12results about How to "Lower launch costs" patented technology

This invention discloses a space radiating plate based on a large vapor chamber and its manufacturing method, belonging to the field of spacecraft thermal control technology. The radiating plate includes a one-piece sealed metal cavity, the interior of which is a vacuum vapor chamber. The cavity has opposing first and second sidewalls; the outer surface of the first sidewall forms a radiating surface, and the inner side of the second sidewall integrates cooling channels. A micro-pillar array structure, directly formed integrally through micromachining, is disposed on the inner wall surface of the vacuum vapor chamber, and the cavity is filled with a phase change working fluid. This invention highly integrates the vapor chamber, channels, and radiating surface, utilizing phase change heat transfer to achieve ultra-high thermal conductivity and temperature uniformity, solving the problems of low heat dissipation efficiency, anisotropy, large weight, and poor adaptability to the space environment in existing technologies. It is suitable for thermal management of high-power spacecraft such as satellites, space stations, and space data centers.

The utility model discloses an inclined nozzle angle engine shell suitable for a high-altitude aircraft, which comprises a shell assembly, a top cover assembly is matched with the top of the shell assembly, a bottom sealing cover is matched with the bottom of the shell assembly, four inclined nozzles are arranged on one side of the bottom sealing cover, and the four inclined nozzles are uniformly distributed. According to the engine shell with the inclined spray pipe angle suitable for the high-altitude aircraft, the spray pipe is obliquely arranged, the direction of hot airflow sprayed by the spray head is changed, the hot airflow sprayed downwards originally is sprayed towards the side face, and therefore the hot airflow sprayed towards the side face is changed into the hot airflow sprayed towards the side face. Damage of sprayed hot air flow to the launching pad and surrounding objects can be greatly reduced, launching cost is reduced, meanwhile, the reflecting plate is arranged below the spraying pipe, most of heat emitted by the spraying pipe can be reflected to the outer side perpendicular to the spraying pipe, and the influence of the heat of the spraying pipe on the surrounding objects is further reduced.

The present application relates to the field of spacecraft thermal control technology, and provides a two-phase fluid loop system for spacecraft heat dissipation, comprising an evaporation unit, a pump-driven two-phase fluid loop and a heat pump loop, the evaporation unit is provided with a liquid accumulator for storing heat transfer working medium, a micro-channel cold plate, the micro-channel cold plate comprises a micro-channel cold plate inlet, a micro-channel cold plate liquid phase outlet and a micro-channel cold plate vapor phase outlet; the pump-driven two-phase fluid loop is provided with a mechanical pump and a first radiator, the outlet of the first radiator is connected with the liquid accumulator liquid phase fluid inlet of the liquid accumulator; the heat pump loop is provided with a compressor and a second radiator, and the second radiator is connected with the liquid accumulator through a throttle valve. The present application can effectively reduce the weight of the system and the launch cost, and the micro-channel cold plate has the function of vapor-liquid separation in a closed system, which reduces the internal flow instability, improves the heat exchange capacity and the critical heat flux density, and solves the problems of mechanical pump cavitation and compressor liquid knock.

The utility model discloses a dual-wavelength optical path structure capable of reducing ASE noise, which belongs to the field of aerospace optical amplifiers and is used for reducing the ASE of light output by an optical amplifier, improving the OSNR and improving the sensitivity of a system. The device mainly comprises a jumper IN, a secondary amplification light path and an optical fiber ER, the jumper IN is connected with a wavelength division multiplexer M1, the wavelength division multiplexer M1 is connected with the input end of a wavelength division multiplexer W2 and the input end of a wavelength division multiplexer W5, the wavelength division multiplexer M1 screens two kinds of light with different wavelengths to form two light paths, and the two light paths are connected with the optical fiber ER. The output end of the wavelength division multiplexer W2 and the output end of the wavelength division multiplexer W5 are sequentially connected with the optical fiber ER and the filters BP, the output ends of the two filters BP are connected with the wavelength division multiplexer M2, the two optical paths are converged on the wavelength division multiplexer M2, and the wavelength division multiplexer M2 is connected with the secondary amplification optical path. The utility model is mainly used for spaceflight optical amplifiers.

A launch system and method for a launch vehicle that can save initial launch fuel includes a support structure, characterized in that: the support structure consists of a horizontal support frame (3) and an ascending section support frame (4), a horizontal track section (5) is installed on the horizontal support frame (3), an inclined track section (6) connected to the horizontal track section (5) is installed on the ascending section support frame (4), a launch acceleration pipe (10) is fixedly installed on the horizontal support frame (3) and the ascending section support frame (4), an electromagnetic catapult (14) is installed on the electromagnetic catapult slide rail (13), a rocket trolley (8) is installed on the horizontal track section (5) and the arc track section (6), and a rocket bracket (9) is provided on the rocket trolley (8). The present invention has a reasonable structure, can obtain sufficient initial velocity before launch ignition, can save launch fuel, can increase the effective payload, and is conducive to saving launch costs.

The application provides a power-oriented electric propulsion gas supply self-learning closed-loop control system and method, which is composed of a storage and supply unit, a power processing unit, a control unit and an electric thruster; the working current of the electric thruster is monitored, if the current is lower than the lower limit, the valve is opened to supplement the gas for a certain time length, the current value of the electric thruster working maximum 3s sliding average within a certain time length is monitored after the gas supplement is completed, and the gas supplement time length is adjusted according to the maximum value with a certain step distance, and the step distance is corrected according to the number of adjustments. The application solves the working condition drift problem of the electric thruster caused by the state change of the spacecraft or the erosion of the thruster itself in the prior art by using the multi-stage self-learning adjustment mode, under the premise of limited on-orbit computing resources, and collects the working parameters for data accumulation and open-loop control of the fault mode.

The invention relates to the field of space remote sensing measurement, in particular to a double-line-scan digital camera optical system based on multiplexing of a primary mirror and a secondary mirror, which comprises the primary mirror, the secondary mirror, a first third mirror, a second third mirror, a rear line-of-sight array focal plane and a front line-of-sight array focal plane, for a rear-view array imaging light path, a light beam is reflected by a first reflection area of a primary mirror, a secondary mirror, a second reflection area of the primary mirror and a first third mirror in sequence and then enters a rear-view array focal plane; for a front-sight array imaging light path, a light beam is reflected by the second reflection area of the primary mirror, the secondary mirror, the first reflection area of the primary mirror and the second third mirror in sequence and then enters a front-sight array focal plane. Through multiplexing of the primary mirror and the secondary mirror, two independent three mirrors are adopted to respectively adjust observation angles corresponding to a double-linear-array imaging light path, the system can flexibly adapt to asymmetric observation angles, and is excellent in image quality, compact in system structure and light in weight under the condition of a large relative aperture. And the method has important engineering significance for promoting the miniaturization and lightweight development of the double-linear-array surveying and mapping camera.

The application relates to a manufacturing method of a head structure of a carbon fiber composite material-based aerodynamic push rod, which comprises the following steps: dividing the head structure into five parts, i.e. a circular truncated cone-shaped upper shell, an inverted circular truncated cone-shaped lower shell, a supporting pipe, a flange seat and a circular truncated cone-shaped protective sleeve; setting a first thickened area and a second thickened area in the upper shell; setting a third thickened area and a fourth thickened area in the lower shell; adopting carbon fiber prepreg to carry out multi-layer pasting on the upper shell, and then inserting multi-layer carbon fiber prepreg into the multi-layer interval of the first thickened area and the second thickened area for pasting; adopting carbon fiber prepreg to carry out multi-layer pasting on the lower shell, and then inserting multi-layer carbon fiber prepreg into the multi-layer interval of the third thickened area and the fourth thickened area for pasting; adopting carbon fiber prepreg to carry out multi-layer pasting on the supporting pipe, and then carrying out curing and forming through a hot pressing tank process after the pasting is completed. According to the head structure requirements, reasonable optimization design of prepreg layer angle, layer sequence and layer proportion is carried out, the head structure bearing requirements are met, and overall weight reduction is realized.

The application discloses an inductance energy storage type pulse power supply with improved energy efficiency and belongs to the technical field of electromagnetic emission. The pulse power supply comprises a primary power supply module, an intermediate energy storage module, a reverse flow loop module, an energy self-recovery module, a load current opening and shutting module and a voltage polarity recovery module; the primary power supply module provides energy for charging the intermediate energy storage module; the reverse flow loop module is used for disconnecting the primary power supply module when the current of the intermediate module reaches a preset value; the energy self-recovery module is used for converting the energy self-recovery of a capacitor; the load current opening and shutting module is used for turning on and shutting off the load current; and the voltage polarity recovery module is used for recovering the voltage polarity of the primary power supply module. The inductance type pulse power supply disclosed by the application has the advantages of high energy efficiency, recycling of inductance and load current tail wave energy, tail wave cutting function, few circuit elements, good circuit expandability, strong repeatability and the like.

The application discloses a multifunctional outer cabin plate dot matrix structure for a satellite, which mainly comprises an inner cabin plate, an outer cabin plate, a side wall plate, a pre-embedded heat pipe, a rod-shaped regular tetrahedron dot matrix structure and a shear-resistant bearing rod, and a satellite single machine is installed on the inner cabin plate, and the outer cabin plate is directly contacted with a vacuum environment and serves as a heat dissipation plate. The pre-embedded heat pipes are all installed on the inner surface of the inner cabin plate, the pipe fin surface of the pre-embedded heat pipe is closely attached to the inner surface of the inner cabin plate, the contact area is increased, and the heat exchange efficiency is improved. The rod-shaped regular tetrahedron dot matrix structure has a whole regular tetrahedron appearance, and the unit cell structure is similar to a diamond single crystal. The rod-shaped regular tetrahedron dot matrix structure is adopted, a plurality of stable triangular structures are formed between the internal supporting arms of the dot matrix unit cell, the mechanical strength and the structural rigidity of the cabin plate are greatly enhanced, and the cabin plate load bearing capacity is improved. Meanwhile, the heat pipes can be pre-embedded in the gap of the dot matrix structure, and low-density and high-thermal-conductivity materials are filled in the blank part, so that the mechanical properties are ensured, and the heat conduction efficiency is improved.

This invention relates to a space-distributed concentrated solar energy system, comprising: a power generation unit, an energy storage unit, a power management unit, and a space-based intelligent solar tracking mechanism. The power generation unit converts solar energy into electrical energy using concentrated solar power technology; the energy storage unit stores the electrical energy; the power generation unit is connected to the power management unit; the power management unit is connected to the energy storage unit and manages the electrical energy output by the power generation unit for collaborative storage in the energy storage unit; the power management unit is also connected to the space-based intelligent solar tracking mechanism and an external load system for power supply; wherein, the power generation unit includes: multiple sets of power generation modules and a support structure; the multiple sets of power generation modules are mounted on the support structure and connected to the power management unit. The energy system provided by this invention can improve the photoelectric conversion efficiency of solar cells and reduce equipment costs.

The invention discloses a lift fan matrix vertical lifting high-altitude rocket launching platform and a using method, and belongs to the technical field of aerospace. The launching platform comprises at least two platform modules, electric ducted lift fans, a battery pack, a rocket docking clamping lock, a high-pressure push rod push-away system, a platform module docking locking mechanism and a flight control system, wherein the platform modules can be separated and docked, and the electric ducted lift fans are installed in the platform modules in a distributed mode. During use, the platform vertically ascends and descends to carry the rocket to a preset high altitude, the rocket is safely separated from the platform in a mode of pushing away first and then separating / descending, and then the rocket is ignited and launched. The electric lift fan matrix is adopted for providing power, the rocket launching platform has the advantages of being high in loading capacity, high in platform stability, capable of being repeatedly used, low in launching cost and the like, and the rocket launching efficiency can be effectively improved.