1426 results about "Propellant" patented technology

Systems and methods are disclosed for mining lunar and Martian polar permafrost to extract gas propellants. The method can comprise identifying a plurality of near-polar landing sites in craters in which the surface comprises permafrost in perpetual darkness, wherein such landing sites have perpetual sunlight available at altitudes of about 100 to 200 m. A mining outpost can be established in at least one of the sites and a high altitude solar array deployed at the landing site using a lightweight mast tall enough to generate near continuous power for the outpost. Systems and apparatus are disclosed for mining the permafrost at the landing sites using radiant gas dynamic mining procedures. The systems can comprise a rover vehicle with an integrated large area dome for cryotrapping gases released from the surface and multi-wavelength radiant heating systems to provide adjustable heating as a function of depth.

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.

The invention relates to a dynamic modeling method for an attitude control system under a dynamic fault. The method comprises the following steps: determining a modeling hypothesis of a nonlinear dynamic model; establishing a coordinate system used for modeling; rocket configuration parameters and engine related parameters are obtained; building displacement and acceleration descriptions of the three types of structures on the rocket; establishing a rigid body translation equation, a rigid body rotation equation, a rocket body elastic vibration equation, a propellant shaking equation and a sensitive element measurement equation; summarizing the five established equations to form a nonlinear dynamic equation set; and performing small-disturbance linearization expansion on the nonlinear dynamic equation set on the basis of the nominal reconstructed trajectory to obtain a linearized small-deviation equation set suitable for controller design and stability analysis. According to the method, the influence of the thrust drop fault is comprehensively considered, and the model can be used for carrying out relatively accurate dynamic response analysis and reconstruction scheme design under the fault.

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.

An electric pump pressure type liquid oxygen methane rocket main and auxiliary power integrated system adopts an electric pump to pressurize a low-pressure propellant in a rocket storage tank, the pressurized propellant can be controlled to enter an attitude and orbit control engine or return to the rocket storage tank through reversing of a reversing valve, and if the reversing valve continuously supplies the propellant into the attitude and orbit control engine, the propellant can enter the attitude and orbit control engine; continuous and stable work of the attitude and orbit control engine can be realized; if the reversing valve supplies propellants into the attitude and orbit control engine and then immediately switches to the rocket storage tank, pulse work of the attitude and orbit control engine can be achieved through circulation. The reversing valve can keep the electric pump in a normal circulation state at the downstream moment, and the electric pump can work normally and stably no matter whether the attitude and orbit control engine works stably or in a pulse mode. The cavitation pipe is installed in front of the reversing valve, when the attitude and orbit control engine works in a pulse mode, the rotating speed and the power of the electric pump are kept unchanged, the pressure and the flow behind the pump can be constant, the flow entering the attitude and orbit control engine is constant, and it can be guaranteed that thrust generated by the attitude and orbit control engine in each pulse work is constant.

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.

Omnivorous solar thermal thrusters and adjustable cooling structures are disclosed. In one aspect, a solar thermal rocket engine includes a solar thermal thruster configured to receive solar energy and one or more propellants, and heat the one or more propellants using the solar energy to generate thrust. The solar thermal thruster is further configured to use a plurality of different propellant types, either singly or in combination simultaneously. The solar thermal thruster is further configured to use the one or more propellants in both liquid and gaseous states. Related structures can include valves and variable-geometry cooling channels in thermal contact with a thruster wall.

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 invention provides a full-size low-temperature medium outflow test system and an outflow test method. The system comprises a control system, an end socket, a storage tank, a conveying main pipeline, a pressure accumulator, a discharge pipeline, a filling pipeline, an air sealing pipeline, a pressurizing pipeline, a pressurizing tank, a liquid nitrogen filling tank car, a flow meter, a filling port, a regulating valve, a stop valve and an air sealing port. According to the system, a full-size test platform is built, low-temperature liquid nitrogen is used for replacing a low-temperature propellant, the working states of the pressurized conveying system under different flows and pressures are simulated, and the requirement that the system obtains the flowing performance closer to the real working condition through test examination under a low-temperature medium is met.

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 invention discloses a coaxial rotor gun firework launching unmanned aerial vehicle system which comprises a missile-borne shell, an unmanned aerial vehicle and a foldable firework guiding launching system. The missile-borne shell mainly comprises a missile body, a missile tail and a hood; the projectile body and the projectile tail are connected through a fixing mechanism, a shell ejection mechanism is arranged in the fixing mechanism, and shell ejection and separation can be achieved when a separation signal is received. An integral structure formed by the unmanned aerial vehicle and the foldable firework guide launching system is placed in the bomb body in a folded state; the unmanned aerial vehicle is composed of a foldable nose and a fuselage. The foldable firework guiding launching system is composed of a guiding firework launching device and an airborne firework launching frame, the airborne firework launching frame comprises a firework launching tube and a foldable cross-shaped connecting rod, the firework launching tube is connected with the unmanned aerial vehicle through the foldable cross-shaped connecting rod, firework projectiles and propellant powder are contained in the firework launching tube, and the unmanned aerial vehicle carries the firework launching tube to the sky. Remote control is achieved, and fireworks are ignited and launched in an electric ignition mode.

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 a low-temperature propellant filling system for a rocket core-level power system test. The low-temperature propellant filling system comprises a propellant supply subsystem, a ground pipeline precooling and discharging subsystem, a filling port discharging subsystem and a filling port deflating subsystem. The propellant supply subsystem is connected with the rocket body storage tank through a propellant supply pipeline of the propellant supply subsystem, and low-temperature propellant filling can be achieved; the ground pipeline pre-cooling discharge subsystem, the filling port discharge subsystem and the filling port deflation subsystem are respectively and independently connected with a propellant supply pipeline; and the connecting points of the filling port discharge subsystem and the filling port deflation subsystem on the propellant supply pipeline are respectively positioned on the propellant supply pipeline between the ground pipeline pre-cooling discharge subsystem and the rocket body storage tank. The device can prevent a large amount of low-temperature steam from entering the storage box to cause back pressure in the ground filling process, can adapt to impact vibration in the rocket body filling deformation and test ignition stage, and completes remote control to safely discharge the residual propellant of the rocket body, so that the filling index is met.

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.

A method includes: during a first time period, pumping propellant: to a set of moderator coolant channels to cool the moderator; through a set of fuel elements to heat the propellant to a first outlet temperature; out of a thrust nozzle to produce a first thrust; and locating the reflector at a first closed position to reflect neutrons toward the set of fuel elements to maintain a fission rate of the set of fuel elements above a threshold fission rate. Additionally, during a second time period: locating the reflector at a second open position to reduce the fission rate of the set of fuel elements to below the threshold fission rate; and pumping the propellant through the fuel elements to heat the propellant to a second outlet temperature less than the first outlet temperature and out of the thrust nozzle to produce a second thrust less than the first thrust.

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.