1108 results about "Propellant" patented technology

A setting tool fora downhole tool includes a setting housing, a setting assembly, and an integrated igniter. The setting housing has a passage therethrough. The setting assembly is positioned in the passage, and includes a drive portion and a deployment portion. The drive portion has an elongate body with a propellant chamber. The deployment portion has an opening defining a drive cavity shaped to slidingly receive an end of the drive portion. The deployment portion includes a deployment end connectable to the downhole component. The integrated igniter is positioned in the propellant chamber, and includes an integrator housing, an internal propellant, and a switch assembly connected to the internal propellant. Upon triggering the switch, the internal propellant is ignited to release an ignition fluid under ignition pressure to move the deployment portion and the downhole component connected thereto to an activated position.

The invention relates to an SEM (scanning electron microscope) and CT (computed tomography) compatible solid propellant mechanical property testing device and method, belongs to the technical field of precise scientific instruments and material micro-mechanical property testing, and solves the problems that in the prior art, due to the fact that imaging equipment is single, and image quality becomes poor due to creeping accumulation, micro-scale failure information of propellants is incomplete and inaccurate to obtain. In the mechanical property testing device, the in-situ mechanical testing device performs mechanical loading and load holding on the propellant sample, the SEM obtains the surface appearance of the propellant sample, the double-source double-probe industrial CT obtains the internal three-dimensional structure of the propellant sample, and the in-situ mechanical testing device is transmitted between the transfer cabin main cavity and the electron microscope cabin through the transfer channel. And the vacuum transfer cabin containing the in-situ mechanical testing device is transferred between the transfer channel and the scanning turntable of the double-source double-probe industrial CT by the mechanical arm. According to the invention, more comprehensive and accurate mesoscopic failure information of the propellant sample can be obtained while less time cost is paid.

A projectile and method for stopping aerial vehicles. The projectile comprises a shell—and a bunch of filaments. The bunch is wound in a coil and placed inside the shell. The projectile according to the solution may be brought to a target by different methods and utilising different means. For example, the projectile may be fired with compressed air or a propulsion charge, by using a compressed air weapon or rocket launcher. In addition, the projectile may be taken to the target by using a rocket or aerial vehicle.

A ram accelerator for accelerating a projectile is provided. The ram accelerator includes a baffle section configured to start the projectile in the ram accelerator without an obturator. The ram accelerator generally includes a tube having a projectile bore; and a baffle section operably coupled to a proximal end of the tube. The baffle section has an annular baffle wall defining a central bore axially aligned with the projectile bore; and a propellant chamber arranged adjacent to the annular baffle wall, wherein the propellant chamber is configured to enclose a propellant. The propellant can be ignited as the projectile passes through the baffle section by pressure created by the projectile, by an ignition source, or by a flame carried by the projectile to start the ram acceleration of the projectile.

Systems and methods for targeted operational disablement that may utilize, for example, UAV-delivered shaped charges coupled to specific features of a target object. A UAV may navigate to, identify, and attach a disabling device to a critical component of a target such as a barrel of a tank main gun, a crew hatch cover, a tank tread, an engine compartment, or a magazine.

The embodiment of the utility model provides a storage tank bottom structure, a storage tank and a carrier rocket, the storage tank bottom structure comprises a shell and a top cover, the shell is of a hollow arc-shaped revolving body structure which is through in the front-back direction, the two ends of the shell are the small-diameter end and the large-diameter end respectively in the axis direction of the shell, and the small-diameter end of the shell is fixedly connected with the top cover; a propellant is arranged on the top cover; a plurality of protruding structures are evenly distributed on the outer side face of the shell in the circumferential direction in the axis direction of the shell. According to the technical scheme, the bottom structure of the thrust bearing type storage tank is simple in process, easy to manufacture in batches and beneficial to achieving light weight of a carrier rocket.

The utility model discloses a wall breaking device for breaking a perforation compaction layer of an oil and gas well, and relates to the technical field of petroleum perforation, the wall breaking device comprises a perforating gun, a bullet rack, a perforating bullet, an acting device and a detonating cord, the acting device is detachably installed at an opening of the perforating bullet, the head of the acting device is filled with energy particles arranged in multiple layers, and the detonating cord is connected with the perforating gun. A positioning clamping ring is installed in the shell and used for adjusting the distance between the energy particles and the blasting powder. Compared with the prior art, the multi-layer charging structure can enable particles to be fully rubbed and impacted in the pore channel, a more violent acting effect is triggered, and the pore channel without compaction pollution is generated. In addition, due to the design of the positioning clamping ring, the perforating bullet can be kept at the end away from the blind hole all the time, larger burst height is obtained for jet flow, and the technical problems that a traditional perforating hole channel is narrow and blocked, and compaction pollution exists in the hole channel are effectively solved.

The invention discloses an integrated closed circulation system for an RBCC engine, and solves the problems that a constant-pressure extrusion mode of an existing RBCC engine propellant supply system is not suitable for variable working condition adjustment of an RBCC rocket thrust chamber, and a high-pressure storage tank is huge in structure, large in weight and not beneficial to application and expansion of the system on the day. According to the system, high-temperature oxygen is generated through catalytic decomposition of hydrogen peroxide and acts on the turbine, then the hydrogen peroxide pump and the coal oil pump are driven to work, integrated pressurization, supply and adjustment of propellants can be carried out on the multiple rocket thrust chambers and the ramjet combustion chamber at the same time, and therefore wide-range fuel oil adjustment is met, and a widely-applied electric pump system can be replaced; and the structural envelope and weight are greatly reduced, so that the power requirement of a longer voyage is met.

A lighted projectile may comprise a switch and a light source. The switch may comprise a moveable portion. The projectile may receive a propellant. The projectile may fluidly couple the propellant to the switch. The projectile may comprise a cavity by which the propellant is fluidly coupled to the moveable portion of the switch. Responsive to receiving the propellant, the moveable portion of the switch may move within the projectile. Responsive to movement of the switch, the light source may be activated. Movement of the switch may close a circuit between the light source and a power supply integrated within the projectile. The moveable portion of the switch may comprise the light source. Responsive to receiving the propellant, the projectile may be deployed from a projectile launcher. The projectile may further comprise a user interface by which a reset force may be applied to the switch to deactivate the light source.

The invention discloses a double-layer artificial debonding heat insulation structure of a solid engine and a forming method, and relates to the field of solid rocket engines. A heat insulation layer, a bottom layer coating layer and a cover layer coating layer are adopted on the inner surface of a combustion chamber shell to form a double-layer manual debonding heat insulation structure, the bottom layer coating layer and the heat insulation layer form an open type manual debonding structure, and the cover layer coating layer and the bottom layer coating layer form a closed type manual debonding structure. The heat insulation layer, the bottom layer coating layer and the cover layer coating layer are sequentially paved with green sheets and are formed through one-time vulcanization. The enough bonding area of the charged powder can be guaranteed, and the stress of the grain can be fully released.

The invention discloses a liquid-propellant rocket engine turbine pump inducer design method and an inducer, and the method comprises the steps: determining the working conditions of the inducer according to the overall design requirements of a liquid-propellant rocket engine; according to design requirements and an empirical formula, basic parameters of the inducer are preliminarily determined; different core parameters are combined according to the core parameter range of the inducer, and a plurality of initial three-dimensional models of the inducer are established; respectively carrying out analogue simulation on the plurality of established initial three-dimensional models of the inducer to obtain the cavitation performance of the inducer; according to a flow analysis result, obtaining an inducer of an optimal blade profile in all parameter combinations; and parameter optimization and adjustment are conducted on the inducer with the optimal blade profile till the inducer with the optimal cavitation performance for the turbine pump of the liquid rocket engine is obtained. By means of the method, the liquid rocket engine turbine pump inducer with the good cavitation performance can be designed.

The present invention provides fluid storage and dispensing tanks and methods for storing reactive fluids, such as component fluids for two-component adhesives, foams, and sealants. The tanks contain a reactive fluid for storage until the fluid is desired to be dispensed. The tank includes a subject fluid storage chamber that contains the subject fluid independent of any propellants or other fluids or gases. When the tank is desired to be used, a propellant, such as a volatile propellant or a compressed gas (e.g. compressed air) are introduced directly into or around the storage chamber to facilitate dispensing the subject fluid from the tank. The tank may include a propellant storage chamber inside the canister body and adjacent to the subject fluid storage chamber. The propellant chamber is selectively punctured when the tank is to be used. The tank may include a flexible bladder defining the storage chamber.

There is disclosed a rocket engine arranged to deliver one or more propellants to a combustion chamber of the rocket engine for generating motive thrust. The rocket engine comprises an electrical transmission system, a turbine generator unit arranged to receive a propellant flow through the turbine generator unit and to generate electrical power and to pass the generated electrical power to the electrical transmission system, and a motor pump unit arranged to receive electrical power from the electrical transmission system and to use the received electrical power to pump a propellant flow though the motor pump unit.

The embodiment of the invention provides a rocket vertical recovery method and system and a rocket, and the method comprises the steps: after a first-stage rocket is separated, the thrust of a central engine of the first-stage rocket is adjusted to a preset proportion of rated thrust so as to provide positive overload for a propellant in a storage tank; the propellant can be gathered to an outlet in the bottom of the storage tank through positive overload; the flying direction of the first-stage rocket deviates from the current flying direction by performing bidirectional swinging on the central engine, and the posture of downward diving of the first-stage rocket towards the earth direction is kept; tangential force in two different directions tangent to the current flight direction of the first-stage rocket can be generated at the same time by swinging the center engine in two directions, the posture of the first-stage rocket is adjusted to the upward flight posture from the downward diving posture, and speed reduction is conducted till the first-stage rocket lands on the earth. The thrust of the center engine of the first-stage rocket is adjusted to be small, and attitude control and propellant management of first-stage rocket recovery are achieved.

In a MSCJ warhead detonation of the main charge is controlled to provide elevated pressure at multiple locations on the back surface of the liner to cut the liner and to form and propel forward a plurality of SCJs. An initiation system is configured for multi-point initiation of a plurality of booster charges to detonate the main charge to produce a plurality of detonation waves that constructively interfere at multiple locations on the back surface of the liner to form pressure hot spots that cut the liner and to form and propel forward a plurality of SCJs. In different embodiments, the elevated pressures are between 110% and 200% of the detonation pressure at the front of an individual detonation wave. The liner may, for example, include a plurality of recesses such as shallow dimples or deeper conical structures in which case the boosters are aligned to the center of the recessed structures.

The invention provides a conventional and detonation dual-mode spontaneous combustion propellant space engine, and belongs to the technical field of spacecraft propulsion systems. The conventional and detonation dual-mode spontaneous combustion propellant space engine comprises a propellant control valve, an engine head, an engine body and an external excitation device. The propellant control valve, the engine head part and the engine body part are connected in sequence, and the external excitation device is embedded into the outer wall of the engine body part; according to the invention, by adopting an external excitation mode, the space engine in a detonation mode is realized; by means of a selective excitation mode, the space engine capable of selecting a conventional mode and a detonation mode is achieved. The detonation combustion is applied to the spontaneous combustion propellant space engine, the upper limit of the combustion efficiency of the space engine of a constant-pressure thermodynamic cycle is broken through, and the combustion efficiency is further improved on the original basis.

The invention provides a solid propellant combustion performance prediction method based on an AI model, and the method comprises the following steps: 1, carrying out the multi-modal parameter coupling design: carrying out the multi-modal parameter coupling design of a solid propellant, determining the combustion performance parameters of the propellant to be predicted and the formula component parameters influencing the combustion performance parameters, and carrying out the multi-modal parameter coupling design of the solid propellant; establishing a multi-modal parameter coupling prediction system which takes the formula component parameters as input parameters and takes the combustion performance parameters as output parameters; step 2, acquiring sample data; step 3, establishing a differentiated normalization standard; step 4, constructing a combustion performance prediction model: establishing an AI model for predicting the combustion performance of the propellant, and carrying out model training to obtain the combustion performance prediction model meeting the precision requirement; the AI model is a neural network model added with a high-voltage compensation sub-network; and step 5, predicting.

The invention discloses an experimental method of a gas production device for adjusting the water outlet posture of a navigation body, and relates to the technical field of water outlet navigation body model experiments. In order to solve the technical defect of inaccurate control of a high-pressure air source in an existing experimental method for attitude adjustment requirements under a complex water outlet working condition in the prior art, the technical scheme provided by the invention comprises the following steps: calculating and obtaining a function relationship between the sectional area of a spray throat and the internal pressure of a high-pressure air cavity; in combination with the geometric arrangement model of the propellant, deducing the rule that the gas production area of the propellant changes along with time, and establishing a gas production speed model; predicting the exhaust characteristics of the air cavity; high-pressure air cavity pressure and exhaust flow data in the actual exhaust process are collected through experiments, experimental data and prediction results are compared and analyzed, and parameters of the propellant arrangement model and the pressure prediction model are corrected according to difference feedback. The device is suitable for experiment research related to attitude adjustment and cavitation suppression in the high-speed water outlet process of the navigation body.

A rocket motor comprising a non-permeable outer shell, an intra-wall channel between the non-permeable outer shell and an annular section, and a center chamber formed at least in part by a first propellant is disclosed. The first propellant is made up of a multi-grained fuel grain, and the center chamber is further formed by an annular section. The intra-wall channel is configured to permit combustion gases to flux between the non-permeable outer shell and the annular section. The annular section configured to permit combustion gases to flux through pores embedded within the annular section.

The invention relates to the technical field of computer simulation, and provides a numerical simulation method and system for the working process of an ADN-based space engine, and the method comprises the steps: collecting historical physical and chemical attribute data of an ADN-based propellant under different combustion conditions, constructing a combustion process simulation model containing a chemical reaction kinetic action relationship and a fluid mechanics action relationship; inputting initial state data of a propellant into the model, simulating an energy release process in a combustion chamber after ignition triggering, and obtaining a simulation analysis result containing a flow state and a temperature gradient influence factor; generating an energy release state spectrum of the propellant in the ignition stage according to the simulation analysis result so as to represent energy release intensity distribution at different moments; and excavating a spatio-temporal evolution rule of propellant energy release based on the energy release state map, and outputting an energy release characteristic analysis report to guide engine combustion efficiency optimization.

The invention relates to a fluid supply system and method, in particular to a precise densified cryogenic propellant supply system and method, and aims to solve the problem that the follow-up test run precision is affected due to the phenomena that an existing densified cryogenic propellant is difficult to store and uneven in propellant flow. The system comprises a main container group, a heat exchange device, a first flow meter, a second flow meter, a liquid nitrogen supply valve, a liquid nitrogen flow regulating valve, a liquid nitrogen flow regulating valve bypass valve, a nitrogen emptying valve, a second valve, a first valve, a third valve, a front filling valve, a front filling bypass valve, a rear filling valve and a first low-temperature propellant discharge valve, a first low-temperature propellant discharge valve, a second low-temperature propellant discharge valve, a first temperature sensor, a second temperature sensor, a third temperature sensor and a fourth temperature sensor. The heat exchange temperature of the low-temperature propellant can be dynamically adjusted, and subsequent storage and uniform and stable supply are facilitated.

An air rifle device for using a cylinder of carbon dioxide as a propellant includes a housing having a front end, a rear end, a top side, and a bottom side defining a chamber extending through the housing. A barrel coupled to the front end is in fluid communication with the chamber. A clip removably couplable to the housing holds a plurality of projectiles. A canister removably mounted to the housing holds a propellant. The canister comprises a neck and a body. A collar configured to receive the neck is mounted to the housing. A cradle configured to receive the body is coupled to the housing. The cradle is aligned with the collar to facilitate coupling the canister to the collar. A trigger mechanism mounted to the housing expels a quantity of the propellant into the barrel sufficient to propel one of the plurality of projectiles outwardly through the barrel.

The combustion turbulence device comprises a speed-adjustable fan system, a speed-adjustable fan system placing platform, a porous air distribution plate and a particle filtering device, the speed-adjustable fan system placing platform is placed above the particle filtering device; a porous air distribution plate is arranged at the upper part of the speed-adjustable fan system placing platform; the speed-adjustable fan system placing platform comprises a cylinder, a circle of annular positioning boss is arranged at a cylinder opening of the lower end of the cylinder, and an installation disc is arranged on the outer wall of the lower end of the cylinder and installed on a lower flange end cover through a supporting rod and a platform fixing nut. A groove is formed in the upper end opening of the cylinder; the porous air distribution plate comprises an annular disc, a medicine strip clamp small cylinder is arranged on the inner ring of the annular disc, a plurality of uniform gas circulation holes are distributed in the annular disc, and a boss is arranged on the edge of the annular disc. The device adsorbs condensed phase particles generated by combustion of solid propellant medicine strips, clean internal circulation flowing gas is provided in the high-pressure combustion chamber, and the problem that smoke particles are unfavorable for observation of the combustion phenomenon when the solid propellant is combusted can be effectively solved.

The present invention relates to a spacecraft electric thruster particularly suitable for both propellant-fed and air-breathing electric space propulsion comprising an ionization chamber (11) for plasma generation which develops along a central development axis (X), wherein the ionization chamber (11) has a chamber inlet (12) configured to allow gas inflow and a chamber outlet (13) configured to allow extraction of ionized gas particles; a power injection assembly (30) configured to perform electron heating for plasma generation inside the ionization chamber (11); a magnetic assembly (20) configured to generate a magnetic field for plasma confinement in the ionization chamber (11); and an ion extraction and acceleration assembly (40) configured to extract ionized gas particles from the ionization chamber (11) through the chamber outlet (13) and accelerate them along an ion acceleration direction (Z), wherein the power injection assembly (30) comprises a microwave power delivery system (31) configured to deliver an electric field oscillating at a selected microwave frequency (f) and to inject the generated oscillating electric field in the ionization chamber (11), and wherein the ionization chamber (11) comprises a plurality of walls (11a, 15,16,41) closed to the passage of microwaves apart from a portion of at least one lateral wall (11a) of the plurality of walls (11a, 15,16,41) at which the power delivery system (31) is coupled to inject the generated oscillating electric field in the ionization chamber (11), and is resonant for electromagnetic fields oscillating at the selected microwave frequency.

The invention provides a spacecraft attitude control method, a spacecraft attitude control device and a spacecraft, and the spacecraft attitude control method comprises the following steps: calculating a pitch attitude angular acceleration at a starting moment of designing a pitch attitude angular velocity instruction by adopting a parabola attitude adjustment mode; determining a pitch attitude angular velocity instruction under each working condition according to the relationship between a pitch attitude angular velocity target value at the satellite-rocket separation moment and 0 and the magnitude relationship between the pitch attitude angular velocity target value and the pitch attitude angular velocity at the attitude adjustment starting moment of the pitch attitude angular velocity instruction; and tracking the pitch attitude angular velocity instruction of the current control period, and controlling the angular velocity of the yaw attitude and the angular velocity of the rolling attitude to be 0 degree / s so as to realize real-time control of the flight attitude of the carrier rocket. The reliability of the carrier rocket can be improved, meanwhile, the propellant consumption of the auxiliary power system is reduced, and the carrying capacity of the carrier rocket is improved.

The invention discloses a mixed time-varying reliability analysis method for a solid rocket engine shell structure. Uncertain parameters and distribution are determined according to complex characteristics such as high-temperature and high-pressure time-varying load borne by a solid rocket engine shell structure, geometric sudden change in a star hole charging combustion stage, a throat ablation non-stationary interval process and a time-varying correction coefficient random process, and the random and interval processes are processed in a unified mode through equivalent time-varying uncertainty conversion. Input dimension explosion is avoided, specifically, a random process vector and an interval process vector are converted into an equivalent random vector, so that a time-varying limit state function is equivalently converted into a static function, the input dimension of the RBF combination model is effectively prevented from being increased, the advantages of multiple basis functions are fused through the radial basis function RBF combination model, the calculation cost is reduced, and the calculation efficiency is improved. The modeling precision of the strong nonlinear limit state function of the solid rocket engine shell is improved, the nonlinear mixed time-varying reliability analysis problem is efficiently solved, and the calculation precision is high.

The invention relates to a jet pump (100) for conveying a suction fluid (242) in a refrigerant circuit (200) of a temperature control system using a propellant fluid (231) and utilizing the Bernoulli effect, wherein the jet pump (100) has a primary nozzle (103) for accelerating the propellant fluid, a secondary connection interface (102) for supplying the suction fluid, a mixing tube (104) downstream of the primary nozzle (103) and downstream of the secondary connection interface (102), and a diffuser (105) downstream of the mixing tube (104), wherein in an end section (140) of the mixing tube (104), which is located at an end of the mixing tube (104) furthest from the primary nozzle (103) and closer to the diffuser (105), a thrust edge (130) is provided which is configured to act in a supersonic The flow in the mixing tube (104) is intended to trigger a compression pulse in the final section (140). Furthermore, a refrigerant circuit (200) with such a jet pump (100) is proposed.

An apparatus and methods are provided for a suppressor to be coupled with a muzzle end of a barrel of a firearm to reduce muzzle blast and muzzle flash. The suppressor comprises a housing having a proximal end and a distal end. A front portion within the housing comprises a series of cylindrical gas expansion chambers for attenuating the temperature and energy of propellant gases accompanying a projectile fired from the firearm. An annular gas expansion chamber surrounds the cylindrical gas expansion chambers and directs a portion of the propellant gases from a rear portion of the suppressor to peripheral vents disposed at the distal end. Lateral chambers within the rear portion deflect and rebound a portion of the propellant gases before passing them into the annular gas expansion chamber. Ledges within the annular gas expansion chamber direct the propellant gases distally through suppressor toward the peripheral vents.

The invention relates to the technical field of deep space exploration, in particular to a cryogenic propellant on-orbit low-loss storage system and method for a deep space probe. After the low-temperature propellant is injected into the storage tank through the injection pipe, along with continuous accumulation of external leaked heat, the first pressure sensor detects that the pressure in the storage tank is continuously increased, and when the numerical value of the first pressure sensor exceeds the preset pressure, the cooling assembly is started to cool the low-temperature propellant in the storage tank, so that the external leaked heat is consumed; before the engine needs to supply the low-temperature propellant, the low-temperature propellant in the storage tank is in a supercritical state, and the storage temperature is increased in a high-pressure storage mode, so that the temperature difference between the storage tank and the outside and the heat leakage amount are reduced, and the requirement for the refrigerating capacity of the cooling assembly is reduced; when the low-temperature propellant is conveyed, due to the fact that the low-temperature propellant is in a supercritical state and free of gas-liquid phase separation, the interior of the whole storage box is in a homogeneous state, the stable conveying flow of the low-temperature propellant can be guaranteed, and stable output of an engine is also facilitated.

Carbide-based fuel assembly includes outer structural member of ceramic matrix composite material, the interior surface of which is lined in higher temperature regions with an insulation layer of porous refractory ceramic material. A continuous insulation layer extends the length of the fuel assembly or separate insulation layer sections have a thickness increasing step-wise along the length of the fuel assembly from upper (inlet) section towards bottom (outlet) section. Fuel element positioned inward of the insulation layer and between support meshes has a fuel composition including HALEU and the form of a plurality of individual elongated fuel bodies or one or more fuel monolith bodies containing coolant flow channels. Fuel assemblies are distributively arranged in a moderator block, with upper end of the outer structural member attached to an inlet for propellant and lower end of the outer structural member operatively interfaced with a nozzle forming a nuclear thermal propulsion reactor.