127 results about "Specific impulse" patented technology

The invention discloses an orbit control safety shutdown time calculation method suitable for deep space exploration, and belongs to the field of deep space exploration. The method comprises the steps that firstly, the theoretical speed increment of orbit control is calculated, and a sinking engine, an orbit control engine, an attitude control engine and related engine thrust and specific impulse parameters are determined; secondly, the actual orbit control speed increment is obtained, and the orbit control nominal starting time length is determined; thirdly, calculating time error deviation, and giving out engine shutdown forbidding time and forced shutdown time in the orbit control process; if the task is a conventional orbit control task, taking the two times as final safe shutdown time; for the orbit control task related to safety, multiple groups of shutdown forbidding time and forced shutdown time under the nominal working condition and the pulling deviation working condition are given, and finally the safety shutdown time of the orbit control task related to safety is determined. According to the method, the track safety and task safety of the detector are ensured by selecting proper shutdown forbidding time and forced shutdown time.

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.

According to the light helium storage spray pipe with the integrated efficiency of efficiency increasing, pushing changing, phase change heat prevention and infrared stealth, by means of the high phase change latent heat characteristic of liquid helium, the liquid helium is directly injected into the spray pipe to be mixed with high-temperature fuel gas; the temperature of fuel gas and thermal stress borne by a spray pipe material are remarkably reduced through phase change heat absorption, heat convection and gas film cooling, efficient thermal protection is achieved, the infrared radiation intensity of exhaust plume is effectively restrained, and the infrared stealth effect is achieved; the characteristic that helium is small in molecular weight is utilized, the expansion power capability of mixed gas and the propulsive efficiency and specific impulse of the engine are improved, meanwhile, through real-time regulation and control of helium injection flow, thrust adjustment to a certain degree can be achieved, and the maneuverability of the engine is enhanced. According to the designed composite interlayer structure, the thickness of the wall face of the spray pipe can be greatly reduced, the structural strength of the spray pipe is improved, and then the overall mass of an engine is reduced.

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 invention discloses a time-of-flight mass spectrometry device for characterizing different-angle performance of an electrospray thruster, and belongs to the field of testing of emission performance of the electrospray thruster. The device comprises a flight time mass spectrum device, a thruster rotating system, a thruster, a power supply system and a data acquisition device. The time-of-flight mass spectrometry device comprises an electrostatic door device, a particle flight tube section and a collection device. The thruster device comprises an electrospray thruster, an electrospray thruster rotating system and a fixing device. The collecting device is a stainless steel round pipe with a flange, and an insulating spacer, a metal grid mesh and a collecting plate are placed in the round pipe. According to the invention, the wire mesh wraps the transmission line for transmitting the current signal on the collection plate, so that electromagnetic interference is reduced. The emission angle of the electrospray thruster is controlled through the electric rotating table outside the vacuum cabin, the rotating angle of the thruster is accurately controlled, plume velocities of the electrospray thruster at different angles are measured, and the thrust and specific impulse of the electrospray thruster at different angles are calculated through measured data.

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.

Electric resistance thermal propulsion systems that are configured to heat propellants for expulsion from a thruster using a heat exchanger that includes ultra-high temperature ceramics, such as hafnium carbide (HfC), are disclosed. The heating chamber may also be constructed from such materials. In operation, the electric resistance thermal propulsion system raises the temperature of the propellants that exit the heating chamber to expand in the nozzle to generate thrust. HfC, for example, has both metal and ceramic properties, and thus, it can conduct the electrical current. The higher the temperature, the higher the specific impulse that can be generated, and the electric resistance thermal propulsion systems can tolerate a much higher temperature than has been possible before with this type of thruster.

The present invention relates to a field-reversed configuration pulse plasma thruster, comprising: a spiral wave antenna, a preionization cavity, a main discharge cavity, a rotating magnetic field antenna group, a bias magnetic field coil group and a gas distributor. The cylindrical preionization cavity and the conical main discharge cavity are made of high-temperature resistant insulating materials. The spiral wave antenna is installed outside the preionization cavity, and preionizes the gas working medium under the excitation of radio frequency voltage. The rotating magnetic field antenna group is arranged outside the main discharge cavity, and generates a rotating magnetic field in the main discharge cavity under the drive of a high-frequency alternating current with a certain phase difference. The bias magnetic field coil group is used to generate an axial bias magnetic field in the preionization cavity and the main discharge cavity. The present invention uses spiral wave discharge to preionize the gas working medium, and ionizes and accelerates the plasma group through an angular rotating magnetic field and an axial bias magnetic field, so as to realize electrodeless electromagnetic acceleration, and at the same time has the advantages of high specific impulse, high efficiency, good power scalability, and diversified propellants.

The utility model relates to the technical field of space propulsion, in particular to a gas propulsion device. The gas propulsion device comprises a shell, a partition plate, a control assembly, a first channel, a thruster and a grid body. A filling port and a mounting assembly are arranged on one side of the shell, an outlet is formed in the other side of the shell, and an inner cavity is defined by the shell and used for storing liquid gas; the spacing plate is arranged on the inner wall of the shell and divides the inner cavity into a first cavity close to the filling port and a second cavity close to the outlet, and small through holes are fully distributed in the spacing plate; the control assembly is arranged on the mounting assembly and communicates with the first cavity, and liquid gas in the shell is gasified and then decompressed to preset pressure through the control assembly; the first channel is communicated with the control assembly and the outlet, and gas-liquid mixed fluid decompressed by the control assembly flows through the first channel. According to the highly-integrated gas propelling device, gas-liquid separation can be better achieved in a weightless environment, liquid gas is fully subjected to phase change, and the gas utilization efficiency and specific impulse of the device are improved.

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 present invention belongs to the technical field of tactical weapon system layout and engine structure design, and specifically relates to a solid fuel ramjet rotating detonation combined power system, comprising: an air inlet, an isolation section, a main charge, a pre-detonator tube, and a rotating detonation combustion chamber. To address the current difficulty in organizing mixed combustion in the afterburner, the present invention proposes using a rotating detonation combustion chamber instead of a traditional isobaric combustion chamber, thereby improving the engine cycle thermal efficiency and making the engine structure more compact. Theoretical calculations show that this invention can increase the engine specific impulse by more than 5%. Furthermore, the present invention provides a systematic design solution in conjunction with a solid fuel ramjet rotating detonation combined engine.

The invention 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.

The invention 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-current mutual-striking type inspirator, and the combustion chamber comprises a linear section, a convergence section, a throat part and an expansion section which are connected in sequence; the direct-current mutual-striking type injector is arranged at the end of the straight line section, the length-diameter ratio of the straight line section is 1.2-1.5, the characteristic length of the straight line section is 290-350 mm, a first central area injection hole, a second central area injection hole and an edge area injection hole are formed in the direct-current mutual-striking type injector, an oxidizing agent is sprayed out through the first central area injection hole, and the edge area injection hole is communicated with the second central area injection hole. Fuel is sprayed out through the second injection holes in the center area and the injection holes in the edge area. The adaptive working condition range of the thrust chamber is greatly widened, and the problems that the throat temperature of a space engine with high specific impulse performance is too high and is greatly influenced by working condition deviation are solved.

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 application provides an air-breathing space thruster and a specific impulse optimization method thereof. The thruster comprises an oxygen and air mixing supply module, a catalytic chamber, a flame arrester, a combustion chamber and a nozzle. The specific impulse optimization method comprises the following steps: 1) setting an initial oxygen / fuel mass ratio e1 for the air-breathing space thruster; 2) introducing air into the oxygen for the air-breathing space thruster to reduce the oxygen concentration; 3) controlling the total flow of the propellant, i.e. the total mass flow of the oxidizer and the fuel, to be stable before or under the choked condition; 4) measuring the thrust of the air-breathing space thruster by a test platform, and calculating the test specific impulse and the theoretical specific impulse without considering the air mass; and 5) obtaining the oxygen concentration corresponding to a higher specific impulse, i.e. the flow ratio of the oxygen and the air, according to the change of the thruster specific impulse with the oxygen concentration. The application adjusts the oxygen concentration in the oxidizer by introducing the thin air, and improves the specific impulse performance of the thruster without changing the fuel flow, thereby improving the overall propulsion efficiency.

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 present invention discloses a microwave detonation dual-mode space thruster, which includes a propellant supply system, a large-thrust propulsion module and a micro-thrust propulsion module coaxially arranged from the inside to the outside in sequence; the large-thrust propulsion module includes an injector, a plasma ignition combustor, a detonation chamber, a tail nozzle, etc. The micro-thrust propulsion module includes a microwave supply system, magnetic poles, a gas collection chamber, a discharge chamber and a grid; the microwave supply system includes a microwave connector and an antenna; the antenna is a cylindrical antenna, and a cyclotron cavity is formed between the antenna and the magnetic poles; the outer electrons in the working medium molecules located in the cyclotron cavity perform cyclotron motion in the cyclotron cavity, and resonance occurs when the cyclotron angular frequency is equal to the microwave frequency. The present invention combines microwave discharge technology, plasma combustion assistance technology and pulse detonation combustion technology, and realizes efficient and convenient switching of working modes through circuit regulation, enabling the thruster to meet the stringent requirements of large thrust, high specific impulse, long life, etc., and can be applied to diverse space missions.

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 utility model discloses a single-component micro-thruster based on an MEMS (micro-electromechanical system) technology, which is characterized in that nitrous oxide is used as a propellant, a reaction cavity is arranged on one surface of a thruster base, and a gas inlet area and a Laval nozzle which are communicated with each other are arranged in the reaction cavity; a plurality of columnar bodies are arranged at the position, close to the gas inlet area, of the reaction cavity, and the columnar bodies are arranged in a mutually separated mode; a plurality of accommodating holes are covered on the surface of the reaction cavity between the plurality of cylindrical bodies and the Laval nozzle, and the plurality of accommodating holes are filled with an iridium catalyst; the catalyst can be decomposed under the action of the catalyst and high temperature, the reaction releases heat, decomposition is self-maintained, specific impulse and thrust are improved, and energy is saved. And moreover, after the reaction cavity is designed to be porous, the catalytic decomposition efficiency can be improved, and the specific impulse of the thruster is further improved, so that the problem that an existing single-component micro-thruster is difficult to realize the balance of safety, low cost and high performance of a propellant is practically solved.

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 invention discloses a direct-drive liquid rocket engine which comprises an engine body, a combustion chamber and a thrust chamber. The combustion chamber is installed at the upper end of the interior of the engine body, and the thrust chamber is located below the combustion chamber. According to the simplified system structure, the requirements for part precision and the manufacturing process are reduced, the number of parts is reduced, and the material cost and the manufacturing cost are reduced; the oxidant is fully gasified before reaching the nozzle, the fuel is fully gasified into fuel gas through multi-point heat exchange in the flowing process, the oxidant and the fuel are mixed with the fuel more uniformly, the combustion reaction is more sufficient, the combustion efficiency of the engine is effectively improved, and the thrust and specific impulse performance of the engine are further improved; compared with a staged combustion rocket engine, due to the fact that the structure is simple and the number of transmission links is small, a control instruction can be responded more quickly during starting, shutdown, thrust adjustment and other operations.

The invention relates to the technical field of target damage assessment, in particular to a building component explosion damage rapid calculation method based on uncertainty quantification, and the method comprises the steps: obtaining explosive power field parameters according to the proportional distance from the explosion center to the center of a component, the standard deviation of the shock wave overpressure peak value of an explosive and the standard deviation of the specific impulse; obtaining component size and material characteristic parameters according to the standard deviation of the size of the building component and the standard deviation of the material characteristic; obtaining an average damage result of the component according to the explosive power field parameter, the component size and the material characteristic parameter; and obtaining a building component damage probability density distribution function according to the average damage result. The method has the advantages that random factors such as explosive power fields, building structure sizes, material uncertainty and explosion source positions are quantified, building component explosion damage rapid calculation based on uncertainty quantification is achieved, and probability distribution of damage results can be given.

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.