472 results about "Propellant" patented technology

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 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.

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.

The invention relates to a spaceflight electric propulsion integrated pressure closed-loop proportional storage and supply system and a control method thereof. The system comprises a flow channel integration body, wherein a fluid channel used for containing a propellant is formed in the flow channel integration body; the multiple functional components are detachably installed on the flow channel integration body and communicate with the fluid channel, and each functional component comprises a first pressure sensor configured to sense the pressure of the propellant, a second pressure sensor configured to sense the pressure of the propellant, and a third pressure sensor configured to sense the pressure of the propellant; and at least one proportional valve configured to adjust the pressure of the propellant; and the control circuit is configured to drive the proportional valve according to a signal of the first pressure sensor so as to carry out closed-loop proportional control on the pressure of the propellant. Control hardware and software are integrated in the storage and supply system, real plug and play is achieved, different electric propulsion system users do not need to redesign external control circuits and software for adapting to the system, and the universality, the response speed and the integration degree of the system are greatly improved.

The invention relates to a zero-loss methane propellant supercooling filling system and method, and the system comprises a liquid nitrogen supply unit, a methane supply unit, a supercooling methane storage tank, a heat exchange pipeline unit and a temperature control unit; the heat exchange pipeline unit comprises a first heat exchanger, a second heat exchanger, a third heat exchanger and a separation and recovery module; the liquid nitrogen supply unit is connected with the cold end of the first heat exchanger. The cold ends of the first heat exchanger, the second heat exchanger and the third heat exchanger are sequentially connected in series. Methane output by the methane supply unit exchanges heat with the third heat exchanger and then is input into the supercooled methane storage tank; methane output by the exhaust end of the supercooled methane storage tank exchanges heat with the second heat exchanger and then is input into the separation and recovery module; and the separation and recovery module is connected with the supercooled methane storage tank and the methane supply unit and is used for converging the recovered methane into the supercooled methane storage tank and / or returning the recovered methane to the methane supply unit. According to the method, freezing in methane supercooling filling is effectively avoided, and lossless filling is achieved.

The invention belongs to the technical field of solid propellants, and particularly relates to a propellant containing active metal fuel and a preparation method of the propellant. 3.0-4.0% of a plasticizer; 0.3-0.4% of a curing agent; 0.1%-2.0% of a combustion rate regulator; 0.01%-0.5% of a bonding agent; 15-19% of an active metal fuel; 10-18% of a nitramine explosive; the metal fuel is an aluminum-gallium metal alloy, and the aluminum-gallium metal alloy is an aluminum-gallium metal alloy of which the outer layer is coated with an aluminum film. The propellant provided by the invention has the advantage of high combustion efficiency, and can meet the requirement of a solid rocket engine on the high energy performance of the propellant.

The invention discloses an intermittent emission online control system of a low-temperature engine. The intermittent emission online control system comprises a control gas cylinder, an electric gas valve and a pre-cooling discharge valve which are sequentially connected through a pipeline, the precooling discharge valve is respectively communicated with the electric gas valve, the liquid oxygen main valve, the rocket liquid oxygen storage tank and a discharge pipeline through pipelines; the system further comprises a rocket-borne computer and a temperature sensor, and the temperature sensor is installed on a low-temperature pipeline connected with the engine. The temperature sensor is connected with the rocket-borne computer through a measuring cable; and the electric air valve is connected with the rocket-borne computer through a control cable. The invention further discloses a low-temperature engine intermittent emission online control method. According to the method, through real-time monitoring and dynamic optimization of the intermittent emission time sequence, the adaptability of an existing emission control method to different flight tasks is improved, the propellant emission consumption is reduced, and the success rate and economical efficiency of the rocket flight tasks are improved.

The present invention relates to a missile (10) with a firing device (20) having several barrels (11) extending in the direction of a longitudinal axis (A) of the missile (10) and arranged inclined or skew relative to the longitudinal axis (A), wherein projectiles (12) provided with propellant charges (13) for ejection from the barrels (11) can be loaded into the barrels (20) and a method for operating such a missile (10).

The application provides a small satellite platform structure which carries a large amount of propellant and is connected and separated in point mode, belongs to the technical field of spacecraft structure, and comprises a frame, a bottom plate, a lower cabin partition plate, a storage tank support rod, a middle layer plate, an upper cabin partition plate, a top plate and a side plate, so that the side plate surrounds the parallelly arranged bottom plate, middle layer plate and top plate into a closed box body, the bottom plate is connected to the frame, the storage tanks are arranged diagonally in the round holes on the bottom plate and the middle layer plate, and the storage tank support rods are uniformly distributed around the round holes; the middle layer plate separates the box body into an upper cabin and a lower cabin, the upper cabin partition plate is connected in the upper cabin, and the lower cabin partition plate is connected in the lower cabin. The application proposes a configuration of "point connection + storage tank sinking", meets the demand of the small satellite platform for carrying a large amount of propellant, effectively reduces the whole satellite center of mass, and improves the satellite dynamic environment; meanwhile, the bottom main load-bearing frame cooperates with a detachable adapter block, can adapt to different mechanical interfaces of launch vehicles, and realizes point connection and separation.

This invention relates to an on-orbit high-density storage and supply device and method for electric propulsion working propellant, belonging to the field of spacecraft deep space exploration and on-orbit servicing technology, which can effectively save space and mass resources of spacecraft. The cooling circuit consists of a storage tank, a refrigerator, a pump, and an injection pipe connected sequentially; the pressurization circuit consists of a storage tank, a heating evaporator, and a third solenoid valve connected in a loop; and the gaseous working propellant circuit consists of a storage tank, a vaporizer, and a buffer chamber connected sequentially. This invention simplifies the complexity of the propulsion system and reduces potential failure risks, effectively supporting the application of high-power, high-specific-impulse, and long-life new electric propulsion technologies, providing important support for deep space exploration and on-orbit servicing missions.

An ejection device (100) for a gas generator, the ejection device (100) for ejecting a combustion product, comprising: a detonator (101) with an ignition charge and an ignition area (102) designed to burn the ignition charge; a pyrotechnic device (121) arranged to be burned by burning the ignition charge; a cup (104) made of metal, wherein the cup (104) has a base (107) and a peripheral wall (105), wherein the bottom surface (107) is arranged opposite the ignition area (102), the peripheral wall (105) adjoins a peripheral edge of the bottom surface (107) and is arranged to surround the ignition area (102), wherein the cup (104) further comprises a receiving space (120) which is defined by the peripheral wall (105) and the floor area (107) is formed and receives the pyrotechnic agent (121), wherein the bottom surface (107) of the cup (104) contains a through-hole (108) through which the receiving chamber (120) and an exterior area of ​​the cup (104) are connected, wherein the through-hole (108) is formed by several peripheral through-hole edges (108a, 108b, 108c, 108d), and the adjacent peripheral through-hole edges (108a, 108b, 108c, 108d) are connected to each other at a predetermined connection point, the through-bore (108) contains a narrow part (115) in which the width of the through-bore (108), formed by each of parts of the two peripheral through-bore edges (108a, 108b, 108c, 108d) of the several peripheral through-bore edges (108a, 108b, 108c, 108d), gradually decreases with increasing distance from a center of the bottom surface (107) to the peripheral edge of the bottom surface (107), and the narrow part (115) is closer to the peripheral edge of the bottom surface (107) than a non-narrow part (125) formed by each of the other parts of the two peripheral through-bore edges (108a, 108b, 108c, 108d), and the narrow part (115) is a predetermined part in which stress due to the emission of the combustion product is concentrated, and the bottom surface (107) is designed to be split starting from the narrow part (115), and the peripheral wall (105) is designed to be split.

The invention relates to a low-temperature propellant temperature control method based on real-time pressure feedback of a storage tank on a rocket, which comprises the following steps of: acquiring a pressure value, a temperature value and flow of the storage tank in real time to determine whether the pressure value is within a set pressure range; the pressure value of the storage box is adjusted by executing control strategies such as adjusting the temperature set value, and the first set temperature is obtained; when the pressure value is within a set pressure range, determining an optimization target of the state prediction model, inputting the pressure value, the temperature value and the flow into the state prediction model, obtaining a target control variable and a second set temperature of the state prediction model when the optimization target is met, determining a target set temperature according to the first set temperature or the second set temperature, and performing state prediction according to the target set temperature. The target adjustment information including the target set temperature is obtained, and the temperature set value in the filling process is adjusted according to the target adjustment information, so that the low-temperature propellant is always in the optimal temperature state in the filling process, and the filling quality and efficiency are improved.

The invention discloses an electrolytic preheating-electrolytic ignition method for a three-electrode hydroxylamine nitrate liquid propellant, and belongs to the field of liquid rocket engines and electrochemistry. Three electrolysis electrodes are arranged in the electrolysis chamber and comprise two positive electrodes and one negative electrode, so that two groups of electrolysis electrode pairs are formed. The method comprises the following steps: injecting a nitric acid hydroxylamine-based liquid propellant into an electrolysis chamber, and electrifying a group of electrolysis electrodes to start electrolysis preheating; after the preset switching time is up, closing the current working electrolysis electrode pair, and switching to another group of non-working electrolysis electrodes to continue electrolysis; and low-voltage and rapid electrolytic ignition is realized. According to the three-electrode electrolysis structure, the time for ignition and reaction to reach the highest temperature can be shortened, and the system response speed is increased; the preheating and ignition voltage can be changed, and the large voltage adjusting and optimizing space is achieved; meanwhile, the electrolytic ignition current is relatively stable, and the instantaneous power consumption is small; and electrolysis power consumption can be saved, electrolysis electrode loss is reduced, and reuse of the electrode is facilitated.

The present invention relates to a fire suppression composition and system designed to extinguish challenging fires, such as Class B fires involving lithium batteries as well as Class A, B, and D fire scenarios. The composition includes a mineral component, a water-based liquid propellant, and a synergistic blend of additives. This innovative combination enables the composition to float on liquid surfaces and adhere to burning materials, effectively suppressing and extinguishing fires. The composition's ability to form a barrier on burning materials prevents fire spread and re-ignition, making it highly effective for both vehicle fires and industrial settings. This advancement offers a versatile and robust solution to modern fire hazards, suitable for use in various delivery systems.

This invention relates to the field of solid propellant measurement technology, specifically to an interfacial strain measurement device under temperature and pressure synergy conditions. It includes an experimental chamber with mirror-distributed sliding doors rotatably connected to it. A controller is installed on one side of each sliding door. Inside the experimental chamber are a first fixed frame and an interfacial strain measurement module electrically connected to the controller. An electric actuator electrically connected to the controller is fixed to the first fixed frame. A second fixed frame is fixed to the telescopic end of the electric actuator. A drive motor electrically connected to the controller is located on the side of the second fixed frame away from the telescopic end of the electric actuator. This invention simulates the phenomenon of a propellant gradually shortening in length during combustion by cutting the propellant grain. The interfacial strain measurement module then performs layered measurements of the propellant grain, achieving the measurement of interfacial strain under different layered conditions without burning the propellant grain.

The invention discloses a propellant fuel gas sampling system and a fuel gas composition sampling test method, and the propellant fuel gas sampling system comprises a drainage device, an absorption device and a gas collection device which are connected in sequence, or comprises a drainage device, an absorption device, a drying device and a gas collection device which are connected in sequence. The fuel gas composition sampling test method performs sampling by means of the sampling system, and comprises the following steps: S1, performing gas replacement or emptying treatment on part or all devices of the sampling system by using replacement gas before sampling; s2, fuel gas generated in the working process of the engine or in the target collection interval in the working process is collected; s3, after sampling, purging the sampling system by using replacement gas, so that all fuel gas components are sequentially absorbed, dried (if the sampling system comprises a drying device) and collected; and S4, measuring the components of the fuel gas. The device can adapt to engines with different flows to realize gas collection and component testing and improve the accuracy of test results.

An automatic uncoupled charging device includes a control system, a power system, a charging system, and an auxiliary system. The control system is equipped with a programmable circuit to control a screw speed. Explosives are charged into a container through manual handling or flowing from a mixed explosive vehicle. Explosive bags with different diameters are installed at the bottom of a discharge port. The device is set to start with one click, and the screw drives the explosive into the bag, achieving uncoupled charging. An automatic uncoupled charging method applying to the automatic uncoupled charging device not only reduces labor intensity, achieves fast charging, improves labor efficiency, but also achieves the goal of “mechanization reducing personnel”, further advancing towards the essential safety of blasting construction.

A rocket propulsion engine includes a propellant vessel, a sensitizer vessel, a combustion chamber, an exhaust nozzle, and a controllable sensitizer injector. The combustion chamber combusts combustible propellant received from the propellant vessel and the combustion gas produced therefrom is exhausted through the exhaust nozzle. When activated, the controllable sensitizer injector introduce sensitizer from the sensitizer vessel into the propellant vessel to form an explosive mixture of the combustible propellant and the sensitizer. A method of operating a rocket propulsion engine includes supplying a combustible propellant from the propellant vessel into the combustion chamber, combusting the combustible propellant within the combustion chamber, and exhausting combustion gas produced through the exhaust nozzle. The method also includes introducing the sensitizer from the sensitizer vessel into the propellant vessel to form an explosive mixture within the propellant vessel.

The invention belongs to the technical field of deep underground target penetration, and particularly relates to an energy gathering / kinetic energy series penetration structural body which comprises a charging shell, the front end of the charging shell is provided with an opening, a shaped charge liner is arranged at the opening of the front end of the charging shell in an adhesive mode, the rear end of the charging shell is fixedly connected with a rear-stage connecting part, and the rear-stage connecting part is fixedly connected with the charging shell. The rear end of the rear-stage connecting part is connected with a rear-stage main projectile body, a charging buffer layer is arranged at the front end of an inner cavity of the rear-stage main projectile body, a rear-stage fuze is installed at the rear end of the rear-stage main projectile body, and the charging shell and the rear-stage main projectile body are both filled with gunpowder. Wide-speed-range penetration is achieved through the series connection structure, the rear-stage connecting part extends into the charging shell to be in threaded connection with the charging shell, the rear-stage connecting part is protected against damage by utilizing three-direction pressure stress generated by explosive detonation, meanwhile, the length of the rear-stage connecting part can be adjusted, the rear-stage connecting part is used for protecting a penetration projectile body structure under the condition of different penetration target-hitting speeds, and the penetration target-hitting speed is improved. And therefore, the purpose of increasing the penetration power is achieved.

The embodiment of the invention relates to the technical field of refrigeration, in particular to a low-temperature propellant storage system and method. The system is characterized in that a compressor compresses helium, a first part of compressed helium flows into a circulating refrigeration assembly, and a second part of compressed helium flows into a heating assembly; the circulating refrigeration assembly is used for cooling the first part of helium, liquid hydrogen stored in the liquid hydrogen storage tank and liquid oxygen stored in the liquid oxygen storage tank are cooled through the cooled first part of helium, and the cooled first part of helium flows back to the compressor; and the heating assembly is used for heating liquid oxygen flowing out of the liquid oxygen storage tank and liquid hydrogen flowing out of the liquid hydrogen storage tank through the second part of helium to obtain gaseous oxygen and gaseous hydrogen, the gaseous oxygen and the gaseous hydrogen are conveyed to the engine, and the second part of helium flows back to the compressor after heating treatment. According to the technical scheme, high-quality cooling capacity in the storage box can be effectively utilized, and the circulating refrigeration efficiency is improved.

A gas generator comprising: - a gas chamber (10) with a first orifice (21) having a first passage surface, - a main cover (51) closing the first orifice (21), - pressurized gases stored in the gas chamber (10), - a pyrotechnic chamber (30) arranged within the gas chamber (10), - an ignition device (40), characterized in that the gas generator comprises at least one partition wall (33) of the gas chamber (10) with at least one second orifice (22) having a second passage surface smaller than the first passage surface, and in that the pyrotechnic chamber (30) is arranged between the partition wall (33) and the first orifice (21). Figure for the abbreviation: Fig. 1

An apparatus and methods are provided for a front plate having a diverging central bore for firearm sound suppressors that improves noise and flash characteristics during firing a weapon. The central bore is disposed between a back surface and a front surface of the front plate. An untapered portion of the central bore extends from the back surface to a diverging portion that opens toward the front surface. The diverging portion includes a curvature profile configured to allow for more controlled expansion of high-pressure propellant gases exiting of the suppressor through the central bore. The curvature profile provides an included angle of the central bore that decreases secondary flash events accompanying the expulsion of propellant gases accompanying a fired bullet exiting the suppressor through the central bore. The curvature profile exhibits a cross-sectional area of the central bore that is proportional to a distance along the diverging portion.

A Hall effect thruster system (12) includes a thruster body (52) and a diffuser (1) configured to apply a torque (Td) to the thruster body (52) during operation. The diffuser (1) applies the torque (Td) by ejecting the propellant in a non-axial direction, such as a direction tangent to helical or curvilinear channels (8) formed within a body (10) of the diffuser (1). The applied torque (Td) can be used to counteract a swirl torque (Tswirl) that is induced on the thruster body (52) by the ionizing Hall current flowing in an annular channel (16) of the thruster body (52). This effective counter-torque (Td) is useful in deep space applications outside the Earth's magnetic field.

A mandrel assembly (100) for net-shape casting of an undercut finocyl grain (133) with aft-end counterbore (140) in a solid propellant rocket case (101) comprising a core (105) capable of forming central cylindrical cavities in the grain and a plurality of fins (106) capable of forming a plurality of forward-swept longitudinal cavities circular patterned around the core. The method of assembling and disassembling the mandrel (100) comprises the acts of hanging the plurality of fins (106) from the core (105) using a fin hanger (107) with a plurality of long and flexible fin cables (108), hoisting the mandrel assembly (100) over the case aft-end aperture (104), bundling the fins (106) and lowering the mandrel assembly (100) inside the case (101), fastening of all the fins (106) to the core (105) and the core (105) to the case (101), inverting the total assembly, casting propellant slurry through case fore-end aperture (103), curing the propellant, inverting the total assembly back, removing core (105) first, assembling fin decoring tool (134) and removing fins (106) one by one, to obtain net-shape cast undercut finocyl grain (133) with aft-end counterbore (140) without employing squeeze casting or propellant machining.

The application discloses a kind of composite solid propellant slurry extrusion friction safety characteristic testing device, comprising: from bottom to top respectively base, operating platform, integral sensor, barrel, pillar, piston and motor;Wherein, base contains hydraulic system, for providing extrusion force load;Motor is used to provide rotary load, encoder is installed on motor, for measuring the rotational speed of motor;Barrel and piston are directly contacted with solid propellant slurry, barrel is used to fill solid propellant slurry, piston is used to apply force load to slurry;Integral sensor can measure torque and pressure, for measuring the torque and extrusion force of slurry in test process.The technical scheme of the application, the load applied can cover the load range in mixing process, the load condition of friction sensitivity test can be realized and the load condition of actual working condition of slurry is consistent.

The wire embedding tool comprises a base, a metal wire guiding unit is coaxially arranged at the top of the base, the metal wire guiding unit is coaxially sleeved with a pouring barrel, the bottom of the pouring barrel is in sealed connection with the base, and the top of the pouring barrel is connected with an upper clamping ring which is coaxially arranged; and the top of the upper clamping ring is connected with an upper disc which is coaxially arranged. The metal wire guiding unit comprises a long-tube-shaped metal wire guiding unit body, a plurality of axially-through wire passing grooves are formed in the outer side wall of the metal wire guiding unit body, the multiple wire passing grooves are evenly distributed in the circumferential direction, and metal wires are distributed in the wire passing grooves. A metal wire transition guide block is arranged at the bottom of the metal wire guide unit body, a plurality of metal wire temporary fixing columns are arranged at the top of the metal wire guide unit body, and the metal wire temporary fixing columns correspond to the wire passing grooves one to one. According to the wire embedding tool, the problem that the operation space in a powder pouring barrel corresponding to a small-diameter metal wire embedding grain is limited is solved, and the quality of the metal wire embedding grain is ensured.

The utility model relates to the technical field of on-site explosive loading, in particular to a blast hole filling length measuring device based on an on-site mixed loading ammonium nitrate fuel oil explosive truck, which is used for the mixed loading explosive truck and comprises a top spiral cylinder connected with the tail end of a top spiral of the mixed loading explosive truck. A pulley is connected to the mounting base through a pin shaft, an inner groove is formed in the periphery of the pulley, a rope slides in the inner groove, the two ends of the rope are connected with a balancing weight and a tension sensor respectively, a depth sensor is arranged in the balancing weight, and after the balancing weight is arranged at a preset position in a blast hole, explosive can be fed into the blast hole in real time; compared with the conventional mode of manually measuring the filling length of the blast hole after filling the explosive, the method provided by the utility model has the advantage that the measuring efficiency is obviously improved.

The invention discloses an automatic continuous charging device for vertical blast holes, which belongs to the technical field of blasting engineering equipment and comprises an advancing mechanism, a positioning and clamping mechanism and a plane sliding mechanism. The advancing mechanism is used for moving and coarse positioning on the ground; the positioning and clamping mechanism is composed of a barrel shell, long-strip-shaped cartridge clamping blocks in an annular array and a supporting bottom block and can integrally clamp multiple cartridges of pre-loaded explosives, and a positioning pin shaft at the lower end of the supporting bottom block abuts against the wall of an explosion hole to achieve accurate positioning of the axis of a cartridge set and the axis of the blast hole; the plane sliding mechanism passively slides on the supporting stand column through the transverse moving disc, the horizontal freedom degree is provided for centering adjustment, after the positioning pin shaft is finely adjusted and centered, the supporting bottom block and the cartridge clamping block are loosened in sequence, and the cartridge set vertically falls into the bottom of a hole under the action of gravity. According to the device, overall clamping and one-time throwing of multiple cartridges of explosives are achieved, the explosive charging efficiency and centering precision of deep hole blasting are remarkably improved, cartridge clamping stagnation is effectively avoided, and the operation safety and the automation level are improved.

The invention discloses a propulsion system and spaceflight equipment, and relates to the technical field of spaceflight power systems, the propulsion system comprises a nuclear power source, a storage feeding device, an atomization device and an acceleration device, the nuclear power source can supply power to the atomization device and the acceleration device, the storage feeding device is used for storing a magnetic fluid propellant, and the atomization device is used for supplying power to the acceleration device. The storage feeding device communicates with the atomization device, the storage feeding device can convey the magnetic fluid propellant to the atomization device, and the atomization device can atomize the magnetic fluid propellant and convey the magnetic fluid propellant to the acceleration device; the accelerating device can generate a magnetic field and accelerate the magnetofluid propellant in the atomized state. The propelling system and the aerospace equipment are suitable for being used in the environment lacking solar energy, and power is more sufficient.

The invention provides a power cabin for releasing pressurized deformation of a storage tank, which comprises a power cabin body provided with a column section, and a storage tank upper mounting plate and a storage tank lower mounting plate which are positioned on the inner side of the column section; the storage box is arranged in the power cabin body, a storage box upper flange and a storage box lower flange are arranged on the outer side of the storage box, the storage box upper flange is matched with the storage box upper mounting plate, and the storage box lower flange is matched with the storage box lower mounting plate; an adjusting gasket is arranged between the storage box upper mounting plate and the storage box upper flange, after the storage box lower mounting plate is connected with the storage box lower flange, the storage box is pressurized to set pressure, the storage box upper mounting plate is connected with the storage box upper flange, and the storage box upper flange is connected with the storage box lower flange. And the adjusting gasket is used for adjusting a gap between the storage box upper mounting plate and the storage box upper flange. The invention further provides a method for installing the storage tank to the engine compartment body. According to the invention, the requirements of a large manned spacecraft on a general storage tank and a heavy-weight propellant are met at the cost of lower weight of the cabin body structure.