38 results about "Scramjet" patented technology

A new dual-mode ramjet combustor used for operation over a wide flight Mach number range is described. Subsonic combustion mode is usable to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle throat. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated.

The Hypersonic Orbital Fighter and the methods of flying it are disclosed. The Hypersonic Orbital Fighter is characterized by a winged, modified lifting body shape with the forward ventral section comprising an X-30 type compressive air inlet, throttleable supersonic combustion ramjet array under the center of gravity and half-cone exhaust outlet at the back, fore mounted canards, cockpit / payload area top front and mid section and one or more tailplanes at the top back. Flight profile comprising runway takeoff, high subsonic climb up to 65,000 feet, alternatively full power climb from zero to 65,000 feet, with progressive acceleration from there up to Mach 24 @ required orbital altitude. Able to intercept missiles from early through post-boost phase from chosen orbit or suborbital arc, to intercept aircraft and to interdict surface targets. Winged dynamic soaring capability in reentry followed by powered runway landing.

The invention belongs to the technical field of sub-combustion ramjet engines, and discloses a combustion chamber of sub-combustion ramjet engines, which include sequentially arranged and sequentially connected: an isolation section, a sudden expansion section, a concave cavity, an equal straight section and an exhaust nozzle; the isolation The distance between the respective bottom plates of the section, the sudden expansion section, and the concave cavity increases sequentially from the top; the support plate structure is provided with a support plate cavity, and the support plate structure is provided with a fuel inflow port, a high-temperature bleed air port and a A plurality of injection holes communicated with the cavity of the support plate. The invention adopts the combination of the sudden expansion section, the concave cavity and the support plate structure, so that the mixing effect of the fuel and the incoming air is good; the support plate structure is arranged obliquely, and the fuel mixture sprayed by the support plate structure has a wide distribution range in the flow channel; The fuel and the high-temperature bleed air flow are pre-contacted in the support plate cavity, so that the fuel can gain energy, which is convenient and fast to participate in combustion, and the ignition speed is fast; the concave cavity can form a continuous and stable flame, which is helpful for spreading and stabilizing the flame.

The invention relates to the technical field of ramjets, in particular to a device for testing fuel combustion performance of a ramjet. The device for testing the fuel combustion performance of the ramjet comprises a combustion system, a pyrolysis system and a flame data acquisition system, wherein the combustion system is used for atomizing liquid fuel, mixing the liquid fuel with gaseous fuel and igniting the mixed fuel; the pyrolysis system is communicated with the combustion system and used for providing controllable liquid fuel and gaseous fuel; the flame data acquisition system is connected with the combustion system and used for detecting the combustion performance of mixed combustion of the gaseous fuel and the liquid fuel. The device for testing the fuel combustion performance of the ramjet is mainly used for detecting the combustion performance of endothermic hydrocarbon fuel, can simulate the working state of a scramjet of a power device of a hypersonic vehicle, detects and reasonably observes the fuel combustion performance of the ramjet, can detect the combustion performance of both pyrolysis gas and the liquid fuel, and has the advantages that the device is concise and convenient to operate, the safety is high, a flame is convenient to observe and temperature measurement point positions are more flexible.

The invention provides a scramjet combustor supporting plate with right-angled triangular prism grooves. Compared with a traditional trailing edge downstream spray supporting plate, the scramjet combustor supporting plate is provided with the right-angled triangular prism groove structures, fuel injectors move forwards to the top faces of the right-angled triangular prism grooves from the trailing edge, and the direction of side edges of the right-angled triangular prism grooves and the fuel injection direction are parallel to the leading edge back airflow direction of the supporting plate. The supporting plate has the advantage that downstream spray has few losses compared with transverse spray; by arranging the fuel injectors on the front portion, the problem that downstream spray of fuel is short in atomization, evaporation and mixing time is solved; by fully utilizing streamwise vortexes generated at the right-angled junctions of the triangular prism grooves and the supporting plate, the fuel atomization performance and the fuel and oxidizing agent mixing performance are improved at a low flow loss cost; by fully utilizing the spanwise vortexes at the trailing edge of the supporting plate are fully utilized, mixing of fuel and an oxidizing agent is further prompted, and distribution of fuel in the spanwise direction is improved; and a fuel runner is distributed at the leading edge of the supporting plate, so that the thermal load capacity of the supporting plate is effectively improved.

The invention belongs to the scramjet technology and particularly relates to a supersonic lower jaw air intake duct front blocking cap. The supersonic lower jaw air intake duct front blocking cap is mounted on the head (1) and an air intake duct lip (3) of an aircraft and used for closing and opening an air intake duct cavity (2). The supersonic lower jaw air intake duct front blocking cap comprises a plurality of first adjustment sheets (5), sliding pins (6), connecting rods (7), a plurality of second adjustment sheets (8) and fixing pins (9), wherein the first adjustment sheets (5) and the second adjustment sheets (8) are arranged in an alternately spaced mode; the fixing pins (9) and the sliding pins (6) are arranged on the first adjustment sheets (5) and the second adjustment sheets (8) correspondingly; the sliding pins (6) on the adjacent adjustment sheets are movably connected with the fixing pins (9) through the connecting rods (7); the sliding pins (6) are arranged in sliding grooves of the head (1) of the aircraft and can slide freely in the sliding grooves (4). By using the supersonic lower jaw air intake duct front blocking cap, the problem that an integrated jettisonable front blocking cap threatens an aircraft carrying platform is avoided; meanwhile, the research and development cost can be minimized; the supersonic lower jaw air intake duct front blocking cap canbe repeated used at the research and development stage; and the reliability and safety are improved.

The invention discloses a scramjet high-temperature pure air test system, which comprises a low-pressure air supply unit, an oxygen supply unit, a nitrogen supply unit, a high-pressure air supply unit, a natural gas supply unit, a gas generator, and a regenerative heater , circulating cooling water supply unit, gas discharge unit, direct connection test device and test exhaust discharge unit; through the combined use of the above units, the three stages of gas preheating, pressurized heat exchange and direct connection test operation have been realized for a long time. Safe and reliable run.