35 results about "Rocket-based combined cycle" patented technology

The invention discloses a variable working condition type primary rocket system of a rocket based combined cycle engine. The variable working condition type primary rocket system comprises a propellant supply system and a primary rocket engine thrust chamber which are connected through a pipeline, wherein the propellant supply system comprises an oxidant supply system and a fuel supply system, the oxidant supply system comprises an oxygen flow rate control system, the oxygen flow rate control system comprises a plurality of oxygen conveying pipelines, each oxygen conveying pipeline is provided with an oxygen pressure relief valve for opening and closing the corresponding oxygen conveying pipeline, an outlet of each oxygen conveying pipeline is connected with an oxygen perforating plate, the fuel supply system comprises a plurality of fuel conveying pipelines, an outlet of each fuel conveying pipeline is provided with a cavitation venturi pipe, and an inlet of each fuel conveying pipeline is connected with a nitrogen pressure relief device. The variable working condition type primary rocket system of the rocket based combined cycle engine has the advantages that the flow rate of a propellant can be accurately, controllably, stably and quickly adjusted, the complexity is low, and the operation is simple.

The present disclosure generally pertains to rocket based combined cycle (RBCC) propulsion units. In one exemplary embodiment, at least one rocket thruster is integrated with a jet engine but is external to the flow path of the jet engine, forming an altitude compensating plug nozzle. Since the rocket thruster is external to such flow path, the rocket flow from the rocket thruster interacts with the jet flow from the jet engine aft of the nozzle of the jet engine. Such interaction occurs without a significant performance penalty in the operation of the jet engine. In fact, it is possible that the interaction of the rocket flow with the jet flow may actually improve the efficiency of the jet engine under some conditions. Moreover, having the rocket thrusters positioned external to the flow path of the jet engine helps to avoid many of the problems plaguing conventional RBCC propulsion units.

The invention provides a hydrogen-peroxide-based RBCC (Rocket Based Combined Cycle) engine pressure pumping and squeezing integrated fuel system. The hydrogen-peroxide-based RBCC engine pressure pumping and squeezing integrated fuel system comprises a pressurization unit, a supplying unit, an engine body and a connecting pipeline unit; the pressurization unit is used for performing pressurization on the supplying unit; the supplying unit which is connected with the engine body through the connecting pipeline unit supplies propellant to the engine body. The hydrogen-peroxide-based RBCC engine pressure pumping and squeezing integrated fuel system has the advantages of overcoming disadvantages of a complex and huge system due to the fact that a rocket thrust chamber and a stamping combustion chamber separately supply fuel, synthesizing respective advantages of the hydrogen peroxide and kerosene propellant, being high in density specific impulse, satisfying combination engine design requirements for integrated fuel supplying and an integrated cooling structure and being mature in utilization and maintenance technology, relatively simple in system composition and easy to achieve.

The invention provides a rocket-based combined cycle engine thrust calculation method. Firstly, one-dimensional geometrical parameters of an inner flow channel of the rocket-based combined cycle engine are determined and include an air inlet channel inlet sectional area, an air inlet channel compression angle, an e isolation section sectional area, a sub-expansion section expansion ratio, a combustion chamber expansion ratio, a rocket combustion chamber sectional area, a rocket nozzle size and an engine exhaust nozzle expansion ratio; the flight height, the flight Mach number and the mainstream flow are input; and then thrust estimation is conducted on the rocket-based combined cycle engine, and the final thrust of an outlet of an exhaust nozzle of the rocket-based combined cycle engine and the specific impulse of the engine are obtained. According to the method, rapid estimation of thrust and specific impulse performance of the rocket-based combined cycle engine is achieved, and a newtool is provided for initial design of the engine and trajectory planning of an aircraft.

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 discloses a fuel supporting plate for improving non-premixed combustion of a RBCC (rocket-based combined cycle) bimodal combustion chamber. The streamwise vortex fuel supporting plate is in a structure with unilateral wedged bosses, and two fuel supporting plate bodies are symmetrically distributed on two sides of a rocket. The front portion of the fuel supporting plate is wedged, the rear portion of the fuel supporting plate is straight, and three wedged bosses at equal intervals are arranged on the outer side of the front portion and fixedly connected with the supporting plate. The upper portion of a fuel jet block is provided with two screw holes connected with fuel jet hoses, fuel passage holes in the fuel jet block are communicated with the screw holes, and the lower portion of the fuel jet block is connected with a passage on the fuel supporting plate. The fuel jet block is fixed onto an engine, and jet positions can be changed in different experimental operating conditions. By means of the structure of the fuel supporting plate, guide vortexes of fluids on the lateral sides of the supporting plate can be intensified to enable fuel mixing to be more effective, a regular shear layer on the rear side of the supporting plate is reasonably destroyed to achieve higher turbulence level of the shear layer on the rear side of the supporting plate, the area of a combustion boundary layer is broadened, and the fuel can be better mixed with an incoming flow to form a larger combustion area.

The invention provides an ejector rocket for rocket-based combined cycle engine. The ejector rocket includes a jet pipe. The jet pipe is a plug type jet pipe. According to the invention, in the ejector rocket, the plug type jet pipe is adopted instead of a conical or bell-shaped jet pipe in the prior art so as to improve the pressure at the bottom of the jet pipe of the ejector rocket in a stamping mode and reduce the bottom resistance. Therefore, a recirculation zone at the bottom of the ejector rocket is reduced and the total pressure loss is reduced. Meanwhile, in the stamping mode, the plug type jet pipe can reduce the area of a high temperature zone and reduce thermal protection requirements.

The invention relates to a rocket-based combined cycle engine. The rocket-based combined cycle engine comprises a shell body, an air inlet, a central cone, a rocket combustion chamber, a supersonic-combustion combustion chamber and a supersonic-combustion spray pipe. The air inlet and the supersonic-combustion combustion chamber are arranged in the shell body sequentially, the central cone is arranged in the air inlet, the tail end of the central cone is connected with the rocket combustion chamber, and the tail end of the rocket combustion chamber is connected with the plug-type spray pipe. The rocket combustion chamber comprises a cylindrical shell and a rocket combustion chamber inner core arranged in the cylindrical shell, wherein one end of the cylindrical shell is connected with thecentral cone and is in smooth transition connection with the central cone; one end of the rocket combustion chamber inner core is connected with the central cone; and the rocket combustion chamber inner core coincides with the central axis of the cylindrical shell. Incoming flow air enters the supersonic-combustion combustion chamber from an inlet of the air inlet through the air inlet, and the tail end of the supersonic-combustion combustion chamber is connected with the supersonic-combustion spray pipe. According to the rocket-based combined cycle engine, a traditional rocket engine bell-shaped spray pipe is changed into the plug-type spray pipe, and it is guaranteed that the engine has higher performance in a wide range by utilizing the automatic matching characteristic of the plug-typespray pipe and ambient pressure.

The invention discloses structural design of an inlet isolation segment of a rocket-based-combined-cycle engine. Under the restraint that the minimal section of the inlet isolation segment is located in a throat segment, the sectional area of an effective circulation area of the isolation segment is improved, so that the sectional area of the inlet throat segment is increased to be equal to the sectional area of an intake of a combustion chamber, effective circulation area equal-sectional-area circulation is achieved on an inlet rocket supporting plate segment, and an equal straight segment behind the inlet rocket supporting plate segment is improved into a contraction segment so as to be matched with the section of the intake of the combustion chamber. It is ensured that the minimal section area of the inlet isolation segment is located in the throat segment. By means of the inlet isolation segment, the start mach number of an inlet is lowered, and the flow coefficients of the inlet in an injection mode and a ramjet mode and other performance parameters are increased. The design of the inlet isolation segment of the rocket-based-combined-cycle engine is high in development performance, the inlet isolation segment is designed according to different sectional sizes of the intake of the combustion chamber, the work range of the inlet can be effectively expanded, and the overall performance of the engine can be improved.

The invention discloses a rocket based combined cycle (RBCC) combustion chamber adopting a slope-type rocket layout mode, and a design method of the RBCC combustion chamber. The combustion chamber issequentially provided with an isolation section, a rocket slope mounting section and an expansion section from an air inlet to an outlet of gas generated by combustion. The isolation section, the rocket slope mounting section and the expansion section are sequentially arranged from the air inlet to the outlet of the gas generated by combustion. A rocket thrust chamber is mounted at the rocket slope mounting section, so that a rocket jet flow is in three-side contact with a stamping main flow and is located in front of a front peak of a fire flame of the expansion section, ignition and flame stability of the combustion chamber are achieved through the rocket jet flow, and the problem that through a step rocket layout mode, rocket high-temperature and high-enthalpy activity jet flow ignitionand flame stability cannot be effectively utilized is avoided.

The invention discloses a motive sealing structure of a variable-geometry combustor of a rocket based combined cycle engine. The mode that contact type sealing and non-contact type sealing are combined is adopted, so that rotary motive sealing of the joint of hinges and transverse motive sealing of all top plates and side plates of the combustor are achieved. A sealing groove is formed in the end face of a connection flange between the fixed section and the variable section of the combustor. Sealing grooves are formed in the two side edges of a bottom plate of an inner cavity of the combustor. The top plates of the inner cavity of the variable section of the combustor are connected with arc grooves to form mechanical sealing. Graphite bars in longitudinal grooves and a rotation shaft are sealed through packing. The top plates of the variable section of the combustor are provided with labyrinth channels. The top plates and the side plates are in transverse motive sealing. Red copper bars in the horizontal grooves formed in the side faces of the top plates and the side plates are sealed through packing. The red copper bars of every two top plates are attached to the joints of the top plates to achieve mechanical sealing. The sealing structure can meet the sealing requirement of the engine and also has the advantages of being simple in structure, convenient to machine, and easy to implement.

The invention discloses a variable-structure air inlet channel of a rocket-based-combined-cycle engine. The variable-structure air inlet channel comprises a variable-structure air inlet channel inner compression section, and a bottom plate of the variable-structure air inlet channel inner compression section comprises an inner contraction section bottom plate and a throat section bottom plate which are connected through a second rotating device from front to back; the front end of the air inlet channel inner compression section communicates with an air inlet channel front body section, a bottom plate of the air inlet channel front body section is an air inlet channel front body bottom plate, and the rear end of the air inlet channel front body bottom plate is connected to the front end of the inner contraction section bottom plate through a first rotating device; when the mach number of the free jet is increased, the inner contraction section bottom plate rotates towards a cavity of the air inlet channel inner compression section by taking the first rotating device as a rotating shaft; and meanwhile, the throat section bottom plate takes the second rotating device as a rotating shaft and rotates in the direction far away from the cavity of the air inlet channel inner compression section. According to the variable-structure air inlet channel of the rocket-based-combined-cycle engine, the working performance under the injection and subsonic-combustion modes is effectively improved, and meanwhile, the performance loss of the air inlet channel under the supersonic-combustion mode is not caused.

The invention discloses a rocket based combined cycle engine rocket ejector T-type layout comprising a central supporting plate rocket ejector and a plurality of side-wall rocket ejectors. The side-wall rocket ejectors are located above the central supporting plate rocket ejector. The front portions of the side-wall rocket ejector are connected to rocket combustion chamber pipelines, the central supporting plate rocket ejector is connected to a combustion chamber through upper-side pipelines, and the section of two sets of rockets is the interface of an insulating segment and a mixing segment. By means of limitation on T-type rocket ejector layout size, and a layout manner with smaller duty ratio than a full throughout central supporting plate layout manner and larger wetted perimeter than a side-wall layout manner is provided; congestion effect on the insulating segment of a full throughout supporting plate type rocket ejector and flight drag of scramjet mode are reduced, and cutting area of a primary jet flow and secondary jet flow are increased; the scalability of the rocket ejector T-type layout is high, and the rocket ejector layout is designed according to different mixing segment inlet section sizes so as to acquire excellent engine performance.

The invention discloses a rocket based combined cycle engine combustion chamber which comprises a combustion chamber cavity. Two fuel support plates are vertically arranged in the combustion chamber cavity, and the height of the fuel support plates is consistent with that of the combustion chamber cavity. The rear ends of the two fuel support plates are mounted in the combustion chamber cavity through rotary shafts. When the support plate rocket combustion chamber works, the two fuel support plates are consistent with the incoming flow direction, the front ends of the two fuel support plates are on the corresponding same cross section, and the rear ends of the two fuel support plates are on the corresponding same cross section. When the support plate rocket combustion chamber stops working, the two fuel support plates rotate around the rotary shafts in different directions to be perpendicular to the incoming flow direction, and the two fuel support plates are on the same straight line so as to divide the combustion chamber cavity into a first-stage return flow zone in front of the fuel support plates and a second-stage return flow zone behind the fuel support plates. The rocket based combined cycle engine combustion chamber meets air flow matching under different incoming flow conditions.

The invention discloses a curved-compression variable geometric gas inlet channel of a rocket-based-combined-cycle engine, the curved-compression variable geometric gas inlet channel comprises a precursor section of the gas inlet channel, an inner contraction section of the gas inlet channel, and an isolation section of the gas inlet channel, which are connected in sequence from front to back; theisolation section of the gas inlet channel is provided with a lip cover plate and a throat plate which are connected with each other. The throat plate comprises a front throat plate and a back throatplate from front to back. The front throat plate is fixed on a frame, and the back throat plate and the lip cover plate can slide back and force along the side wall of the gas inlet channel. Under the premise of improving the startability of the gas inlet channel, the working performance under ejector and sub-combustion modes is effectively improved, and the performance loss of the gas inlet channel in a super-combustion mode is not caused at the same time.

The invention provides an integrated rocket ejector engine assembly. The integrated rocket ejector engine assembly comprises a stamping gas inlet path, three support plates and a rocket ejector, wherein the three support plates are separately an oxidant support plate, a fuel support plate and a voltage test circuit board; and the three support plates are used for hanging the rocket ejector in thestamping gas inlet path. The invention further provides a rocket based combined cycle engine. The integrated rocket ejector engine assembly can reduce the structural dimension, so that influences on air incoming flow are reduced.