18 results about "Burn rate" patented technology

This invention discloses a one-dimensional internal ballistic calculation method for multi-propellant rocket engines that considers differences in gas properties. The method includes: establishing a physical model of a solid rocket motor; establishing a transient control equation set to calculate the density of the mixed gas, combustion chamber temperature, gas velocity, and combustion chamber pressure at the next time step; obtaining the remaining gas volume of different propellants at the next time step; updating the weighted average of physical property parameters; calculating the burning rate and burning surface area at the next time step; repeating the above operations until the propellant grain is completely burned out, thus completing the one-dimensional internal ballistic calculation. This invention considers the mixing process of gases generated by different propellants during the combustion of a mixed-charge solid rocket engine, resulting in accurate calculation results.

The invention discloses a design method of a wedge-shaped grain for testing a combustion rate by an impulse method based on a pressurization rate boundary. The problem that in an impulse method experiment, when an inner hole surface-increased combustion tubular grain is adopted for measuring the dynamic combustion speed of a solid propellant under different pressure intensities, the pressurization rate of a high-pressure-index propellant in the later stage of combustion is too high, and consequently the combustion speed cannot respond to pressure intensity changes in time is mainly solved. The method has the innovation points that (1) a mathematical relationship between the combustion surface area size and the change rate and the influence of the nozzle throat diameter on the pressurization rate is established, and a design method for controlling the pressurization process by regulating and controlling the combustion surface area is provided; (2) an outer wedge-shaped grain is designed on the basis of a traditional tubular grain, the combustion surface area and the change rate of the combustion surface area in the later stage of combustion are reduced, and the pressurization rate in the later stage of combustion is inhibited; and (3) in the design process of the wedge-shaped grain, the specific design process of the wedge angle of the wedge-shaped grain, the thickness of the cylinder section and the sizes of other grains is given in combination with the pressurization rate, and the design process of the wedge-shaped grain is standardized.

The application discloses a high-precision high-speed prediction method for the burning rate of hydroxyl-terminated propellant, which comprises the following steps: establishing a database containing a plurality of groups of data, wherein each group of data comprises the burning rate of the propellant and its influencing factors; standardizing the database to obtain a sample library, wherein each sample in the sample library comprises the standardized burning rate of the hydroxyl-terminated propellant and its standardized influencing factors; dividing the sample library into a training set and a test set; training a plurality of algorithm models by using the training set to obtain a plurality of propellant burning rate models; preliminarily screening the plurality of propellant burning rate models by using prior knowledge; finely screening the plurality of propellant burning rate models obtained through the preliminary screening; fusing the plurality of propellant burning rate models obtained through the fine screening to obtain a propellant burning rate fusion prediction model; and predicting the burning rate of the propellant by using the propellant burning rate fusion prediction model. The application is convenient to use, simple to operate, fast in calculation, high in calculation precision and high in engineering application value.

The invention discloses a method for improving the burning rate of an end-combustion solid propellant. The method comprises the following steps: preparing a pure silver mesh belt; cutting the prepared pure silver mesh belt to obtain a rectangular strip-shaped pure silver mesh belt; the rectangular strip-shaped pure silver mesh belt is pretreated; embedding the pretreated rectangular strip-shaped pure silver mesh belt into the end combustion solid propellant in the direction perpendicular to the combustion surface; and after demolding, shearing and removing redundant pure silver mesh belts at the two ends of the end-combustion solid propellant to obtain the end-combustion solid propellant embedded with the pure silver mesh belts so as to improve the combustion speed of the end-combustion solid propellant. According to the method, the pure silver mesh belt is embedded in the end-combustion solid propellant, the area of a rapid heat dissipation path is increased, and heat is transferred from a combusted part to an uncombusted part, so that the combustion surface of the end-combustion solid propellant is enlarged, the combustion speed of the end-combustion solid propellant is increased, and the service life of the end-combustion solid propellant is prolonged. The problems that in an existing technical scheme, the burning speed of an end-burning solid propellant is limited in increasing amplitude, burning is uneven, and a single silver wire is prone to being snapped are solved.

The present application provides a method for determining the boundary condition of the pressurization rate in the impulse method of testing the burning rate; it belongs to the technical field of testing the combustion performance of solid propellant; the present application defines the pressure burning rate as the burning rate of propellant under a certain constant pressure, which is only related to the pressure; when the dynamic pressure burning rate changes rapidly, the burning rate depends not only on the current pressure, but also on the pressure change rate; the present application starts from the “burning rate lag caused by pressure change”, and obtains that when the product of the burning rate change rate and the response time is much smaller than the burning rate under the current constant pressure, it is considered that “the increasing pressure of propellant combustion is instantaneously responsive”; according to the burning rate response time, a pressurization rate relational expression containing a time term is established, and when the burning rate response time is less than 5 ms, the expression of the pressurization rate limiting value of the burning rate timely responding to the pressure change is given, which is used for judging the effectiveness of the pressure curve in the impulse method of testing the burning rate, that is, the effectiveness of the original data.

The invention discloses a CuCr2O4 / PPy composite catalyst for a GATo propellant and a preparation method of the CuCr2O4 / PPy composite catalyst. The CuCr2O4 / PPy composite catalyst is a binary compound of spinel type CuCr2O4 nanoparticles and a polypyrrole three-dimensional network. The preparation method comprises the following two steps: preparing spinel type CuCr2O4 nanoparticles by a hydrothermal method; and reacting the CuCr2O4 / PPy composite catalyst with a pyrrole monomer at 0-3 DEG C by an in-situ polymerization method to obtain the CuCr2O4 / PPy composite catalyst. According to the composite catalyst, the burning speed of the GATo propellant can be synchronously increased by 15%-44%, the electrostatic spark sensitivity negative electrode is increased from 1.78 J to 2.45 J, the friction sensitivity burning explosion probability is reduced from 95% to 30%, the impact sensitivity characteristic drop height is increased from 34.5 cm to 52.5 cm, the burning speed increase and the sensitivity reduction of the GATo propellant can be synchronously achieved, the process conditions are mild, industrialization is easy, and the method is suitable for industrial production. The method is suitable for performance regulation and control of high-energy solid propellants in the fields of space carrying, weapon systems and the like

The present application belongs to the technical field of static burning rate test of energetic materials, and specifically provides a target line-like burning rate test device and method based on acoustic emission technology. A sample is prepared by cutting and coating the measured propellant strip using a sample preparation cutter and a coating device, the sample is loaded into an acoustic emission method burning rate instrument, ignition is initiated, the acoustic emission signal strength-time curve is detected, recorded and identified by the acoustic emission method burning rate instrument, the pulse signals emitted during the combustion of the high-energy propellant tablets at the beginning and end of the curve are obtained, the burning rate of the measured propellant strip is obtained from the length of the cut measured propellant strip and the interval time length of the first and last pulse signals, and the burning rate test of the measured propellant strip is completed. The present application solves the problems of the existing target line method burning rate test, such as the complicated process of threading the target line and binding the propellant, and the poor nitrogen acoustic emission method burning rate test signal, which cannot effectively capture the complete signal. The present application does not need to punch, thread and burn the support to bind the target line, simplifies the target line method burning rate test, and makes the burning rate test in a high-pressure gas environment more easily automated.

The invention provides a method for determining the size of a tubular grain for testing the burning rate through the impulse method, and the method comprises the following steps: establishing a relationship between the pressurization rate and the burning characteristic of a propellant and a mathematical relationship between the pressurization rate and a tubular charging size parameter, and revealing the influence of the inner hole burning tubular charging size on the pressure intensity change. Different requirements of impulse method combustion rate tests of propellants with different combustion characteristics on the sizes of the explosive types are determined. And in combination with constraints such as pressurization rate limitation, pressure intensity test range and erosion-resistant combustion, a tubular charging design criterion considering both test precision and safety is provided.

The invention belongs to the technical field of aerospace, and relates to a burning rate testing device for a metal wire embedded propellant. A combustion surface can be directly observed through the visual shell, the whole combustion process of the propellant is recorded through the high-speed camera, and the real combustion speed of the metal wire embedded propellant is measured. In addition, the device is easy to mount and dismount, and the experiment efficiency can be greatly improved. Meanwhile, the propellant dosage required by a single test of the device is small, and the experiment cost can be saved.

The invention discloses a solid propellant burning rate measuring method, device and equipment and a storage medium, and relates to the technical field of fixed propellant burning tests.The method comprises the steps that an image sequence of the solid propellant burning process is obtained; for a plurality of frames of images of the image sequence, based on local gray statistical characteristics of the images, adaptive threshold matrixes are generated for pixels in different areas in the images, so that edge judgment is carried out, and a combustion end face contour and a combustion background are distinguished; and measuring the dynamic burning speed of the solid propellant according to the burning end face profile of the multi-frame image. According to the method, the combustion end face contour can be accurately extracted, and then the dynamic combustion speed of the propellant in the stable combustion stage is calculated.

The invention discloses an engine for testing the combustion rate of a solid propellant with a large increase area ratio and a combustion rate calculation method of the engine. The engine for testing the combustion rate comprises a combustion chamber, a grain, a spray pipe, an ignition charge, a wire, a blanking cap, a flange, a sealing groove, a pressure measuring hole and a sealing ring, the grain is arranged in the combustion chamber; the tail of the combustion chamber is connected with the front edge of the nozzle through a flange; a radial sealing groove is formed in the flange, and a sealing ring is arranged in the radial sealing groove; the blanking cap is arranged in the expansion section of the spray pipe; the ignition charge is arranged in an inner hole of the grain; one end of the lead is connected with the ignition charge, and the other end is led out of the blanking cap; the pressure measuring hole is formed in the nozzle convergent section. According to the scheme, the loading thickness is increased, the test working time is prolonged, a dynamic burning rate test within a certain pressure range can be completed through one test, the test cost is reduced, the test precision is improved, and the method is suitable for the burning rate test of a high-pressure-index propellant required by a solid attitude and orbit control engine and the like.

The application provides a propellant safe service life evaluation method based on characteristic parameters, which takes effective stabilizer content and burning rate as characteristic parameters; the method comprises the following steps: step one, measuring original data of characteristic parameters; step two, performing constant temperature accelerated aging test; step three, arranging detection time points; step four, predicting and arranging detection time points; step five, testing and recording data; step six, truncation; step seven, establishing characteristic parameter-time curve; step eight, life prediction taking effective stabilizer content as characteristic parameter; and step nine, life prediction taking burning rate as characteristic parameter. The service life evaluated by the method is more accurate, and the propellant can meet the safe use requirement in the whole life cycle, and the life limit of the propellant can be guaranteed to be safe in storage and use.

The invention relates to a simplified simulation method, system, equipment and medium for solid propellant combustion surface retrogression based on isothermal surface equivalence, and the method comprises the steps: processing a propellant model, and generating a reference model; hybrid grid division and local gradient encryption are carried out on the reference model, and a simulation model is obtained through grid distortion constraint; after the simulation model is imported into a heat transfer simulation system, the thermal diffusion coefficient of the propellant is adjusted based on the burning rate, the isothermal surface migration rate of the propellant is made to be synchronous with the burning rate, constant value temperature excitation and adiabatic constraint are applied, and a collaborative configuration model is formed; and based on the collaborative configuration model, adaptively selecting an implicit fully coupled or explicit separated solution architecture to execute transient heat conduction simulation, and when a preset convergence condition is met, outputting a morphological evolution result of the isothermal surface. Therefore, according to the method, the burning rate is mapped through the thermal diffusion coefficient, the burning surface is replaced by the isothermal surface, and the solid propellant burning surface regression simulation efficiency and the configuration adaptability are remarkably improved.

The invention provides a method for determining a supercharging rate boundary condition in an impulse method burning rate test. Belongs to the technical field of solid propellant combustion performance testing. The pressure intensity combustion speed is defined as the combustion speed of the propellant under a certain constant pressure intensity and is only related to the pressure intensity; when the dynamic pressure intensity combustion speed changes rapidly, the combustion speed not only depends on the current pressure intensity, but also is influenced by the pressure intensity change rate; according to the method, starting from combustion speed lag caused by pressure change, when the product of the combustion speed change rate and the response time is far smaller than the combustion speed under the current constant pressure when the pressure in a certain micro period is increased, it is considered that progressively increasing of the pressure is instantaneously responded by combustion of propellants; a pressurization rate relational expression containing a time item is established according to the combustion rate response time, and a pressurization rate limit value expression that the combustion rate can respond to pressure change in time when the combustion rate response time is less than 5ms is given and used for judging the effectiveness of a pressure curve in an impulse method combustion rate test, namely the effectiveness of original data.

The invention provides a spherical particle solid fuel and a high-frequency pulse power device applying the spherical particle solid fuel, and belongs to the technical field of aerospace propulsion. The spherical particle solid fuel specifically comprises 60% of a high-energy propellant, 20% of an adhesive, 10% of aluminum powder and 10% of a burning rate regulator. The high-frequency pulse power device comprises a fuel storage component, a fuel conveying component, a fuel supply component and a fuel combustion component which are connected. Spherical particle solid fuel stored in the storage box enters the fuel channel through the fuel pump and the fuel channel, and the solid fuel is pushed by the emission motor to enter the fuel chamber at a certain speed. When the fuel moves to a set position, the igniter ignites the surface part of the fuel, and the process of turning from slow combustion to violent combustion is initiated. The high-frequency pulse power device can work for a long time without the help of an air source, the internal space of the aircraft is optimized, the working frequency of an engine can be effectively improved, and the overall performance of the aircraft can be improved.

In a rocket motor in which burn rate enhancement (BRE) wires are used to accelerate the burn rate of the solid propellent, embedded charge assemblies (EGAs) are configured as support structures for the BRE wires. Each EGA includes an energetic material that is configured to bum along with the solid propellent to produce thmst and, upon detonation, to break up the solid propellent to terminate thrust. The detonation may also be initiated as a part of process to prevent a higher-order reaction, such as in reaction to heating from a fire or other cause. By being located inside the casing, the energetic material and EGAs do not adversely affect aerodynamics of the flight vehicle of which the rocket motor is a part, such as a missile.

The present application relates to solid propellant combustion technology, aiming at providing a lattice Boltzmann simulation prediction method for aluminum-based solid propellant combustion agglomeration. It comprises: using LBM method for solid propellant combustion process simulation and aluminum particle agglomeration behavior research, through mesoscale numerical simulation of solid propellant combustion process quantitative calculation and prediction, using the statistical physics principle of LBM to couple multiple field distribution, realize the simulation of complex aluminum particle combustion. The present application can more naturally simulate the drag force, thermophoretic force and other effects of particles in fluid, and is closer to the physical reality; can realize the reduction of calculation complexity, realizes the efficient simulation of tens of thousands of aluminum particle agglomeration process, significantly improves the calculation efficiency of large-scale particle system simulation; can deeply reveal the internal mechanism of pressure, burning rate, particle size, oxidizing gas concentration and other factors affecting agglomeration; realize the coupling of multiple physical fields, and truly reflect the complex environment in the combustion chamber.