32results about "Space shuttles" patented technology

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).

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.

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 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.

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).

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 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 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.

The invention belongs to the technical field of unmanned aerial vehicles, and discloses a dual-rotor Mars unmanned aerial vehicle and a control method. The Mars unmanned aerial vehicle is composed of a front rotor wing, a rear rotor wing, a front rotor wing control system, a rear rotor wing control system, a front power motor, a rear power motor, a vehicle body, an equipment cabin, an undercarriage and a solar cell panel. The fuselage adopts a carbon fiber truss type structure; the Mars unmanned aerial vehicle adopts a tandem double-rotor layout, and a front rotor and a rear rotor are the same in rotating speed and opposite in rotating direction; a front rotor wing and a rear rotor wing of the Mars unmanned aerial vehicle are vertically arranged in a staggered mode, and the front rotor wing and the rear rotor wing are both rigid rotor wings and free of waving deformation. The front rotor control system is located below the front rotor, and the rear rotor control system is located below the rear rotor; the front power motor and the rear power motor respectively drive the front rotor wing and the rear rotor wing to provide pulling force; the undercarriage is a strut-type undercarriage, and the equipment compartment is fixed at the bottom of the fuselage.

The invention discloses a sub-orbital reusable manned tourist aircraft, and relates to the technical field of spaceflight, the sub-orbital reusable manned tourist aircraft comprises a booster and a load cabin; the booster comprises a multifunctional cabin, a liquid oxygen storage tank, an inter-tank section, a kerosene storage tank, a transition section and a tail section, three engines are arranged on the transition section, one engine is located in the center of the transition section, the two engines are arranged in an axial symmetry mode with the engine located in the center of the transition section as the center, and the three engines are arranged in a straight line after being connected; an environmental control life support system is arranged in the load cabin; four reverse thrust engines are arranged at the bottom of the load cabin; a parachute cabin is arranged at the top of the load cabin, and a main parachute device and a backup parachute device are arranged in the parachute cabin. By changing the layout of the engine, the propellant type and the working mode, the recovery pneumatic speed reduction and attitude control mode is adjusted, and the device has the advantages of being simple in structure, efficient, practical, safe and reliable.

The application discloses a method and system for quickly eliminating nutation, wherein the method comprises the following steps: S100, installing a thruster and a gyroscope; S200, establishing a composite model of a sine of a lateral angular velocity and a constant term; S300, setting an optimization index form, a variable to be optimized, a constraint condition of the variable to be optimized and particle swarm algorithm parameters; S400, obtaining a preliminary estimation result of the variable to be optimized by using the particle swarm algorithm; S500, obtaining a final estimation result of the variable to be optimized by using a sequential quadratic programming; S600, obtaining an angular velocity half-cone angle according to the final estimation result of the variable to be optimized; S700, if the angular velocity half-cone angle is smaller than a preset value, ending, otherwise, turning to step S800; S800, obtaining a thruster starting period; S900, if the thruster starting period is in a starting period, starting the thruster, if the thruster starting period is not in the starting period, turning to step S400. The application is simple in configuration and is beneficial to practical application.

The application discloses a land-air amphibious Mars exploration platform and belongs to the technical field of exploration platforms. The platform comprises a platform, a moving mechanism and an adjusting mechanism. The moving mechanism is arranged below the platform. The moving mechanism comprises a front wheel assembly and two rear wheel assemblies which are symmetrically arranged relative to the front wheel assembly. The adjusting mechanism is connected between the moving mechanism and the platform. The adjusting mechanism is used for adjusting the directions of the front wheel assembly and the rear wheel assemblies, and switching between a land driving mode and an air flight mode. The exploration platform can maximize the use of existing space, effectively reduce the width of the platform in the land driving mode, improve the passability of the land-air amphibious motion platform in a narrow environment, effectively reduce air resistance in the flight mode, and effectively solve the problems that the exploration platform cannot realize land-air amphibious motion, has poor cross-country performance and has a large size.

A rocket ship consists of several fuel tanks and is surrounded with two levels of rockets, one from the top part for the detachable dome and one from the bottom part of the rocket ship for its main body. This structure of the rocket ship is detachable and will help it to travel longer distance in space. The broad-based rocket ship has its own loading and unloading mechanism with an underground accessway for the infrastructures and other cargo and supplies to be mobilized into the cargo bay. These cargos, supplies and infrastructure will be transported and be used in the rocket ship space travel. Specifically for the infrastructures loaded into the rocket ship, these will be unloaded into the space to serve another function of offering space and work tourism. With the functions mentioned, this does not limit the endless possibilities of broad-based rocket ship.

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.

A vehicle configured for in-space and atmospheric reentry operations includes a pressure-fed propulsion engine that uses cryogenic fuel and cryogenic oxidizer as propellants. In some embodiments, the vehicle includes a first conduit configured to provide a first propellant to the pressure-fed propulsion engine, a second conduit configured to provide a second propellant to a heat shield heat exchanger disposed relative to a heat shield wall defining an outer surface of the vehicle, and a pump located along the second conduit. In such embodiments, the first conduit is configured to provide the first propellant to the pressure-fed propulsion engine at a first mass flow rate, the second conduit and the pump are configured to provide the second propellant to the heat shield heat exchanger at a second mass flow rate, and the second mass flow rate is substantially less than the first mass flow rate. A related method is also disclosed.

This application belongs to the field of aerospace technology and relates to an auxiliary control method for spacecraft. It is based on a towed weather vane auxiliary rudder, the auxiliary rudder body being a symmetrical dart shape, towed to the rear of the spacecraft by a rigging system. A release cable controls the release of the auxiliary rudder body from the pod to a predetermined position, a control cable controls the pitch deflection of the auxiliary rudder, and a static cable maintains the relative attitude stability between the auxiliary rudder and the spacecraft. The control method described in this application can improve the handling and stability characteristics of the spacecraft in certain speed and altitude ranges at a relatively low cost, freeing up overall design constraints. Designers can focus more on optimizing performance indicators in other speed ranges, improving the overall performance of the spacecraft. This method can also be applied as an emergency control measure to existing spacecraft, improving the controllable flight envelope of existing spacecraft in emergency situations and contributing to improved safety.

The invention provides a lightweight prestressed cable-rod panel supporting structure which comprises a main structure framework, an inhaul cable assembly and a panel, the main structure framework is of a prism structure formed by rigidly connecting a plurality of truss rod pieces through a plurality of multi-way wrapping and sleeving pieces, each multi-way wrapping and sleeving piece comprises a plurality of sleeves and a fixing part, and the truss rod pieces are installed in the matched sleeves; the inhaul cable assembly comprises a plurality of inhaul cables, two sliding rings and a connecting rod, one ends of the inhaul cables are connected with the multi-way package fixing parts on the upper surface and the lower surface of the main structure frame respectively, the other ends of the inhaul cables are divided into two sets to be connected with the sliding rings respectively, and the two sliding rings are fixed to the connecting rod. The panels are fixedly connected to the upper surface and the lower surface of the main structure framework, the face, facing the inhaul cable assembly, of each panel comprises a plurality of plate ribs, and the two ends of each connecting rod are connected with the plate ribs through connecting structures. According to the invention, the out-of-plane static and dynamic stiffness of the large-breadth panel in wing and rudder parts is effectively improved, the structure is simple and efficient, and the lightweight design is favorably realized.

The invention provides a lightweight prestressed inflatable bag-cable panel supporting structure which is characterized in that a main structure framework is of a prism structure formed by rigidly connecting a plurality of truss rod pieces through a plurality of multi-way sheathing pieces, each multi-way sheathing piece comprises a plurality of sleeves and a fixing part, and the truss rod pieces are mounted in the matched sleeves; the panels are connected to the upper and lower surfaces of the main structure framework; the inhaul cable assembly comprises 2n stay cables, a vertical inhaul cable and two first connecting structures, the n stay cables are connected with the fixing parts of the n multi-way wrapping and sleeving parts on the upper surface of the main structure framework and the first connecting structures I correspondingly, and the other n stay cables are connected with the fixing parts of the n multi-way wrapping and sleeving parts on the lower surface of the main structure framework and the second connecting structures I correspondingly. The vertical inhaul cable is connected with the first and second connecting structures I; and the air bags are filled between the panels and the adjacent stay cables and are used for supporting the panels. According to the invention, the out-of-plane static and dynamic stiffness of a large-format panel in wing and rudder parts can be effectively improved, the structure is simple and efficient, and the lightweight design can be realized.

According to an aspect of the present invention, there is provided 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. According to another aspect of the present invention, there is provided 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.

This invention discloses a variable-diameter frame cable-driven mechanism for the nose cone tip of an aircraft, comprising multiple frame branches, a top plate, and a base platform. Each frame branch includes a first frame section, a second frame section, a third frame section, a fourth frame section, a fifth frame section, and a sixth frame section. The base platform is stepped and has three layers of truncated cones, with multiple frame branches evenly distributed in parallel on the base platform. A pulley is provided in the first frame section to fix the end of the steel cable. Two pulleys are provided in the second, third, fourth, fifth, and sixth frame sections for positioning the steel cable around each frame section. A pulley is located below each positioning section of the base platform to ensure that the initial position of the steel cable remains unchanged when controlling the movement of the frame sections. This invention can vary the nose cone diameter according to the different requirements of the aircraft's optimal aerodynamic layout during the ascent and reentry phases.

The invention provides a distributed type driving aircraft morphing nose cone capable of achieving stretching and bending, and belongs to the technical field of morphing aircrafts. Comprising a shell assembly, a driving chain and a driven chain. The shell assembly comprises N shells which are sequentially arranged from bottom to top, N is a natural number larger than or equal to 4, and the diameters of the shells are gradually decreased from bottom to top to form a blunt conical shape; the driving chains are of RRR unit structures, are distributed on the left side and the right side in the shell assembly and are connected with the adjacent shells; the driven chains are of UPU unit structures, are distributed on the front side and the rear side in the shell assembly and are connected with the adjacent shells; the driving chain is used for driving the variant nose cone to realize telescopic deformation and bending deformation; and the driven chain moves along with the deformation of the nose cone, so that motion constraint and stability are provided for the deformation of the variant nose cone. Two deformation modes of stretching and bending of the nose cone are achieved through independent control of the driving chains on the left side and the right side, the driven chains on the front side and the rear side provide guiding and stable supporting, and the device has the advantages of being simple in structure, stable in movement and flexible to control.

The invention provides a lightweight prestressed cable-rod substructure which comprises a main structure framework, an inhaul cable assembly and a screw and bolt kit, the main structure framework is of a prism structure formed by rigidly connecting a plurality of truss rod pieces through a plurality of multi-way wrapping kit pieces, each multi-way wrapping kit piece comprises a plurality of casing pipes and a fixing part, and the truss rod pieces are installed in the matched casing pipes; the inhaul cable assembly comprises a plurality of inhaul cables and two tie plates, one ends of the inhaul cables are connected with the multi-way sheath fixing parts at different positions, and the other ends of the inhaul cables are connected with the two tie plates correspondingly; the screw and bolt kit comprises a screw and bolts, the screw penetrates through the centers of the two tie plates, the two bolts are installed at the two ends of the screw, the positions of the tie plates on the screw are adjusted by adjusting the positions of the bolts, and prestress adjustment of the inhaul cable assembly is achieved. According to the invention, the out-of-plane static and dynamic stiffness of the large-breadth panel in wing and rudder parts is effectively improved, the structure is simple and efficient, and the lightweight design is favorably realized.

Embodiments provide a spacecraft airlock system. Embodiments provide a method and apparatus for attaching space exposed payloads to a space station. The spacecraft airlock system provides a defined volume of space payload to the international space station. The airlock further includes a means of attaching to a space station, a closed structure attached to said means, said means of attaching is capable of robotic manipulation, and a cooling system for cooling payload components within said closed structure.

The application relates to an aircraft fuselage heat dissipation system. The system comprises a heat exchange section arranged on an aircraft, wherein a first heat exchange section is arranged on the windward surface of the aircraft, and a second heat exchange section is arranged on the leeward surface of the aircraft; the first heat exchange section and the second heat exchange section are communicated through a communication assembly; the first heat exchange section is filled with cooling working medium; and the cooling working medium flows to the second heat exchange section through the communication assembly after heat absorption phase change. The application can realize efficient heat conduction, solve the destructive problems caused by the deformation of the fuselage structure due to thermal stress, and also plays an optimizing space role.

A vehicle for supersonic or hypersonic flight comprises a thermal rocket engine (lb) with a nozzle (2) and an active cooling system (8). The active cooling system cools a heat shield (6, 7). A working fluid absorbs heat inside the active cooling system and the heated working fluid expands through the nozzle to create thrust. Such a vehicle is suitable to fly a multi-skip trajectory, a boost-glide trajectory, a trajectory with a cruise phase or a re-entry into an atmosphere, for example.