded surface of a flow channel of the exhaust nozzle is in an expansion type and corresponds to a dual-mode scramjet mode; and in the second state, the adjusting mechanism protrudes towards the inner cavity of the exhaust nozzle, and the exhaust nozzle runner comprises a first expansion section, a first contraction section and a second expansion section in the airflow direction and corresponds to the equivalent subsonic combustion stamping mode. The working range of a conventional speed range scramjet engine can be widened to Ma2.5-7.0, and the conventional speed range scramjet engine can be directly combined with a common turbine engine.

The invention discloses a thermal photovoltaic Brayton cycle power generation system for an aircraft and a control method. The aircraft comprises a combustion chamber and a cooling channel, the system comprises a fuel assembly, the fuel assembly comprises a fuel storage tank and a first diverter valve in pipeline connection with the fuel storage tank, an outlet pipeline of the first diverter valve is connected with a cooling pipeline, and an outlet pipeline of the cooling pipeline is connected to the combustion chamber; the thermophotovoltaic power generation assembly is arranged outside the cooling channel; the cooling pipeline is arranged in the cooling channel, and fuel in the cooling pipeline is used for cooling the cooling channel and the photovoltaic power generation assembly; and the Brayton cycle power generation assembly comprises a cooler, and the cooling pipeline is connected with a first inlet pipeline of the cooler. On the basis of Brayton cycle power generation, a thermal photovoltaic power generation system is coupled, so that the generating capacity can be increased while cooling of a cooling channel is ensured, and the thermoelectric conversion efficiency of the system is improved.

The invention belongs to the field of scramjet engines, and particularly relates to an ammonia-hydrogen power wide-speed-range scramjet engine and a combustion method thereof. The ammonia-hydrogen power wide-speed-range scramjet engine comprises a combustion chamber, a liquid ammonia storage device and a hydrogen storage device; a concave cavity is formed in the combustion chamber, a fuel injection section upstream injection hole is formed in a fuel injection section of the concave cavity, the concave cavity is located at the downstream of the fuel injection section, a fuel injection section upstream micro-jet injection hole is formed in the upstream of the fuel injection section upstream injection hole, and a concave cavity inner micro-jet injection hole is formed in the concave cavity; the liquid ammonia storage device is connected with an upstream injection hole of the fuel injection section through a cooling structure; and the hydrogen storage device is connected with the upstream micro-jet injection hole of the fuel injection section and the micro-jet injection hole in the concave cavity. According to the invention, liquid ammonia and hydrogen are introduced as fuels to realize wide-range combustion organization, so that wide-range efficient combustion is realized, the fuel heat absorption requirement of high-Mach-number active cooling thermal protection is met, and zero carbon emission can be realized in the whole process.

The invention belongs to the technical field of design of turbo-ramjet combination engines, and particularly relates to a turbo-ramjet combination engine and a control method thereof. According to the design, a complex outer turbine channel is separated from a smooth ramjet channel, and an afterburner and a ramjet combustion chamber are designed to share the space; the designed center stamping channel has multiple functions of a turbine engine stator force bearing frame, a center deflation channel, a subsonic combustion stamping rectification channel, a scramjet channel and the like. Self-adaptive switching of four working modes including a turbine engine conventional working mode, a turbine engine center deflation working mode, a subsonic combustion stamping working mode and a scramjet working mode is achieved, the space of a center stamping channel is utilized to the maximum extent, the working speed range of the turbine modes is widened, the cross sectional area is greatly reduced, the flight resistance is effectively reduced, and the service life of the turbine engine is prolonged. And wide-speed-range work with the Mach number of 0-7 is realized.

The invention provides a powder fuel-based cross-medium ramjet engine which comprises a powder fuel cabin, a combustion chamber and a spray pipe are fixedly connected to the axial front end of the powder fuel cabin, a high-pressure gas tank is fixedly connected to the axial rear end of the powder fuel cabin, and a plurality of gas generators are fixedly arranged on the inner wall of the combustion chamber; a plurality of air inlet channels are arranged on the outer wall of the combustion chamber; the high-pressure gas tank is communicated with the powder fuel cabin through a high-pressure gas flow channel and a fluidization pipeline, and is communicated with the gas inlet channel through a gas inlet channel pressurizing pipeline; a liquid medium cavity is formed in the air inlet channel and communicates with the combustion chamber through a liquid medium main supply pipeline, an upstream liquid supply pipeline and a downstream liquid supply pipeline. Combustion energy release of the powder fuel in the air mode and the underwater mode of the cross-medium ramjet engine is effectively promoted, and the method is suitable for the powder fuel cross-medium ramjet engine.

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 method for improving the mixing efficiency of a ramjet engine through zero-mass jet flow. The method comprises the steps that a zero-mass jet flow exciter is arranged on the wall face of a combustion chamber, close to one side of supersonic airflow, of the ramjet engine; the zero-mass jet flow exciter comprises a cavity, a driver and a piston. Wherein the driver drives the piston to reciprocate in the cavity; the end of the cavity communicates with the interior of the combustion chamber, and the movement direction of the piston is perpendicular to the central axis of the combustion chamber. The movement frequency and the movement amplitude of the piston in the cavity are configured through the driver; wherein the motion frequency is the same as the unsteady frequency of the turbulence structure; unsteady dynamic oblique shock waves moving at the movement frequency are generated in the mixing layer through the change of the position of the piston and oscillate back and forth in the mixing layer, then the resonance phenomenon is generated with the turbulent flow structure, and therefore the fuel and air mixing efficiency is improved.

By manipulating the fluid flow in the proximity of an object such as a fuselage, a wing, or the hull of ship, the wave drag associated with this object can be substantially reduced. This can be accomplished by locally changing both the fluid flow velocity and the pressure of the fluid flow.

The invention discloses a multistage self-adaptive enhanced mixing device for a solid fuel ramjet engine, which belongs to the technical field of solid fuel ramjet engines and comprises a front flange, a rear flange, an afterburning chamber barrel, a mixing partition plate, a heat insulating layer, a pressure measuring seat and a temperature measuring seat. The front flange and the rear flange are welded with the afterburning chamber barrel, a water pressure test is carried out after welding is completed, the mixing partition plates are bonded to the inner wall of the afterburning chamber barrel through high-temperature-resistant glue, and the mixing partition plates are distributed in a multi-stage mode, are separated through a heat insulation layer and are bonded through the high-temperature-resistant glue. In order to verify the effect of the multi-stage self-adaptive enhanced mixing device for the solid fuel ramjet engine, a pressure measuring point and a temperature measuring point are arranged on an afterburning chamber barrel and a mixing partition plate; according to the multi-stage self-adaptive enhanced mixing device, the combustion efficiency of a solid fuel ramjet can be remarkably improved, and the efficiency is further improved along with the increase of the number of the mixing partition plates.

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 technical field of ramjet engines, in particular to a wide-speed-range scramjet combustion chamber, a scramjet engine and a combined power device. The wide-speed-range scramjet combustion chamber comprises a front combustion section and a rear combustion section; the front combustion section comprises a front-section runner wall surface, a front-section concave cavity flame stabilizer, a front-section wall surface fuel injector and a front-section multi-mode flame stabilizer; the rear combustion section comprises a rear section runner wall surface, a rear section concave cavity flame stabilizer, a rear section wall surface fuel injector and a rear section multi-mode flame stabilizer; the front-section wall surface fuel injector and the rear-section wall surface fuel injector are fixedly mounted on the wall surfaces in front of the front-section concave cavity flame stabilizer and the rear-section concave cavity flame stabilizer respectively, and the front-section multi-mode flame stabilizer and the rear-section multi-mode flame stabilizer are located in the middle of a front-section flow channel and the middle of a rear-section flow channel respectively and located at the same axial position as the front-section concave cavity flame stabilizer and the rear-section concave cavity flame stabilizer; the lower limit of the working speed domain of the scramjet combustion chamber in the conventional speed domain of Ma 4.0-7.0 can be widened to Ma 3.0-3.5.

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.

The invention belongs to the field of scramjet engine isolation section suction control, and relates to a scramjet engine isolation section boundary layer suction control system and method. The control system comprises a scramjet engine, a fuzzy control system and a boundary layer suction system. The fuzzy control system comprises a wall surface pressure sensor and an FPGA (Field Programmable Gate Array) development board for setting a shock wave string leading edge position detection algorithm and a fuzzy controller, and the boundary layer suction system comprises a suction box, a suction device and a driving circuit. Scramjet engine isolation section boundary layer suction self-adaptive control is achieved through a fuzzy controller, a wall surface pressure sensor and a shock wave string front edge position detection algorithm. A scramjet engine boundary layer active suction control system which can be practically applied is provided; the proposed control system solves the problems of high flow loss and fixed suction parameters faced by passive suction.

The present invention proposes a flow control method for improving the starting performance of the air inlet by using a pressure relief slit. First, the flow field of the initial configuration of the air inlet during the starting and non-starting process is solved and the starting performance of the air inlet is evaluated; secondly, according to the flow field characteristics of the initial configuration of the air inlet during the starting and non-starting process obtained by the solution, the arrangement scheme of the pressure relief slit is designed, including determining the separation point and reattachment point positions of the main separation zone on the compression surface in the non-starting flow field of the air inlet, and arranging slits behind the separation point of the main separation zone, in the middle of the separation zone, and in front of the reattachment point of the separation zone; finally, the arrangement scheme of the pressure relief slit is determined to improve the performance of the air inlet and widen the working Mach number range of the air inlet. The present invention can effectively improve the starting performance of the air inlet, widen the working range of the air inlet, and has extremely small performance loss and flow field changes under cruise conditions.

The invention belongs to the technical field of scramjet engine control, and particularly relates to a regenerative cooling ramjet engine temperature control system which comprises a novel cooling oil supply system with an oil drain valve as a core and an engine thrust and temperature decoupling control method. The working boundary of the engine is widened, and it is guaranteed that in the temperature control process, the temperature is controllable, and the thrust of the engine is kept constant.

The invention provides a solid ramjet gas flow self-adaptive control method based on variable rate integral gain scheduling, which comprises the following steps: S1, constructing a gas flow control inner trajectory model of a solid ramjet, and designing a gain scheduling main controller; s2, considering negative tuning characteristic suppression, and designing a tracking instruction pre-filter; s3, considering the uncertainty of the dynamic response rate of the solid ramjet, and designing a variable rate integral control strategy; and S4, combining the steps S1 to S3 to construct a solid ramjet gas flow comprehensive controller, and performing solid ramjet gas flow regulation control digital simulation with strong time varying and non-minimum phase characteristics. In order to improve the fuel gas flow control performance of the solid ramjet, on the basis of a fuel gas flow adjustment inner trajectory model, the adaptive capacity of a fuel gas flow controller of the solid ramjet to the parameter time-varying characteristic and the inhibiting capacity of the fuel gas flow controller to the negative adjustment characteristic brought by the non-minimum phase are improved, the control precision of the solid ramjet is guaranteed, and the fuel gas flow control efficiency is improved. And the robustness of engine control is improved.

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.

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 turbine-based combined ramjet engine is characterized in that the turbine-based combined ramjet engine comprises an inner culvert channel, and the inner culvert channel is provided with a core diffuser, a core engine, a low-pressure turbine, an afterburner and an injection spray pipe; the core diffuser, the core machine, the low-pressure turbine, the afterburner and the injection spray pipe are sequentially arranged in the airflow direction. The outer culvert channel surrounds the radial outer side of the inner culvert channel, the outer culvert channel is provided with a mode selection valve, an outer culvert diffuser, a ramjet combustion chamber and a single-side expansion spray pipe, and the mode selection valve, the outer culvert diffuser, the ramjet combustion chamber and the single-side expansion spray pipe are sequentially arranged in the airflow direction; the mode selection valve is used for controlling opening and closing of the outer culvert channel; and the supersonic through-flow fan is arranged on the upstream sides of the inner culvert channel and the outer culvert channel.

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.

The separable supporting plate structure comprises two unit supporting plate structures and two sets of hydraulic telescopic mechanisms, the two unit supporting plate structures are used for being arranged in the combustion chamber of the engine, the combustion chamber of the engine is provided with an inlet and an outlet, and the two sets of hydraulic telescopic mechanisms are arranged in the combustion chamber of the engine. The two unit supporting plate structures are arranged in a mirror image mode with the connecting line of the inlet and the outlet as the center line, each unit supporting plate structure is provided with a first face, a second face and a third face, the first face is parallel to the center line, and the second face is arranged on the side, close to the inlet, of the first face and forms an included angle alpha with the first face. The third face is arranged on the side, close to the outlet, of the first face and forms an included angle beta with the extension line of the first face. The two hydraulic telescopic mechanisms are connected with the side walls of the unit supporting plate structures in a one-to-one correspondence mode and used for controlling the two unit supporting plate structures to be separated and combined in the vertical direction. Flames in the combustion chamber can be controlled to be extinguished, and conversion to an oblique detonation mode is facilitated.

A passive bypass for an inlet to a supersonic or hypersonic air-breathing engine allows airflow in the inlet to exit through the cowling when the inlet supplies more airflow than the air-breathing engine demands. The air-breathing engine may be the only form of propulsion or a secondary form of propulsion to reach higher speeds. The passive bypass includes a plurality of lower channels in the cowling that are operatively coupled to the inlet diffuser at an inner surface of the cowling and swept forward towards the throat, a plenum in the cowling operatively coupled to the plurality of lower openings and a plurality of upper channels in the cowling that are operatively coupled to the plenum and swept back away from the throat to an outer surface of the cowling. A serpentine path through the plurality of lower openings, the plenum and the plurality of upper openings allows airflow in the inlet to exit through the cowling when the inlet supplies more airflow than the air-breathing engine demands.

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 belongs to the technical field of design of turbine-based ramjet combination engines, and particularly relates to a turbine-based ramjet combination engine and a design method thereof. A turbine engine and a subsonic combustion ramjet engine are designed to be connected in parallel to intake air and share a thrust augmentation / ramjet combustion chamber; main advantages of a series-connection type turbine-based ramjet combination engine and a parallel-connection type turbine-based ramjet combination engine are integrated, and compared with the series-connection type turbine-based ramjet combination engine, mode selection valves can be arranged at inlets of an air inlet channel low-speed channel and an air inlet channel high-speed channel when the series-connection type turbine-based ramjet combination engine works in a ramjet engine mode; the high Mach number working reliability is good, the efficiency is high, compared with a parallel type turbine-based ramjet combined engine, a set of ramjet combustion chamber and a combined spray pipe channel are reduced, the weight can be reduced by 12%-15%, the mission load and transonic speed performance of an aircraft can be improved, and the control difficulty is reduced.