91 results about "Space launch" patented technology

A method of launching space vehicles by towing them aloft, then twirling them around a large transport aircraft (40) at the center of a formation (AA) of other tow aircraft (28, 34) and other devices of the invention. A lengthy, semi-rigid tow pipeline (14) serves as a conduit for the transfer of fuels and oxidizers, as the tow cable, and as an energy storage device that reflexes efficiently when it is flexed. The flexing of tow pipeline (14) is caused by a parachute (22) acting in conjunction with all the aircraft making the tighest turn they are capable of doing. Tow aircraft in certain arrays (28) are joined to tow pipeline (14) by sliding trollies (26) that also host canard rotor wings for the aerodynamic support of the main tube (12). The tow trollies (26) aid the sliding tow aircraft arrays (28) in gaining mechanical advantage to accelerate the space vehicle. The space vehicle may also burn its own motors for a longer-than-usual time as it gains in angular velocity because its fuels are replenished by a pipeline that automatically increases the pressure and volume of fuel flow as the angular acceleration increases. The tow pipeline (14) features a micro-hole laminar lift foil (50) on the top surface of the wing and on the bottom. By balancing the vacuum level in the plurality of chamber underneath the plurality of lift foil 50, the pilot or autopilot has an effective way to control the flight characteristics of the tow pipeline (14). In some embodiments the micro-hole laminar lift foil (50) on the top and bottom will made of transparent hollow wire segments or other material so arranged as to pass ambient light through to photo-voltaic cells below that convert sunlight to electrical energy.

The invention discloses a satellite and launch vehicle integrated aerospace system, and relates to a satellite and rocket integrated aerospace system. The system aims to solve the problems that a small satellite is difficult to flexibly adapt to the requirements of different loads of emergency disaster monitoring and that the maximum capability of a small space launch vehicle is restricted in terms of the orbit injection quality of the launched satellite. The system comprises a mission payload bay and a launch vehicle. A high-resolution camera and a data measurement and control module are arranged in the mission payload bay. The launch vehicle comprises a first-stage launch rocket, a second-stage launch rocket and a third-stage launch rocket. The system further comprises a shared service capsule and a controller area network (CAN) bus. A data acquisition and control instruction module, a propelling module, a posture and orbit control module, a power module, an apparatus mounting module and a thermal control module are arranged in the shared service capsule. The data acquisition and control instruction module controls the mission payload bay and the launch vehicle through the CAN bus. The system is applied in the field of aerospace.

The invention discloses a landing point forecasting system for space launch rocket. According to the invention, the landing point forecasting system is connected to a spaceflight test and control network. Rocket debris and rocket flight position and speed information are extracted from data provided by a radar data processing system. Rocket debris separation time is extracted according to information provided by a remote measurement data processing system. Initial separation point data is obtained through fitting simulation. A rocket debris flight dynamics model is established through combination of flight environment mathematics modeling and aerodynamic characteristics analysis. Through landing point influence factor analysis, landing point accurate forecasting is performed, so that a landing area search range is reduced and rocket debris recycling efficiency is improved.

Propellant management systems and methods are provided for controlling the delivery of liquid propellants in a space launch vehicle utilizing multiple rockets. The propellant management systems and methods may be configured to enable substantial simultaneous depletion of liquid propellants in each of a plurality of active rockets during operation of various booster stages of the launch vehicle.

A system and a related method are disclosed for matching space launch vehicle, satellite, and mission hosted payload opportunities utilizing multi-level, priority algorithms. The potential matches include numerous mission-matching scenarios, e.g., 1) matching hosted payload and host spacecraft, and 2) matching spacecraft(s) and launch opportunities, to include rideshare opportunities. The matching opportunities can be pursued by any interested project team, e.g., hosted payload, spacecraft, and space launch vehicle.

A landing buffer device of a reusable launch vehicle relates to the field of space launch vehicles and comprises a rocket body and four support legs, wherein the rocket body is of a vertically-arranged cylinder structure; the four support legs are symmetrically fixedly mounted to the side wall of the axial bottom end of the rocket body in a mutually perpendicular manner; the four support legs actsimultaneously and rotate along the joint between the support legs and the side wall of the rocket body to achieve the folding or unfolding of the four support legs; the diameter of the rocket body is120-140 mm; the height of the rocket body is 510-530 mm; the total mass of the rocket body is 1.2-1.3 kg; and the included angle between the rocket body and the vertical direction after the landing of the rocket body is no more than 10 degrees. The landing buffer device of the reusable launch vehicle is capable of realizing the repeated folding, unfolding and deformation by the four-legged motionto make support preparation for the landing of the launch vehicle, and is of great significance to the fields of adaptation to a landing motion sequence, self-adaptive attitude adjustment, stable landing achievement, efficient buffer against ground impact and so forth of the landing device for the reusable launch vehicle.

The invention discloses a preparation method of calcium hydroxide nano particles with high adsorption activity. The preparation method is characterized by comprising the following steps: putting CaCl2 in a three-necked bottle; adding deionized water to dissolve the CaCl2; stirring with a magnetic stirrer; weighing SDS or PEG, wherein the mass ratio is as follows:CaCl2:SDS is equal to 22.2:(0.47-2.34) or CaCl2:SDS is equal to 22.2:(2.65-14.44); adding SDS or PEG into the three-necked bottle; stirring to dissolve; dissolving NaOH in the deionized water, wherein mass ratio of the CaCl2 to the NaOH is 22.2:8;dropwise adding an obtained solution to the three-necked bottle at the speed of 40ml/h; stirring to react for 0.5-2.5 hours to obtain slurry; performing suction filtering on the slurry; and washing and drying. According to the preparation method, the prepared Ca(OH)2 power body is uniform in particles and are in single dispersion; the particle granularity is 200-300nm and the specific surface area is 88m<2>/g; the maximum N2 adsorption amount can reach 80ml/g; meanwhile the removal rate of NOx can reach 98%; the preparation method has important practical significance on leakage treatment of liquid propellant N2O4 at a space launching site.

A space vehicle and its launcher are capable of low-cost launch, safe maneuvering, plus powered go-around for landing.

The invention discloses a spiral belt winding guide concentric launching cylinder, and belongs to the field of space launching. The spiral belt winding guide concentric launching cylinder comprises aninner cylinder, an outer cylinder and a guide cone; and spiral winding belts are encircled in an annular cavity between the inner cylinder and the outer cylinder, so that guided missile gas enteringthe annular cavity after passing through the guide cone is exhausted around a spiral cavity. Axial guide holes are formed in the ends, near the guide cone, of the spiral winding belts; and the guide holes in the same spiral winding belt are peripherally arranged at equal intervals. Through arrangement of the spiral winding belts, the time of thoroughly discharging gunpowder gas from the annular cavity is delayed; and the tail gas is exhausted in a direction deviated from the axis, so that the ablation influence on the guided missile wall surface during launching is prevented, and improvement of the guided missile launching safety is facilitated.

The invention discloses a heat-proof device for a swing type rocket engine nozzle, and belongs to the technical field of heat-proof and insulation devices for space launch. The heat-proof device for the swing type rocket engine nozzle comprises a nozzle flange, a cabin shell rear end cover and a heat-proof telescopic apron board; the nozzle flange is connected with an inner ring end surface of theheat-proof telescopic apron board, and the cabin shell rear end cover is connected with an outer ring end surface of the heat-proof telescopic apron board; the section of the heat-proof telescopic apron board is of a corrugated structure; and the heat-proof telescopic apron board is made of flexible ablation-resistant heat-proof materials. The problems that the structure is cracked due to poor ductility, and internal equipment is burnt by flame backflow since the apron board is of a flat plate type or conical surface heat insulation structure made of metal or composite materials are solved. The heat-proof telescopic apron board is made of flexible ablation-resistant heat-proof materials, and the section is of the corrugated structure, the flame of an engine can be prevented from flowing back to the interior of a cabin body, the returned flame is blocked from entering the cabin body, and the heat-proof device has high barrier property on ablation property and gas flow.

The invention provides an unfolding antenna basic unit, unfolding antenna and folding method based on origami. The invention relates to the field of aeronautical antennas, which solves the problems ofcomplicated deployment mechanical structure, small deployment area ratio and high cost of the traditional space deployable antenna structure. The invention comprises a shape memory material connecting piece, a heating device and a plurality of substrates. A shape memory material connector is used as a fold mark, and a heating device is arranged on the shape memory material connector, and the heating device is powered by a power supply. The expanded area of the basic unit of the expanded antenna is 4n times the folded area, n >= 1, n is an integer. The ratio of the unfolded area to the foldedarea is 4n, which can save the loading space to a great extent, thereby reducing the cost of aerospace emission. The plurality of unfolded antenna basic units of the invention can be combined and connected to realize different unfolded and folded area ratios. The invention is applicable to space launch.

A system and methods are disclosed for an upper stage space launch vehicle that uses gases from the propellant tanks to power an internal combustion engine that produces mechanical power for driving other components including a generator for generation of electrical current for operating compressors and fluid pumps and for charging batteries. These components and others comprise a thermodynamic system from which system enthalpy may be leveraged by extracting and moving heat to increase the efficient use of propellant and the longevity and performance of the launch vehicle.

The invention provides a carrier rocket previous level recycling system and method and a carrier rocket recycling system, and belongs to the technical field of carrier rocket recycling. The carrier rocket previous level recycling system comprises a counteraction control system, a controllable propelling system provided with a rocket engine arranged at the tail of a previous level, at least two duck wings symmetrically arranged at the head position of the previous level, at least two tail wings symmetrically arranged at the tail position of the previous level, and landing supporting legs arranged at the tail position of the previous level, wherein the counteraction control system is provided with a plurality of thrusters arranged on the side wall of the previous level, the multiple thrusters are evenly arranged on the side wall circumference, close to the head and tail positions, of the previous level. According to the carrier rocket previous level recycling system, carrier rocket previous level recycling is achieved, the higher carrying capability, the more advanced technology and other space launch market requirements are met.

The invention discloses a three-stage orbit injection space launch vehicle. The three-stage orbit injection space launch vehicle comprises a first-stage RBCC power aircraft, a second-stage RBCC power aircraft and a third-stage rocket power aircraft, wherein the second-stage RBCC power aircraft is arranged at the back of the first-stage RBCC power aircraft, and the third-stage rocket power aircraft is arranged in the second-stage RBCC power aircraft. According to the three-stage orbit injection space launch vehicle, the requirement for takeoff thrust is low, and the performance of engines and the whole efficiency are enhanced.

The invention relates to an automatic docking device and method for a right angle coordinate flexibility non-contact type connector used for space launch technology and in the transportation process of liquefied natural gas (LNG). The invention aims at providing the docking device and method for the connector, which have the advantages of high automation degree, strong adaptive capacity, simple structure, safety, reliability and the like. The automatic docking device for the right angle coordinate flexibility non-contact type connector comprises a pedestal, an advancing and retreating seat, atransverse movement seat, a lifting seat, a flexible clamping propulsion mechanism, a non-contact detection system and a connector component, wherein the advancing and retreating seat is installed above the pedestal, the transverse movement seat is installed above the advancing and retreating seat, the lifting seat is installed above the transverse movement seat, the flexible clamping propulsion mechanism is installed above the lifting seat, the non-contact detection system comprises a target plate and a non-contact sensor, the non-contact sensor is installed on the lifting seat, and the target plate is installed on a surface of a storage box; and the connector component comprises a connector, and the connector is used for connecting with an interface.

This invention relates to a propulsion bay to be transported, at least temporarily, in a space launch vehicle and comprising an adapter that co-operates with at least one system located, at least temporarily, on board the bay, said system comprising an electrical power supply. The bay is characterized in that it also comprises at least one electric space propulsion engine that can be powered by the power supply of the system.

The invention relates to a space launch site comprehensive drill system, comprising a master control subsystem and one or more role subsystems for respectively reading a predefined comprehensive drillscript control file. The script control file includes one or more operation steps. The master control subsystem selects the first operation step and its associated role subsystem, and sends the execution instruction of the first operation step to the role subsystem; the post operator is prompted to execute the first operation step; in response to the execution of the first operation step by the post operator, the role subsystem performs a preconfigured action and sends the execution result to the master control subsystem; the master control subsystem broadcasts the execution result to the oneor more role subsystems; the one or more character subsystems update the execution status of the role subsystem; and the master control subsystem selects the next operation step and repeats the abovesteps until all the operation steps in the script control file are completed.

The invention discloses a liquid hydrogen supercooling degree acquisition system for a space launching site. The system involves a liquid hydrogen storage tank, wherein a top pressurization opening ofthe liquid hydrogen storage tank is connected with an outlet of a high-pressure helium tank through a valve; a bottom outlet of the liquid hydrogen storage tank is connected with a liquid hydrogen side inlet of a helium-hydrogen heat exchanger through a circulating pump, and a liquid hydrogen side outlet of the helium-hydrogen heat exchanger is connected with a top reflux inlet of the liquid hydrogen storage tank through a reflux check valve; a helium side outlet of the helium-hydrogen heat exchanger is connected with a helium side inlet of a liquid hydrogen bath-type heat exchanger through acold helium compressor, and a helium side outlet of the liquid hydrogen bath-type heat exchanger is connected with a helium side inlet of the helium-hydrogen heat exchanger through a throttling valve; and a liquid hydrogen inlet of the liquid hydrogen bath-type heat exchanger is connected with the bottom outlet of the liquid hydrogen storage tank through a float regulating valve, and a hydrogen outlet of the liquid hydrogen bath-type heat exchanger is emptied through a hydrogen discharge valve. According to the system, a large number of liquid hydrogen cold sources existing in the launching site are used for precooling, and large-scale on-site preparation of the launching site with the large supercooling degree of liquid hydrogen can be achieved in combination with the cold helium compressor; and the equipment investment and transformation cost is low, and engineering implementation is easy.

The invention relates to a combined spacecraft. The combined spacecraft comprises a hard helium airship and a spacecraft body. The airship is connected with the spacecraft body in the manner that the airship can be mounted and unlocked. The airship is used for launching the spacecraft body at the high altitude in a sliding orbit launching or delivery launching manner after carrying the spacecraft body to the stratosphere. The spacecraft body is provided with a rocket engine capable of sending the spacecraft body into an orbit and achieving deceleration in a return voyage and attitude control. By the adoption of the combined spacecraft, the launching and recycling safety, reliability and convenience can be improved, the repeatable launch can be achieved, and therefore the space launching and recycling cost can be reduced greatly.