31results about How to "High specific impulse performance" patented technology

The invention discloses an orbit and attitude control integrated space marching system based on an electric pump. The orbit and attitude control integrated space marching system comprises a high-pressure air bottle, an oxidizing agent storage tank, a fuel storage tank, an electric pump system, an oxidizing agent pressure accumulator, a fuel pressure accumulator, an orbit-control engine and attitude-control engines; the high-pressure air bottle is connected with the oxidizing agent storage tank and the fuel storage tank through a pipeline I, the oxidizing agent storage tank and the fuel storagetank are connected with the electric pump system through a pipeline II and a pipeline III correspondingly, the electric pump system is connected with the orbit-control engine and the oxidizing agentpressure accumulator and the fuel pressure accumulator through a pipeline V and a pipeline VI correspondingly, and the oxidizing agent pressure accumulator and the fuel pressure accumulator are connected with the attitude-control engines through a pipeline VII and a pipeline VIII correspondingly. The electric pump system controls the rotating speed and the lift of a motor pump through an actuator,the pressure and flow of an oxidizing agent and fuel supplied for the engines are controlled, the adjustment is convenient, and the supplying requirements for a propellant of large-range thrust adjustment and multi-time starting and stopping of the orbit-control engine can be met.

The invention provides an electric pump pressure type liquid oxygen methane space propelling system. The electric pump pressure type liquid oxygen methane space propelling system comprises a high-pressure gas cylinder, a liquid oxygen storage tank, a methane storage tank, an electric pump system, an oxygen heat exchange injection rod, a methane heat exchange injection rod and a track control engine; the high-pressure gas cylinder is connected with the liquid oxygen storage tank and the methane storage tank through a first pipeline; the liquid oxygen storage tank and the methane storage tank are connected with an electric pump system through a second pipeline and a third pipeline; the oxygen heat exchange injection rod is arranged in the liquid oxygen storage tank; the methane heat exchangeinjection rod is arranged in the methane storage tank; the electric pump system is connected with the oxygen heat exchange injection rod and the track control engine through a fourth pipeline; and the electric pump system is connected with the methane heat exchange injection rod and the track control engine through a fifth pipeline. The electric pump pressure type liquid oxygen methane space propelling system is suitable for a space aircraft propulsion system, and particularly has obvious application advantages on a space propelling system which has large propellant filling amount and whose track control engine needs to be started for a plurality of times..

The invention relates to a sample machine on the basis of the principle of a nitrous oxide mono-component thruster and a use method thereof. The invention utilizes the characteristics of the nitrous oxide gas such as no toxicity, no pollution, high catalytic decomposition temperature and capability of self-maintaining decomposition reaction, and applies the nitrous oxide gas in the propulsion field of aerospace micro-satellites. Compared with the existing commonly-used hydrazine mono-component thruster, the invention can effectively reduce the risk and cost of land disposal with a propellant, and also can realize higher specific impulse performance compared with the cold gas thruster. The invention can conduct principle demonstration verification for mono-component thruster adopting the nitrous oxide as the propellant, and provide technical support for research and development of a non-toxic micro-thruster applied in posture control of the micro-satellite.

The invention discloses a variable structure combustion chamber of a rocket-based-combined-cycle engine. The variable structure combustion chamber is composed of a drive system, a back pressure cavity and a combustion chamber body. The drive system is located at the upper portion of the back pressure cavity. The back pressure cavity is fixed to the portion above variable sections of the combustion chamber body and forms an integral structure with the combustion chamber body. A combustion chamber fixed section is sequentially connected with variable sections of the combustion chamber body. The variable sections are connected through cooperation of ball head hinge convex faces and ball head hinge concave faces. Two sets of hydraulic drive mechanisms drive the first variable section and the third variable section of the combustion chamber body through drive rods and connecting rods. Axial sealing grooves of the variable sections are filled with graphite strips so as to isolate high-temperature fuel gas in the combustion chamber body, and sealing is guaranteed. According to the variable structure combustion chamber, stable transition among different modes is achieved through grade adjustment. In the injection mode and the subsonic combustion mode, the variable sections of the combustion chamber body are adjusted by a hydraulic drive mechanism to generate different divergence angles and throat heights, so that the combustion chamber body is in the optimal work state, and the thrusting force and specific impulse performance of the combustion chamber body are obviously improved.

The invention belongs to the technical field of air turbine rocket combined engines, and particularly relates to a novel high-pressure ratio large-flow turbocharging device for air turbine rockets. When the novel high-pressure ratio large-flow turbocharging device for the air turbine rockets works, the fuel gas generated by a fuel gas generator is discharged through a convergent-divergent type drilling hole inclined nozzle to drive a symmetrical blade type pure impact turbine to rotate, and a working wheel drives a compressor impeller to pressurize the incoming flow air in an axial air inlet;and then the air flows through a diffuser in a compressor shell assembly, and the direction of the air is changed from radial direction to axial direction by a turning section in the outer wall of thediffuser, and finally the air flow is rectified by a flow-guiding cascade in the diffuser. The air flow discharged from the turbine working wheel and the compressor flow-guiding cascade is mixed andcombusted behind the turbo charger. The novel high-pressure ratio large-flow turbocharging device for the air turbine rockets provides a core power device capable of generating large thrust and high specific impulse performance for the air turbine rocket combined engines.

The invention provides a novel pre-cooling air combined engine. The engine comprises an air combustion path (1), a helium cooling path (2), and a hydrogen energy supply path (3); and the ultralow temperature characteristic of the liquid hydrogen fuel is utilized to cool the high-temperature incoming flow air after being stagnated through an air inlet channel, the helium cooling path (2) is introduced to be used as an intermediate circulation to carry out energy transfer and conversion between the hydrogen energy supply path (3) and the air combustion path (1), and thus the fuel required by cooling and the fuel combustion rate are reduced; and meanwhile, the concurrent heating power of a heater is reduced. The engine has the characteristics that the flight envelope is large, the mode conversion is simple and convenient, the air pre-cooling efficiency is high, and the system specific impulse is high; and the engine can be used as one stage of a horizontal take-off and landing repeatabletwo-stage to-orbit vehicle and as a power system of near-air hypersonic velocity launching platform.

The invention relates to a micro-electrolysis water thruster which can utilize water or water solution, belonging to the micro-mechanical technique. It comprises: (1) a micro-electrolysis water thruster formed by three layered plates as a top plate, a bottom plate and spray plate with a one-way valve; (2) and a micro-electrolysis water thruster formed by layering top plate and bottom plate with high depth-width rate electrode or plane electrode without one-way valve. Wherein, the electrolysis reaction chamber is between the top and bottom plates which comprises an electrolysis anode and a electrolysis cathode; the burning chamber has a fire electrode; the hydrophobic film is plated on the inner surfaces of electrolysis reaction chamber and the burning chamber; and the one-way valve is between the burning chamber and the micro-spray pipe. The invention can generate hydrogen-oxygen mixed gas via the electrolyte of water to be fired to generate high-temperature high-pressure steam to be sprayed outside the spray pipe to form pushing power, which can improve the specific thrust property, thermal efficiency, responding speed and control accuracy, while its structure is compatible with MEMS to apply the micromation and integration.

The invention discloses a rocket ramjet wide-speed-range engine based on annular pressurization center body tail confluence. The engine comprises a center body, an air inlet channel, a combustion chamber with a concave cavity and a spray pipe. An embedded rocket in a conventional rocket ramjet engine is arranged on the outer side of the center body, the rocket gas and air are mixed in an annular mixing channel, and the contact area of rocket jet flow and the air is increased, so that the mixing distance of two air flows can be effectively shortened. Meanwhile, annular airflow is converted into central pipe flow through a central body contraction section at the spray pipe, the outlet area of the engine spray pipe can be reduced, and therefore the structural size of the engine is reduced, and the overall thrust performance of the rocket ramjet engine can be effectively improved by adopting the scheme.

The invention relates to the technical field of aerospace, in particular to a control system for a reciprocating type positive displacement pump of a solid-liquid rocket engine. The system comprises astorage mechanism and a liquid pressurizing mechanism; the storage mechanism is used for storing propellant; and the liquid pressurizing mechanism comprises a reciprocating type positive displacementpump, the storage mechanism is communicated with the reciprocating type positive displacement pump, and the reciprocating type positive displacement pump is used for pressurizing propellant and supplying the pressurized propellant to the attitude and orbit control engine mechanism. The liquid pressurizing mechanism is different from a traditional extrusion type or turbine pump pressure type conveying system, a high-pressure gas cylinder and a turbine pump assembly are not used any more, the propellant flowing out of the storage mechanism is pressurized through the light reciprocating type positive displacement pump, and the propellant can have high specific impulse performance.

The invention discloses a liquid-film cooled thrust chamber for a rail posture control engine. A first propellant chamber is arranged on the outer side of the head of the thrust chamber, and a secondpropellant chamber is arranged on the outer side of the middle position of the body of the thrust chamber. A plurality of first spray holes communicating with the first propellant chamber are formed in the shell wall of the head of the thrust chamber. A plurality of combustion jet holes and a plurality of liquid film jet holes are evenly formed in the shell wall of the body of the thrust chamber in the circumferential direction. A second propellant is sprayed to the head and the throat of the thrust chamber through the combustion jet holes and the multiple liquid film jet holes respectively, aliquid film is sprayed towards the throat of the thrust chamber, and a tangential liquid feeding mode is adopted, so that loss of the liquid film can be reduced. The cooling performance of the liquidfilm is given full play to. In addition, the combustion jet holes face the head of the thrust chamber, a tangential liquid feeding mode is adopted, the cooling performance of the liquid film can be achieved, meanwhile, the liquid film can react with the first propellant, therefore the proportion of the liquid film in the propellant can be reduced, and whole-body cooling of the thrust chamber is achieved.

The invention relates to a solid-liquid dual-mode attitude and orbit control power system and a control method thereof in the field of space propulsion systems. The solid-liquid dual-mode attitude and orbit control power system comprises a controller, a fuel gas generator, a gas collection tank, a propellant storage tank, a fuel gas attitude control thruster and a liquid orbit control engine. A solid grain is arranged in the fuel gas generator, the fuel gas generator, the gas collecting tank, the propellant storage tank and the liquid orbit control engine are sequentially communicated through pipelines, the gas collecting tank is communicated with the fuel gas attitude control thruster through a pipeline, and the controller is used for controlling the system. The controller controls the fuel gas generator to ignite the solid grain, the solid grain combusts to generate high-pressure fuel gas, the high-pressure fuel gas is filled into the gas collecting tank, one path of the high-pressure fuel gas in the gas collecting tank enters the fuel gas attitude control thruster, and the other path of the high-pressure fuel gas in the gas collecting tank enters the propellant storage box to extrude a propellant into the liquid orbit control engine. The advantages of a solid gas pressurization system and a liquid propelling system are combined, normal-pressure prepackaging storage of the propelling system is achieved, and the safety of prepackaging storage is improved.