28 results about "Ramjet" patented technology

FIELD: rocket science.SUBSTANCE: invention concerns the design of a ramjet engine. The ramjet engine body is a structure consisting of several layers, including: an internal heat-protection layer made of a composite material based on a thread frame of heat-resistant fibers impregnated with an elastic polymer binder, a power body based on a high-temperature polymer composite material and an external heat-protection layer made of a composite material based on an elastic polymer binder and a finely dispersed mineral filler. Thanks to the use of layers of external and internal thermal protection, it becomes possible to use a polymer composite material for the power casing.EFFECT: reducing the labor intensity of manufacturing and weight of the product and increased energy-mass perfection of the propulsion system.1 cl, 2 dwg

This invention discloses a method and system for gas-thermal coupling control of multiple fuel cell stacks in a rotary-wing unmanned aerial vehicle (UAV), relating to the fields of UAV power systems and fuel cell thermal management technology. The system captures the rotor downwash airflow as a source by setting an integrated annular ramjet inlet on each arm, and supplies it to distributed fuel cell stacks after adjustment by adaptive valves and ejectors. Exhaust gases from each stack converge into a central thermal energy aggregation chamber and can be drawn back by ejectors via branch lines. The method intelligently switches operating modes based on stack temperature: at low temperatures, a thermal circulation mode is activated, utilizing waste heat from the exhaust gas to preheat and humidify the intake air; at normal / high temperatures, it switches to a forced exhaust mode for rapid heat dissipation. This invention utilizes rotor airflow to achieve heat dissipation and intake with zero parasitic power consumption, solving the problems of high power consumption, difficult low-temperature start-up, and large temperature differences among multiple stacks in traditional systems, significantly improving the environmental adaptability and endurance of the UAV.

The application discloses a hyperplane climbing track group intelligent optimization method, which comprises the following steps: constructing a hyperplane climbing section dynamics equation, and obtaining to-be-optimized parameters in the hyperplane climbing section dynamics equation; designing a hierarchical cost function considering multiple constraints and turbojet, subsonic ramjet and supersonic ramjet three power mode switching Mach numbers; and optimizing the cost function through a chicken swarm algorithm, thereby completing the hyperplane climbing track group optimization. The hyperplane climbing track parameter optimization is performed through the improved chicken swarm algorithm, the suitable power mode switching Mach number is selected while the complex constraints of the hyperplane climbing section are met, the climbing scheme capable of significantly shortening the climbing time and total fuel consumption is obtained, and the problem that the track is difficult to optimize under the hyperplane cross-power mode and long-time window climbing condition is solved.

This invention belongs to the field of ramjet engine direct-drive testing technology. It provides a ramjet engine direct-drive testing method based on operating condition adaptation and control, comprising: during the ramjet engine direct-drive test, gradually adjusting test parameters, performing operating condition scanning, recording combustion chamber pressure fluctuation data under each operating condition, and determining oscillation-sensitive areas; performing multiple repeated tests on each operating condition within the oscillation-sensitive area to identify the probability of pressure oscillation occurring under each operating condition within each oscillation-sensitive area, and determining high-probability pressure oscillation operating conditions; extracting high-probability pressure oscillation operating conditions, and using Fast Fourier Transform to extract the dominant frequency of the direct-drive test oscillation; this invention can effectively avoid engine shutdown, structural damage, and other failures caused by oscillations in actual flight, ensuring engine flight safety and operational stability.

A Ramjet solid fuel comprising up to 50% metal by weight selected from the group consisting of aluminum, boron, boron carbide, magnesium, titanium, and titanium boride; between 2.0 and 9.0 weight percent fluorinated combustion aids selected from the group consisting of 3-hydroxybenzo trifluoride (3HBTF), hydroxyfluoro-2-propanol (HF2P), polytetrafluoroethylene, and trifluoroethanol (TFE); and up to 48% by weight of the remaining component being a hydroxyl terminated, epoxidized polybutadiene polymer (Poly BD 605E) (HO—CH2-CH═CH—CH2-CH2-CH—O—CH—CH2-CH2-CH—(CH═CH2)-CH2-CH═CH—CH2-CH2-CH—O—CH—CH2-OH).

The application provides a rotating detonation afterburner of an aero-engine, and relates to the technical field of gas turbine engines, and specifically comprises a rotating detonation afterburner connected to an exhaust end of a casing of the engine, an air inlet of the rotating detonation afterburner is connected to an exhaust port of the casing, and an airflow injection channel is arranged on the exhaust end of the casing and / or the air inlet of the rotating detonation afterburner, the airflow injection channel is connected to an internal space and an external space of the air inlet of the rotating detonation afterburner, and when exhaust gas of the exhaust end of the engine passes through the air inlet of the rotating detonation afterburner, the air in the external space of the connection part between the rotating detonation afterburner and the casing is introduced into the rotating detonation afterburner through the airflow injection channel. Through the processing scheme, the stability of the detonation wave is stabilized, and the stability of the rotating detonation afterburner is improved.

The application provides a series type double combustion chamber rotating detonation ramjet engine and a working method, which comprises a combustion chamber outer shell, an air inlet channel, a separation section, a first stage combustion chamber, a first stage contraction section, a second stage combustion chamber, a second stage contraction section, an outlet nozzle, a combustion chamber center body, a first stage fuel injection cavity, a second stage fuel injection cavity, a first stage fuel injection hole, a second stage fuel injection hole, a center body front support and a center body rear support. The application is aimed at the rotating detonation combustion organization flow characteristics under high or low Mach number working conditions, realizes the compatibility of high or low Mach number rotating detonation combustion working conditions in the same flow channel, and thus widens the working Mach number range of the rotating detonation ramjet engine. Considering the large difference in the combustion back pressure transmission distance under the high / low incoming flow Mach number working conditions, the two combustion chambers are arranged in series, and the combustion organization working mode is reasonably arranged, so that the combustion back pressure transmission problem under the wide incoming flow Mach number condition can be solved.

This invention discloses an active cooling structure and application method for optimizing kerosene heat sink utilization, relating to the design of scramjet engines. The structure includes variable cross-section passage zones I, II, and III, a flow restrictor, a local manifold, and a spacer. By providing different heat exchange efficiencies through the variable cross-section passage zones I, II, and III, the invention adjusts the kerosene heat sink utilization rate according to the actual thermal load on the cooling panel, improving the temperature uniformity of the cooling panel structure and effectively guiding the engine's thermal structure design.

A health monitoring system and adaptive adjustment method for a scramjet engine cooling channel, relating to the field of aero-engine management and condition monitoring technology, is disclosed. The system includes: a condition perception and monitoring module, which uses fiber optic measuring points deployed within the cooling channel to perceive multi-physical field information such as temperature, pressure, and vibration in real time, achieving in-situ, passive, and interference-resistant measurement; a data transmission and conversion module, which converts optical signals into digital signals; an intelligent analysis and decision-making module, which analyzes the cooling channel condition based on real-time data and historical models, identifies abnormal conditions such as overheating, coking, thermal stress concentration, and flow instability, and generates corresponding pump and valve adjustment commands; and a strategy execution and adjustment module, which dynamically adjusts the cooling medium flow rate by regulating pump speed or valve opening, achieving closed-loop adaptive adjustment of the cooling channel condition and improving engine thermal protection efficiency and operational reliability. This invention is applicable to the integrated and intelligent development needs of future hypersonic aircraft.

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

This invention discloses a subsonic ramjet engine based on counter-rotating rotary detonation flame stabilization ignition, belonging to the field of subsonic ramjet engine flame stabilization ignition technology. It includes an intake cooling channel, an integrated fuel supply device, and a symmetrical rotary detonation flame stabilization device. The intake cooling channel cools the engine wall temperature; the integrated fuel supply device achieves high system integration, low flow resistance loss, and lightweight and convenient design; the symmetrical rotary detonation flame stabilization device uses a counter-rotating jet to assist in creating a stable low-speed recirculation zone in the main combustion chamber, thereby achieving flame stabilization. Furthermore, the continuously operating rotary detonation generator acts as a pre-combustion tube, providing forced ignition to the main combustion chamber. This highly integrated invention effectively cools the combustion chamber wall temperature of the subsonic ramjet engine, improves combustion stability, and significantly increases the ignition success rate of the main combustion chamber, thus providing an effective technical approach for subsonic ramjet engines to operate within a wide range of conditions.

This invention discloses an experimental apparatus and method for studying the ignition and combustion mechanism of condensed-phase particles under supersonic airflow, relating to the field of solid scramjet engine technology. The apparatus includes an air heater, a nozzle, an isolation section, and a combustion chamber connected in sequence. The airflow, after being rectified by the isolation section, enters the combustion chamber and reacts with gaseous fuel injected through the gaseous fuel inlet, providing a high-temperature, high-pressure combustion environment for the condensed-phase particles. A high-speed camera records the ignition and combustion process of the condensed-phase particles in detail through a high-speed camera lens.

The application discloses a composite device applied to a combustion chamber and a ramjet engine, and relates to the technical field of ramjet engines, to solve the problem of how to improve the mixing efficiency of fuel and air. The composite device applied to the combustion chamber comprises a support plate injector, a cover plate provided with a first through hole, and a pipeline, and the cover plate covers one open end of the pipeline. Two sides of one support plate injector are respectively provided with one evaporation type flame stabilizer, the evaporation type flame stabilizer is in a blocking state in the composite device applied to the combustion chamber. The evaporation type flame stabilizer comprises a first pipeline in communication with the pipeline, a second pipeline connected with the side wall of the pipeline, and the first pipeline is located in the second pipeline. The first pipeline is provided with a second through hole for transmitting a first fluid in the first pipeline to the second pipeline, the second pipeline comprises a first connecting plate provided with an air bleeding port for transmitting a second fluid to the second pipeline, and a second connecting plate provided with a third through hole.

The present disclosure relates to the technical field of engines, and particularly provides a flow passage variable-width speed range ramjet engine, a combined power device and an aircraft. The engine body is provided with a first flow passage and an adjusting assembly. The bottom wall of the first flow passage is fixed, and the upper wall is adjustable. The upper wall has two states. In the first state, the first flow passage is a double-mode scramjet engine flow passage. In the second state, the first flow passage is a common subsonic ramjet engine flow passage. The upper wall comprises a variable-length sleeve plate, a front adjusting plate, a middle adjusting plate and a rear adjusting plate which are sequentially hinged. One end of the variable-length sleeve plate away from the front adjusting plate is hinged to the engine body. One end of the rear adjusting plate away from the middle adjusting plate is hinged to a connecting piece. The connecting piece is slidingly installed on a first sliding rail. The first sliding rail is fixedly arranged on the engine body. The present disclosure greatly widens the lower limit of the effective working speed range of the ramjet engine and reduces the technical difficulty of combination with a turbine engine.

The application relates to the field of aerospace technology, in particular to a high-Mach number ramjet engine with an afterburner. The high-Mach number ramjet engine with the afterburner comprises an aircraft body, an engine lower wall surface and an air inlet; the aircraft body is internally provided with a fuel tank; an isolation section, a first combustion chamber and a nozzle section are arranged between the aircraft body and the engine lower wall surface; the side, away from the first combustion chamber, of the isolation section is provided with the air inlet; the high-Mach number ramjet engine with the afterburner further comprises a powder fuel injection assembly and a second combustion chamber; the powder fuel injection assembly is arranged on the aircraft body, and the second combustion chamber is formed between the first combustion chamber and the nozzle section; the powder fuel injection assembly can inject metal fuel into the second combustion chamber; the metal fuel can be subjected to secondary combustion with combustion products in the second combustion chamber; and the energy generated by the secondary combustion is used to improve the thrust of the aircraft.

This invention discloses a drag reduction structure for a ramjet engine and its active-passive coupling drag reduction method, comprising: a transverse groove structure arranged in the high-friction region of the inner wall of the ramjet engine isolation section, and a stepped horizontal injector arranged at the front end of the combustion chamber inlet; the transverse groove structure is used to induce a stable turbulent boundary layer under incoming flow conditions; the injection direction of the stepped horizontal injector is parallel to the mainstream direction, used to horizontally inject premixed fuel into the turbulent boundary layer; this invention combines passive drag reduction via grooves with active drag reduction via boundary layer combustion, and significantly improves fuel mixing efficiency and combustion stability by utilizing the stable turbulent boundary layer induced by the grooves, solving the problems of flame instability and easy flameout in simple boundary layer combustion drag reduction, while achieving efficient and controllable flow channel drag reduction effect.

This invention relates to a compact dual-combustion chamber and engine of a turbo-ramjet combined engine. The dual combustion chamber includes an outer bypass casing with multiple sets of casing fuel supply channels running through its circular outer wall at equal angles. The area between the intermediate casing and the outer bypass casing forms an annular outer bypass duct. The outer wall of the intermediate casing has an outer concave cavity, and the ramjet combustion chamber is formed between the outer concave cavity and the outer bypass casing. The turbine combustion chamber is located inside the intermediate casing and has a flow divider ring assembly. The area between its outer shell and the inner wall of the intermediate casing forms an annular inner duct. Its outer shell has an inner concave cavity, and the turbine combustion chamber and the ramjet combustion chamber share an integrated fuel supply device, which is connected to the casing fuel supply channels. This invention, with its combined combustion chambers sharing an integrated fuel supply device and a portion of the casing being shared, can effectively improve the engine's compactness, reduce overall weight, and increase thrust-to-weight ratio.

This invention belongs to the field of high-temperature wind tunnel testing technology, and discloses a thermal deformation isolation force transmission device and an aerodynamic force measurement method for high-temperature wind tunnels. The thermal deformation isolation force transmission device includes a high-temperature end adapter plate and a low-temperature end adapter plate that are parallel to each other vertically. A force transmission unit is set between the high-temperature end adapter plate and the low-temperature end adapter plate. The end force transmission unit fixes several longitudinal force transmission plates located on the outer side and several transverse force transmission plates located on the inner side through an end force transmission base. The internal force transmission unit fixes several transverse force transmission plates symmetrically arranged front and rear through an internal force transmission base. The aerodynamic force measurement method includes designing the thermal deformation isolation force transmission device; establishing a six-component aerodynamic force measurement test system for high-temperature wind tunnels; and conducting a six-component aerodynamic force measurement test in a high-temperature wind tunnel. The thermal deformation isolation force transmission device and the aerodynamic force measurement method effectively suppress the interference of thermal deformation of the ramjet engine test model on the wind tunnel balance, improving the accuracy of aerodynamic force measurement.

The application discloses a test structure of a modular two-phase ramjet propulsion device, which comprises an inner flow channel assembly, an air inlet assembly, a front assembly and a tail nozzle assembly; the air inlet assembly comprises a central passage, the inner flow channel assembly is coaxially embedded in the central passage in a detachable manner, one end of the front assembly is connected with the front end surface of the inner flow channel assembly and the air inlet assembly, and one end of the tail nozzle assembly is detachably connected with the rear end surface of the air inlet assembly, and the tail nozzle assembly comprises an inner flow channel coaxial with the central passage. The test structure of the modular two-phase ramjet propulsion device can effectively reduce the machining difficulty and improve the machining quality by decomposing the structure of the two-phase ramjet propulsion device, can realize the independent replacement of components, can reduce the machining quantity of the sample in the multi-parameter comparison and verification test process, can reduce the test cost, and can improve the research and development efficiency.

The application provides a real-time active oxygen supplementing regulation method and device for combustion of a scramjet engine, comprising the following steps: acquiring wall surface pressure data and flame self-luminous radiation data in a combustion chamber in real time; determining a current flame combustion state based on a flame combustion state discrimination criterion and the current acquired wall surface pressure data and flame self-luminous radiation data in the combustion chamber; and generating a corresponding control signal for the combustion chamber according to the current flame combustion state to realize active oxygen supplementing and combustion active regulation of the oxygen injection valve of the combustion chamber. The application can strengthen the combustion effect, improve the combustion efficiency, and improve the precision and stability of combustion control.

The present application relates to solid fuel-rich propellant for hypersonic ramjet engine, and particularly discloses a boron-containing fuel-rich propellant and a preparation method thereof, which comprises, by mass percentage, 15-24% of a binder system, 28-32% of an oxidizer, 40-50% of composite boron powder, 5-10% of auxiliary metal fuel, and 0.1-2% of a performance regulator; wherein the binder system comprises a plasticizer, and the plasticizer is ethylene glycol perfluorodecanoate. The propellant uses the composite boron powder to replace the ordinary boron powder, the process performance of the propellant is greatly improved, the boron powder content in the propellant is greatly increased, and the energy level of the propellant is effectively improved; meanwhile, the ethylene glycol perfluorodecanoate (FOE) is used to replace the ordinary plasticizer DOS, the phase change energy loss caused by B2O3 generated after the propellant is burned is reduced, and the energy performance of the propellant is further improved.

This invention discloses a fuel injection method for accelerating the stabilization of liquid fuel rotational detonation initiation, relating to the design of ignition schemes for rotational detonation ramjet engines. In the first stage of fuel pipeline pressure build-up and the second stage of the development of rotational detonation after initiation, the fuel injection method is limited by short-term fuel flow rate regulation to obtain the optimal fuel-air mixture state suitable for rotational detonation initiation. The short-term fuel flow rate regulation is achieved through short-term fuel injection bypass or short-term fuel pressure regulation. This invention, without significantly altering the original supply system, utilizes short-term fuel regulation to control the fuel injection pressure during the initiation stage, achieving a better combustible gas mixing effect and distribution, improving initiation reliability, and shortening the evolution time for the development of a stable rotational detonation flow field after initiation.

A concentric annular plasma fuel activation device for the injector rod of a ramjet combustion chamber is provided, comprising a combustion chamber housing (1), an injector rod (2), a flame stabilizer (3), an outer ring low-pressure electrode (4), and an inner ring high-pressure electrode (5). The outer ring low-pressure electrode (4) consists of a circular outer ring body (6), an outer ring fixing rod (7), and an outer ring discharge contact (8). The inner ring high-pressure electrode (5) consists of an inner ring high-pressure terminal (9), an inner ring body (10), and an inner ring discharge contact (11). To achieve insulation between the inner ring high-pressure terminal (9) and the combustion chamber housing (1), a hollow cylindrical inner ring fixing sleeve (12) is designed. A method for activating concentric annular plasma fuel for the injector rod of a ramjet combustion chamber is also provided. This invention addresses the problems of difficult ignition and low combustion efficiency of ramjet combustion chambers under low operating conditions. This invention employs an innovative sliding arc plasma excitation device after the fuel injector rod in the ramjet combustion chamber, which can optimize the working performance of the ramjet combustion chamber over a wide range of incoming flow parameters without the need for an external geometrically adjustable mechanism.

This invention provides a direct-injection swirl-type mixing nozzle for a water-jet engine and a water-jet engine itself, relating to the field of water-jet engine technology, to solve one of the technical problems of low nozzle flow rate, poor atomization effect, and short penetration depth in water-jet engines. The direct-injection swirl-type mixing nozzle for a water-jet engine includes a core and a cavity enclosed by a shell, the shell having opposing inlet and outlet ends; along the direction from the inlet end to the outlet end, the cavity sequentially has a first cavity section and a second cavity section, the core being disposed in the first cavity section, and the core having at least one channel for water flow. This nozzle, used in water-jet engines, can improve atomization effect while maintaining a large flow rate.