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

The invention discloses an electric pump-based rail attitude control integrated space propulsion system, which includes a high-pressure gas cylinder, an oxidant storage tank, a fuel storage tank, an electric pump system, an oxidant pressure accumulator, a fuel pressure accumulator, an orbit control engine and The attitude control engine, the high-pressure gas cylinder is connected to the oxidant storage tank and the fuel storage tank through pipeline 1, and the oxidant storage tank and fuel storage tank are respectively connected to the electric pump system through pipeline 2 and pipeline 3, and the electric pump system is respectively connected through pipeline 5 and pipeline 6 It is connected with the rail control engine, the oxidant pressure accumulator and the fuel pressure accumulator, and the oxidant pressure accumulator and the fuel pressure accumulator are respectively connected with the attitude control engine through pipeline seven and eight. The electric pump system controls the speed and head of the motor pump through the driver, and then controls the pressure and flow of the oxidizer and fuel supplied to the engine. It is easy to adjust and can meet the propellant supply requirements of a wide range of thrust adjustment and multiple starts and stops of the rail control engine.

The invention relates to a single component micro-propulsion module device based on additive manufacturing technology. The device is characterized by comprising a micro-thruster assembly, a gas addingand discharging valve, a liquid adding and discharging valve, a solid nitrogen generator, a temperature sensor, a pressure sensor, a bursting disc safety device, a tank assembly and a control drive circuit; the gas adding and discharging valve, the solid nitrogen generator, the temperature sensor, the pressure sensor, the tank assembly, the liquid adding and discharging valve, the bursting disc safety device and the micro-thruster assembly are installed in series in a module gas room. The invention provides a novel device and method for safety isolation of a propulsion system, the high integration is facilitated, the system complexity is reduced, the reliable operation of an entire system is guaranteed, and the work performance of the entire system is improved.

An expansion cycle liquid oxygen methane upper stage engine system including a propellant supply system, an ignition system and a thrust chamber. The propellant supply system includes a methane supply system and a liquid oxygen supply system. The methane supply system pressurizes the cryogenic liquid methane from the external storage tank and supplies it to the thrust chamber; the liquid oxygen supply system pressurizes the cryogenic liquid oxygen from the external storage tank. After pressurization, it is provided to the thrust chamber; the ignition system is located at the head of the thrust chamber, and is ignited under the control of the control system to ignite the liquid oxygen and methane entering the thrust chamber, and the high-temperature gas generated is ejected from the nozzle of the thrust chamber to generate thrust. The present invention uses liquid oxygen and liquid methane as a propellant combination, and adopts a closed expansion cycle system scheme. The system is simple, has high inherent reliability, high specific impulse performance, and is easy to realize multiple starts.

The invention discloses a thrust chamber of a liquid film cooling rail attitude control engine. A first propellant chamber is arranged on the outside of the head of the thrust chamber, and a second propellant chamber is arranged on the outside of the body of the thrust chamber and is located outside the middle position; the thrust chamber The shell wall of the head is provided with a plurality of first injection holes communicating with the first propellant chamber; the shell wall of the thrust chamber body is evenly provided with a plurality of combustion injection holes and a plurality of liquid film injection holes along the circumferential direction ; The second propellant is sprayed to the head and throat of the thrust chamber through the combustion injection hole and a plurality of liquid film injection holes, because the liquid film injection is directed towards the throat of the thrust chamber, and the tangential liquid inlet method is adopted , so it can reduce the loss of the liquid film and give full play to the cooling performance of the liquid film; at the same time, the combustion injection hole faces the head of the thrust chamber, and adopts a tangential liquid inlet method, which can exert the cooling performance of the liquid film. The reaction of the first propellant not only reduces the proportion of liquid film to propellant, but also realizes the whole body cooling of the thrust chamber.

The present invention proposes a composite mechanism full-flow cycle supersonic propulsion system and its working method, including an air inlet, a compressor, a diffuser duct, a shaft, an open pre-combustion driver, a closed pre-combustion driver, a turbine, a post-combustion Thruster, tail nozzle; the inlet, compressor, shaft, turbine, afterburning thruster, and tail nozzle are connected in sequence; the compressor and turbine are connected through a shaft; the turbine is located on the side of the inlet of the afterburning thruster, The tail nozzle communicates with the outlet of the afterburner thruster; the open pre-combustion driver and the closed pre-combustion driver are located behind the compressor, and the outlet is facing the turbine; the open pre-combustion driver and the closed pre-combustion driver are alternately distributed on the outer ring of the shaft Into the expansion duct of the space. The system maximizes the propulsion advantages of each subsystem in terms of specific impulse and thrust-to-weight ratio, so as to realize the rapid self-starting of the propulsion system, multi-degree-of-freedom, compound working mechanism, efficient and economical supersonic combined propulsion system.

The invention discloses a nozzleless solid rocket motor, comprising: a shell, which is in the shape of a column, propellant grains filled in the inner cavity of the shell, a gas channel is opened along the central line, a support tube, which is in the shape of a tube, and is sleeved on the In the gas channel and close to the outlet of the gas flow, a plurality of inner cavity electromagnetic field mechanisms and a plurality of inner cavity high-voltage thermoelectric mechanisms are sequentially arranged on the inner wall, and the inner cavity electromagnetic field mechanisms and the inner cavity high-voltage thermoelectric mechanisms cooperate with each other for the gas to pass through. When supporting the gas channel in the inner cavity of the tube, a millivolt-level thermoelectric potential is generated due to the thermoelectric effect, and then a high-voltage thermoelectric potential is formed. The thermoelectric potential greatly accelerates the gas containing a large number of ionized products, and is discharged through the inner cavity of the supporting tube to propel the rocket. Operation; the invention cancels the nozzle structure of the conventional solid rocket motor, greatly reduces the structural quality of the motor, and reduces the probability of failure.

The invention provides a stepped combustion chamber and an inner wall of a nozzleless engine. The inner wall of the combustion chamber includes a radial wall surface and an axial wall surface of the head; , the initial step wall and the pressure-stabilizing step wall are the same plane and perpendicular to the engine axis, the transition wall has a certain slope, and the initial step wall and the stabilizing step wall are connected as a whole; the axial wall is connected with the outer edge of the initial step wall and along the The combustion chamber extends axially; the axial wall and the radial wall of the head jointly form the combustion chamber.

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.

A new type of pre-cooled air combined engine, including air combustion path (1), helium cooling path (2), hydrogen energy supply path (3), by using the ultra-low temperature characteristics of liquid hydrogen fuel to cool the stagnant air through the intake port The final high-temperature incoming air is introduced into the helium cooling circuit (2) as an intermediate cycle to carry out energy transfer and conversion between the hydrogen energy supply circuit (3) and the air combustion circuit (1), reducing the fuel and cooling requirements Fuel combustion rate, while reducing the heating power of the heater, has the characteristics of large flight envelope, simple mode conversion, high air pre-cooling efficiency and high system specific impulse, and can be used as a repeatable two-stage orbiting aircraft for horizontal take-off and landing The power system of the first stage and adjacent air hypersonic launch platform.

The present application relates to the field of aerospace technology, in particular to a control system for a reciprocating displacement pump of a solid-liquid rocket engine. The system includes a storage mechanism and a liquid pressurization mechanism; the storage mechanism is used to store the propellant; the liquid pressurization mechanism includes a reciprocating positive displacement pump, the storage mechanism communicates with the reciprocating positive displacement pump, and the reciprocating positive displacement pump is used to pressurize the propellant, And it is used to supply the pressurized propellant to the attitude and orbit control engine mechanism. The liquid pressurization mechanism is different from the traditional extrusion or turbo pump delivery system. This application no longer uses high-pressure gas cylinders and turbo pump components, and uses a lightweight reciprocating positive displacement pump to boost the propellant flowing out of the storage mechanism. The pressure can make the propellant have higher specific impulse performance.

The invention discloses a rocket-based combined cycle ejector mode performance test engine. A main rocket combustor is fixed in the front of a secondary combustor; intake ducts are mounted above and below the main rocket combustor respectively; intake ends of the intake ducts are fixed in the rear of the main rocket combustor through support lugs; the rear ends of the intake ducts are connected with the secondary combustor through an intake duct flange and a second combustor flange; the secondary combustor is in a divergent straight structure; the main rocket combustor is fitly connected with the secondary combustor; a flue of a nozzle divergent segment gradually changes from round to square. The main rocket combustor is charged through the combination of seven-pointed star shape and inner holes by close-to-wall pouring; after ignition, the flow of gas in a main rocket is kept basically unchanged; variation pattern of the rocket-based combined cycle ejector mode performance along with flight Mach number and height is verified; the basis is provided for passage design. The rocket-based combined cycle ejector mode performance test engine is simple in structure and convenient to maintain, and the influence of gas composition upon engine performance and trajectory climb can be extensively researched.