1242 results about "Propellant" patented technology

The invention discloses a parabola type explosion action and static action dual-energy-gathering device and a blasting method, and relates to the technical fields of directional fracture control blasting, building demolition and the like in mine and tunnel engineering. The device comprises a main body pipe which is a hollow tubular circular ring body and forms a closed channel along the annular direction; the energy-gathered jet flow groove is formed in the outer side wall of the main body pipe and distributed in the axial direction, the cross section contour of the energy-gathered jet flow groove is in a parabola shape, and the energy-gathered jet flow groove is used for guiding explosion stress waves to be intensively propagated and released in the preset direction; the energy gathering seam is formed in the bottom of the energy gathering jet flow groove and penetrates through the inner wall of the main body pipe; and the explosive loading cavity is used for fixedly placing explosives. By combining the dual energy gathering principle of explosion action and static action, energy is concentrated in the energy gathering direction in the explosion process, so that explosion energy is accurately released along a parabola-shaped track, the explosion wave focusing effect is enhanced conveniently, and invalid energy loss is reduced.

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 discloses a free filling type metal wire embedded solid propellant forming device and process, and belongs to the field of solid propellants, the free filling type metal wire embedded solid propellant forming device comprises a bottom plate, a base, a mold cylinder and a top cover are sequentially arranged on the bottom plate, and a plurality of pressing structures are arranged on the upper portion of the top cover; a solid propellant coating barrel is arranged on the inner wall of the mold barrel; the lower end of the metal wire is fixed on the bottom plate, the upper end of the metal wire is fixed on the first fastener of the fastener mounting block on the top cover, and the first fastener is screwed into the threaded hole, so that the metal wire can be tensioned, and the internal stress of the metal wire can be adjusted by adjusting the expansion amount of the metal wire; according to the invention, the collimation of the metal wire in the propellant is ensured, the proper stress of the metal wire in the slurry casting process is also ensured, the problem that the metal wire is broken due to overlarge stress is avoided, and the technical problem that the yield of the solid propellant is low due to the fact that the metal wire is easy to break in the propellant in the prior art is solved.

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 invention provides an injector based on a liquid oxygen methane two-component propellant and an attitude control engine, and the injector comprises an injector body, a centrifugal nozzle assembly and a spark plug; the centrifugal nozzle assembly is arranged on the injector body, and the injector body is provided with a liquid collecting cavity; a propellant in the storage box is guided to the liquid collecting cavity and is sprayed to the spraying pipe through the centrifugal nozzle assembly; the sparking plug is arranged on the injector body, and an ignition part of the sparking plug is arranged close to an outlet of the centrifugal nozzle assembly so as to ignite the propellant sprayed into the spray pipe by the centrifugal nozzle assembly. The injector is more compact in structure, and the working reliability of the injector can be improved.

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 a fire extinguishing water spraying intensity detection method for propellant fire in the field of propellant fire extinguishing. A superspeed deluge system, an ignition system, a charging device and a weighing device are included. The ultra-high-speed deluge system comprises a high-speed detector, a control system, a quick opening valve, a sprinkler head, a water supply pipe, a pressure water supply device and an air supply device; the ignition system comprises a resistance wire, an ignition circuit and an ignition device.The fire extinguishing water spraying intensity detection method for the propellant fire is provided for solving the problems that the propellant fire cannot be extinguished through the water spraying intensity of an existing deluge system, and an ultra-high-speed deluge system cannot be effectively applied to a propellant production site. The problem that an existing propellant powder production place does not have a key parameter of fire extinguishing water spraying intensity is solved, effective application of the ultra-high-speed rain extinguishing technology in the propellant powder production place is achieved, and safety of production personnel and equipment is protected.

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.

The present invention describes a safe-and-arm device, mechanism or assembly (160) for a projectile (100,102) equipped with a second-stage propellant (124). The second-stage propellant (124) is ignited after the projectile has been ejected out from a launcher barrel. A lockpin (190), being urged by a spring (194), is responsive to ignition of the second-stage propellant; after the lockpin (190) and spring (194) sense and respond to the second propulsion, an unbalanced rotor (164) forming part of the safe-and-arm device or assembly (160), is released to rotate from a “safe” state to an “armed” state, as the projectile continues traveling along its trajectory to a target. In one embodiment, the projectile is configured with a 40 mm cartridge containing a first propellant (122). An impact sensor may trigger an electric detonator or a point detonator may trigger a stab detonator to set off explosives disposed in the projectile.

The invention relates to the technical field of low-temperature propellant filling, in particular to a liquid oxygen propellant filling system and a pre-cooling pump starting method thereof, and provides the liquid oxygen propellant filling system which cancels a filling bypass, reduces the bypass leakage risk, enables the whole filling system to be more simplified, and improves the filling efficiency. The invention further provides a pre-cooling pump starting method of the liquid oxygen propellant filling system. The pre-cooling pump starting method comprises the following steps that the pipeline is pre-cooled, the filling valve at the liquid oxygen storage tank end is opened, the filling valve at the rocket liquid oxygen storage tank end is closed, the exhaust valve is opened, liquid oxygen flows through the pipeline, and the pipeline is pre-cooled. Pre-cooling the storage tank, namely pre-cooling the storage tank after the pipeline is pre-cooled; starting the pump for pre-cooling, starting the liquid oxygen filling pump at low frequency, and gradually increasing the frequency of the liquid oxygen filling pump. Storage tank filling is carried out through the high-position pump and the low-position pump, liquid oxygen does not need to be extruded into the storage tank through bypass pressure increasing, the precooling flow process does not need to be interrupted, procedure conversion of liquid oxygen precooling and large-flow filling is not needed, the operation difficulty is reduced, and the filling time is shortened.

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 relates to a combined type pre-cooling system for a ground ejection and pressurization coupled liquid rocket basic-stage engine. The combined type pre-cooling system comprises a low-temperature propellant storage tank, a storage tank exhaust valve, an oxygen conveying pipeline, an oxygen backflow pipeline, an ejector, a pre-cooling discharge valve, an oxygen discharge pipeline, a pre-cooling discharge valve control pipeline, an ejection pressurization pipeline, an oxygen discharge pipeline and an exhaust valve control pipeline. The low-temperature propellant storage tank is communicated with the storage tank exhaust valve, and the bottom end of the low-temperature propellant storage tank is communicated with the oxygen conveying pipeline; the oxygen conveying pipeline is communicated with the oxygen conveying pipeline; the precooling discharge valve is communicated with the engine and the ejector; the ejector is communicated with the oxygen return pipeline; the oxygen return pipeline is communicated with the low-temperature propellant storage tank; the oxygen discharge pipeline is communicated with the pre-cooling discharge valve; the pre-cooling discharge valve control pipeline is communicated with the pre-cooling discharge valve control air port; the ejection pressurizing pipeline is communicated with the ejector; the oxygen exhaust pipeline is communicated with the storage tank exhaust valve; and the exhaust valve control pipeline is communicated with the storage box exhaust valve control air port. According to the application, random switching between circulation and discharge pre-cooling can be realized.

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 belongs to a flow control method, and provides an electric propulsion flow control method and system aiming at the technical problems that a flow control unit in an existing electric propulsion system is long in reaction time, small in flow adjusting range, complex in structure and high in control method complexity. Wherein two branches serve as an anode and a cathode of the thruster, and each branch performs closed-loop control on the electric propulsion flow according to the real-time pressure detection results of the first pressure detection piece and the second pressure detection piece in combination with the pressure reduction degree of the propellant, the parameters of the first electromagnetic valve and the parameters of the second electromagnetic valve. The control accuracy is high, the flow adjusting range is wide, and light weight is achieved.

Solar thermal and chemical hybrid rocket configurations for mining and other space applications are disclosed. One aspect is a rocket propulsion system configured to provide rocket thrust, including a solar absorber, a rocket nozzle, and a solar power collection system configured to collect solar energy from the sun, generate an energy beam from the collected sunlight, heat the solar absorber to transfer heat to one or more pressurized propulsive gases, and expel the heated pressurized propulsive gases through a rocket nozzle. A solar absorber can be formed from a granular collection or agglomeration of solids (e.g., of beads), which can be layered with more transparent layer(s) above and more absorbing layer(s) below to create a temperature profile in propellant(s) flowing through the absorber. A hybrid motor can provide an energy (e.g., solar) absorber for absorbing and transferring radiative energy as well as a combustion area. Multiple propellants can be present in a single chamber and be forced from a nozzle to produce thrust. Pressure in a rocket can be achieved from heating inert gasses, and alternatively or simultaneously, from mixing and igniting non-inert gasses.

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 channel type flow guide plate based on an inverted-cone-shaped pore plate structure. The channel type flow guide plate comprises a capillary porous plate and a groove structure. And the channel type guide plate is connected with a liquid accumulator at the bottom of the storage box through a bracket. The channel type flow guide plate is parallel to the inner wall face of the storage box and extends to the top of the storage box along the inner wall face. The cross section of the channel type guide plate is of a trapezoidal structure, the channel type guide plate is hollow, and a closed channel is formed for propellant flowing. A capillary perforated plate is arranged on the side, close to the wall face of the storage box, of the channel type flow guide plate, a plurality of micron-sized inverted-cone-shaped capillary pores are formed in the capillary perforated plate in a single-side laser drilling mode, and the functions of storing liquid and serving as a gas barrier can be achieved. The channel type guide plate is suitable for a satellite propulsion system with strong maneuverability and high acceleration, has the advantages of simple structure, portability, high reliability, strong flow conductivity and the like, and can provide guidance for the development of a third-generation surface tension storage tank PMD.

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 provides a combustion analog simulation analysis method and system based on an ADN-based liquid propellant, and the method comprises the steps: carrying out the prediction of combustion parameters through employing a joint learning network, obtaining X single-parameter prediction networks during the network adjustment, and carrying out the adjustment of first reasoning layers corresponding to assembly sets in the single-parameter prediction networks. Fusing the component sets at the a level in the X single-parameter prediction networks to obtain a combustion parameter prediction network, and setting X second reasoning layers for the a fused component set; taking a real combustion parameter corresponding to combustion condition training data of the ADN-based liquid propellant sample as a prior mark, and taking a predicted combustion parameter of the first reasoning layer as a prior mark of a predicted combustion parameter of the second reasoning layer; the prediction combustion parameter of the single-parameter prediction network is used as a priori mark to adjust the combustion parameter prediction network, so that the precision of combustion parameter prediction of the combustion parameter prediction network can be improved, and the precision of combustion analog simulation is improved.