9 results about "Dual-thrust" patented technology

The invention discloses a double-thrust-chamber pump rear-swinging liquid rocket engine, relates to the technical field of rocket engines, and provides a technical scheme capable of effectively making up for the defects of a traditional single-turbine pump and single-thrust-chamber structural layout in a high-thrust or heavy-duty liquid rocket engine. The invention discloses a double-thrust-chamber pump rear swing liquid rocket engine which comprises a rack, two fuel gas swing devices, two thrust chambers and a fuel gas assembly. The fuel gas assembly communicates with the corresponding thrust chambers through the two gas swing devices and is used for providing fuel for the two thrust chambers correspondingly. The first ends of the two fuel gas swing devices are fixedly arranged on the two sides of the rack correspondingly, the second ends of the two fuel gas swing devices are fixedly connected with the corresponding thrust chambers correspondingly, and the two fuel gas swing devices and the two thrust chambers are symmetrically distributed along the center line of the rack. The fuel gas assembly is fixedly arranged in the middle area of the rack.

The application belongs to the technical field of rocket engines, and discloses a double-thrust solid-liquid rocket engine system based on a combustible spoiler and a working method thereof, which comprises the following steps: a front propellant column, a four-fan-hole combustible spoiler and a rear propellant column are sequentially arranged in a combustion chamber shell; a chemical bed component is used for decomposing injected oxidants; the four-fan-hole combustible spoiler, the front propellant column and the rear propellant column are combusted under the action of high-temperature and high-pressure gas generated by the decomposition of the oxidants; the four-fan-hole combustible spoiler arranged in a combustion chamber passage of the engine can change the original flame layer structure, so that the oxidant flow and the pyrolysis products of the fuel can be fully mixed and reacted, and the combustion efficiency of the engine can be improved. The four-fan-hole combustible spoiler simultaneously strengthens the turbulent effect, so that the burning rate of the rear fuel can be several times higher than that of the front fuel. The application has a simple structure and is easy to implement.

This utility model relates to a satellite electric propulsion system, addressing the problems of high cost, poor control precision, and low efficiency and reliability of existing chemical thrusters for micro and nanosatellites, as well as the issues of low efficiency and low reliability of resistance-heated thrusters. The system includes a propellant tank, a power supply and control module, a dual-path pressure regulation module, and a dual-thrust module. The propellant tank uses a full TC4 titanium alloy spherical structure to store liquid ammonia. The power supply and control module controls the high-voltage arc heating of the working fluid via a CAN bus. The dual-path pressure regulation module achieves redundant pressure reduction through solenoid valves, preheaters, and pressure reducing valves. The dual thrusters are arranged vertically at their centers of mass, providing mutual cold backup. This design reduces size and weight through modular integration, improves electrothermal conversion efficiency, orbit climbing mission success rate, and working fluid utilization efficiency, effectively solving the technical bottlenecks of existing propulsion systems and is suitable for high-precision orbit control of micro and nanosatellites.

The invention discloses a thrust measuring device and a thrust measuring method for improving the thrust measuring precision and stability of an engine. The measuring device comprises a fixed frame, a movable frame, a rear spring piece, a front spring piece, a left thrust sensor, a left measuring cable, a right thrust sensor, a right measuring cable, a sensor power supply, a data collector and a computer. The movable frame is hung on the fixed frame through the two front spring pieces and the two rear spring pieces, a left thrust sensor and a right thrust sensor are fixedly installed on the fixed frame, and the left thrust sensor and the right thrust sensor are symmetrically arranged about the axis of the aero-engine to be tested. When the thrust is measured, the measuring method that double thrust sensors are configured with double-channel data acquisition and superposition is adopted, the stability of thrust measurement is improved, the influence of interference is reduced, the fluctuation of the thrust is reduced, and the measuring precision is improved.

The application discloses a dual-thruster rotation attitude control system and method, which comprises a first thruster distributed in a first mounting surface perpendicular to a longitudinal axis of a flight body and a second thruster distributed in a second mounting surface perpendicular to the longitudinal axis of the flight body; the first thruster and the second thruster are distributed in a radial symmetry; the first thruster and the second thruster are used for outputting attitude control force according to a set rotation attitude control strategy, so as to realize complete decoupling of a roll channel; the set rotation attitude control strategy comprises five control positions, i.e., a middle position, a positive pitch control position, a negative pitch control position, a positive yaw control position and a negative yaw control position; the dual-thruster rotation attitude control system provided by the application has the advantages of small number of thrusters, high reliability, simple structure and strong attitude control adaptability.

The invention discloses a device and a method for calibrating an aero-engine thrust measuring rack. The calibration and calibration device comprises a fixed frame, a movable frame, a right thrust measuring sensor, a right adjustable ejector rod, a standard thrust sensor, an adjustable ejector rod, a thrust loading hydraulic cylinder, a hydraulic pipe, a hydraulic station, a left adjustable ejector rod, a left thrust measuring sensor, a spring piece, a standard thrust scale header, a computer, a controller, an engine and a right thrust measuring cable. And a left thrust measuring cable and a rack data acquisition system. According to the calibration and calibration method, a traditional thrust single-sensor measurement method is improved into a measurement mode of double-thrust sensor and double-channel data acquisition and then superposition, the double-thrust measurement sensor is calibrated and calibrated twice, the thrust axis of a movable frame and a fixed frame of a calibration rack can be consistent with the axis of an engine during working, and the calibration precision is improved. And the purpose of accurately measuring the thrust is achieved.

The application discloses a single-chamber double-thrust solid rocket engine charging structure and process, which comprises a grain, a heat insulation layer and a combustion chamber shell. The inner surface of the combustion chamber shell is subjected to sand blasting treatment and uniformly sprayed with adhesive, and then is bonded with the heat insulation layer. The inner surface of the heat insulation layer is roughened and uniformly sprayed with a lining layer. The grain is integrally solidified on the inner surface of the heat insulation layer by adopting a vacuum wall pasting pouring process. The solid rocket engine combustion chamber charging structure of the application adopts a variable cross-section integrated design, improves the complexity of the traditional series combination type single-chamber double-thrust star hole grain forming process, avoids quality problems such as debonding, slag inclusion, air bubbles and cracks of the combination type star hole grain in the forming process of the transition section, guarantees the charging structure integrity and interior ballistic performance, realizes the minimum initial thrust required for the missile to leave the rack and the characteristic requirement of the front position of the missile mass center, effectively solves the problem of missile flight instability, improves the missile flight reliability, and is successfully applied in a type.

The invention belongs to the technical field of rotary kilns, and particularly relates to a high-temperature rotary kiln for calcining gibbsite, which is characterized in that the diameter of a rotary kiln cylinder is 3.8-4.5 m, the length of the kiln is 105-120 m, and the inclination is 35: 1000; the transmission device is a bilaterally symmetrical dual-driving mechanism, each group of driving mechanism is meshed with the gear ring through a gear, and the rotating speed of the rotary kiln barrel is 1.2-1.5 r / min; and the gear and the gear ring form a helical gear pair. The rotary kiln has the beneficial effects that (1) the stability of the rotary kiln barrel in the rotating process can be guaranteed through the double-thrust bearing; the double-catch-wheel device can prevent the barrel from moving downwards along the axis, and the equipment safety is good; the heat transfer efficiency can be improved through the material raising section, meanwhile, the effect of partially crushing the raw materials is achieved, calcination of the follow-up high-temperature section can be improved, the perfect sintering effect is achieved, high-quality high-aluminum clinker is obtained, and the requirement for processing the high-quality gibbsite raw materials is fully met.

The application relates to the technical field of solid engine, in particular to a single-chamber double-thrust solid engine and a nozzle thereof, which comprises a nozzle shell, a tailpipe heat insulation sleeve, an expansion section heat insulation sleeve and a throat liner. The throat liner comprises a front section throat liner, an ablation throat liner, a middle section throat liner and a rear section throat liner. The thermal conductivities of the front section throat liner and the rear section throat liner are lower than that of the middle section throat liner. The application can guarantee the two-stage thrust ratio requirement of the solid engine.