151 results about "Propellant tank" patented technology

An In Orbit Transportation & Recovery System (IOSTAR™) (10) One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

An in-space spacecraft servicing system (10) includes a servicing spacecraft (22) and a propellant module (24). The servicing spacecraft includes a client servicing system (136), as well as navigation avionics (108) for independent flight operation and a servicing propellant tank (170). The propellant module moves the servicing module from an upper stage drop off location and releases it in proximity to a client spacecraft (16) for a servicing mission. It has a propellant tank (172) with capacity for multiple missions and is used to refill the servicing spacecraft's propellant tanks between missions. Either or both the servicing spacecraft and the propellant module may have navigation avionics. The servicing spacecraft also has a universal docking adaptor (70) for different client spacecraft, and can convert a client spacecraft from non-cooperative to cooperative.

An in-space spacecraft servicing system (10) includes a servicing spacecraft (22) and a propellant module (24). The servicing spacecraft includes a client servicing system (136), as well as navigation avionics (108) for independent flight operation and a servicing propellant tank (170). The propellant module moves the servicing module from an upper stage drop off location and releases it in proximity to a client spacecraft (16) for a servicing mission. It has a propellant tank (172) with capacity for multiple missions and is used to refill the servicing spacecraft's propellant tanks between missions. Either or both the servicing spacecraft and the propellant module may have navigation avionics. The servicing spacecraft also has a universal docking adaptor (70) for different client spacecraft, and can convert a client spacecraft from non-cooperative to cooperative.

An In Orbit Transportation & Recovery System (IOSTAR™) (10). One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

A system for transportation and storage of cargo in space includes one or more platforms. The platforms are operable to hold multiple removable propellant tanks and rendezvous with one or more other platforms in space. The platforms also include one or more thruster systems and positioners. Each positioner has an effector operable to grip a band disposed around each propellant tank. Each positioner is operable to facilitate the transfer of the multiple removable propellant tanks between platforms during the rendezvous.

An In Orbit Transportation & Recovery System (IOSTAR(TM)) (10) is disclosed. One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR(TM) includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR(TM) (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

An active satellite dispenser is preferably attachable to a reusable launch vehicle for deployment of one or more satellites into one or more desired orbits. The active satellite dispenser includes a center mast that releasably receives the satellite(s), a liquid propellant rocket, and an orbital control system on an avionics pallet. In the preferred embodiment, a pressurized gas selectively pressurizes the propellant tanks (which may include fuel and oxidizer tanks), to provide propellant to the rocket. In operation, the launch vehicle releases the satellite dispenser in a first deployment orbit. The active dispenser rocket and orbital control system then transport the active dispenser and satellite(s) into the final deployment orbit. In the preferred embodiment the active dispenser can operate multiple times to place individual satellites in different orbits.

The invention belongs to the technical field of space propulsion. The technical problem to be solved is to provide a realization method of the space propulsion system, which can effectively overcome the defects of the existing technology, and can adapt to the development requirements of long service life and high safety of space vehicles in orbit. The present invention has designed a "well" structure gas path supply system and a cross-supply propellant supply system, and can realize repeated filling / leaking of the propellant on the ground metal membrane box type storage tank, and the present invention brings the following beneficial effects: 1) The air cavities of the two propellant tanks are physically isolated and not connected to each other; 2) Solve the fault isolation of gas leakage, and the system will not fail to work due to the leakage of a gas cylinder; 3) Meet the long-term on-orbit aircraft to adapt to multiple times Supplemented tank design requirements.

An In Orbit Transportation & Recovery System (IOSTAR™) (10) is disclosed. One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

An In Orbit Transportation & Recovery System (IOSTAR™) (10). One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

A field-ionization based electrical ion thruster utilizes a single propellant that can be used in either a high specific-impulse mode, i.e., ion-thruster mode, or a low-specific-impulse mode, i.e., a cold-gas thruster mode. In one embodiment, the high specific impulse mode (ion thruster mode) utilizes a miniaturized positive-ion field-ionization chamber including a permeable substrate infused with properly oriented carbon nanotubes (CNTs), which is fed propellant from a porous injection plug made from permeable carbon or equivalent material. In one embodiment, field-electron emission from a neutralizer, such as a carbon nanotube array neutralizer, positioned after one or more accelerator grids is used for ion neutralization. In one embodiment, the low specific-impulse (cold-gas mode) uses a conventional supersonic nozzle-expansion to generate thrust. In one embodiment, both the high specific-impulse mode and the low specific-impulse mode thruster embodiments share the same gas propellant, storage propellant tank, and delivery equipment.

The invention disclosed here is a pressurization system for pressure-fed launch vehicles. The system employs a hot pressurant gas to minimize the mass of pressurant gas required. It also employs a set of baffles within the propellant tank to reduce heat transfer between the propellant and the hot pressurant gas. The baffles keep the pressurant gas flowing uniformly in one direction as the propellant is expelled, and inhibit mixing of the cold gas near the propellant with the hot gas being blown into the tank. The baffles also prevent large-scale sloshing of the propellant. The metal baffles are rigidly attached to one end of the tank and attached with a mount which allows travel in the longitudinal direction but not in the two lateral directions which minimizes stess on the pressure vessel from differences in thermal expansion.

Methods for making skin panels suitable for use in aircraft structures and fuel tank structures, such as fuselages and reusable launch vehicle propellant tanks. A first method includes the following steps: machining a starting plate to a final part geometry; and explosive forming the machined plate to at least one of a uniaxial curvature and a biaxial curvature. A second method includes the following steps: explosive forming a starting plate to at least one of a uniaxial curvature and a biaxial curvature; and machining the curved plate to a final part geometry. Accordingly, the invention enables making integrally stiffened unitized axial load carrying skin panels that are either uniaxially or biaxially curved.

A cryocooler is located on a spacecraft bus, such as a bus box, separate from the cryogenic propellant tanks disposed on a separable and distinct propellant cartridge system spacecraft docked to the spacecraft bus. In operation, propellant may be continuously pumped from the tanks through the cryocooler cold heat exchanger and then back to the tanks on the separable propellant cartridge system spacecraft through temporarily couplable lines. After the propellant tanks are depleted, the propellant cartridge system is then undocked from the bus and typically discarded. A new propellant cartridge system spacecraft comprising a full set of tanks may then be docked to the bus and the cryocooler supply / return lines coupled. The remote cryocooler may function as part of a larger space depot for spacecraft resupply, although it is not limited to such applications.

The present disclosure relates to a launch system, a launch vehicle for use with the launch system, and methods of launching a payload utilizing the launch vehicle and / or the launch system. The disclosure can provide for delivery of the payload at a terrestrial location, an Earth orbital location, or an extraorbital location. The launch vehicle can comprise a payload, a propellant tank, an electrical heater wherein propellant, such as a light gas (e.g., hydrogen) is electrically heated to significantly high temperatures, and an exhaust nozzle from which the heated propellant expands to provide an exhaust velocity of, for example, 7-16 km / sec. The launch vehicle can be utilized with the launch system, which can further comprise a launch tube formed of at least one tube, which can be electrically conductive and which can be combined with at least one insulator tube. An electrical energy source, such as a battery bank and associated inductor, can be provided.

A platform for storage and transfer of propellant tanks in space includes frames configured to releasably support plural propellant tanks in storage positions, and transfer mechanisms configured to transfer propellant tanks to and from the storage positions. The platform includes a central structure to which the plural frames are connected at predetermined spaced intervals. The transfer mechanisms include extendable telescoping arms mounted to the central structure. The arms have end effectors configured to grip bands on propellant tanks. Two platforms may be positioned adjacent each other for the transfer of propellant tanks therebetween.

An orbital stage system has an orbital stage and one or more launch stages. The orbital stage incorporates an orbital maneuvering system (OMS) and an abort propulsion system which both utilize the same propellants, propellant tankage, and propellant pressurization system, but which employ radically different engines. The OMS engines are comprised of at least two engines which have a combined thrust in the neighborhood of 1 / 10 the weight of the orbital stage, an area ratio of 50 or more and an operating life of many hundred seconds, preferably many thousands of seconds or more. The abort engine may be a single engine and typically has a thrust of three, four, or more times the weight of the vehicle and an area ratio in the neighborhood of two and an operating life of at most a few tens of seconds.

A propulsion system for a rocket engine and a method of cooling a rocket engine includes a propellant tank fluidically coupled to the rocket engine to hold a pressurized propellant, a coolant tank to hold a coolant, a first heat exchanger thermally coupled to the rocket engine and fluidically coupled to the coolant tank, a second heat exchanger thermally coupled to the propellant tank and fluidically coupled to the first heat exchanger, and a third heat exchanger disposed inside the propellant tank to thermally couple a propellant withdrawn from the tank for combustion to a propellant disposed inside the tank. The coolant flows from the coolant tank to the first heat exchanger and through the first heat exchanger to cool the rocket engine. The propellant withdrawn from the propellant tank receives heat from the propellant disposed inside the tank through the third heat exchanger to convert to a gaseous propellant when withdrawn from the propellant tank as a liquid propellant. The coolant flows from the first heat exchanger to the second heat exchanger and through the second heat exchanger to heat the propellant disposed inside the propellant tank.

The invention discloses a power system for a carrier rocket upper stage and an orbit transfer vehicle. The power system for the carrier rocket upper stage and the orbit transfer vehicle solves the existing problems that a main engine and an attitude control engine are relatively independent, the carrying burden is relatively heavy,a propellant used in the attitude control engine is rather toxic,and the environment and safety of a human body are seriously threatened. The power system for the carrier rocket upper stage and the orbit transfer vehicle comprises the main engine and the attitude control engine; a propellant tank of the main engine can simultaneously deliver a propellant to a thrust chamber of the main engine and a thrust chamber of the attitude control engine under the action ofa propellant supply system of the propellant tank, and liquid oxygen is used as an oxidant of the propellant; kerosene, or liquid hydrogen or liquid methane is used as an incendiary agent; the detoxication of the attitude control engine is achieved, and integration of the homologous propellant of the main engine and the attitude control engine is achieved; a propellant tank of the attitude control engine can be supplemented after operation through the main engine; and the propellant dosage carried by the attitude control engine during flight can be reduced, thus the weight of the overall system is lowered, and the payload of a carrier is greatly increased.

A method of regulating pressure within a first propellant tank of a rocket engine having a first propellant tank containing a first propellant and a second propellant tank containing a second propellant, and a regulator device for regulating pressure within the first tank, the regulator device comprising a gas generator and a heat exchanger co-operating with the gas generator so as to vaporize at least part of the first propellant prior to reintroducing it into the first tank (16), the gas generator and the heat exchanger both being fed with the first propellant by a single first motor-driven pump, while the gas generator is fed with the second propellant by a single second motor-driven pump, wherein the flow rate of the first motor-driven pump is controlled as a function of a first parameter, while the flow rate of the second motor-driven pump is controlled as a function of a second parameter.

A compact single-propellant unitary propulsion system for placing a satellite into an orbit and subsequently correcting the orbit so that the satellite is stabilized on three axes comprises a liquid propellant tank secured to a satellite platform and having a reinforced bottom wall with an outlet and filter element, a distribution block welded to the reinforced bottom wall, at least one filling / emptying value mounted on the distribution block, and a set of at least two thrusters mounted on the distribution block and fed directly from the distribution block without additional pipework. The set of at least two thrusters point substantially along the axis of the tank, which is aligned with the axis of the satellite. The propulsion system may be fully assembled before it is integrated with the satellite.

The invention discloses a hydraulic accumulator for a propellant tank in a spacecraft. The hydraulic accumulator comprises an air guide tube, a cover plate, an outer shell, blades, a support column, a base, channel window pressing plates, a channel window meshes, hydraulic accumulator meshes, a fixing block and hydraulic accumulator mesh pressing plates, wherein the blades are uniformly distributed on and fixed to the support column in the radial direction to form a whole structure, and the whole structure is mounted on and fixed to a circular partition plate in the base; the hydraulic accumulator meshes are pressed on the circular partition plate in the base through the hydraulic accumulator mesh pressing plates and then fixed; the channel window meshes are pressed on the outer side of a cylinder wall of the base through the channel window pressing plates and then fixed; the outer shell is fixed at the top end of the base, the cover plate is fixed on the outer shell, and the air guide tube is fixed on the cover plate through the fixing block. The hydraulic accumulator disclosed by the invention is simple in structure, high in strength and reliability, large in liquid storage quantity and wide in work application range, has the function of repeatedly and circularly filling liquids to the hydraulic accumulator and discharging the liquids from the hydraulic accumulator, is large in liquid filling and discharging rate, and meanwhile has the capacity of realizing ground repeated filling.

The present invention relates to propellant tanks for space platforms, launchers and every sort of space transport craft.

An aerospace frame accommodates propellant tanks for facilitating propellant operations in space. The frame includes at least two plates for supporting the propellant tanks; at least one brace supporting the plates; and a cavity between two plates that accommodates a removable propellant tank.

A tank suitable for containing cryogenic propellant includes an inner liner structure comprised of two layers forming a gap. An intermediary layer comprised of micro-particles is disposed in the gap formed between the two layers of the inner liner structure. A structural overwrap layer covers the inner liner structure in a manner that enables the inner liner structure to float within the structural overwrap layer. The inner floating liner structure can be formed by two thin interlocking (outer and inner) shell s formed in a cylindrical, end dome geometry. The inner and outer shells are both sectioned beyond the tank midpoints. The micro-particles can be comprised of PTFE. An outer structural overwrap comprised of graphite / epoxy covers the double shell liner. A pair of separately fabricated polar bosses seal against the inner of the two interlocking floating liner shells are then each bonded to the tank, fittings, and piping.

A satellite system and method features one or more electric propulsion thrusters optimized to have a specific impulse for station keeping and connected to a propellant tank for ionizing a propellant during station keeping maneuvers and receiving power from a solar photovoltaic array via a power management and distribution system during station keeping maneuvers; one or more electric propulsion thrusters are optimized to have a thrust for orbit raising and are connected to the propellant tank for ionizing the propellant during orbit raising maneuvers and receiving power directly from the solar photovoltaic array during orbit raising maneuvers or from the power management and distribution system