256results about "Launch systems" patented technology

Systems and methods for providing a space launch service platform (SLSP) that integrates artificial intelligence and data analytics to support launch operations. The SLSP may connect to multiple data sources, collect data, and standardize the collected data according to regulatory and operational standards. The SLSP may evaluate launch safety and risks using standardized data and generate a situational analysis for decision-making. The SLSP may provide graphical overlays and decision-support tools to highlight optimal launch windows and potential risks.

The system and method for decelerating space objects moving at high velocities are provided. The system comprises at least one closed volume containing a combination of gas, liquid, solid particles, or a mixture thereof. The volume can be transported to the targeted location using a range of means, including a chemical gun, light gas gun, electromagnetic coil gun, superconducting quench gun, a rocket, or a combination thereof. The volume is strategically positioned on the trajectory of the moving object. Upon penetrating the walls of the volume, the object passes through it, experiencing deceleration.

Implementations of the disclosed subject matter may provide an adaptive thrust structure for a space launch vehicle that may include an interface block that is fluidically coupled to a fuel supply, a fuel manifold, an oxidizer supply, and an oxidizer manifold. A first adapter plate may be coupled to a first engine assembly. First adapter lines may be fluidically coupled to the first engine assembly and the interface block, via the first adapter plate. A second engine assembly, second adapter lines, and a second adapter plate are capable of replacing the first engine assembly, first adapter lines, and first adapter plate. Implementations of the disclosed subject matter may provide a reusable and reconfigurable launch vehicle that may include a reusable second stage, and a reusable and reconfigurable transporter having a plurality of zones that are configurable to transport cargo, one or more passengers, and / or one or more satellites.

An aircraft has a boom, a propulsion assembly coupled to a first end of the boom, and a first wing coupled to a second end of the boom. The propulsion assembly is coupled to the boom by a rotating joint. A second wing is optionally coupled to the rotating joint. The first wing is coupled to the boom by a rotating joint. The first wing is coupled to the rotating joint by a hinge. A vehicle with roll, pitch, and yaw maneuverability able to mirror the aircraft movements may be coupled to the second end of the boom. The vehicle body may be picked up with a vehicle chassis disconnected from the vehicle body. The boom houses an energy source to power the propulsion assembly. A rudder is coupled to the second end of the boom. A paddle is disposed between the propulsion assembly and the boom.

Combined aerospace system is a multi-stage technology to deliver payloads to orbit and to midair. The system comprises a carrier aircraft, a tow aircraft, a towed spaceplane or a rocket plane or a glider. The carrier aircraft, coupled with the spaceplane and / or the rocket plane or several gliders, takeoff from a runway and ascends to a staging altitude and speed under the carrier aircraft power. After staging, the tow aircraft will pick up the spaceplane or the rocket plane or the glider from the carrier aircraft in flight. The towed spaceplane or the rocket plane or the glider continues ascending and accelerating by the tow aircraft. The spaceplane or the rocket plane will disconnect from the tow aircraft and continue to ascend and accelerate in an atmosphere by using the ramjet engine. The spaceplane or the rocket plane equipped with a propulsion unit is used to place the spacecraft into the earth orbit. The glider will disconnect from the tow aircraft and lands in designated place. Gliders can be used for instantly delivering of correspondence, parcels, and civil goods from a supplier to wide variety of individual consumers or corporate customers.

The present invention discloses a mobile rocket launch vehicle, including a tractor, a chassis, a heat preservation cabin, an erection system, a launch platform, a temperature control system, a hydraulic system, a vehicle control system and a power supply system. A traction seat is installed on the tractor, and the traction seat is connected to the chassis through a traction pin. The heat preservation cabin is fixedly installed on the chassis, the erection system is hinged at the rear end of the chassis, the launch platform is hinged at the rear end of the chassis, and the launch platform is coaxial with the erection system. The temperature control system is installed in front of and inside the heat preservation cabin, the hydraulic system is installed below the chassis, the vehicle control system is installed inside the heat preservation cabin, the operation panel is located outside the side wall of the heat preservation cabin, and the power supply system is installed at the front of the chassis and inside the heat preservation cabin. The present invention is ingenious in conception and reasonable in structural layout, which greatly shortens the long-distance transportation time, ensures the stability during transportation, improves the shock absorption capacity, and the heat preservation cabin has good sealing performance. At the same time, the fully open design meets the requirements of erection and launch.

The present application relates to a rotary launch device and method capable of continuously launching satellites. The device comprises a base mechanism, a launch cabin, a rotary support mechanism, a plurality of release mechanisms, a loading mechanism and an accelerator. The base mechanism comprises a base and a connecting shaft. The launch cabin comprises a launch cabin and a launch tube connected to the launch cabin, and the launch cabin is rotatably connected to the connecting shaft. The rotary support mechanism comprises a plurality of support arms arranged radially in a uniform array along a circumference with a rotation point as the center. A plurality of release devices are respectively arranged at the tail end of a support arm, and are used to lock / unlock the corresponding carrier at the tail end of the support arm. The loading device is located in the launch cabin and is used to load the carriers to the tail ends of the plurality of support arms. The accelerator is fixed on the connecting shaft and is used to drive the rotary support mechanism to rotate. The embodiment of the present application provides a rotary launch device capable of continuously launching satellites. The carrier carrying the satellite is launched to a predetermined orbit by using centrifugal force. The device is reusable and has low launch cost.

A method for rendezvous with an orbiting object comprising: launching a tug and a servicer into a client orbit; separating the servicer from the tug; and docking the servicer with a client. A system for rendezvous with an orbiting object comprising: a first spacecraft comprising a tug capable of towing a second spacecraft, wherein the second spacecraft is a servicer configured to dock with a tumbling client orbiting object.

Stackable satellites with a large diameter antenna that are configured to stack within each other like fully overlapping shallow bowls and to fit tightly within the area in a pay load fairing. The bowl-shaped satellites allow the satellites to nest within each other so that many satellites can be stacked within the height limits of the launch vehicle pay load fairing. A cradle is used to hold the stack, and tensions cables keeps the satellites in compression during launch. The satellite is an antenna with the bus integrated into the antenna structure. The antenna occupies the entire diameter / footprint of the satellite and thus can be matched up with the cross section of the fairing. This allows the largest possible antenna size to be launched without a complex deployment mechanism.

The invention relates to the technical field of spaceflight launching, and provides a swarm-type spaceflight launching system and method, and the system comprises a honeycomb system, a nectar source system, and a bee passage system. The honeycomb system is a technical area for spaceflight launching, is connected with the nectar source system through the bee passage system and is used for completing a satellite-rocket combination; the plurality of nectar source systems are arranged around the honeycomb system, and each nectar source system comprises a plurality of emission positions; the bee passage system is connected with the transmitting position and the honeycomb system and comprises a plurality of traffic roads. According to the scheme, a large number of satellites with different orbits can be launched in a short time, and the satellite launching efficiency is improved.

A system and method are disclosed for softly capturing, protecting, and landing a spacecraft returning from space using a VTOL (vertical takeoff and landing) aircraft including drones and helicopters. The spacecraft is decelerated by a parachute. One or more VTOL drones transport a waterproof pocket that catches and captures the descending spacecraft in the air. In this way, the spacecraft is protected inside the pocket and, after continuing to descend, makes a soft landing in a body of water. In another embodiment, a recovery helicopter, which is a type of VTOL aircraft with heavy-lift capabilities, is used to directly catch the returning spacecraft. One or more VTOL drones are attached to the lower end of a recovery cable suspended from the helicopter. These drones carry a clutch that directly, quickly, and accurately catches the descending spacecraft without interfering with the parachute. In this way, the spacecraft is caught and protected by the helicopter while maintaining the floating function of the parachute.

A controlled reactor comprises a reactor core thermally coupled to one or more heat pipes and an active cooling loop. A fluid may be circulated through the active cooling loop. A heat exchanger is thermally coupled to the active cooling loop and extracts heat from the fluid as the fluid is circulated through the active cooling loop. A heating system may be provided to deliver the heat extracted by the heat exchanger to a community. A thermoelectric generator may be provided to convert heat extracted by the heat pipes to electricity for delivery to the community.

A satellite apparatus is disclosed, including a housing having first and second opposing walls, and a support structure inside the housing spanning the first and second walls. The support structure is structurally connected to the housing only at the first and second walls, and an end portion of the support structure is configured for connection to a launch vehicle by a separation system.

A propulsion stage for a launch rocket, the propulsion stage comprising: a rocket body having a longitudinal axis; at least one recoil propulsion unit acting substantially parallel to the longitudinal axis, wherein a liquid propellant is supplyable to the at least one recoil propulsion unit from at least one propellant tank through at least one fuel pump; and a plurality of rotor assemblies, each drivable by a rotor drive including at least one motor, wherein the at least one motor of at least one of the rotor drives is configured to selectively drive at least one of the rotor assemblies and / or the at least one fuel pump of the at least one recoil propulsion unit.

A launch system and method for orbital or suborbital air-launch of a payload involving releasably coupling a launch vehicle with a towed aircraft via an articulatable carriage to form an air-launch assembly, towing the air-launch assembly via a tow aircraft and interconnected tow cable to a first altitude, releasing the air-launch assembly from tow at or above the first altitude, activating the towed aircraft propulsion system and initiating a pull-up and climb maneuver of the towed aircraft to a second altitude, articulating the articulatable carriage to shift the air-launch assembly from a stowed position to a deployed position with the launch vehicle spaced from the towed aircraft, releasing the launch vehicle from the articulatable carriage and thus from the towed aircraft, and activating the launch vehicle propulsion system for further altitude gain or to meet specific mission requirements.

A satellite system may have a constellation of communications satellites in orbits such as highly inclined eccentric geosynchronous orbits and low earth orbits. To place satellites in inclined eccentric geosynchronous orbits, a series of launch vehicles may be launched. Each launch vehicle may be used to place a set of satellites, such as a set of three satellites, into a common orbital plane with distinct longitude of ascending node values. To place satellites in low earth orbits, a series of launch vehicles may be launched, each of which releases satellites in sequence from a stack of satellites into a common orbital plane. After desired separations have been produced between the released satellites, circularization procedures may be performed using the propulsion systems of the satellites to place the satellites into final orbit.

A reusable orbital vehicle (14) for a space transport system (10) comprises a fuselage (22) the shape of which widens in a direction extending from the front end (24) of the reusable orbital vehicle to a rear end (26) of said fuselage (22) and inside which there is defined a housing (32), and a crew-evacuation vehicle (34) housed in said housing (32) and oriented in such a way as to be extracted from the housing (32) in a direction (D1) directed from said rear end (26) of the fuselage (22) toward said front end (24) of the reusable orbital vehicle (14).

The invention relates to the technical field of spaceflight, and particularly discloses a satellite combination structure and a satellite launching control method. The satellite combination structure comprises a first satellite and at least two second satellites; the second satellites are arranged on the same side of the first satellite side by side, the second satellites are connected with the first satellite through separation mechanisms, and every two adjacent second satellites are connected through a separation mechanism; the first satellite comprises a propelling system and a first force bearing structure, the propelling system comprises a main propeller and a fuel tank, the fuel tank is located in the first force bearing structure, and the main propeller is connected to the outer side of the bottom of the first force bearing structure. Therefore, complex structures such as an upper stage or a propulsion power cabin do not need to be additionally arranged, under the condition that the satellite function and the on-orbit deployment condition are guaranteed, the carrying capacity is effectively saved, new space junk is avoided, meanwhile, the satellite launching cost is saved, and the system construction cost is reduced.

Systems and methods using VTOL (vertical take-off and landing) aircrafts including drones and helicopters for soft capture, preserving, and landing of a returning spacecraft from space are disclosed. The spacecraft is decelerated by parachutes. One or multiple VTOL drones transport a water impermeable pocket meeting and capturing the descending spacecraft in the air. The spacecraft is thus preserved inside the pocket and keeps descending and then softly lands in a body of water. In another embodiment, a recovery helicopter, one type of VTOL aircraft with heavy payload lifting capacity, is used to directly catch the returning spacecraft. One or multiple VTOL drones are coupled to the bottom end of a recovery cable hung from the helicopter. These drones bring a clutch quickly and precisely catching the descending spacecraft directly without interrupting the parachutes. The spacecraft is thus caught and preserved by the helicopter with lifting function of the parachutes maintained.

To provide a system, apparatus and method for a satellite thermal enclosure.SOLUTION: A satellite assembly is disclosed, including a satellite and a shroud. The satellite is stowed in a launch vehicle, and the shroud 200 includes a frame 210 supporting a flexible thermal blanket 212 enclosing the satellite.SELECTED DRAWING: Figure 3

Provided is a core material that may be smoothly produced as a result of improvements in the texture and strength of the base material thereby having excellent light weight property and strength. The present invention is a core material having a structure including a collective body of hollow columnar cells compartmentally formed by cell walls. The cell walls include a sheet material in which discontinuous reinforcing fibers (Cf) are dispersed in a form of monofilament.

A device comprises a tube with a tube HTS solenoid, wherein a projectile in a sabot comprising a sabot HTS solenoid. A method comprises disposing a seed HTS crystal on a growing crystal in contact with an a-b plane of the seed HTS crystal to grow the growing crystal, wherein the a-b plane is perpendicular to a c-axis. A method comprises disposing a seed HTS crystal on a growing crystal in contact with a b-c plane of the seed HTS crystal to grow the growing crystal, wherein the b-c plane is perpendicular to an a-axis. A device comprises a reinforced HTS material in a graphene casing, wherein the HTS in the graphene casing includes a cooling channel and a return channel.

A propulsion-based payload transporter (PPT) and methods of its operation are provided. The PPT may operate in a gravitational field such as that on the lunar surface. Via rocket engines that produce thrust, the PPT is configured to descend onto and over a payload to be moved. Once in position with respect to the payload, latches engage with the payload so as to secure the payload in a payload holding bay of the PPT. The rocket engines produce thrust with an exhaust plume that is angled away from the payload to avoid payload damage. The rocket engines, which are located at the sides of the payload holding bay, produce thrust to propel the PPT and the payload in a substantially horizontal trajectory with respect to gravity and toward a landing destination.

The invention relates to the area of space science, in particular, to the area of industrialization of outer space and, directly, to the facilities for inserting and moving various objects in circumplanetary cosmic space. It is intended to solve geocosmic problems in industrial-scale volumes - to carry out research and development, special-purpose, touristic and other types of works and services in outer space and stabilize global climate. Artificial orbital ring complex by Yunitski is made in the form of a ring satellite (1), located in a circular orbit (5) around a natural cosmic body (2) in the equatorial (4) plane (3), which comprises: structural frame (7), utilities (8) and communications (9), power units (10), transition galleries (11) positioned in annular housing (12) with protective shell (13), gate chambers (14) with docking units (15), configured to dock with space vehicles (16), residential (19) and research-production (20) units, equipped with systems of control, life support, active and passive (17) protection against the harmful effects of space (18). The structural frame (7) is longitudinally tensioned, and the ring satellite (1) is in circular motion in orbit around the natural cosmic body (2) with velocity V0, m / sec, defined by the ratio: 1≤V0 / V1H≤1.01, where V1H, m / sec, - first space velocity for the equatorial circular orbit of the location of the ring line connecting the centers of mass of the cross sections of the artificial orbital ring complex, at altitude of H0, m, defined by the ratio: 0.02≤H0 / R0≤0.5, where R0, m, - radius of natural cosmic body (2) in equatorial (4) plane (3).

This invention is embodied in a system and device for transporting assets (small satellites, commodities, or other payloads) in space. The preferred system comprises a scaffold of relay stations strategically orbiting one or more celestial bodies. The orbits are complimentary to each other so that the relay stations can transfer mass from one station to another. In this way, the system can support different logistical networks connecting the space between orbits such as Low Earth orbit (LEO), geostationary orbit (GEO), and Low Lunar orbit (LLO), and even between Earth and Mars.

A reusable payload module (2; 2') as part of an upper stage (12) of a space rocket (1) is provided with a fairing (20A) which, when the payload module (2) is closed, surrounds a payload receiving space (21) and forms an outer shell (20; 20') of a fuselage (29) of the payload module (2; 2'), which is rotationally symmetrical to a payload module longitudinal axis (XN) which runs coaxially with or parallel to a longitudinal axis (XR) of the space rocket (1), and is characterized in that the fairing (20A) forming the outer shell (20; 20') has at least two surface sections (23, 24) which can each be pivoted outwards about a pivot axis (X23, X24).

The invention relates to a method for manufacturing a cylindrical, thermoplastic composite part, the method comprising: - positioning a cylinder (3) made of thermoplastic pre-impregnated fibrous material in a vacuum moulding tooling (10), at least one vacuum bag (12) being situated facing the cylinder, and - consolidating the cylinder in the vacuum moulding tooling, during which a vacuum is drawn so that the at least one bag applies a shaping pressure to the cylinder, the cylinder being subjected to a heat treatment comprising (i) sweeping (B) its circumference with a local heating, of extent (E24) limited to a sector of the cylinder, in a relative movement in relation to the cylinder about its axis, and (ii) cooling, after this local heating, so as to fix the shape of the cylinder. Fig. 1

The present invention is embodied in systems and devices for transporting assets (small satellites, goods, or other payloads) in outer space. A preferred system comprises a scaffold of relay stations strategically orbiting one or more celestial bodies. Since the orbits are complementary to each other, the relay stations can transport mass objects from one station to another. In this way, the system can support various logistics networks connecting space between orbits such as low Earth orbit (LEO), geostationary orbit (GEO), and low lunar orbit (LLO), as well as space between Earth and Mars.

The invention provides a flat plate satellite mixed stacking layout structure and a releasing method, and the layout structure comprises a large flat plate satellite and a small flat plate satellite; the small flat satellites are stacked and placed in a supporting cabin of the rocket to form a first vertical layer; each layer of the first vertical layer comprises two small-sized plate satellites which are laid flatly; the large flat plate satellites are stacked at the top end of the first vertical layer to form a second vertical layer; the two side faces, away from each other in the tiling direction, of the two small flat plate satellites of each layer of the first vertical layer are aligned with the two corresponding side faces of the large flat plate satellite respectively. By means of the layout structure, efficient utilization of the space of the fairing can be achieved.

The present invention relates to a system and method of mobile orbital launchers. The mobile launcher of the present invention mainly consists of the launch vehicle and launch related subsystems. All the systems will fit on a trailer which is capable of going anywhere via road or sea. It will carry the vehicle in a horizontal position and after reaching the required location, it will articulate and make the vehicle vertical to get ready for launch. The system and method of the present invention have all the subsystems which enable the launch without any hindrance.