292results about How to "Lower launch costs" patented technology

The invention relates to an anchor positioning system for detecting a lander, in particular to the anchor positioning system for detecting a planetoid lander. The anchor positioning system solves the following problems of a telescopic sleeve connecting and positioning system: an anchor bill is inclined when in incidence and transverse impact load is acted on the lander, so as to damage instruments at the inner part of the lander; a moving part of a telescopic sleeve is changed in quality during the incidence, so as to cause difficulty in estimating initial emission energy. An advancing mechanism is fixedly arranged on the side face of a cable box body, a winding mechanism is fixedly arranged on a rear end cover of the cable box body, a locking and unlocking mechanism is fixedly arranged on a right shell body of the winding mechanism, a connecting rope is stored in the cable box body, one end of the connecting rope is fixed on an anchor rod pin, and the other end of the connecting ropeis fixedly connected with a winding reel. A winding motor and an unlocking motor are connected with a pin socket by a lead wire. The anchor positioning system can also be used not only for detecting the lander of a small celestial body with weak surface gravitation such as the comet and the like, but also for automatic operation of a robot in dangerous environments.

A single-stage hypersonic vehicle is comprised of a low-speed and a high-speed propulsion system. The low-speed propulsion system propels the single-stage vehicle to a threshold velocity, after which the high-speed propulsion system then takes over. The low-speed propulsion system includes a combined-cycle engine featuring a swirl generator that is integrated into a turbojet engine to provide a compact turbojet and swirl afterburner-ramjet propulsion system. The high-speed propulsion system includes a hypersonic engine that is operable at the threshold takeover velocity and beyond. In various embodiments, the high-speed propulsion system comprises a scramjet, rocket, or scramjet/rocket engine depending requirements. Benefits of the swirl generator design include its ability to rapidly and efficiently atomize, vaporize, mix and burn the fuel and oxidizer for the low speed propulsion system, significantly reduce length, weight, cooling requirements and complexity for both propulsion systems, while maintaining high propulsion performance and reducing propulsion and launch costs.

The invention discloses a folding lightweight landing mechanism and relates to the landing mechanism of a lander. The invention aims at solving the problems of large volume, complexity in a buffer structure and high cost of the existing landing mechanism of the lander during the launching process. The folding lightweight landing mechanism comprises a triangular space truss type support, a buffer component, an instrument platform and three landing legs, wherein the triangular space truss type support comprises a triangular support, a star-shaped support, a space truss connecting rod component, three groups of main landing leg connecting components and three groups of auxiliary landing leg connecting components, the triangular support and the star-shaped support are arranged at the upper part and the lower part, the triangular support is connected with the star-shaped support through the connecting rod component, the three landing legs are distributed and connected on the periphery of the triangular space truss type support through the main landing leg connecting components and the auxiliary landing leg connecting components, and the triangular space truss type support is connected with the instrument platform through the buffer component. The folding lightweight landing mechanism is applicable to lander exploration activities for asteroids, comets, the moon, the Mars and the like.

The invention relates to a magnetic liquid damp shock absorber and belongs to the field of mechanical engineering vibration. The magnetic liquid damp shock absorber solves the problems that under the condition that external amplitude is larger than a gap between a shell body and a permanent magnet of an existing magnetic liquid damp shock absorber, the permanent magnet collides with an inner wall of the shell body and thus service life of the shock absorber is shortened and shock absorption performance is influenced. The permanent magnet is large in specific gravity and difficult to process to a required shape. A cylindrical permanent magnet of the magnetic liquid damp shock absorber is a rubber cylindrical permanent magnet (7). A wedge angle gasket (5) is arranged at the bottom inside the shell body (4) and a V-shaped included angle of the wedge angle gasket (5) faces upwards. The rubber cylindrical permanent magnet (7) absorbing magnetic liquid (6) is arranged in the shell body (4) and is magnetized along the axial direction. The axis of the rubber cylindrical permanent magnet (7) is perpendicular to a horizontal plane. The V-shaped included angle of the wedge angle gasket (5) is 60 degrees to 170 degrees. The gap between the inner wall of the shell body (4) and the rubber cylindrical permanent magnet (7) is smaller than or equal to the amplitude of shock absorption required by the shock absorber. The magnetic liquid damp shock absorber is simple, reliable, obvious in shock absorption effect and long in service life.

The invention is the curved self-separating bullet pallet, and relates to the improvement of the pallet structure. The invention aims to solve the unstable separation of the straight bullet pallet. The sections of the left distinguish 1 and the right distinguish 2 are two corresponding semicircles; the upper part of one side of the inner face 11 of the left distinguish 1 opens the left bullet hole 3; the inner face 11 of the left distinguish 1 has left tooth profile 5; the upper part of one side of the inner face 12 of the right distinguish 2 opens the right bullet hole 4 that is corresponding to the left bullet hoe 3; the inner face 12 of the right distinguish 2 has the right tooth profile 6 that engages with the left tooth profile 5; the left distinguish 1 is the curved body 9 and the right distinguish 2 is the curved body 10 corresponding to the left distinguish. The invention improves the launching efficiency and lowers the cost.

The invention discloses a separation modular satellite system and a method based on MEO data relay. The system comprises a separation modular satellite cluster, an MEO data relay system and an earth station, wherein the separation modular satellite cluster generates data after onboard processing and sends the data to the MEO data relay system; the MEO data relay system performs routing on the data and sends the data to the earth station in a transparent forwarding mode; and the earth station performs further analysis processing on the received data to obtain a final result. According to the system and the method, end-to-end delay of the separation modular satellite system can be effectively reduced, and therefore real-time data transmission of the separation modular satellite system is achieved.

On the basis of defining the concept of a spacecraft solar panel conformal antenna, the invention discloses S1, a spacecraft solar panel conformal antenna structure; S2, a spacecraft solar panel conformal antenna array structure; S3, a spacecraft solar panel conformal antenna array feeding mode; S4, a spacecraft solar panel conformal antenna beam control network; S5, a spacecraft solar panel conformal antenna microwave signal transmission line; S6, a connecting method of the spacecraft solar panel conformal antenna and a spacecraft. The spacecraft solar panel conformal antenna has the advantages of reducing the spacecraft launching cost, prolonging the in-orbit service life of the spacecraft, reducing the fairing requirement, facilitating appearance attractiveness of the spacecraft, improving spacecraft maneuverability and safety, and being high in replicability and good in expandability. Multiple solar panel conformal antennas with different orbit heights and types can be applied through networking, and cover with a bigger bandwidth range and a wider spatial range can be achieved through mission planning. Besides, the conformal antenna thought can be expanded to the field of interstellar probe antennas. The spacecraft solar panel conformal antenna has a wide application prospect.

The invention discloses a solar saucer-shaped floating aircraft provided with a shrinkable airbag. The solar saucer-shaped floating aircraft provided with the shrinkable airbag comprises a saucer-shaped aircraft body which is transformed from a saucer-shaped motor, and a propeller is arranged in the middle of a rotor of the saucer-shaped aircraft body. The shrinkage airbag is installed on the edge of the saucer-shaped aircraft body, and a solar power generation device is installed on the skin of the upper portion of the airbag. The lifting force of the aircraft body can be provided by air in the airbag and the propeller in the middle of the aircraft body independently or jointly, and power, in the horizontal direction, of the aircraft body is provided by a steering engine. The shrinkage airbag is mainly composed of airbag skin, a pull pressing rod system in the airbag and a telescopic system supporting the skin. The solar power generation device is in a fan shape, is circularly distributed on the skin in an attached mode, and moves along with the stretching or retracting of the skin.

The invention discloses a method for preparing a fiber-reinforced ceramic matrix composite material with a low thermal expansion coefficient, comprising: firstly, carrying out structural weaving of a fiber preform, and the material of the fiber preform is carbon fiber or silicon carbide fiber, or a combination of the two ; secondly, the interfacial phase is prepared on the surface of the fiber preform; thirdly, the chemical vapor deposition process is used first, and then the fiber preform is densified by the impregnation and cracking process of the organic precursor. The invention can obtain the fiber-reinforced ceramic matrix composite material with low coefficient of thermal expansion.

The invention discloses a yawing maneuvering control method based on a sinusoidal yawing guidance principle. The sinusoidal yawing guidance principle is designed in the yaw axis maneuvering method of an aircraft, thus, a yaw axis follows a sinusoid and the rotating speed of a sailboard is calculated based on the sinusoid at the same time, so as to ensure that an inclined angle between the sun direction and the normal of the sailboard is minimum in this mode. With the adoption of the sinusoidal yawing guidance principle, the energy of the aircraft is ensured, the demand of original yawing maneuvering mode to the capacity of an angular momentum exchange device is reduced, the demand on the capacity of the angular momentum of the aircraft is reduced, and a method is provided for reducing the weight of a control system.

The invention discloses a power density adjustable film reflecting and condensing space solar energy collecting station, which comprises a spacecraft platform, an optical system and a light path adjusting mechanism, wherein the optical system comprises a main condenser, an auxiliary condenser and a plane mirror; the main condenser and the auxiliary condenser are arranged in opposite directions, optical axes of the main condenser and the auxiliary condenser are overlapped, and a normal distance between the main condenser and the auxiliary condenser is adjusted through the light path adjusting mechanism to change the power density of output light beams; the angle of the plane mirror is adjusted through the light path adjusting mechanism to change the direction of the output light beams; sunlight is condensed and reflected to the auxiliary condenser through the main condenser, the reflected light rays are condensed and reflected to the plane mirror through the auxiliary condenser, and the light beams are reflected to a target through the plane mirror.

The invention discloses an efficient isotope battery based on a gas radioactive source. The isotope battery comprises a battery base with an embedment groove, a radioactive light source support put in the embedment groove of the battery base, at least one radioactive source radiation light-emitting unit put on the radioactive light source support, a lens buckled on the embedment groove of the battery base and isolating the inside and the outside of the embedment groove, an output end support connected with the battery base, and a photovoltaic conversion unit put on the output end support, matched with the lens in position and used for outputting electric energy, wherein light rays emitted by all the radioactive source radiation light-emitting units are projected to the photovoltaic conversion unit after being gathered by the lens. A gas isotope radioactive source is ingeniously used for stimulating a radioactive light-emitting material to emit light, light rays are gathered on the photovoltaic conversion unit by the lens through the combination of an optics principle so that electric energy can be output, the light intensity on the photovoltaic conversion unit is greatly improved, the efficiency of the photovoltaic conversion unit is greatly improved, and the output power and the energy conversion efficiency of the isotope battery are improved.

The invention discloses a connecting and separating device for satellite batch launching. The connecting and separating device comprises a base, docking mechanisms and a locking mechanism. The dockingmechanisms are connected with the satellite body, the multiple docking mechanisms are arranged in an overlapped mode, elastic mechanisms are arranged between two adjacent docking mechanisms and between the bottommost docking mechanism and the base, the two adjacent docking mechanisms are connected in an abutting mode through the corresponding elastic mechanism, and the bottommost docking mechanism is connected with the base in an abutting mode through the corresponding elastic mechanism. The base is connected with the docking mechanism through the locking mechanism, and the locking mechanismis used for fixing the docking mechanism to the base. By adopting the scheme provided by the invention, large-batch satellite launching can be simultaneously realized, the launching cost is greatly reduced, and guarantee is provided for satellite large-batch launching and constellation construction; the design meets the expandability of one-rocket multi-satellite launching and the uniformity of satellite design and manufacturing, all satellites are convenient to assemble and loose in requirement before launching, the safety requirement is met in the launching process, and on-orbit unlocking issafe and convenient.

The invention discloses a spacecraft attitude control method for controlling moment gyro singularity avoidance, specifically comprising the steps as follows: S1, solving an SGCMG (single gimbal control moment gyro) group angular momentum set; S2, deciding the angular momentum amplitude of an SGCMG group and spacecraft system; S3, determining the optimal control performance index of a spacecraft attitude control system; S4, determining a linear model of the SGCMG group and spacecraft system; S5, determining the linearization range of the SGCMG group and spacecraft system; S6, determining the linearization range meeting the constraints; and S7, using nonlinear predictive control to implement spacecraft attitude control. By using the method, the attitude of a spacecraft can be controlled precisely, the spacecraft launching cost is reduced, and the in-orbit operation life of spacecrafts is increased.

The invention discloses a light mechanical collision type space debris capturing web claw. The light mechanical collision type space debris capturing web claw comprises a bottom plate and a pluralityof web claw teeth, wherein the plurality of web claw teeth are mounted on the bottom plate; toothed claw switch module for driving the web claw teeth to be closed when the web claw teeth have collisions with a space debris is also mounted on the bottom plate; the web claw teeth consist of claw teeth with spokes on the two sides; and when the web claw teeth are closed, the plurality of web claw teeth and the spokes on the two sides of the plurality of web claw teeth form a semi-closed capturing envelope to prevent a capturing target from escaping. According to the light mechanical collision type space debris capturing web claw, at the moment of approaching the space debris in the capturing process, the capturing envelope is formed by triggering the switch through triggering a collision triggering switch of the bottom plate and releasing a torsion spring to drive the closed claw teeth; and therefore, the motion of a non-cooperative target is restricted and limited to finally realize thecapturing.