150results about How to "Improve specific impulse" patented technology

The invention discloses an ejector-assistant turbine-based combined cycle engine, and relates to a reusable combined cycle engine applied to aircrafts and spacecrafts. The engine mainly consists of a turbine engine, rocket ejectors and a dual-mode scramjet engine, wherein the turbine engine and the scramjet engine work in parallel, and the rocket ejectors and the scramjet engine are in integrally designed. When the engine is converted from a turbine engine mode to a scramjet engine mode, ejection fuel-gas streams are ejected into the scramjet engine from support plates or side walls of the scramjet engine; inlet air of the scramjet engine is mixed with the ejection fuel-gas streams, and afterburning is performed. The fuel gas ejected by the rocket ejectors can overcome the defect that thrust is not enough when the turbine engine mode is converted into the scramjet engine mode, so that the combined engine disclosed by the invention can has the advantage of smooth transition during mode conversion.

The present invention discloses a turbo-rocket combined ramjet engine, which comprises an air turbine ramjet and a rocket based combined cycle engine that share an air intake passage and an exhaust nozzle. An air intake passage guiding device is disposed on a duct of the air intake passage. The air intake passage guiding device is used for turning to seal an inlet of the rocket based combined cycle engine at low-speed start of the air turbine ramjet to achieve internal communication of the air intake passage and the air turbine ramjet, and further used for turning to seal an inlet of the air turbine ramjet when a speed of the air turbine ramjet is greater than or equal to Mach, thereby achieving internal communication of the air intake passage and the rocket based combined cycle engine to start the ramjet. A solution is provided for the problem that two existing power systems cannot independently and completely meet the power requirements of modern warfare on weapon systems in such aspects as ultrahigh acoustic velocity, superaltitude, high maneuverability and high airspace.

The invention discloses a Hall thruster adopting a lightweight integrated anode. The Hall thruster comprises a bottom plate, wherein an inner magnetic core wound with an inner magnetic coil is configured at the centre of the bottom plate; a plurality of outer magnetic cores wound with outer magnetic coils are configured at the periphery of the bottom plate; the lightweight integrated anode is arranged between the inner magnetic core and the outer magnetic cores; a protective cover is arranged outside the anode; and the protective cover and the anode are insulated and supported by ceramic. An inner ceramic ring and an outer ceramic ring are arranged at the outlet of the lightweight integrated anode, and the inner ceramic ring and the outer ceramic ring are compacted by the outer edge of the inner magnetic core and a front panel separately; and all the parts are installed on the bottom plate, and an insulation sleeve sleeves the gas pipe of the lightweight integrated anode. According to the Hall thruster disclosed by the invention, an integrated anode structure is adopted, and the flowing and distribution characteristics of a gas are improved, so that the thruster is better in ionization effect, better in performance, lower in thermal load, higher in the capacity of resisting to severe mechanical environments, more compact in structure, smaller in size, and lighter in weight.

The invention discloses sulfur-free ring stabilizing agent propellant powder which consists of the following components in parts by weight: 70 parts of an oxidant, 34 parts of a burnable agent, 6 parts of an adhesive, 2 parts of a stabilizing agent, 2 parts of an anti-aging agent, 1 part of a catalyst and 2 parts of an antismoke agent which are uniformly mixed according to a general method, wherein the oxidant is a mixture consisting of potassium perchlorate, amine perchlorate and barium nitrate; the burnable agent is a mixture consisting of potassium hydrogen terephthalate, potassium acid phthalate, nitrocellulose, purified terephthalic acid and hemp stem carbon powder; the stabilizing agent adopts calcium stearate; the anti-aging agent adopts calcium phthalate; the catalyst adopts copper oxide; the antismoke agent adopts aluminum hydroxide; the adhesive is selected from one or a mixture of multiple of phenolic resin, strong adhesive powder, glutinous rice flour and polyvinyl alcohol according to any ratio. The sulfur-free ring stabilizing agent propellant powder is high in power capability and excellent in safety performance indexes and has excellent storage stability, and the environment-friendly performance is greatly improved.

The invention discloses a pinhole array miniature electrostatic electric thruster. The pinhole array miniature electrostatic electric thruster comprises a neutralizer and a thruster body which are mutually connected, wherein the thruster body sequentially comprises a grid component, a pinhole array component and a propellant storage chamber which are cooperatively connected from top to bottom; the pinhole array component is formed by conveying plates and a pinhole array plate; the conveying plates are arranged on the lower surface of the pinhole array plate; pinhole arrays are uniformly distributed in the pinhole array plate; the grid component consists of an extraction grid, an acceleration grid and an insulation bracket; the extraction grid, the acceleration grid and the pinhole array plate are sequentially connected through the insulation bracket; and coaxial pinholes corresponding to the pinhole arrays are formed in the extraction grid and the acceleration grid; and heaters and the propellant are all located in the propellant storage chamber. The pinhole array miniature electrostatic electric thruster has the advantages that the structure is simple; the size is small; the specific impulse can reach 700-2000 s; the propellant storage function and the conveying function are integrated; the propellant is stored in a pressure-free mode; no movable components exist; and the like. The pinhole array miniature electrostatic electric thruster can be packaged in a chip mode, and can be applied to microsatellites as a standard thrusting module.

The invention discloses a bi-order grid spiral wave ion propulsion device which comprises a discharge chamber, a spiral wave excitation antenna, a metal sleeve, an electromagnetic coil and a bi-order grid system. The spiral wave excitation antenna is connected with a radio-frequency power source and is arranged outside the discharge chamber in a sleeving mode, the spiral wave excitation antenna and the discharge chamber are both arranged inside the metal sleeve, and the metal sleeve is sleeved with the electromagnetic coil. The bi-order grid system is arranged at the tail end of the discharge chamber. The bi-order grid system comprises a screen grid, an extraction grid, an accelerating grid and a decelerating grid which are sequentially arranged away from the discharge chamber from near to far, and the screen grid, the extraction grid, the accelerating grid and the decelerating grid are all provided with grid holes. According to the bi-order grid spiral wave ion propulsion device, the mode that a spiral wave plasma source and the bi-order grid system are combined is adopted, the propulsion device can work under high power, and thus the reliable performance of high specific impulse and high propulsion capability can be achieved.

The invention discloses a solid-liquid scramjet engine, and belongs to the technical field of engines. The engine comprises an air inlet channel, a combustion chamber, a tail spray pipe, a fuel gas generator, an oil tank system, a control system and the like. The oil tank system comprises a high-pressure gas cylinder, a pressure reducing valve, a safety valve, fuel oil, an oil filter, a leather bag and the like, and the control system provides fuel oil flow meeting requirements for the scramjet engine according to set requirements. The fuel gas generator is mainly composed of a heat insulation combustion chamber, a wrapped grain, a long tail pipe spray pipe component and an igniter. The fuel gas generator is fixedly arranged in the oil tank system, and a long tail pipe of the fuel gas generator penetrates through the head portion of the combustion chamber and stretches into the combustion chamber. The solid-liquid scramjet engine can keep the advantages that the specific impulse of the liquid fuel scramjet engine is high, and the solid fuel scramjet engine is not prone to being shut down under a large attack angle, and meanwhile the defects that the liquid scramjet engine is prone to being shut down under the large attack angle, the thrust of the solid fuel scramjet engine is hard to adjust, and the specific impulse is low are overcome.

The invention provides an anode and cold air thruster combined air supply cusped magnetic field plasma thruster and belongs to the field of the space technology and low-temperature plasma. The plasmathruster is provided for solving the problem that an existing cusped magnetic field thruster can hardly generate the thrust less than 100 micronewtons. The plasma thruster comprises a discharge channel, a magnetic conducting ring, a first permanent magnet, an anode, a magnetic conducting ring heat conducting plate, a second permanent magnet and a shell. The anode is arranged at the bottom of the discharge channel. The two permanent magnets are sequentially arranged between the outer sidewall of the discharge channel and the shell from the bottom of a discharge outlet. The magnetic conducting ring and the magnetic conducting ring heat conducting plate are arranged between the two permanent magnets. A gap is reserved between the two permanent magnets and the outer sidewall of the discharge channel. The plasma thruster further comprises a cold air thruster. The anode is of a hollow structure. The cold air thruster is arranged in an inner cavity of the anode. The cold air thruster can operate independently and is used for generating the micronewton-scale thrust. The discharge channel is a variable cross-section channel. The radium of the channel gradually decreases from the discharge outlet to the bottom. The longitudinal section of the sidewall of the discharge channel is a right trapezoid.

The invention discloses an axial injection end combustion solid-liquid rocket engine structure which comprises an engine head cover, a solid propellant shell, a solid propellant heat insulating layer, end combustion solid propellants, a spray pipe shell and a spray pipe heat insulating layer. The engine head cover is mounted at the front end of the solid propellant shell, the spray pipe shell is mounted at the rear end of the solid propellant shell, the engine head cover is provided with an air inlet channel, the end combustion solid propellants are coaxially arranged in the solid propellant shell, a constant volume cavity is formed between the front end faces of the end combustion solid propellants and the rear end face of the engine head cover, each end combustion solid propellant is provided with a plurality of axial injection holes with small diameters, the flow rate of oxidizing agents in the injection holes is high, and combustion reaction only occurs at the end faces of the end combustion solid propellants; the joint between the solid propellant shell and the engine head cover and the joint between the solid propellant shell and the spray pipe shell are respectively provided with a sealing structure, so that air impermeability of an engine is guaranteed; the heat insulating layers are arranged in the solid propellant shell and the spray pipe shell, so that thermal protection performance of the engine during operation is guaranteed. Thus, performance of the engine is improved, meanwhile the filling fraction of the engine is improved, the slenderness ratio of the engine is reduced, and the application rage of engine is widened.

The invention belongs to a space electric propulsion device, and particularly discloses a bipolar solid ablation type plasma accelerator. A positive pole of a direct-current power supply of the accelerator is connected with one end of each of a primary high-voltage control switch and a secondary high-voltage control switch; the other end of the primary high-voltage control switch is connected with one end of each of a primary discharge capacitor and a primary electrode; the other end of the primary discharge capacitor is connected with each of a negative pole of the direct-current power supply and a cathode of a primary silicon controlled rectifier; an anode of the primary silicon controlled rectifier is connected with the other end of the primary electrode; the other end of the secondary high-voltage control switch is connected with one end of each of a secondary discharge capacitor and a secondary electrode; the other end of the secondary discharge capacitor is connected with each of the negative pole of the direct-current power supply and a cathode of the secondary silicon controlled rectifier; and an anode of the secondary silicon controlled rectifier is connected with the other end of the secondary electrode. According to the device, the problem of low efficiency of the solid ablation type plasma accelerator can be effectively solved.

The invention discloses an ion and Hall hybrid type electric thruster. By means of the ion and Hall hybrid type electric thruster, an ion thruster and a Hall thruster can be effectively combined, the integral design is formed, and the thruster can fully inherit the advantages of the ion thruster and the advantages of the Hall thruster. The ion and Hall hybrid type electric thruster comprises an inner ion ring, an outer Hall ring, a cathode and a magnetic shielding structure, wherein a direct current ion thruster structure is adopted in the inner ion ring; a steady state plasma Hall thruster structure is adopted in the outer Hall ring; the inner ion ring and the outer Hall ring are concentrically arranged, and the outer Hall ring is embedded in the outer side of the inner ion ring; the annular magnetic shielding structure made of permanent magnet materials is arranged between the inner ion ring and the outer Hall ring. The cathode is of a hollow cathode structure and is located on the outer side of the outer Hall ring.

The invention relates to a preparation method for a rhenium-iridium combustion chamber. The preparation method comprises the following steps: preparing a rhenium combustion chamber by adopting a powder metallurgy near-net shape forming technology; preparing iridium coatings at inner and outer surfaces of the rhenium combustion chamber by adopting an electric-arc deposition technology; preparing a platinum-rhodium ring by adopting a spinning technology and assembling the platinum-rhodium ring on the inner wall surface of the combustion chamber; and preparing a metal oxide high-radiation coating on the outer surface of the combustion chamber by adopting a plasma spraying technology. According to the preparation method, by virtue of an optimized design of process methods and process conditions at different stages of a preparation process, the prepared rhenium-iridium combustion chamber is applied to an engine, so that a use and working environment of the engine can be remarkably increased, specific impulse can be improved for the engine, and technical support is provided for prolonging flight life of an aircraft, increasing effective load of the aircraft or increasing range of the aircraft.

The invention provides a magnetic plasma thrustor supported through lasers. The magnetic plasma comprises a thrustor power source for supplying power to all electric equipment in the thrustor; satellite platform electric energy is converted through the thrustor power source to be then supplied to the thrustor; the magnetic plasma thrustor further comprises a working medium supply device, a negative pole, a positive pole and a laser generation mechanism, a discharge channel is formed between the negative pole and the positive pole, the working medium feeding device used for fixing and conveying a working medium is arranged outside the discharge channel, and the working medium is a solid working medium and extends into the discharge channel; and the laser generation mechanism emits laser beams into the discharge channel to conduct ablation on the working medium, the laser beams conduct ablation on the working medium continuously to generate partial ionized gas and particulate matter, and the partial ionized gas and the particulate matter enter the discharge channel. The magnetic plasma thrustor solves the problems that an air supply system is complex, dynamic response is slow, negative pole ablation is serious, and performance reduction is caused by large-current vibration, and the magnetic plasma thrustor which is supported through the lasers and is high in dynamic response, high in working medium using rate and high in energy conversion efficiency is provided for a spacecraft.

The invention provides an engine combined of pulse detonation ejection and supersonic combustion ramjet, and the engine is mainly composed of the components of a pulse detonation burner, an ejector and the like. The ejector comprises an ejector inlet valve, a mixing isolation segment, a secondary burning chamber and ejector spray pipes. The engine combined of the pulse detonation ejection and thesupersonic combustion ramjet comprises two working modes of a pulse detonation ejection mode and a supersonic combustion ramjet mode. When the flight mach (Ma) number is equal to 0-5, the engine worksin the pulse detonation ejection mode, under the ejecting action of a combustion ejection flow jetted from the detonation burner, more air is ejected into an air inlet passage, the total air pressurerises, the combustion ejection flow is mixed with the ejected air, fuel is jetted to the mixed air flow, secondary detonation is performed, and at this time, the thrust of the engine is mainly provided by the pulse detonation burner and a secondary pulse detonation system; and when the Ma is larger that 5, the engine is switched from the pulse detonation ejection mode to the supersonic combustionramjet mode, and an aircraft continues to accelerate under the push of the supersonic combustion ramjet engine.

The invention provides a combined cycle engine. The combined cycle engine comprises a supersonic air inlet mechanism, a continuous rotation detonation engine body and a turbojet engine body, wherein the supersonic air inlet mechanism is provided with a first gas inlet which communicates with the outside and is opened in the axial direction, and the continuous rotation detonation engine body and the turbojet engine body are fixedly connected to the lower portion of the supersonic air inlet mechanism. The continuous rotation detonation engine body and the turbojet engine body both communicate with the supersonic air inlet mechanism in a controlled mode in order to work in the mode that external gas is introduced into the continuous rotation detonation engine body and the turbojet engine body through the first gas inlet in the controlled mode. According to the combined cycle engine, mode transition can be completed within an operating envelope range, the stable thrust is provided for an aircraft, and then the aircraft hypersonically flies at 5.0 Mach or above in the atmosphere and can land and take off horizontally; in addition, the specific impulse of the engine is improved on the basis of adopting the continuous rotation detonation technique, the combustion heat efficiency is improved, and the economy and the reliability of the engine are improved.

The invention provides gun, artillery and rocket jet engines, steam turbines, internal combustion engines and air vehicles with an additional magnetic device. A magnet is included, translational motion, through Lorentz force, rotation and vibration of fluid molecules near the magnet in the direction of being perpendicular to the micro-flow direction of fluid are strongly inhibited, and pressure and temperature of the fluid molecules in the direction of being perpendicular to the micro-flow direction of the fluid are significantly lowered, so that the speed that the fluid molecules diffuse in the needed micro-flow direction is significantly increased. Viscous friction, turbulent shear stress, friction force and heat conduction in a boundary layer and shock waves, resistance generated by the shock waves and negative influences of heating on relative movement between air flow and an object are eliminated fundamentally in the microcosmic level or at least weakened greatly. By equipping the magnet on various gun, artillery and rocket engines, jet engines, turbines, gasoline engines, diesel engines and the like, radial pressure and temperature must be significantly lowered, the initial speed of pills is significantly increased, the power and the speed of the rock engines and the jet engines are increased, the energy utilization rate of the gas turbines, the turbines and the internal combustion engines is increased, and various resistances borne by the air vehicles and heating of the air vehicles are significantly reduced.

The invention discloses an array type capacity coupling radio frequency plasma micro thruster. The thruster comprises an air source, valves, antennas, thrust chambers and a power supply. The air source is directly connected with the thrust chambers, the valves can control thrust chamber switches respectively, and the antennas wind around the exteriors of the thrust chambers and are connected withthe power supply. The space size and overall quality of a pocket rocket are extremely small, the impulse pulse is low, thrust from micro-newton to millinewton can be achieved, and the average power dissipation is low; meanwhile, precise vector control can be achieved, the thrust is suitable for equipping microsatellites to be used for high precision posture control and position keeping relative totracks.

The invention relates to a reflecting type laser-electromagnetic field coupling thrustor which comprises a laser system, an electromagnetic accelerating electrode assembly and a static electric fieldassembly. The electromagnetic accelerating electrode assembly comprises an electromagnetic accelerating cathode, an electromagnetic accelerating anode and an electromagnetic accelerating electrode power source; the static electric field assembly comprises a static electric field accelerating electrode, a working medium and a static electric field accelerating power source, the static electric field accelerating power source provides power for the working medium and the static electric field accelerating electrode, and a potential difference is formed between the working medium and the static electric field accelerating electrode to form a static electric field; laser plasmas generated by ablation of the working medium by laser beams enter the static electric field, when the laser plasmas accelerated by the static electric field move to the position between the electromagnetic accelerating cathode and the electromagnetic accelerating anode, discharge arcs can be induced to be generated,the laser plasmas are heated by the discharge arcs and are further ionized, the laser plasmas are sprayed out in an accelerated manner under the joint effect of lorentz force and aerodynamic force, and thrust is accordingly generated. The reflecting type laser-electromagnetic field coupling thrustor has the beneficial effects that the thrust is adjustable, the thrust efficiency is high, and the thrust density is large.

The invention discloses a solid rocket powder supersonic combustion ramjet engine. The solid rocket powder supersonic combustion ramjet engine comprises a gas inlet path, a fuel gas generator, a supersonic-velocity combustion chamber and a tail spray tube, and further comprises a powder supply system, wherein a powder spray-out hole which communicates with the powder supply system is formed in theinner wall of the gas inlet path; powder fuels output by the powder supply system are pushed to the gas inlet path through a powder spray-out hole; supersonic-velocity inflow air in the gas inlet path rolls and absorbs powder fuels, so that mixing of the fuel powder and the supersonic-velocity inflow air is realized, and the mixed fuel powder and supersonic-velocity inflow air enters the supersonic-velocity combustion chamber; and high-temperature high-pressure gas, generated in the supersonic-velocity combustion chamber, of the powder fuels and combustion-rich fuel gas does work by tail spray tube expanding to generate thrust force. The solid rocket powder supersonic combustion ramjet engine can realize active regulation on parameters such as supply amount and supply speed of the powderfuels through the powder supply device, has the advantage of adjustable thrust force, and enables an aircraft to have better maneuverability and a wider airspace range.

The invention discloses a micropulse plasma thruster which is common anode type; cathode plates are symmetrically arranged at the periphery of an anode of the micropulse plasma thruster, a propellant is arranged between the anode and the cathode plates, and discharge chambers are formed in a space among the anode, the cathode plates and the propellant; anti-creeping ceramic insulating plates are arranged at the sides of the propellant; a cathode of a capacitor is connected with the cathode plates; an anode of the capacitor is connected with the anode; and spark plugs are arranged on the cathode plates. The invention provides a concept of a parallel-plate common-anode type micropulse plasma thruster based on an existing relatively-mature pulse plasma thruster technology for feeding the propellant on the tails of the parallel plates, and the meaning of the concept is that singe cathode plate is changed into four same cathode plates which are symmetrically arranged at the periphery of an anode plate with a hollow structure on the basis of the existing mature pulse plasma ion thruster design, so that four same discharge chamber structural designs are formed, impulse bit is improved while high specific impulse is kept, and therefore, the purpose of improving total impulse is realized.

The invention discloses a fullerene ethylenediamine nitrate as well as a preparation method and an application thereof. Firstly, fullerene C60 and ethylenediamine are taken as raw materials, fullerene ethylenediamine is prepared by the aid of a nucleophilic addition principle, and the fullerene ethylenediamine reacts with nitric acid so that the fullerene ethylenediamine nitrate is obtained. The fullerene ethylenediamine nitrate has the advantages of oxidant ammonium nitrate (AN), can avoid disadvantages of hygroscopicity, and crystal transformation of AN, and can be used as an oxidant to be applied in a novel solid rocket propellant or explosive.

Te invention discloses piperazine cation-base ionic liquid containing ammonia and boron groups and a preparation method thereof. The preparation method comprises the steps of introducing the ammonia and boron groups into piperazine cation-base ionic liquid, and generating replacement reaction by virtue of a piperazine onium-salt borane complex and sodium salt or silver salt of anions, so as to obtain the piperazine cation-base ionic liquid containing the ammonia and boron groups. The product has very high combustion gas discharging amount and has potential application values in the field of natural liquid propellants; and the preparation method is simple, mild and easy to implement.