14 results about "Monopropellant" patented technology

A system for generating hot, high-pressure exhaust gas for thrust includes an energetic powder monopropellant, wherein each particle of powder is comprised of a combination of oxidizer and fuel in a ratio suitable for combustion; and a rocket motor for combusting the monopropellant, said rocket motor comprising a container for storing the monopropellant, and a unitary feeder-combustor, wherein the unitary feeder-combustor is configured to feed the powder monopropellant from the container and to combust the powder monopropellant therein without consumption of air, thereby generating hot, high-pressure exhaust gas, generating thrust.

The application discloses an on-orbit use method of a green single-component thruster, and the single-component thruster structure comprises five parts of an injector, an anti-inversion functional material, a catalytic bed, a nozzle and a thermal control assembly, wherein the thermal control assembly comprises an armored heater, an armored thermistor and a multilayer thermal insulation assembly. The application provides the on-orbit use method of the green single-component thruster containing the anti-inversion functional material, and solves the technical problem of the warm start life of the green single-component thruster through reasonable working condition design, and the green single-component thruster is successfully applied to micro-nano satellites.

The present application relates to the technical field of data analysis, and provides a method and system for analyzing the ignition state of a space monopropellant engine. The method includes: obtaining the initial temperature parameter, initial pressure parameter, and propellant flow parameter of the combustion chamber of the space monopropellant engine in the ignition state; generating a flow domain grid model including the characteristics of porous medium effect and chemical reaction effect based on the initial temperature parameter, initial pressure parameter, and propellant flow parameter of the combustion chamber; setting dynamic boundary conditions corresponding to the ignition transient process in the flow domain grid model, and performing a full-cycle numerical simulation on the internal temperature field distribution of the space monopropellant engine to generate a temperature field dynamic change sequence in the ignition state; extracting the thermal stress evolution data of the key components of the space monopropellant engine according to the temperature gradient distribution characteristics at different moments in the temperature field dynamic change sequence, and combining the thermal stress evolution data with a preset overheat failure threshold to output the performance stability evaluation result in the ignition state.

A system for generating hot, high-pressure exhaust gas for thrust includes an energetic powder monopropellant, wherein each particle of powder is comprised of a combination of oxidizer and fuel in a ratio suitable for combustion; and a rocket motor for combusting the monopropellant, said rocket motor comprising a container for storing the monopropellant, and a unitary feeder-combustor, wherein the unitary feeder-combustor is configured to feed the powder monopropellant from the container and to combust the powder monopropellant therein without consumption of air, thereby generating hot, high-pressure exhaust gas, generating thrust.

The invention provides a low-freezing-point propellant supply system and a using method. The low-freezing-point propellant supply system comprises a propellant storage unit, a propellant conveying unit and a pipeline local cooling unit. The propellant storage unit and the propellant conveying unit are in butt joint in a threaded pipe mode, and the spraying range of a first cold blowing ring, a second cold blowing ring and a third cold blowing ring in the pipeline local cooling unit covers a local cold leakage area on a pipeline through which propellant conveying passes in the propellant conveying unit. The overall structure and the using method solve the problems that in the prior art, the cooling capacity is weak, the temperature control precision is low, and the application range is narrow.

The invention provides a regenerative cooling nitrous oxide single-component thruster thermal management method and system, and relates to the technical field of aerospace propulsion, and the method comprises the steps: obtaining catalytic bed temperature distribution and propellant flow data, obtaining a thermal state parameter through nonlinear mapping, and obtaining a thermal state parameter; the rotational flow intensity is determined, main inlet axial flow and auxiliary inlet tangential flow are coordinated to form a rotational flow field, a thermal boundary layer is destroyed through rotational flow, so that heat exchange is enhanced, and the heat exchange state is determined; the flow distribution proportion of the main inlet and the auxiliary inlet is dynamically adjusted through a self-adaption PID control algorithm in combination with the inner wall face heat flow and a catalytic bed temperature signal, and the real-time flow proportion is obtained; and predicting a thermal runaway risk level through a fuzzy logic decision maker based on the ratio, updating control parameters and re-executing flow regulation according to the thermal runaway risk level, forming closed-loop management, and realizing efficient and reliable thermal management of the thruster. The thermal management accuracy and reliability of the thruster under variable working conditions are improved.

The invention discloses a high-performance monopropellant engine which comprises a head flange, an injection disc, a catalytic bed front bed, a catalytic bed rear bed and a spray pipe. The head flange is connected with an injection disc which is connected with a catalytic bed, and the rear end of the catalytic bed is connected with a spray pipe; a plurality of circles of injection holes are uniformly distributed in the injection disc; the catalytic bed comprises a front catalytic bed body and a rear catalytic bed body, the front catalytic bed body and the rear catalytic bed body are separated through a rhenium-iridium alloy partition plate, and a plurality of through holes are formed in the rhenium-iridium alloy partition plate; and a heating wire is wound outside the catalytic bed. The catalytic bed material combination of the niobium-tungsten alloy outer shell and the rhenium-iridium alloy lining is adopted, the high-temperature working condition requirement is met, and the decomposition characteristic of the propellant is enhanced through the form design of the front bed and the rear bed. The engine can be suitable for high-performance single-component non-toxic propellants, and is simple in structure, high in reliability, non-toxic and pollution-free.

The invention relates to the technical field of spaceflight electric propulsion systems, in particular to a flow adjusting device. The inlet connector is used for inputting working medium gas, one end of the inlet connector is located outside the first end of the valve seat, and the other end of the inlet connector is embedded into the first end of the valve seat; the flow sensor is installed on the upper surface of the valve seat, connected to the inlet connector through a first pipeline located in the valve seat and used for measuring the working medium gas flow of the first pipeline; the proportional electromagnetic valve is installed on the upper surface of the valve seat and connected to the flow sensor through a second pipeline located in the valve seat, and the proportional electromagnetic valve adjusts the opening degree of the valve according to the control signal; the outlet connector is used for outputting working medium gas, one end of the outlet connector is located outside the second end of the valve seat, and the outlet connector is connected to the proportional electromagnetic valve through a third pipeline located in the valve seat. According to the invention, the direct measurement of the propellant flow of the storage and supply unit is realized, and the influence of upstream pressure on the measurement precision is avoided.

The application provides a hydroxylamine nitrate-based monopropellant and a preparation method thereof, and comprises the following components with mass percentage: HAN: 50%-70%; alcohol: 14%-25%; functional additive: 14%-35%; and water: 1%-5%; wherein the functional additive comprises high-energy additive and cosolvent additive, the mass percentage of the high-energy additive is 13.5%-34.8%, and the mass percentage of the cosolvent additive is 0.2%-0.5%. The hydroxylamine nitrate-based monopropellant provided by the application has the characteristics of green, low toxicity, high specific impulse performance, low sensitivity, clean and non-polluted combustion products and the like.

A system for generating hot, high-pressure exhaust gas for thrust includes an energetic powder monopropellant, wherein each particle of powder is comprised of a combination of oxidizer and fuel in a ratio suitable for combustion; and a rocket motor for combusting the monopropellant, said rocket motor comprising a container for storing the monopropellant, and a unitary feeder-combustor, wherein the unitary feeder-combustor is configured to feed the powder monopropellant from the container and to combust the powder monopropellant therein without consumption of air, thereby generating hot, high-pressure exhaust gas, generating thrust.

The invention discloses an automatic catalyst filling system for a single-component thrust chamber. The automatic catalyst filling system comprises a material taking device, a catalyst storing and taking device and a vibration device, wherein the material taking device takes out a thrust chamber and decomposition chamber assembly and places the thrust chamber and decomposition chamber assembly on the catalyst storing and taking device or the vibration device; the catalyst access device pours a catalyst with a set weight into the thrust chamber and decomposition chamber assembly; and the vibration device is used for knocking and vibrating the thrust chamber decomposition chamber assembly. According to the invention, automatic filling of the catalyst in the thrust chamber is realized, so that the production efficiency is improved, the labor intensity is reduced, and the quality consistency of products is improved.

The present invention provides a staged non-catalytic ignition non-toxic monopropellant engine and a spacecraft. The staged non-catalytic ignition non-toxic monopropellant engine comprises a branch ignition module, a main injector, a main combustion chamber and an engine nozzle; the end of the branch ignition module is located in the main combustion chamber, and a plurality of through holes are arranged in the middle and lower part of the branch ignition module, and the middle and lower part of the branch ignition module is connected with the main combustion chamber through the through holes. The branch ignition module comprises a branch injector, an igniter and a branch cavity, and the other end of the branch cavity is defined as a branch combustion chamber; the present invention overcomes the problem that large-flow non-toxic monopropellant is difficult to ignite by adopting a staged non-catalytic ignition scheme through the design of the branch combustion chamber and the main combustion chamber, and the engine adopts a non-catalytic ignition scheme, and does not need to use precious metal catalysts, which can save the cost of expensive catalysts, and can avoid the problems of catalyst temperature resistance limitation and the difficulty of self-control of precious metal raw materials.

A monopropellant thruster according to an exemplary aspect of the present disclosure includes, among other things, a first part having a catalyst bed, a thrust chamber, and a nozzle. The first part is integrally formed via a single additive manufacturing process. The thruster further includes a second part, which is a closeout. A method is also disclosed.

The invention provides an anti-backfire single-component liquid engine injector structure which comprises a collector, a support, an injection plate, a capillary tube, an injection head, a fire retardant element and the like. The support is used for supporting and connecting the collector and the injection plate, one end of the capillary tube is fixedly connected to the collector, the other end of the capillary tube is fixedly connected to the injection plate, the injection head is in contact with the injection plate, and the fire retardant element is installed in the injection head and the collector. According to the single-component liquid engine injector, the fire retardant elements are installed in the injection head and the collector, secondary fire retardance is achieved, the fire retardant capacity of the single-component liquid engine injector is improved, and the problem that the engine injector for some oxygen-combustion integrated single-component propellants is prone to tempering is solved. The structure is simple, the system cost is remarkably reduced, and engineering application of a green high-performance propellant is further promoted.