75 results about "Specific impulse" patented technology

The invention provides a Hall and ion hybrid electric propulsion system. The Hall and ion hybrid electric propulsion system comprises a composite material gas cylinder, an inflation valve, a pressure adjusting module, a Hall flow adjusting module, an ion flow adjusting module, a Hall thruster, an ion thruster, an electric propulsion control unit, a power processing unit and a power supply processing and control unit, the pressure adjusting module supplies gas to the Hall flow adjusting module and the ion flow adjusting module in a time-sharing manner; the Hall flow regulation module and the ion flow regulation module are used for carrying out flow control on the Hall thruster and the ion thruster respectively; the electric push control unit performs time-sharing power supply control on the power processing unit and the power supply processing and control unit; the power processing unit and the power supply processing and control unit respectively provide electric energy for the Hall thruster and the ion thruster; the Hall thruster and the ion thruster provide thrust or impulse for the spacecraft; according to the Hall and ion hybrid electric propulsion system, the integrated design of air supply, power supply and autonomous process control of the Hall and ion thrusters is achieved, system air path and circuit configuration is simplified, and the requirement for a spacecraft with high specific impulse and high thrust electric propulsion systems at the same time is met.

The invention relates to the technical field of plasma physical simulation, in particular to an SOLPS simulation method suitable for a variable specific impulse magnetic plasma electric propulsion device. According to the technical scheme, the method comprises the following steps of parameter obtaining, wherein simulation parameters of the electric propulsion device are obtained; file generation: based on the simulation parameters, generating input files required by SOLPS simulation through an external program, including a device geometric file, a fluid grid file, a power source file and a particle source file; and SOLPS calculation: importing the input file into an SOLPS program, generating a calculation grid by the SOLPS program, and performing plasma simulation calculation. According to the method, the adaptability problem of application of the SOLPS program in the variable specific impulse magnetic plasma linear electric propulsion device is solved, end-to-end integration from parameter input to result visualization is achieved, energy deposition distribution can be accurately described, and the simulation accuracy and efficiency are remarkably improved.

The invention discloses an attitude control power system based on carbon dioxide phase change gas storage, which belongs to the technical field of attitude control, and comprises a thruster, a phase change energy tube, a storage shunting mechanism and a control assembly. The phase change energy tube comprises a shell and an excitation assembly. A sealing cavity is formed in the shell and filled with liquid carbon dioxide. And excitation powder of the excitation assembly is arranged in the sealed cavity. The shell is provided with a sealed air outlet capable of being opened and closed. The storage and distribution mechanism comprises a storage tank and a distribution channel, and the storage tank communicates with the phase change energy pipe. The storage tank is communicated with the flow dividing channel, and the flow dividing channel is provided with an air supply branch pipe which can be opened and closed and corresponds to each thruster. According to the attitude control power system, the attitude adjustment cost can be reduced as much as possible and the system operation convenience can be improved while the rapid response and high specific impulse of attitude adjustment are ensured, so that the accurate regulation and control of the attitude of the aircraft can be realized.

The application discloses an energetic material, a preparation method thereof and application of the energetic material as explosives and self-ignition propellants. I C 12 H 12 N 10 or M Ⅱ C8H6N8; wherein, M I is selected from Co, Ni and Cu; M Ⅱ is selected from Zn. The energetic material is convenient to synthesize, has good thermal stability, excellent mechanical sensitivity, excellent propelling performance and specific impulse, has a significant energy density advantage, can effectively solve the problem of continuous evaporation of hydrazine propellants during storage, avoid the risk of internal pressure increase of a fuel tank, and overcome the problems of liquid fuel in the storage and transportation process. Therefore, the innovative high-performance energetic material has excellent performance and wide application potential, and has great commercial value in the fields of aviation, aerospace and military.

The application discloses a kind of plasma propulsion devices, belong to plasma electric propulsion technical field, comprising: discharge chamber, spiral antenna I, center arc structure, spiral antenna II, center compensation structure, radio frequency power supply and electromagnetic coil;Center arc structure is located in discharge chamber, spiral antenna I spiral is wound in the outer wall of discharge chamber, with the position opposite of center arc structure;Center compensation structure is located in discharge chamber, spiral antenna II spiral is wound in the outer wall of discharge chamber, with the position opposite of center compensation structure;Spiral antenna I and spiral antenna II are electrically connected with radio frequency power supply respectively;Two electromagnetic coils are wound and set in the outer wall of discharge chamber, respectively located in the two ends of spiral antenna II;Center compensation structure can release neutral particles under the action of heat;The application can improve ion number density, realize high specific impulse output of propulsion device, improve propulsion efficiency.

The present application relates to a kind of adjustable thrust and specific impulse of multi-working substance mixed effect laser plasma micro-propeller, solid working substance supply system includes multiple solid working substance, wherein one solid working substance is located in the middle, the rest is uniformly arranged in the periphery of the solid working substance in the middle along the circumferential direction, each solid working substance is fixed on the installation base by fixed assembly B, the material of each solid working substance is two or more than two in metal, high molecular polymer, energy material;Laser system includes control power supply and multiple laser assembly, the number of laser assembly is same with the number of solid working substance, one-to-one correspondence, wherein one laser assembly is located in the middle, the rest is uniformly arranged in the periphery of the laser assembly in the middle along the circumferential direction.The present application can make the thrust and specific impulse generated by propeller can be adjusted in a large range by adjusting the independent action of laser and single solid working substance or the coupling effect of laser and multiple types of solid working substance, satisfy more complex space task.

The invention belongs to the technical field of hypersonic flight vehicle cruise flight range performance determination, and particularly relates to a hypersonic flight vehicle cruise flight range performance optimization determination method, which comprises the following steps: step 1, calculating an effective mass correction coefficient of an aircraft, where V is the speed of the aircraft, g is the gravitational acceleration, and r is the radius of the earth; 2, calculating the maximum lift-drag ratio of the aircraft; 3, calculating the fuel weight coefficient epsilon of the aircraft; and 4, calculating the optimal voyage of the aircraft, wherein Isp is the specific impulse of the aircraft.

The invention provides an attitude and orbit control two-component liquid rocket engine and a unified type establishing method thereof. The attitude and orbit control two-component liquid rocket engine comprises targeted design of an electromagnetic valve, an injector, a combustion chamber section and a spray pipe extension section. The electromagnetic valve is a pneumatic electromagnetic valve and works in a double-voltage driving mode, and an electrical interface is a double-point double-wire interface; the head cavity of the injector is small in volume and is provided with a goblet, a V-shaped groove and a heating sheet mounting groove structure; the characteristic length of the combustion chamber section is small; the area ratio of the combustion chamber section is iteratively designed according to an airflow separation theory; the spray pipe extending section is combined with the combustion chamber section in a welding or screwing mode, different area ratios of the spray pipe extending section are obtained, and engines with different thrust, specific impulse and sizes are designed. According to the method, the working characteristics and requirements of the attitude and orbit control engine are extracted, the performance which must be met is designed in a targeted mode, and the purpose of sharing components of the engine is achieved, so that the product complexity is reduced, the research and development, test, manufacturing and management cost is reduced, and the reliability and maintainability of the product are improved.

The invention discloses an ultra-low orbit satellite electric propeller two-stage cascade power overall planning control method, and relates to the technical field of satellite electric propellers, and the method comprises the steps: constructing a thrust-radio frequency power matching model; utilizing an extended Kalman filtering algorithm to carry out radio frequency source power dynamic estimation technology research, and constructing a state observer; based on a thrust-radio frequency power matching model and a state observer, constructing a self-adaptive predictive power control framework based on measured data and dynamic state estimation; determining a power coupling mechanism between the spiral wave radio frequency source and the ion cyclotron resonance heating radio frequency source, and constructing a front-stage and back-stage power linkage relation mathematical model; based on a front-stage and back-stage power linkage mathematical model, a dynamic power optimal distribution method with optimal efficiency, thrust and specific impulse under power constraint is established, and a feedback controller based on delay compensation linear quadratic Gaussian control is constructed. And the performance of the ultra-low orbit satellite electric propeller is optimized.

This invention belongs to the field of engines, specifically relating to a fuel supply structure and a solid-liquid hybrid rocket engine, including an oxygen-enriched propellant supply mechanism and a gel fuel supply mechanism. The oxygen-enriched propellant supply mechanism includes a coating column, which includes a coating sleeve with an opening facing the combustion chamber, and the coating sleeve is used to hold the oxygen-enriched propellant. The gel fuel supply mechanism includes a gel fuel supply device, an annular channel, and an atomizer arranged sequentially. The annular channel is arranged around the outer edge of the coating sleeve, with its downstream end protruding downstream of the coating sleeve. The atomizer is located at the end of the annular channel protruding downstream of the coating sleeve, and the annular channel and the coating sleeve together seal the upstream end of the combustion chamber. This invention's engine adopts a novel combination of gel fuel and oxygen-enriched propellant, combining some of the advantages of both to achieve the effects of increased specific impulse, adjustable thrust, and multiple starts.

The present application relates to the technical field of aerospace engine, in particular to a liquid fuel engine thrust chamber needle bolt head regenerative cooling structure, comprising: a thrust chamber and a needle injection injector, the injector shell of the needle injection injector is arranged in the thrust chamber from the top of the thrust chamber, and the injector shell is located in the thrust chamber, at least one week main injection hole is arranged on the radial of the injector shell, the annular gap is formed between the injector shell and the top shell of the thrust chamber, and the end of the injector shell is provided with an inner recess structure in the thrust chamber; The first cooling flow channel and the second cooling flow channel are arranged in the injector shell above and below the main injection hole, and the outlet of the second cooling flow channel is connected with the rotational flow tangential hole; wherein, the liquid propellant in the propellant combination is injected into the injector shell, and the annular gap between the injector shell and the top shell of the thrust chamber is used for injecting another propellant in the propellant combination. The present application improves the combustion efficiency of the thrust chamber, and improves the specific impulse and the thrust-to-weight ratio of the engine.

The invention provides a space propulsion method of a nanosecond-femtosecond composite light filament, which is used for solving the problem of low long-distance laser propulsion efficiency. The method specifically comprises the following steps: forming a kilometer-level length filament through a space filament forming principle of femtosecond laser, stabilizing the power density at 1013-1015W / cm < 2 > magnitude by utilizing a light intensity clamping effect, and realizing dozens of kilometer-level low-loss energy transmission; meanwhile, the transmission direction of the nanosecond laser is adjusted through the optical lens group, the formed nanosecond filament and the femtosecond laser are coaxially transmitted or transmitted at a certain angle and act on the same position, a plasma channel generated by the filament enables the transmission loss of the high-energy nanosecond laser to be greatly reduced, and the energy absorption of the target material to the nanosecond laser is remarkably improved; therefore, the surface of the target material is melted, gasified and even ionized in an extremely short time, a large number of high-speed reverse-spraying plasma plumes are generated, and high specific impulse and thrust are obtained. The method is suitable for various target materials, and precise control over the thrust, the specific impulse and the propelling direction can be achieved by regulating and controlling laser parameters.

The present invention discloses a flow control device and method for pulse injection mixing enhancement. The device includes an internal flow channel in the engine, an incoming flow control mechanism, and a fuel injection mechanism. The incoming flow control mechanism is located at the inlet of the internal flow channel in the engine and is used to fan the steady-state axial air inflow to form an unsteady, highly turbulent axial pulsating air inflow. The fuel injection mechanism is located on the inner wall of the internal flow channel in the engine and is used to sequentially inject fuel in intervals of pulses to form an unsteady, highly turbulent transverse pulsating fuel jet. The pulsation frequency of the axial pulsating air inflow and the transverse pulsating fuel jet are the same, and the pulsation phase difference is less than or equal to π / 2. The present invention is applied to the field of non-uniform fuel mixing and is used to achieve short-range, rapid, efficient and uniform mixing of reactants, improve the combustion efficiency, combustion intensity, and pressure gain of the combustion chamber, and thus realize an efficient supercharged combustion process with high-frequency self-sustaining engine and high specific impulse output.

The present application relates to the technical field of spacecraft liquid rocket engine thrust chamber design, and provides a wide working condition liquid rocket engine thrust chamber, which comprises a combustion chamber and a direct impingement injector, the combustion chamber comprises a straight section, a converging section, a throat and an expansion section connected in sequence, the direct impingement injector is arranged at the end of the straight section, the length-diameter ratio of the straight section is 1.2-1.5, the characteristic length is 290-350 mm, the direct impingement injector is provided with a central zone first injection hole, a central zone second injection hole and a peripheral zone injection hole, the oxidizer is sprayed out through the central zone first injection hole, and the fuel is sprayed out through the central zone second injection hole and the peripheral zone injection hole.The present application greatly widens the working condition range that the thrust chamber can adapt to, and solves the problems that the throat temperature of a high specific impulse performance space engine is too high and is greatly affected by working condition deviation.

The invention relates to an efficient miniature ion thruster based on microwave ionization enhancement, and relates to the field of ion thrusters, a permanent magnet group comprises a plurality of annular permanent magnets which are distributed at intervals along the axis direction of a shell, and cusped magnetic fields are formed between adjacent permanent magnets to reduce the loss of plasmas on the wall surface; the tungsten filament cathode is fixedly connected to the middle of one side of the shell and the axis coincides with the axis of the shell; the annular anode is fixedly connected in the shell on the periphery of the tungsten filament cathode, the axis of the anode coincides with the axis of the tungsten filament cathode, and the gas supply assembly is fixedly connected to the shell on one side of the anode so as to supply a gaseous propellant into the shell; the device has the advantages that microwave energy coupling, efficient magnetic confinement of a cusped magnetic field, ionization area expansion and high specific impulse acceleration of a grid system are combined, and therefore efficient ionization and high specific impulse acceleration of low-density working medium gas can be achieved under the low-power condition; the device has the advantages of compact structure, high efficiency and high specific impulse.

The application discloses a kind of far-field uniform explosion load simulation loading device and test method, including explosion chamber, suspension piece, rectifier tube, target structure, pressure sensor, laser vibration meter and display terminal;Explosion chamber one end is open, the other end is connected rectifier tube, explosion chamber inner cavity and rectifier tube inner cavity are through;Suspension piece is set in explosion chamber top, and suspension piece can move back and forth along rectifier tube axis direction;Rectifier tube is used to shape explosion shock wave;Target structure is connected at the end of rectifier tube, pressure sensor is used to detect the pressure load of target structure surface, laser vibration meter is aimed at target structure, for measuring the structure response of target structure;The loading device of the application has the function of adjustable pressure peak value and specific impulse, can be used to carry out the damage effect research of target structure under the combined action of different pressure peak value, specific impulse, can provide support for the damage research of typical target structure under large equivalent charge far-field explosion load.

The application discloses a performance analysis method of a mixed fuel multi-working condition oblique detonation engine, and the oblique detonation engine comprises an air inlet channel section, a fuel mixing section, a combustion chamber section and a tail nozzle section, and the method comprises the following steps: according to the sequential characteristics of the working process of the oblique detonation engine, sequentially performing parameter solving of an oblique shock wave compression module of the air inlet channel section, parameter solving of a mixing module of the fuel mixing section, parameter solving of an oblique detonation module of the combustion chamber section and parameter solving of an isentropic expansion module of the tail nozzle section; and combining corresponding parameters of the solving processes, including temperature, pressure and velocity, performing theoretical thrust calculation of the oblique detonation engine, and calculating thrust performance including specific impulse. The method sets fuel composition, air inlet working condition, mixing working condition, combustion working condition and expansion working condition of the oblique detonation engine, and can quickly calculate specific impulse of the oblique detonation engine and each physical quantity along the way.

The invention discloses a composite vortex generator for increasing high-speed air inlet flow and a design method, and belongs to the field of high-speed aircraft air inlet flow control. The vortex generator consists of a reverse slope type vortex generator and a streamline type vortex generator. The reverse slope type vortex generator comprises a slope, two side faces and a bottom face. The streamline vortex generators are symmetrical on the left side and the right side, and each streamline vortex generator on one side comprises an inclined face, an outer side face, two inner side faces and a bottom face. The left side and the right side of each streamline vortex generator are spaced by a certain installation distance, and the streamline vortex generators are installed on the two sides of the reverse slope type vortex generator. The front edge of the vortex generator is close to the incoming flow direction and arranged in front of the high-speed air inlet channel, and the vortex generator is used for improving the flow capturing capacity of the high-speed air inlet channel. The reverse slope type vortex generator and the streamline vortex generator work together, high-density fluid is conveyed to a boundary layer in a wider range in the circumferential direction, and the unit thrust and specific impulse of the combined engine are improved.

This utility model relates to the field of satellite propulsion equipment technology, and specifically provides an integrated gas-fuel hybrid propulsion system, including an integrated tank, gas generator, gas container, and thruster. The tank contains a flexible metal diaphragm, with one side forming a gas chamber and the other a liquid chamber. The gas chamber has a gas inlet, and the liquid chamber has a liquid inlet. The liquid inlet's output pipe is connected to the gas generator, the gas generator's output pipe is connected to the gas container, and the gas container's output pipe is connected to the thruster. This utility model solves the problems of low specific impulse in cold gas propulsion systems, heavy thruster mass and large envelope size in monocomponent chemical propulsion systems, and the difficulty in achieving low thrust output in bicomponent chemical propulsion systems, thus meeting the application requirements of satellites.

This invention relates to a multiphysics-coupled optimization design method and equipment for a toroidal pebble bed reactor core used for nuclear thermal propulsion. The method utilizes a stochastic elastic filling algorithm for coated particles to obtain the initial geometric parameter combination of fuel elements; constructs a multiphysics-coupled finite element model, and optimizes the final fuel design parameters under multiple constraints; based on this, Monte Carlo neutronics calculations are used to perform criticality optimization on the core, resulting in a neutronics design scheme that meets criticality and safety requirements; based on this core geometry, computational fluid dynamics analysis is used to optimize coolant distribution by adjusting the porosity of the thermally porous plate, and under conditions that meet temperature and pressure drop limits, the core power and hydrogen mass flow rate range that can achieve preset specific impulse, thrust, and thrust-to-weight ratio performance indicators are determined. Compared with existing technologies, this invention achieves a complete and systematic optimization design for a novel spherical metal-ceramic fuel nuclear thermal propulsion pebble bed reactor core, from concept to meeting preset propulsion performance indicators.

The application provides a microwave electrothermal thruster performance prediction method and system based on heat loss power evaluation, which calculates initial specific energy and temperature estimation value by inputting total power and working medium flow. Then, a heat balance equation is constructed to solve the temperature of the inner and outer walls of the resonant cavity, so as to obtain the wall heat loss power and calculate the actual absorbed microwave power. The specific energy and temperature estimation value are recalculated by using the power, and the process is taken as a cycle, and the output temperature is fed back to the heat balance equation for iteration. When the temperature difference of two adjacent iterations is less than a preset threshold, the iteration is stopped, and the final iteration temperature is taken as the actual working medium temperature, so that the thrust and specific impulse performance parameters of the thruster can be accurately predicted.

The invention belongs to the field of engines, and particularly relates to a fuel supply structure and a solid-gel hybrid rocket engine, the fuel supply structure comprises an oxygen-enriched grain propellant supply mechanism and a gel fuel supply mechanism; the oxygen-enriched grain propellant supply mechanism comprises a coating column, the coating column comprises a coating sleeve with an opening facing the combustion chamber, and an oxygen-enriched grain is arranged in the coating sleeve; the gel fuel supply mechanism comprises a gel fuel supply device, an annular channel and an atomizer which are sequentially communicated, the annular channel is wound outside the wrapping sleeve, the downstream of the annular channel protrudes out of the downstream of the wrapping sleeve, the atomizer is arranged at the end, protruding out of the downstream of the wrapping sleeve, of the annular channel, and the annular channel and the wrapping sleeve are combined to block the upstream end of the combustion chamber. The engine adopts a new combination of the gel fuel and the oxygen-enriched grain, has the advantages of the gel fuel and the oxygen-enriched grain, and achieves the effects of increasing specific impulse, adjusting thrust and starting for multiple times.

The invention relates to an air-breathing electric propulsion system ground test system and method, and belongs to the field of ultra-low orbit space electric propulsion, and the system comprises a spiral wave plasma source, an incoming flow parameter measurement device, a tested air-breathing electric propulsion system, a micro-thrust frame, a vacuum system, a gas cylinder and a micro-thrust frame measurement system. A gas cylinder is used as a gas source to generate simulated incoming flow through a spiral wave plasma source, the simulated incoming flow enters a tested air-breathing electric propulsion system, the thrust of the tested air-breathing electric propulsion system is calculated by measuring net thrust, and then specific impulse is calculated. The spiral wave plasma source is used for generating rarefied gas plasma incoming flow equivalent to neutral gas incoming flow, incoming flow with different track heights and different component contents can be simulated by adjusting flow, power, magnetic field intensity and the like, and performance verification of the air suction type electric propulsion system is achieved by measuring net thrust.

The application provides a metal fuel propellant and a preparation method thereof, and belongs to the technical field of solid propellants. 3 Compared with traditional aluminum-based propellants, the zirconium-based propellant has a greater density, and the increase of the density can not only increase the volume specific impulse of the propellant, but also increase the speed increment of a missile, so that the zirconium-based propellant has more excellent energy characteristics and can be applied to small volume engines; the application of zirconium hydride as a fuel additive increases the hydrogen content in the fuel, reduces the flame temperature of the propellant and improves the specific impulse of the propellant; the application of zirconium hydride as a fuel additive has good resistance to direct oxidation of perchlorate and can improve the energy characteristics and combustion performance of the zirconium-based propellant; and the application of aluminum powder as a fuel additive has high combustion heat value and good oxidation performance and can promote the combustion reaction of the zirconium powder.

The invention discloses a mixed fuel multi-working-condition oblique detonation engine performance analysis method, and an oblique detonation engine comprises an air inlet channel section, a fuel mixing section, a combustion chamber section and a tail nozzle section. Sequentially solving parameters of an oblique shock wave compression module of an air inlet section, parameters of a mixing module of a fuel mixing section, parameters of an oblique detonation module of a combustion chamber section and parameters of an isentropic expansion module of a tail nozzle section; and calculating the theoretical thrust of the oblique detonation engine by combining the corresponding parameters of each solving process, including temperature, pressure intensity and speed, and calculating the thrust performance including specific impulse. According to the method, the fuel composition, the air inlet working condition, the mixing working condition, the combustion working condition and the expansion working condition of the oblique detonation engine are set, and the specific impulse and all physical quantities along the way of the oblique detonation engine can be rapidly calculated.

The present invention discloses a controllable scramjet engine and its operating method, belonging to the field of scramjet engines. The engine comprises a supersonic combustion chamber, a controllable gas generator disposed on the outer wall of the supersonic combustion chamber, the output end of the controllable gas generator being connected to the input end of the supersonic combustion chamber via a regenerative cooling pipe disposed on the wall of the tail nozzle of the supersonic combustion chamber; the controllable gas generator comprises a main body, fuel filled within the main body, and an electric pyrolysis mechanism for pyrolyzing the fuel. Using the above-mentioned controllable scramjet engine and its operating method, pyrolysis gas is used as fluidizing gas to carry powdered fuel for injection, effectively enhancing the engine's specific impulse performance. The gas fuel can also serve as a pilot fuel, first burning to increase pressure and temperature, and then carrying a flow to eject high-energy powder particles, thereby solving the problem of short residence time of powdered fuel in the combustion chamber and low combustion efficiency.

The invention discloses a pulse detonation engine for multi-directional attitude adjustment of a rocket. The pulse detonation engine comprises a detonating pipe section, a branch pipe, a spray pipe and the like. The engine is arranged on the transverse plane of the rocket body, the working modes are changed through the switching valves in the branch pipes, clockwise and anticlockwise self-rotation attitude adjustment around the axial direction of the rocket and overall pitching attitude adjustment of the rocket body can be achieved in different working modes, and the engine has the advantages of being fast in thrust response, accurate in pulse response, large in specific impulse, high in reliability and the like. And the rocket attitude adjusting device is suitable for various rocket flight attitude adjusting tasks.

The application relates to the field of aerospace propulsion technology, in particular to a space micro-thrust engine, which comprises an external supporting cylinder, a grid assembly is fixedly connected to one side of the external supporting cylinder, a flange plate is fixedly connected to the other side of the external supporting cylinder, a hollow cathode assembly is fixedly connected to the middle of the flange plate, the hollow cathode assembly is located in the external supporting cylinder, an annular gas pipe is arranged between the grid assembly and the hollow cathode assembly, one end of the annular gas pipe penetrates through the flange plate, a magnetic field part is arranged on the external supporting cylinder, and the magnetic field part is located at the end far from the flange plate. The application can provide a compact structure and high specific impulse space micro-thrust engine.

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.

The invention provides a space engine jet pipe based on multi-material metal ceramic additive manufacturing. The space engine jet pipe comprises a non-high-temperature area, a high-temperature area and a transition area. The non-high-temperature area is made of metal materials which comprise niobium alloy and titanium alloy. The high-temperature area is made of a ceramic material, and the component is SiC; the transition area realizes continuous gradient transition from non-high-temperature area metal to high-temperature area ceramic through multi-material additive manufacturing; the whole spray pipe is of a Laval spray pipe structure, and a non-high-temperature area, a high-temperature area and a transition area are integrally formed through additive manufacturing. The space engine can work in a higher performance interval, the space engine is suitable for combustion of an injector with higher specific impulse and a high-energy propellant, the thermal shock cycle life is prolonged, and meanwhile the weight of the whole spray pipe can be reduced.