81results about "Space shuttles" patented technology

An aerospace panel includes a first skin, a second skin spaced apart from the first skin, and a first truss structure connecting the first skin to the second skin. The first truss structure includes a plurality of truss members. Each truss member is integral with the first skin and the second skin, such that the first skin, the second skin, and the first truss structure collectively form a single monolithic joint-free structure. At least one of the skins also includes at least one lattice region that includes a lattice grid and an array of openings.

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.

The invention discloses a momentum body appearance of an upper atmospheric aircraft. The momentum body appearance comprises an aircraft head, a body, a sweepback airfoil and a wingtip V-shaped tail. The head of the aircraft is a wedge-shaped polyhedron and is in butt joint with a rear body. The body is a multi-stage trapezoidal boss which shrinks backwards, so that the resistance of the body is reduced; the sweepback airfoil is connected with the body to realize static pitching stability; and the two V tails are positioned at the positions of wingtips. According to the invention, the momentum exchange principle that the upper atmospheric layer gas interacts with the object plane is adopted, a polyhedral aerodynamic shape meeting the low-resistance aerodynamic characteristic is designed, the fusion design of the aerodynamic performance and the solar wing is realized, the aerodynamic resistance compensation requirement of the aircraft in the airspace can be effectively reduced, and the hang time of the aircraft is prolonged.

The present invention discloses a clamping release device and a spacecraft, wherein the clamping release device comprises a first mounting base, a second mounting base, a flexible connector, an elastic member and a hot knife assembly; the first mounting base and the second mounting base are respectively fixedly connected to the flexible connector; the hot knife assembly is located between the first mounting base and the second mounting base, and is movably arranged on the second mounting base, and the flexible connector is located on the moving path of the hot knife assembly; the first end of the elastic member is connected to the second mounting base, and the second end of the elastic member is connected to the hot knife assembly, and the elastic member drives the hot knife assembly to abut against the flexible connector; the hot knife assembly is an electric heating structure, and the hot knife assembly fuses the flexible connector when power is turned on. The above scheme can achieve the technical effects of simplifying the structure, reducing the impact load and reducing pollution.

A long range vehicle capable of space travel uses quantum effects such as quantum printing or instantiation to produce consumables during travel. The long range vehicle comprises a transport shell, a propulsor disposed in or on the shell, and at least one Nagel Quantum Effect Device (NQED) operatively coupled to the propulsor, the at least one NQED providing imprinted material to the propulsor.

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).

A fuel tank (100) comprises an interior wall (105), a sump (110), and a baffle (120) that comprises a center fitting (124), a full-length containment petal (132), a partial-length containment petal (134), a dump tube (118). The full-length containment petal (132) comprises a full-length side edge (133), extending radially outward from the center fitting (124). The partial-length containment petal (134) comprises a partial-length side edge (135), extending radially outward from the center fitting (124). The dump tube (118) is connected to the sump (110). The full-length side edge (133) of the full-length containment petal (132) is longer than the partial-length side edge (135) of the partial-length containment petal (134). All of the partial-length side edge (135) of the partial-length containment petal (134) is attached to a linear portion of the full-length side edge (133) of the full-length containment petal (132).

The invention relates to a transport system (1) intended to be installed on a space launcher (100) comprising:an aircraft (2) which comprises propulsion means configured to propel the aircraft (2) in a direction of flight, and at least one seat (4) intended to receive a passenger which is movable in rotation about an axis (0) perpendicular to the direction of flight of the aircraft (2), an acceleration sensor being installed on each at least one seat (4) to measure the acceleration of each at least one seat (4);a passenger interface for each at least one seat (4) which comprises a screen intended to display images to the passenger installed on said at least one seat (4), the screen being coupled to said at least one seat (4) so to remain fixed relative to said at least one seat (4);a control unit (6) which is connected to the acceleration sensor and to said at least one seat (4), the control unit (6) being configured to calculate a load factor experienced by the passenger installed on the at least one seat (4) from the acceleration of said at least one seat (4), the control unit (6) being configured to monitor the rotation of the at least one seat (4) during the operation of the transport system (1) in order to maintain the position of the seat (4) fixed relative to the load factor experienced by the passenger throughout the flight.

This invention relates to a quadcopter Mars spacecraft and its upright deployment and platform takeoff deployment device. The invention addresses the current lack of a suitable structure, deployment, and takeoff deployment scheme for quadcopter Mars spacecraft. The invention includes a lander, an outer envelope assembly, a bottom extension component, a multi-degree-of-freedom robotic arm, and a Mars spacecraft. The bottom extension component and the multi-degree-of-freedom robotic arm are mounted side-by-side on the lander, the outer envelope assembly is mounted on the bottom extension component, and the Mars spacecraft is housed within the outer envelope assembly. This invention belongs to the field of aerospace vehicle technology.

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 discloses an axial series coupling framework aerospace aircraft and carrier aircraft combined aerospace vehicle which comprises an aerospace aircraft and a carrier aircraft with variable sweepback wings. A double-body platform axial series connection design is adopted, full-system reuse is achieved through pneumatic coupling synergy and intelligent power management, and conventional airport take-off and high-altitude separation are supported. The aerospace aircraft and the carrier aircraft both have autonomous return capability, and the power system adopts a three-level switching strategy and supports air combustion supporting and liquid oxygen propelling. And full-process reusable transmission is realized. The carrier aircraft is provided with a variable sweepback wing mechanism, and the aerospace aircraft is integrated with a replaceable heat-resistant protective layer. The modular design supports the kiloton-level carrying capacity, the launching cost is remarkably reduced, and the safety is improved. The method is used for aerospace high-frequency transportation tasks, and fills the technical blank of air-launched aerospace aircraft orbit injection control.

This invention relates to a wheel-rotor composite spacecraft for exploring Titan and its deployment method on a lander, belonging to the technical field of Titan spacecraft. The invention aims to provide a foldable, simple, safe, and reliable wheel-rotor composite spacecraft for exploring Titan, and to provide a deployment method suitable for this spacecraft on a lander to release the spacecraft and provide a takeoff environment. Multiple rotor arms are symmetrically distributed on both sides of the fuselage. One end of each rotor arm is rotatably connected to the fuselage via a corresponding arm motor, and the other end of each rotor arm is rotatably connected to a wheel-type composite rotor via a corresponding rotor motor. A control system is installed inside the fuselage to control the operation of the arm motors and rotor motors. In relatively flat terrain, the spacecraft can use low-power wheel-drive to explore Titan. When encountering more rugged terrain, the spacecraft can also conduct flight exploration.

The light-weight inflatable bag panel supporting structure comprises a main structure framework, a panel and an air bag, the main structure framework is of an oblique quadrangular prism structure formed by rigidly connecting a plurality of truss rod pieces through a plurality of multi-way wrapping and sleeving pieces, and the truss rod pieces are arranged on the two parallelogram side faces of the main structure framework along the short diagonal lines correspondingly; the inclined quadrangular prism structure is divided into two horizontal triangular prism structures, and the multi-way wrapping and sleeving piece comprises a plurality of sleeves for installing truss rod pieces. The panels are fixedly connected to the upper surface and the lower surface of the main structure framework; the space between the panel and the adjacent horizontal triangular prism structure is filled with the air bags, and the air bags are inflated to be used for supporting the panel. According to the panel supporting structure, an inflatable bag structure system is adopted, the out-of-plane static and dynamic rigidity of a large-breadth panel in wing and rudder components is effectively improved, and the panel supporting structure is simple in structure, efficient and beneficial to achieving light-weight design.

The invention is notably directed to an aircraft, which includes an aircraft structure, a thermal exchange system, a thermal insulation material, and fasteners. The thermal exchange system includes a fluid-based cooling circuit and heat exchanger tiles. The latter cover, at least partly, the aircraft structure. They may thus form the skin of the aircraft. The fasteners secure the tiles to the aircraft structure. Each tile of the heat exchanger tiles embeds one or more fluidic channels, which are connected to the fluid-based cooling circuit, whereby the thermal exchange system is adapted to actively cool the tiles, in operation. Remarkably, the thermal insulation material extends between the aircraft structure and the heat exchanger tiles. The thermal insulation material is clamped on either side by the heat exchanger tiles and by the aircraft structure, thanks to the fasteners.

The invention relates to a propulsion stage of a launch vehicle, in particular a reusable propulsion stage, comprising a rocket body (2) having longitudinal axis (Z), wherein the propulsion stage (3) is provided with at least one jet propulsion drive (36; 36A, 36B) acting predominantly parallel to the longitudinal axis (Z), and wherein the propulsion stage (3) has a plurality of rotor arrangements (33) which can each be driven by an electrical rotor drive, which has at least one electric motor (33'; 33'A, 33'B), which is electrically conductively connected to at least one power storage unit (62) for the supply of electrical energy, wherein the at least one power storage unit (62) can be shifted within the propulsion stage, characterised in that, for the purpose of said shifting, the at least one power storage unit (62) can be moved in a direction (R) lying perpendicular to the longitudinal axis (Z) in a plane or having a dominant direction component extending in a radial direction perpendicular to the longitudinal axis (Z).

The invention is notably directed to an aircraft, which includes an aircraft structure, a thermal exchange system, a thermal insulation material, and fasteners. The thermal exchange system includes a fluid-based cooling circuit and heat exchanger tiles. The latter cover, at least partly, the aircraft structure. They may thus form the skin of the aircraft. The fasteners secure the tiles to the aircraft structure. Each tile of the heat exchanger tiles embeds one or more fluidic channels, which are connected to the fluid-based cooling circuit, whereby the thermal exchange system is adapted to actively cool the tiles, in operation. Remarkably, the thermal insulation material extends between the aircraft structure and the heat exchanger tiles. The thermal insulation material is clamped on either side by the heat exchanger tiles and by the aircraft structure, thanks to the fasteners.

Thermal protection systems for aerospace vehicles, such as hypersonic vehicles or space vehicles, are disclosed herein. In some embodiments, the disclosed thermal protection systems can be used to protect propellant tanks (e.g., composite propellant tanks) and vehicle primary structures (e.g., composite primary structures) from exceeding upper design temperature limits during vehicle ascent, reentry into Earth's atmosphere, and / or high velocity flight through an atmosphere. In some embodiments, a thermal protection system can include a framework of elongated composite members forming a plurality of cells. One or more layers of insulation material such as aerogel can be positioned in each of the cells. The cells can be covered by an external skin of the vehicle. In some embodiments, the framework is attached to a propellant tank and the insulation material is in contact with the external surfaced of the propellant tank.

The invention relates to an aircraft fuselage heat diverting system. Comprising heat exchange sections arranged on the aircraft, a first heat exchange section is arranged on the windward side of the aircraft, and a second heat exchange section is arranged on the leeward side of the aircraft; the first heat exchange section communicates with the second heat exchange section through a communicating assembly. The first heat exchange section is filled with a cooling working medium; and after heat absorption phase change, the cooling working medium flows to the second heat exchange section through the communicating assembly. Efficient heat conduction can be achieved, the destructive problems of deformation and the like of a fuselage structure caused by thermal stress are solved, and the effect of optimizing space is achieved.

The reusable orbital vehicle (14) for the space transportation system (10) includes a fuselage (22) having a shape that widens in a direction from a forward end (24) of the reusable orbital vehicle to an aft end (26) of the fuselage (22), a housing (32) defined within the fuselage (22), and a crew escape vehicle (34) housed within the housing (32) in a state oriented for extraction from the housing (32) in a direction (D1) oriented from the aft end (26) of the fuselage (22) toward the forward end (24) of the reusable orbital vehicle (14).

A magnetic position sensor for determining the angular position of a magnet on a rotatable component, comprising: at least one magnetic sensor (30) for determining different vector components of a magnetic field of the magnet; a memory having a look-up table (34) stored therein that is populated with data representative of different angular positions of the magnet, the data representative of each angular position being correlated with data in the look-up table (34) that is representative of the vector components that would be detected by the at least one magnetic sensor at that angular position; wherein the magnetic position sensor is configured to use the vector components determined by the magnetic sensor and logic to determine the angular position of the magnet from the look-up table (34).

Various embodiments of space launch vehicle systems and associated methods of manufacture and use are disclosed herein. In some embodiments, the systems include a reusable, horizontal takeoff / horizontal landing (HTML), ground-assisted singie-stage-to-orbit (SSTO) spaceplane that is capable of providing frequent deliveries of people and / or cargo to Low Earth Orbit (LEO). In some embodiments, the spaceplane can takeoff with the aid of a rocket-powered sled that, in addition to providing additional thrust for takeoff, can also provide propellant for the spaceplane engines during the takeoff run so that the spaceplane launches with full propellant tanks.

The invention relates to a method for manufacturing a crosslinked product having improved carbonization and thermal stability properties in which there is crosslinking of a benzoxazine monomer.

The sub-orbital airship comprises a single-stage rocket and an airship body, the airship body is installed at the top of the single-stage rocket and comprises a load cabin, a propelling cabin is arranged at one end of the load cabin, a consignment cavity is formed in one end of the load cabin, and a supporting cavity is formed in the other end of the load cabin. A consignment cavity is formed in the base, one end of the consignment cavity is connected with an airship nose cone, an air-drop platform is arranged in the consignment cavity, the air-drop platform comprises a first compression releasing machine, one end of the first compression releasing machine is in transmission connection with an air pressure lifting platform, and the airship nose cone comprises a landing cabin body. According to the method, safe recovery and reuse after launching can be achieved, intact recovery of airships and rockets is ensured, the launching cost is greatly reduced, the resource utilization rate is increased, and a solid technical foundation is provided for wide application of sub-orbital aircrafts in the fields of commercial spaceflight, scientific research, emergency response and the like.

The invention is notably directed to an aircraft (10), which includes an aircraft structure (20), a thermal exchange system, a thermal insulation material (26), and fasteners (32, 35). The thermal exchange system includes a fluid-based cooling circuit (70 - 73) and heat exchanger tiles (25). The latter cover, at least partly, the aircraft structure. They may thus form the skin of the aircraft. The fasteners (32, 35) secure the tiles (25) to the aircraft structure. Each tile of the heat exchanger tiles embeds one or more fluidic channels (254), which are connected to the fluid-based cooling circuit, whereby the thermal exchange system is adapted to actively cool the tiles, in operation. Remarkably, the thermal insulation material (26) extends between the aircraft structure and the heat exchanger tiles. The thermal insulation material (26) is clamped on either side by the heat exchanger tiles (25) and by the aircraft structure (20), thanks to the fasteners (32, 35). The latter can for instance be bolted joints, which extend from central areas of the tiles. Some of the fasteners may embed sensors (40The invention is further directed to methods of operating such an aircraft. Such methods operate the fluid-based cooling circuit to circulate a cooling fluid through the fluidic channels (254) of the heat exchanger tiles (25) to cool them down. The present methods may further harness aerodynamic heating to help evaporate a liquid coolant and form the cooling fluid as a gas.

The invention relates to a vehicle, in particular an aircraft or a space vehicle for a supersonic or hypersonic flight, comprising a fuel tank (3), a main propulsion system (4) driven by the fuel, and a cooling system (7, 8) with a cooling channel (7) for cooling a heat source (1). The fuel of the fuel tank acts as a coolant in the cooling system and absorbs heat while being conveyed through the cooling channel (7). The vehicle further comprises a reaction control system (12) with gas thrusters (2). The reaction control system (12) is connected to the cooling system (7, 8), wherein the reaction control system (12) is driven by the heated, gaseous fuel from the cooling system (7, 8).

The present invention relates to a retractable rotor-type Mars manned aircraft overall device and a deployment method thereof, and belongs to the technical field of aerospace Mars rotor-type aircraft. With respect to the existing Mars aircraft overall configuration, there is still a lack of a retractable eight-axis sixteen-rotor Mars manned aircraft overall configuration. A retractable rotor-type Mars manned aircraft overall device and a deployment method thereof are provided, comprising a main frame, a seat, a landing gear, and a rotor assembly. The main frame is connected to the rotor assemblies arranged equidistantly in the circumference. A seat is provided in the middle of the main frame, and a landing gear is provided below the seat. The rotor assembly comprises forward blades, a motor, a rotor arm, and reverse blades. The forward blades and reverse blades are connected to the motor, and the coaxially arranged forward blades and reverse blades are connected to the main frame via the rotor arm. The present invention has a stable structure, simple operation, a small folded envelope size, can complete manned flight, and is of great significance to the overall configuration of multi-rotor Mars manned aircraft.

The invention relates to a near space aircraft adjustable head cone and an adjusting method thereof, and belongs to the technical field of mechanism science and the technical field of aerospace. The adjustable nose cone comprises a first-stage actuator, a second-stage actuator, a third-stage actuator, an aircraft nose cone shell and an aircraft nose cone skin. The aircraft nose cone shell is covered with the aircraft nose cone skin, and the aircraft nose cone shell is connected with the first-stage actuator, the second-stage actuator, the third-stage actuator and the aircraft nose cone skin in a bearing mode and provides structural support for the aircraft nose cone skin. The first-stage actuator, the second-stage actuator and the third-stage actuator are connected in series, can stretch out and draw back along the central axis of the aircraft nose cone and deflect by a preset angle relative to the central axis of the aircraft nose cone, so that the angle and the stretching length of the aircraft nose cone are adjustable.

Implementations of the disclosed subject matter may provide an adaptive thrust structure (400) for a space launch vehicle that may include an interface block (414) that is fluidically coupled to a fuel supply (402), a fuel manifold (406), an oxidizer supply (404), and an oxidizer manifold (408). A first adapter plate (418) may be coupled to a first engine assembly (419). First adapter lines (416) may be fluidically coupled to the first engine assembly (419) and the interface block (414), via the first adapter plate (418). 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 (101), and a reusable and reconfigurable transporter (300) having a plurality of zones that are configurable to transport cargo, one or more passengers, and / or one or more satellites.