663 results about "Ramjet" patented technology

The present invention generally relates to units of engines and more particularly to units containing a unique combined-cycle (combustion-detonation) “counter-rotation, anti-gyration, gyroscopic,” turbine fan-jet/free-piston engine configuration for induced air supercharging and boosting the performance of novel Ramjet engines or Ramjet engine configurations by improving internal air-stream dynamics. These dynamics are the result of co-operative air stream intermixing through convergent, supercharge-attenuated, inducted, compressed, tuned, pre-heated ambient air. Achieved through the varying of the geometric structural form and the utilization of unique engines and air induction and propulsion conformations, aided with supplemental air, fuel, oxygen and optiomal water and electrolyte charging.

The invention discloses turbine-based combined cycle engine super-combustion chamber. Key parts of the super-combustion chamber comprise a shunt ring, a rear area-variable deflation valve, a flow guide tongue piece, a square lobe mixer, a first area fuel ejector, a center cone, a second area fuel injection ring, an evaporation tube flame stabilizer, a heat-isolating screen, a super-combustion chamber cylinder, an outer extending groove-width-variable radial flame stabilizer and an inner extending radial flame stabilizer. An internal flow field of the super-combustion chamber is reasonably distributed, the phenomenon of backflow from a punching duct to a turbofan duct can not occur, flow separation can not occur in the square lobe mixer, and the evaporation tube flame stabilizer can achieve soft ignition of the super-combustion chamber. The whole super-combustion chamber is high in mixing efficiency, little in cold flow resistance, high in combustion efficiency, reasonable in temperature distribution, and capable of meeting use requirements in a whole flight envelope of the super-combustion chamber in each work state and the requirement for boost of an turbofan ramjet.

A single-stage hypersonic vehicle is comprised of a low-speed and a high-speed propulsion system. The low-speed propulsion system propels the single-stage vehicle to a threshold velocity, after which the high-speed propulsion system then takes over. The low-speed propulsion system includes a combined-cycle engine featuring a swirl generator that is integrated into a turbojet engine to provide a compact turbojet and swirl afterburner-ramjet propulsion system. The high-speed propulsion system includes a hypersonic engine that is operable at the threshold takeover velocity and beyond. In various embodiments, the high-speed propulsion system comprises a scramjet, rocket, or scramjet/rocket engine depending requirements. Benefits of the swirl generator design include its ability to rapidly and efficiently atomize, vaporize, mix and burn the fuel and oxidizer for the low speed propulsion system, significantly reduce length, weight, cooling requirements and complexity for both propulsion systems, while maintaining high propulsion performance and reducing propulsion and launch costs.

The invention discloses an electric conical valve type gas flow regulator for solid rocket ramjets. A drive motor is fixed to a connecting box. One end of a coupling is fixedly connected with a shaft of a motor, and a slider is disposed in a slide hole in the other end of the coupling. The slider is fixedly connected with a valve stem through a tension and pressure sensor. A drive nut is fixedly connected to a solid gas generator. The valve stem passes the drive nut and in threaded fit with the drive nut. A conical valve head is fixedly connected with the valve stem in a threaded manner and extends into a spray pipe of the gas generator. The gas generator is in an inner shell structure which is hollow. The valve stem is fitted in the inner shell, and a bushing is fitted between the valve stem and the gas generator. One end of the valve stem, close to the conical valve head, is provided with a peripheral circular groove, and a moving seal is disposed in the circular groove. A pressure sensor is mounted on the rear head of the gas generator. The electric conical valve type gas flow regulator for solid rocket ramjets is simple in structure, small in size, reliable in operation, low in ablation effect, low in cost, and accurately controllable.

A method for modeling a fuselage of a hypersonic aircraft by promoting outline integration comprises the following eight steps: 1, designing a geometric molding surface of a two-dimensional flow channel of a propulsion system; 2, designing a view-from-above contour curve of the fuselage of the aircraft; 3, designing a side-view contour curve of the fuselage of the aircraft; 4, designing section forms of longitudinal main station positions of the fuselage of the aircraft and variation modes of control parameters; 5, generating a three-dimensional fuselage of the aircraft; 6, generating a three-dimensional supersonic combustion ramjet hood; 7, designing an outline drawing of wing surfaces, and generating three-dimensional wing surfaces; and 8, generating a promotion integrated outline of the three-dimensional fuselage by assembling the fuselage, the ramjet hood and all the wing surfaces. The method has the advantage of simple and rapid modeling process, achieves the effect of promoting integration of the fuselage of the hypersonic aircraft, and is capable of improving paces and efficiencies in designing, developing and manufacturing the hypersonic aircraft.

Propulsion from combustion of solid propellant pellet-projectiles for providing a useful propulsion that has the advantages of both solid and liquid propulsion engines, and also can make use of either solid chemical propellants or fissionable nuclear material as the fuel. Preferred methods and systems can include a storage chamber for storing solid propellant pellets, a feeding system having a pellet feeding channel and a pellet feeding mechanism connected to the storage chamber and to a gun assembly, which is positioned along a longitudinal axis to eject the pellets at a certain velocity. A triggering system positioned between gun assembly and thrust chamber can initiate the propellant pellet-projectile, and a thrust chamber having a combustion chamber for combustion of propellant pellet-projectile with an exhaust nozzle. Additionally, an auxiliary power system can be used to power the components and various electrical and electronic systems that may be present in the invention for controlling the engine components. The gun assembly may include an ejector mechanism for ejecting the propellant pellets through at least one barrel. The triggering system can produce a medium creating an ambience for the initiation of propellant pellet-projectiles. Methods and systems can be used for space and rocket crafts, turbojets and ramjets.

An engine for providing rotary power about an output shaft with a high power-to-weight ratio includes a plurality of flow guiding blades mounted on the inner surface of an annular thruster base. The flow guiding blades cooperate with the peripheral surface of a rotor for forming a plurality of ramjet-like thrusters. The configuration of the flow guiding blades allows for optimization of the number of thrusters. The centrifugal forces generated by the rotating components is compensated by an annular reinforcement wall made with high strength materials allowing for downsizing of the rotor and associated components.

The invention discloses a solid-liquid hybrid rocket ramjet for test, which comprises a combustion chamber of a fuel gas generator, a spray pipe of the fuel gas generator, the head of a secondary combustion chamber, the body of the secondary combustion chamber and a tail spray pipe, wherein the combustion chamber of the fuel gas generator, the spray pipe of the fuel gas generator, the head of thesecondary combustion chamber, the body of the secondary combustion chamber and the tail spray pipe are sequentially connected with one another in series by a plurality of flanges; and the flange at the head of the combustion chamber of the fuel gas generator is fixedly connected with a thrust frame. The ramjet adopts an in-series detachable layout, so that each component and part can be conveniently detached and changed, the adjusting requirement of each parameter of the ramjet during test can be met, and the working condition under various structural parameter combination and various operating modes of the solid-liquid hybrid rocket ramjet can be conveniently simulated. A plurality of groups of important measure points are arranged and preset for the ramjet, so that each main parameter required by the test of the solid-liquid hybrid rocket ramjet can be directly or indirectly obtained.

The invention discloses a center rocket type bimodal ramjet with a circular structure. in the bimodal ramjet, a circular-structured combustion chamber is adopted, a center rocket is installed in a separation section, and the high-temperature fuel-rich gas of the center rocket can be used as an effective ignition source and a flame stabilization source; a small support plate group is used to inject secondary fuel so as to perform secondary burning together with incoming flow air tissue. The injection of the secondary fuel is stepwise carried out along the axial direction of the engine, wherein the support plate directly performs injection in a primary combustion chamber, a concave cavity performs forward injection in a secondary combustion chamber, and the support plate performs injection in the secondary combustion chamber as well after the support plate is installed. The engine provided by the invention has simple structure, is easy to process, safe and reliable to operate, is more suitable for application of the bimodal ramjet with larges size structure in an actual project.

An improved torch igniter for use in devices such as thrust augmenters, gas turbine engines, ramjets, combined-cycle engines and industrial burners. The torch igniter includes a housing with a combustion chamber. Fuel and oxidizer are delivered into the combustion chamber and ignited by an electronic ignition source, such as a plasma jet igniter or a spark igniter, so that an upstream recirculation zone and a downstream recirculation zone are created. The upstream recirculation zone stabilizes and pilots combustion within the combustion chamber, while the downstream recirculation zone augments the combustion event. Byproducts of the combustion event within the torch igniter provide a high mass flux with high thermal energy and strong ignition source radicals that are discharged through a neck portion of the housing and are thereafter employed to initiate a primary combustion event in a primary combustor.

A combined-cycle engine having at least one core engine and at least one ramjet engine. The ramjet utilizes a novel swirl generator for rapidly and efficiently atomizing, vaporizing, as necessary, and mixing a fuel into an oxidant. The swirl generator converts an oxidant flow into a turbulent, three-dimensional flowfield into which the fuel is introduced. The swirl generator effects a toroidal outer recirculation zone and an inner central recirculation zone, both of which are configured in a backward-flowing manner that carries heat and combustion byproducts upstream where they are employed to continuously ignite a combustible fuel/oxidizer mixture in adjacent shear layers, which accelerate flame propagation throughout the core flowfield. The swirl generator provides smooth combustion with no instabilities and minimum total pressure losses and enables significant reductions in the L/D ratio of the combustor. Other benefits include simplicity, reliability, wide flammability limits, high combustion efficiency and high specific thrust performance.

The invention discloses a multi-tube pulse detonation combustion chamber and the detonation method. A detonation chamber inlet opening, a detonation jet flow inlet opening, a detonation jet flow outlet opening and a detonation chamber outlet opening are poditioned in sequence on the detonation chamber, the detonation jet flow outlet opening is connected with the detonation jet flow inlet opening of the adjacent detonation unit, and a plurality of detonation units are connected to form a circular combustion chamber with a plurality of detonation chamber inlet opening and a plurality of detonation chamber outlet opening. The detonation starting device is positioned on the detonation chamber of one of the detonation units. When the detonative gas mixture is filled in the whole detonation chamber 2 through the detonation chamber inlet opening 1 of each detonation unit under the same condition, the detonation starting device 6 of the invention begins to work. The invention realizes the multi-tube pulse detonation combustion chamber with very high working frequency without the dependency of the high energy and high frequency ignition system, and the invention can be used to replace the main combustion chamber or booster combustion chamber of the existing gas-turbine engine, or the combustion chamber of the rotating ramjet engine.

The embodiment of the invention provides a micrometer order solid particle ignition burning test device under a jet stream condition. The device comprises four parts of a gas generator part, an afterburning chamber part, a sampling part and an observing part; the test device can change the ratio of fuel to an oxidant in the gas generator and realize stable controllable adjustment on mass and flow of solid particles, thus realizing the simulation of a working environment in the afterburning chamber of a ramjet of a solid rocket under different working conditions; a test process is monitored in real time by equipment such as a sensor and a high speed camera, a burnt product of the high energy solid particles is sampled in the test process, research on an ignition burning process of the micrometer order high energy solid in the afterburning chamber is realized, the test preparation time is shorter, the expense cost is lower, more observing methods can be used for a burning process in the afterburning chamber, more test data are obtained, the test system is good in reutilization, and engineering reference value of the obtained test data is high.

The invention discloses a pre-cooling air-breathing type variable cycle engine and relates to the reusable pre-cooling air-breathing type variable cycle engine applied to aeronautics and astronautics devices. The engine mainly consists of a pre-cooler, a turbine engine, a super combustion chamber and a super combustion ramjet engine, wherein the pre-cooler and the super combustion chamber form a subsonic combustion ramjet engine; a turbine subsonic combustion ramjet engine is connected with the super combustion ramjet engine in parallel; working states of the turbine subsonic combustion ramjet engine and the super combustion ramjet engine are selected by adjusting a switching adjusting device, so that the working modal of the whole engine is controlled. The pre-cooler can reduce an inlet gas flow temperature of a super combustion chamber, can increase the efficiency of the super combustion chamber, can expand the working range of a subsonic combustion punching modal, and can make up insufficient thrust force when turbine subsonic combustion ramjet engine is transferred to the super consumption ramjet engine, so that the pre-cooling air-breathing type variable cycle engine has the advantage that the transfer class of a combined engine is stably transferred.

Propulsion from combustion of solid propellant pellet-projectiles for providing a useful propulsion that has the advantages of both solid and liquid propulsion engines, and also can make use of either solid chemical propellants or fissionable nuclear material as the fuel. Preferred methods and systems can include a storage chamber for storing solid propellant pellets, a feeding system having a pellet feeding channel and a pellet feeding mechanism connected to the storage chamber and to a gun assembly, which is positioned along a longitudinal axis to eject the pellets at a certain velocity. A triggering system positioned between gun assembly and thrust chamber can initiate the propellant pellet-projectile, and a thrust chamber having a combustion chamber for combustion of propellant pellet-projectile with an exhaust nozzle. Additionally, an auxiliary power system can be used to power the components and various electrical and electronic systems that may be present in the invention for controlling the engine components. The gun assembly may include an ejector mechanism for ejecting the propellant pellets through at least one barrel. The triggering system can produce a medium creating an ambience for the initiation of propellant pellet-projectiles. Methods and systems can be used for space and rocket crafts, turbojets and ramjets.

The invention discloses a method for controlling accuracy of the shape of an ultrahigh-strength steel thin-wall cylinder. The method comprises the following steps of: 1) detecting the appearance of a cylinder; 2) mounting a special shape correction clamp; 3) tempering for stable shape; 4) discharging and cooling; 5) removing the special shape correction clamp and detecting the appearance; and 6) performing repeated shape correction. In the method, after a cylinder shell is quenched, the special shape correction clamp performs shape correction of the cylinder shell, and then the corrected shape is stabilized by tempering. The shape accuracy of the obtained cylinder shell is relatively high, and the requirement on the product design accuracy is completely met, wherein after the shape correction, the linearity error is not greater than 2.5 mm, and difference between large and small diameters is not greater than 1.5 mm. The processing method disclosed by the invention can be applied to the shape accuracy control on the cylinder shell of a secondary combustion chamber of a solid rocket ramjet.

A pilot for a scramjet provides a flame front whose arrival at the wall of the scramjet combustor is delayed thereby reducing combustor heat load. By combining in-stream injection of fuel with an interior pilot and a lean (fuel-poor) outer annulus, the bulk of combustion is confined to the scramjet combustor center. This concept, referred to as “core-burning,” further reduces combustor heat load. One such pilot is for a two dimensional scramjet effective to propel a vehicle. This pilot includes a plurality of spaced apart struts separated by ducts and a strut pilot contained within each strut. A second such pilot is for an axisymmetric scramjet engine has, in sequence and in fluid communication, an air intake, an open bore scramjet isolator and a scramjet combustor. This centerbody pilot pod includes a pilot isolator disposed between the air intake and a pilot diffuser, the pilot diffuser disposed between the pilot isolator and a pilot with the pilot disposed between the pilot diffuser and a pilot combustor. The pilot pod is in axis symmetry around a central axis of the scramjet isolator and supported by a plurality of struts extending from an inner wall of the open bore to an outer surface of the centerbody pilot pod.

The invention relates to a hypersonic velocity air-breathing type ramjet pneumatic thrust integrated analysis method. The method comprises the steps of: S1, acquiring the parameters of a to-be-analyzed hypersonic velocity air-breathing type ramjet; S2, calculating the mass flow rate of incoming flow air and the Mach number of an outlet of an admission passage according to the input parameters; S3, calculating the Mach number of an outlet of a combustor heat release section according to the Mach number of the outlet of the admission passage obtained through calculation in S2; S4, calculating the Mach number of an outlet of an exhauster of gas flow according to the Mach number of the outlet of the heat release section obtained through calculation in S3; S5, calculating ramjet thrust data according to the Mach number of the outlet of the exhauster obtained through calculation in S4. The method can perform hypersonic velocity air-breathing type ramjet pneumatic thrust calculation rapidly, provide a powerful tool for the pneumatic layout design optimization of hypersonic aerocrafts, provide optimal analysis results, and save the development time cost and the money cost for numerical simulation and experiment verification in the development of hypersonic aerocrafts.