25results about How to "Increase thrust-to-weight ratio" patented technology

The invention provides a multi-cavity scramjet engine combustion chamber and a control method thereof. The combustion chamber is simple in structure, capable of adapting to different equivalence ratio ranges and different flight heights, low in resistance, high in combustion performance, low in thermal load of the wall face of the combustion chamber, convenient to maintain and capable of facilitating aircraft-engine integrated design. The combustion chamber comprises an isolation section, a first concave cavity group, a second concave cavity group, a third concave cavity group, an injection module and a combustion chamber expansion section, each concave cavity group is provided with two concave cavities which are oppositely arranged up and down; the isolation section is a first square tube and is connected with the first concave cavity group; the first concave cavity group is connected with the second group of concave cavities through a second square tube; the second group of concave cavities is connected with a third concave cavity group through a third square pipe; a combustion chamber expansion section is connected behind the third concave cavity group; the injection unit comprises a front injection device and a rear injection device; front injection devices are oppositely arranged on the upper wall surface and the lower wall surface of the second square pipe; rear injection devices are oppositely arranged on the upper and lower wall surfaces of the third square pipe.

The invention discloses a methane pre-cooling fuel gas generator circulation turbine enhanced combined power system which comprises a fuel turbine, a heat exchanger, a gas generator, a gas generator, a gas generator, a gas generator, a gas generator, a gas generator and a gas generator, and is characterized in that the heat exchanger is used for providing heat for fuel from a fuel pump and conveying high-temperature fuel gas to the fuel turbine so as to drive the fuel turbine to rotate; the oxidant turbine is used for driving the oxidant turbine to rotate by taking the high-temperature fuel gas output by the heat exchanger and the high-temperature fuel gas flowing out of the fuel turbine as power, so that the oxidant turbine drives the oxidant pump to work; high-temperature fuel gas obtained after the fuel from the fuel pump is heated by the heat exchanger is also used for being converged with high-temperature fuel gas sequentially passing through the fuel turbine and the oxidant turbine and then entering the fuel gas generator; liquid methane is adopted for precooling incoming flow, the total temperature of incoming flow behind an air inlet channel is effectively reduced, the upper limit of the working Mach number of the turbine core machine can be expanded to be larger than Ma3.5, a stamping mode matched with the turbine core machine only needs to consider the working speed range larger than Ma3.5, and the working performance of the turbine core machine can be well released.

The invention relates to a low-fault-energy micro-twin synergistic deformation cobalt-based alloy and a preparation method thereof, and belongs to the field of cobalt-based high-temperature alloys. According to the alloy, on the basis of Co-Al-W, Ti and Ta which are smaller than or equal to 4 wt% are added according to calculation of a first principle, and the stacking fault energy is controlled to be-95 mJ / m < 2 > to-89 mJ / m < 2 > (0K). Aiming at the defect of insufficient high-temperature plasticity of the gamma '-phase strengthened cobalt-based alloy, an electronic structure is accurately regulated and controlled through a first principle, intrinsic stacking fault energy is designed to reach an optimal negative value interval, a micro-twin mechanism is activated, dislocation slippage and micro-twin coordination is realized, and the problem of sudden plasticity drop caused by single deformation mechanism (dislocation slippage leading) and grain boundary weakening of a traditional alloy is solved; the high-temperature strength and the plasticity are synchronously improved.

The invention discloses a combustion chamber with a centrifugal ignition nozzle, which comprises a casing shell, one end of the casing shell is provided with an air inlet convenient for introducing high-pressure air, the other end of the casing shell is connected with the tail end of a high-pressure turbine guider, and a flame tube is connected in the casing shell. The head end of the high-pressure turbine guider is arranged in the cartridge receiver shell and connected with the flame tube, an ignition nozzle extending into the flame tube is connected to the cartridge receiver shell and comprises a nozzle mounting seat fixedly connected to the outer side of the cartridge receiver shell, a nozzle rod is fixed to the nozzle mounting seat, one end of the nozzle rod is arranged outside the cartridge receiver shell, and the other end of the nozzle rod is connected with the flame tube. The other end of the nozzle rod extends into the cartridge receiver shell and is connected with the flame tube, the other end of the nozzle rod is connected with a nozzle head, fuel oil is sprayed out from an oil spraying hole of the nozzle head, and the end, away from the nozzle rod, of the nozzle head enters the flame tube; the atomization effect is good in the engine starting state, and the ignition performance of the engine in the starting state is improved.

The invention discloses a ceramic-based turbine rotor of an aero-engine and a preparation process of the ceramic-based turbine rotor, and relates to the technical fields of engine structure design, material application, forming processes and the like. The fiber prefabricated body is formed by stacking common 2D plain fiber cloth from bottom to top, the included angle between fibers of an upper laying layer and fibers of a lower laying layer is fixed to be 30-45 degrees, after the laying layers reach the thickness, the fibers of all the layers are sewn and connected through a puncture method, forming of the turbine rotor prefabricated body is completed, and the fiber prefabricated body is of an integral leaf disc structure; compared with the mode that the independent ceramic-based turbine blades and the turbine disc are connected through tenon grooves, the number of parts is reduced, the part structure is simplified, and the reliability of an engine is improved.

The invention provides a triple-link structure for fixing supersonic monorail rocket sled engines. The triple-link structure mainly solves the problem that a large swing angle occurs when multiple rocket engines are sled. The triple ring structure for fixing the supersonic monorail rocket sled engine comprises a clamping ring, a supporting plate and a bolt. Each clamping ring is of a two-semi-ring structure, the three clamping rings are arranged on the supporting plate in a triangular mode, one semi-ring of each clamping ring is fixed to the supporting plate, and the two semi-rings of each clamping ring are fixed through a bolt; according to the skid body, after the engines are installed through the three-connecting-ring structure, the mass centers of all the engines are relatively concentrated, the rotational inertia of the whole thrust skid is small, the skid body is easier to automatically correct after rolling in the running process, and the stability of the thrust skid in on-orbit running is improved.

The multi-parallel rocket type rotary detonation engine comprises an oxidant cavity and a combustion chamber fixedly connected with the oxidant cavity, and the combustion chamber is internally provided with a plurality of fuel supply systems which communicate with one another, are fixedly connected with the oxidant cavity and are arranged around a rotation center in an array mode. And the ignition device is arranged on one fuel supply system. The multiple rocket-type rotary detonation engines are connected in parallel by sharing the oxidant cavity system, and the combustion chambers and the fuel supply system are coaxially arranged to form combustion chamber channels of the multiple rocket-type rotary detonation engines. The multiple parallel rocket type rotary detonation engines share the ignition device, ignition of the multiple parallel engines is achieved in the mode that windows are formed in the wall face, the system complexity brought by the ignition device can be obviously reduced, the reliability of the parallel rocket type rotary detonation engines can be improved, the structural design of the engines can be greatly simplified, and the cost is reduced. And the influence of the size effect is weakened on the premise of ensuring that larger thrust can be provided.

The invention provides a high-strength wing sled integrated fusion structure. The problems that a single-rail rocket sled is large in vibration overload, and a double-rail rocket sled is large in aerodynamic resistance, small in thrust-weight ratio and high in test cost are mainly solved. According to the scheme, the high-strength wing-sledge integrated fusion structure is a product sledge, thrust sledge and side wing integrated fusion structure and can be applied to end-point effect wing-type rocket sledge tests of various projectile war guiding systems, more stable propulsive force is provided for the whole system, the overall rigidity and strength of a sledge body are improved, the stability of the system is effectively improved, and the service life of the system is prolonged. The quality and complexity of the system are reduced, and the thrust-weight ratio of the system is improved.

The combustion chamber comprises a casing outer ring and a casing inner ring, a radial diffuser with an air inlet is fixed to the head of the casing outer ring, and the head end of the casing inner ring is connected with the inner side of the radial diffuser in the axial direction. An annular mounting cavity is formed among the inner wall of the casing outer ring, the inner side of the radial diffuser in the axial direction and the outer wall of the casing outer ring, and a flame tube is connected into the annular mounting cavity and comprises a first annular flame tube body and a second annular flame tube body. The first flame tube body and the second flame tube body are arranged at an interval in the axial direction and are both made of a ceramic matrix composite material, a main combustion hole and a mixing hole are formed in the first flame tube body, and the flame tube is in floating connection with the casing outer ring and the casing inner ring; the flame tube made of the ceramic matrix composite material is in floating connection with the casing inner ring and the casing outer ring, the flame tube has a free space for thermal expansion, and the problem that thermal expansion of the composite material and thermal expansion of a metal material are not matched is solved.

The invention relates to the field of engines and aircrafts, in particular to a curved tube engine, a curved tube engine aircraft and a control method thereof.The curved tube engine comprises a curved tube and a high-speed fluid generator, the high-speed fluid generator sprays high-speed fluid into the curved tube, and the high-speed fluid acts on the convex face of the inner wall of the curved tube to generate a low-pressure area; the high-speed fluid acts with the concave surface of the inner wall of the curved tube to generate a high-pressure area, and the curved tube engine is subjected to a pushing force F pointing to the concave surface of the inner wall from the convex surface of the inner wall of the curved tube under the action of pressure difference, so that the curved tube engine aircraft is pushed to complete maneuvers such as vertical take-off and landing, hovering and level flight. The engine is large in power load, high in thrust-weight ratio, simple, durable, safe, efficient and convenient to use and maintain, the maximum takeoff weight can exceed that of a large fixed-wing passenger plane with the same power, the engine has the advantages of being large in load, high in speed, long in voyage, high in lift limit, low in noise, convenient to operate and the like, and therefore the problems that an existing engine and an existing aircraft are low in efficiency and not high in safety are solved.

The invention discloses a support plate stabilizer structure for enhancing fuel oil atomization and suitable for an afterburner, which is used for improving the fuel oil atomization effect and the combustion stability, and comprises two support plate outer edge frames, a support plate tail edge frame, a fuel oil injection rod and the like, and a support plate mixing channel is formed between the two support plate outer edge frames; and a plurality of cooling air cavities are formed in the main body part so as to prevent key parts from being ablated in a high-temperature environment. First fuel oil impacts the fan-shaped channels through the oil injection holes to form an oil film, is further atomized under the shearing action of airflow and is mixed with air, and premixed gas is formed and then enters the concave cavity to be combusted. According to the design, flames in the concave cavity can be kept stable under the extreme conditions of low temperature, low pressure and the like, so that the ignition and flameout boundary of the combustion chamber is widened, meanwhile, high-temperature fuel gas in the concave cavity can heat transverse jet flow of second fuel oil, and the evaporation and mixing efficiency of the second fuel oil is improved. By means of the integrated design, the number of components is reduced, the flow resistance in the non-stress-application state is reduced, and meanwhile the thrust-weight ratio is increased.

The present invention belongs to the field of aerodynamic design of aerodynamic ships, and relates to a ducted propeller for aerodynamic ships. The overall shape of the ducted propeller is a rectangular blade, the root of the blade is round, and the blade section from the root to the tip is composed of multiple ARAD airfoil shapes similar to the wing, and the thickness smoothly transitions from thick to thin, and a certain torsion angle is maintained between each section, so that the airfoil of the blade will not stall during operation, and a larger aerodynamic lift and smaller aerodynamic resistance are generated. By designing the continuous change of the airfoil thickness, the airfoil of the blade is continuously thinned from the root to the tip, and the transition is smooth, which ensures that the propeller has sufficient strength and rigidity during high-speed rotation, and can generate strong thrust, so as to ensure that the aerodynamic ship can sail at high speed on the shoal or water surface.

The present invention discloses a tilted combustion chamber scheme with an integrated design of a coupled compressor and a turbine. The mainstream gas inside the annular flame tube of the combustion chamber adopts a flow form that rotates at a certain circumferential tangential angle, which not only increases the residence time of the gas, but also the high-speed gas at the flame tube outlet can directly drive the first-stage turbine blades to do work. The residual vortex airflow at the outlet of the last-stage rotor of the compressor and the combustion chamber head can be directly matched and designed, thereby omitting the compressor outlet guide and the turbine inlet guide structure. Since the present invention omits the compressor outlet guide and the turbine inlet guide respectively and shortens the axial length of the combustion chamber at the same time, the axial length of the engine can be greatly shortened, achieving the purpose of reducing the weight of the entire engine, thereby improving the thrust-to-weight ratio of the engine; at the same time, the turbine cooling air volume can be greatly reduced, reducing the difficulty of turbine design.

The invention relates to a local double-channel stamping rotating detonation engine which comprises an engine body and a cylinder arranged in the engine body. The cylinder is used for constructing local double channels in the engine main body and shunting fluid entering the engine main body, one part of the fluid flows through a flow channel formed between the inner wall of the engine main body and the outer wall of the cylinder, and the other part of the fluid flows through a flow channel defined by the cylinder; the engine main body comprises an isolation section, a combustion chamber, an equal straight section and an expansion section which are connected in sequence; part of the cylinder is arranged in the isolation section, a gap is formed between the cylinder and the isolation section, the rest part of the cylinder extends into the combustion chamber, and the axial length of the part, extending into the combustion chamber, of the cylinder is smaller than that of the combustion chamber. Through the cylinder arranged inside, the local double-channel design is achieved, the compact structure of detonation waves can be achieved, and the size of an engine is reduced, and the thrust-weight ratio is increased.

The present invention discloses an ejection seat and posture correction method using vector detonation propulsion, comprising a seat, a restraining device for restraining a person on the seat, and a locking device for activating the restraining device; a sliding track is arranged on the back of the seat, and the sliding track can slide along the sliding track; the integrated circuit control system is used to control the circuit system arranged on the seat; the fuel tank and the detonation combustion chamber are arranged at the bottom of the seat; the fuel tank and the rocket pack are used to provide a power source. The integrated circuit control system takes ejection modes P1, P2, P3, and P4 as constraint conditions, and different modes set different execution steps. The execution steps are written into the integrated circuit control system through a computer compiler program, and can respond quickly when the ejection is triggered, thereby improving the control effect. At the same time, detonation combustion is used as the power, and the detonation combustion process is equivalent to isochoric combustion, which has higher thermal efficiency and a larger thrust-to-weight ratio. In addition, detonation also has self-pressurization capability, and turbocharging equipment is no longer required in the power propulsion system, which greatly simplifies the system structure and can be well applied to aviation ejection seats.

The present application relates to the technical field of aerospace engine, in particular to a needle injection injector suitable for rich burn generator, an engine and a spacecraft, comprising: a spray valve shell, a first valve core assembly, a second valve core assembly, a blocking assembly and a fluid driving assembly, the fluid driving assembly is used for driving the second valve core assembly to move in the second valve cavity by using fluid medium, so as to drive the blocking assembly to move to open the second communication port and the third communication port, and drive the first valve core assembly to move to open the first communication port. The present application omits the traditional independent valve, part of the pipeline and the joint, compacts the engine layout, reduces the mass of the engine, and improves the thrust-to-weight ratio of the engine.

The utility model relates to the technical field of military instruments, and discloses a novel bulletproof sliding window which comprises an I-shaped plate and further comprises a main frame arranged on one side of the I-shaped plate, a partition plate is arranged in an inner cavity of the main frame, an observation opening is formed in the surface of the partition plate, a guide rail is arranged on one side of the I-shaped plate, and the guide rail is connected with the I-shaped plate. A middle frame is arranged on one side of the partition plate, and glass is arranged in an inner cavity of the middle frame. By means of the sliding strips and the pulleys, the effect of being convenient to use can be achieved, a user can push the sliding window more lightly, a gear adjusting function is achieved, the sliding window composed of the middle frame and the glass can be limited through the guide rails and the sliding bases, the whole structure is made of aluminum alloy materials, and the sliding window is firm, durable, not prone to corrosion and high in practicability. The total mass of the vehicle door window can be greatly reduced, the mass burden of the vehicle is effectively relieved, the thrust-weight ratio of the whole vehicle is increased, and the control performance of the vehicle is improved.

The afterburner hydrogen-kerosene dual-fuel concave cavity supporting plate flame stabilizer comprises a plurality of supporting plates evenly arranged in an outer culvert casing, concave cavities are formed in the sides, away from gas incoming flow, of the supporting plates, the concave cavities are closed through hydrogen injection pieces, the hydrogen injection pieces are provided with splitter plates stretching into the concave cavities, and the splitter plates are connected with the outer culvert casing. The supporting plate is provided with an air entraining inlet penetrating through the air entraining flow channel and an exhaust outlet penetrating through the exhaust flow channels, the supporting plate is provided with a main combustion stage kerosene injection pipe and an on-duty stage kerosene injection pipe which extend into the air entraining flow channel, and the side of the supporting plate is provided with a low-speed backflow area. The main combustion stage kerosene nozzle is located in a low-speed backflow area. Combustion of double fuels of hydrogen and kerosene in the afterburner is achieved, on one hand, the working efficiency of the gas compressor is improved, and combustion is promoted; on the other hand, kerosene serves as an afterburner on-duty stage, hydrogen serves as a working stage, hydrogen ignition is facilitated, and the afterburner can be stably started.

Provided is a turbine-based detonation afterburner engine based on an aerodynamic center body, which is a low bypass ratio turbofan engine, comprising a compressor (1), an air intake duct (2), a main combustion chamber (3), a turbine (4), a rotating detonation afterburner combustion chamber (5), a secondary fuel and oxidant supply device (6), a duty detonator (7) and an adjustable nozzle (8). The present invention eliminates the physical inner barrel, which can avoid the problem of thermal ablation and also greatly reduce the weight of the engine structure. The adjustable aerodynamic center body design timely adjusts the width of the "aerodynamic annular combustion chamber" according to different engine working conditions; the secondary supply of fuel and oxidant can improve the detonation performance of the rotating detonation wave and can be used as a secondary afterburner. The present invention can make full use of the advantages of detonation engines and turbine engines, and is of great significance for reducing fuel consumption, increasing range, achieving higher Mach number flight, and avoiding the "thrust trap" caused by mode conversion.

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.

The present invention proposes an air-breathing pulse detonation combustion chamber and an ignition and detonation method thereof, which belongs to the technical field of detonation engines. The characteristics of the present invention are that the combustion chamber is cylindrical as a whole and is divided into two parts, inner and outer. The main pre-detonation tube and the auxiliary pre-detonation tube are inside the combustion chamber and are connected by a valved flame-crossing tube at the head. The annular main detonation tube is outside and is connected to the tail of the main pre-detonation tube and the auxiliary pre-detonation tube respectively through a valveless flame-crossing tube. The inside of the annular main detonation tube is divided into a plurality of detonation combustion areas by an insulating partition, and the high-temperature combustion gas ignites the next detonation combustion area through the flame transmission notch on the partition. During operation, the ignition device located at the head of the main pre-detonation tube ignites and detonates, and the valveless flame-crossing tube passing through the tail ignites the annular main detonation tube, and the annular main detonation tube ignites the auxiliary pre-detonation tube through the valveless flame-crossing tube to form a cycle. The present invention can realize stable and continuous detonation of air and aviation kerosene under relatively small ignition energy, solve the ignition and detonation problem of the self-breathing pulse detonation combustion chamber, and shorten the length of the air-breathing pulse detonation combustion chamber.

The present invention proposes an axial flux counter-rotating motor, a coaxial propulsion device based on the motor, and an aircraft using the propulsion device. The present invention optimizes the structure of the coaxial propulsion device based on the motor while retaining the advantages of the axial flux dual-rotor motor, so that the coaxial propulsion device based on the axial flux counter-rotating motor has the advantages of simple structure, light weight, high heat dissipation efficiency, and easy maintenance compared with similar devices in the prior art. Specifically, for example, based on a motor with a YASA structure, the connection between the first rotor and the second rotor is cancelled, and the magnetic pole direction of the permanent magnet in the second rotor is changed, so that it rotates independently with the first rotor, and drives two sets of blades in opposite directions to form a new type of coaxial propulsion device.

The invention discloses a turbojet engine control system and method based on a distributed control architecture, a medium and a program product, the control device comprises a main control unit and a plurality of sub-control units, and the main control unit is responsible for controlling system power supply voltage, receiving and processing sub-unit data, executing a control algorithm and transmitting control signals to the sub-units; the subunits are responsible for signal processing, data uploading and closed-loop control of a local execution mechanism. The method is combined with a control algorithm to drive a fuel pump motor to control the fuel supply amount, and starting control, fuel control and steady state control are achieved. According to the invention, a PWM signal is adopted to control a high-power device, and the main unit communicates with the subunits through an I2C bus. According to the control architecture, the performance requirement of the main control unit is lowered, control function decoupling is achieved, the size and weight of a controller are reduced, the thrust-weight ratio of an engine is increased, modularization and standardization of a control system are promoted, and design, production, assembly and test cost is lowered.

The invention relates to the technical field of engine turbines, in particular to a multi-pipeline local air inlet turbine with differentiated stationary blades, which comprises two symmetrically arranged air inlet branches I, two symmetrically arranged air inlet branches II and a turbine assembly for adapting to a wide-range operation condition, the turbine stage number of the turbine assembly comprises but is not limited to three stages, and the turbine assembly comprises first-stage turbine stationary blades I arranged on the paths of the two first air inlet branches correspondingly and first-stage turbine stationary blades II arranged on the paths of the two second air inlet branches correspondingly. Stationary blade types matched with the airflow angle of the current working condition can be adopted under different working conditions, the airflow attack angle can be greatly reduced, and therefore the turbine efficiency is improved, the turbine can adapt to more operation working conditions, the application range of the air inlet turbine is widened, and the service life of the turbine is prolonged. The situation that an existing air inlet turbine needs to select the air inlet degree to be matched with the turbine flow, so that the turbine blade profile cannot adapt to the wide-range operation working condition is avoided, fuel consumption of the flight cycle of an aircraft is reduced, and economical efficiency is improved.

The fluid oscillator is applied to the control method of high-load compressor corner area flow separation, and the structure of the fluid oscillator comprises an air inlet, a porous material, an intermediate transition section, a micro controllable valve, an inlet and an outlet, the air inlet is adjacent to the pressure surface side of the compressor blade, the porous material is connected with the air inlet, one end of the intermediate transition section is connected with the air inlet, the other end of the intermediate transition section is connected with the micro controllable valve, the micro controllable valve is connected with the inlet, the inlet is connected with the outlet, and the outlet is located on the end wall near the compressor cascade corner area. Compared with the traditional technology, the application does not need an external air source, has the advantages of simple structure, small volume and weight, long-term work reliability and strong engineering practicability. Through the method, the flow separation near the compressor corner area can be effectively inhibited, the single-stage compressor can obtain a higher pressure ratio, and the turbomachinery has a larger thrust-to-weight ratio and higher efficiency.