93results about How to "Improve thrust performance" patented technology

The invention discloses a combined cycle engine and a hypersonic aircraft which includes a rotary detonation ramjet engine and a rocket engine arranged therein. The rotary detonation ramjet engine includes a shell and a conical inner body, wherein the shell is in a hollow cylindrical shape and two ends communicate, and the after body of the conical inner body stretches into the shell from an air inlet end of the shell and is connected with the shell. And a gap between the shell and the after body forms a drainage channel for air to be led in the shell. A cavity between the rear end surface ofthe after body and the inner wall of the shell forms a detonation chamber which communicates with the drainage channel, and a first exhaust nozzle communicates with the detonation chamber. The rocketengine includes a combustion chamber arranged in the conical inner body, a second exhaust nozzle communicating with the combustion chamber, and the nozzle of the second exhaust nozzle communicates with the detonation chamber. The outer wall of the shell is machined with an outer nozzle, the two ends of the outer nozzle correspondingly communicates with a fuel supply device and the drainage channel. The outer wall of and / or the after body is machined with multiple internal nozzles, two ends of the internal nozzles correspondingly communicate with the fuel supply device and the drainage channel.

The invention discloses a geometrical variable flame stabilizing device and belongs to the technical field of afterburner flame stabilization of aircraft engines. The geometrical variable flame stabilizing device disclosed by the invention utilizes a V-shape as a basic bluff body structure and consists of a main body part and a transmission power part, wherein the main body part is vertically installed in an afterburner cartridge receiver, the afterburner cartridge receiver comprises a front streamline head part, a blade A with a rotating shaft A, a blade B with a rotating shaft B, a rotating shaft supporting plate and a mounting seat, and the two ends of the streamline head part are fixedly connected between the mounting seat and the rotating shaft supporting plate. The geometrical shape of the flame stabilizing device provided by the invention can be changed to effectively reduce the flow resistance of the flame stabilizing device if necessary, or the flow resistance is increased within a short time to improve the ignition success rate, or the geometrical shape of the flame stabilizing device is adjusted to consistently keep the flame stabilizing device at the optimal combination point of the flame stabilization and flow resistance loss so as to obtain the maximal thrust performance of the afterburner.

The invention provides a jacking-pushing-type final joint for an immersed tunnel and a butt-joint construction method thereof, and belongs to the technical field of immersed tunnel construction. Underwater water stop and push stop of the final joint after jacking pushing can be quickly and effectively achieved, the water stop and push stop effects are good, and the safety is high. The jacking-pushing-type final joint comprises a connecting bearing opening formed in a to-be-mounted pipe segment, further comprises a jacking pushing piece, and further comprises a water stop assembly. A jacking pipe section is sleeved with the connecting bearing opening, a cavity capable of containing the jacking pushing piece is formed in the connecting bearing opening, and a grouting pipe is pre-embedded inthe cavity. The water stop assembly comprises three water stop belts; the first water stop belt is arranged at the end face of the pushed-out end of the jacking pipe section; the second water stop belt and the third water stop belt are both arranged on the periphery of the jacking pipe section; the second water stop belt can extend along with pushing out of the jacking pipe section; one end of thesecond water stop belt is fixedly connected to the jacking pipe section, and the other end of the second water stop belt is fixedly connected to the connecting bearing opening; and the third water stop belt is connected between the jacking pipe section and the connecting bearing opening in a sleeved mode.

The invention relates to a device for a hypersonic pitching dynamic test, which comprises a test piece fastening cone, a connecting rod, an adjusting cone, pitching bearings, an elastic hinge and a bearing baffle plate. The elastic hinge consists of a spring piece of which two ends are provided with platforms and the middle bends into an alpha angle; and beveled edges on two sides are respectively adhered with two resistance strain gages for measuring the pitch angle strain. Due to the adoption of the elastic hinge with a special structure, the axial force of the pitching bearing is unloaded, the anti-thrust capability of the pitching bearing is improved; and when the device is used for measuring the strain capacity and angular movement, the testing accuracy can reach 1 / 60 degree. Due to the adoption of the elastic hinge with the special structure and replaceable thickness and material, the inherent frequency of the test device can be adjusted, and the test can be suitable for wider flying range chain. The test piece is fixed by four pitching bearings, so that pitching movement of the test piece is ensured, and moments of the test piece in the side direction and rolling direction can be unloaded. The test device is simple; and vibration is transmitted through main components, so that the angular movement and damping characteristics of a hypersonic aircraft in the process of dynamic pitching can be accurately provided.

The invention discloses a civil aircraft complex system fault diagnosis method based on a Bayesian network, relates to the field of civil aircraft complex system fault diagnosis, can effectively utilize data collected in the flight process of an aircraft and experience of an engineer to quickly and accurately remove faults, and guarantees arrangement of a normal flight plan of an airline. According to the invention, the problem that civil aircraft complex system route fault diagnosis and isolation consume time and labor is solved; the data acquired by the data recorder only needs to be analyzed after navigation; and then the mapping relation table of each node of the fault and fault presentation layer is input, the network can give out a corresponding diagnosis result, the method is more convenient and quicker than looking up a manual, the diagnosis result obtained through the change of actual data is more scientific, the maintenance cost can be saved for airlines, and the utilizationrate of airplanes is improved.

The invention aims at overcoming the defects in the prior art and provides a control method for a turbine-based double-combustion-chamber punching combined cycle engine to solve the problem that relay of an existing turbine punching combined engine cannot be achieved under the low Mach number. By means of the control method, the double-combustion-chamber punching engine is improved firstly, and combustion chambers arranged in a rectangular parallel connection manner; an air inlet way of the engine is improved, the two-dimensional air inlet way configuration is adopted in the air inlet way, the air inlet way is divided into at least one subsonic combustion runner and at least two supersonic runners through supporting plates, wherein the supersonic runners are evenly arranged on the two sides of the subsonic combustion runner; and inner contraction section adjusting molded surfaces and expansion section adjusting molded surfaces are additionally arranged in the design of all the runners so that the size of the area of a throat of the air inlet way can be controlled, the requirement for air with different compression degrees of the combustion chambers is met, the Mach number range of work of the engine is enlarged, and the performance of the engine is improved.

The invention discloses an air-breathing rocket motor and a hypersonic speed plane. The air-breathing rocket motor and the hypersonic speed plane include an air inlet, a heat exchanger, a gas compressor, a main combustion chamber and an exhaust nozzle, and the air inlet, the heat exchanger, the gas compressor, the main combustion chamber and the exhaust nozzle are arranged in sequence. The gas compressor is provided with a turbine, and the turbine provides driving force for the gas compressor. A wall surface cooling channel is arranged on the outer wall surface of the main combustion chamber and the exhaust nozzle. The air-breathing rocket motor and the hypersonic speed plane also include an oxidizing agent pump, a fuel pump, a pre-burning chamber and an injector. The oxidizing agent pumpcommunicates with the heat exchanger so that oxidizing agent enters the heat exchanger to cool air which enters the air inlet. The fuel pump communicates with the wall surface cooling channel, so thatfuel enters the he wall surface cooling channel to cool the exhaust nozzle and the main combustion chamber. The oxidizing agent after cooling the air and the fuel after cooling the exhaust nozzle andthe main combustion chamber correspondingly enter the pre-burning chamber for combustion to generate rich combustion gas in the pre-burning chamber. The injector is used for spraying the air cooled by a heat exchanger and pressurized by a gas compressor into the main combustion chamber for mixed combustion, as well as the rich combustion gas after driving the turbine.

A counter-rotating 2 cycle piston engine with a combustion chamber and an output shaft rotatably mounted to rotate independently about the same propeller centerline of an aircraft. the Internal combustion power is transmitted to the output shaft by means of a conventional piston, rod, crankshaft configuration journaled to the combustion chamber or engine block in the traditional manner thereby producing concurrent rotation of the shaft and counter-rotation of the engine block. Interior surfaces of the combustion chamber and pistons are angled to control the oil flow inside the crankcase and eliminate unwanted ponding. The pistons are modified to transfer fluid accumulations to the combustion chamber. A throttle is located at the axis of rotation to function normally without the effects of centrifical force. Dual propellers mounted on the shaft and on the block improve thrust performance, balance the torques and moments of inertia of the two counter-rotating masses, All rotational output torque automatically balances itself between counter rotating propellers and is converted into propulsion. No rotational torque is allowed to be transmitted to the airframe through the counter-rotating engine mount.

The invention discloses a double-bell type expansion deflection nozzle. The double-bell type expansion deflection nozzle comprises an expansion deflection stopper cone, a nozzle converging section, adouble-bell type base expanding section and a double-bell type extension expanding section. The expansion deflection stopper cone is partly designed by a plurality of sections of circular curves, andthe expansion deflection stopper cone and an external wall surface form an annular throat, so that an air flow undergoes deflection flowing not parallel to the axial direction of an engine; an inflection point is formed between the double-bell type base expanding section and the double-bell type extension expanding section through a difference between a moulding surface and a nozzle axis chamfer,so that the air flow is separated at the inflection point when the double-bell type expansion deflection nozzle works at a low pressure drop ratio; the expansion deflection nozzle formed by the double-bell type base expanding section and the expansion deflection stopper cone is effectively utilized for altimetric compensation, so that relatively high push force performance is maintained; and at the high pressure drop ratio, the air flow expands along the double-bell type extension expanding section to a nozzle exit all the time, and the area of the nozzle exit is effectively utilized, so thatthe push force performance at the high pressure drop ratio is ensured.

The invention discloses a designing method for a supersonic velocity thrust exhaust nozzle considering inlet parameter unevenness. The designing method includes the following steps that (1), a spiral characteristic line method is adopted for determining distribution areas of an initial line according to inlet parameter distribution of the thrust exhaust nozzle to be designed; according to inlet parameters, the designing pressure ratio and the selected asymmetrical factor G of the thrust exhaust nozzle to be designed, an upper wall face initial expansion angle and a lower wall face initial expansion angle at the throat sharp point of the thrust exhaust nozzle to be designed are respectively determined; (2), all discrete point coordinates and flow field parameters at the throat sharp point of the thrust exhaust nozzle to be designed are firstly determined so that flow parameters at the throat sharp point can be obtained, and then flow parameters of all characteristic lines of the core area of the thrust exhaust nozzle to be designed are determined; (3), a wave absorbing method is adopted for determining an upper wall face curve and a lower wall face curve, and therefore the thrust exhaust nozzle to be designed can be designed. Therefore, the thrust exhaust nozzle considering inlet parameter unevenness can be designed, and good thrust performance can be generated.

The invention discloses a suction type two-stage shock wave focusing ignition engine combustion chamber and a working method thereof, and belongs to the technical field of detonation engines. The combustion chamber comprises a first cavity body, a first combustion chamber shell, a first shock wave focusing concave surface, two rotating valve flaps, a pre-cooling chamber concave surface, a second shock wave focusing concave surface, a divergent section, a second combustion chamber cavity body, a cold oil pipe, a first heat oil pipe, a second heat pipe and a rotary fuel atomization nozzle. By rotating the rotating valve flaps, intermittent jet flow of air is controlled, so that intermittent shock wave focusing ignition detonating mixed combustible gas in a shock wave focusing chamber is realized, and the reliability of detonation of the engine combustion chamber is further improved through the forward and reverse arranged two-stage shock wave focusing concave cavities. According to the combustion chamber and the working method, the combustion chamber can stably realize high-frequency shock-wave-focused ignition while axially air sucking like a traditional engine, the engine combustion chamber quality is effectively improved, the thrust performance of the engine is improved, and a new technical way is provided for developing a practical high-power detonation engine.

The invention discloses a supersonic thrust nozzle inverse design method based on a maximum thrust theory, which comprises the following steps of: presetting physical parameters of a core point, and further determining the physical parameters and a geometric position on a left-hand characteristic line passing through the core point; according to the nozzle inlet parameter distribution and the corepoint parameter, calculating by adopting a whirl characteristic line method and an iterative method to obtain an initial expansion section, and determining a nozzle inlet and expansion surface influence domain according to the inlet parameter and the initial expansion section; determining the last two characteristic lines emitted by the upper expansion surface and the lower expansion surface of the influence domain, and obtaining an intersection point which is a characteristic line intersection point; connecting the core point to the feature line intersection point; according to the law of conservation of flow, the inlet influence domain and the left-hand characteristic line passing through the core point, determining the upper wall surface profile of the thrust nozzle to be designed by adopting a spiral characteristic line method; and determining a lower wall surface profile to finish the design of the thrust nozzle. According to the invention, the performance is maximized under thecondition of meeting geometric and aerodynamic constraint conditions, and the fusion of the nozzle and the fuselage of the aircraft is better.

A fixed sized bell rocket nozzle is lined with a layer of combustible material that is ignited during launch ignition and burns to outgas into the rocket exhaust for spatially variably confining the exhaust and perfecting an effective variably sized altitude compensating exhaust nozzle that maximizes lift during the launch of a spacecraft into orbit.

The invention discloses a continuous staggered pole permanent magnet synchronous linear motor which comprises a primary, a secondary and an air gap formed between the primary and the secondary. The primary comprises multi-tooth primary iron cores and primary armature windings. The secondary comprises a secondary back yoke iron core, a permanent magnet-iron pole array, and a magnetic isolation material. The permanent magnet-iron pole array is arranged at the upper surface of the secondary back yoke iron core and is laterally divided into two groups with a same length, each group of iron pole and secondary back yoke iron core are integratedly formed, and a magnetic isolation material is arranged between the two groups. The polarities of one group are alternately arranged according to a sequence of N-iron pole-N, the polarities of the other group are alternately arranged according to a sequence of iron pole-S-iron pole, permanent magnets used by the two groups have opposite polarities and are arranged in staggered way, the distance between adjacent secondary permanent magnets is 2tau, and an array is formed according to a cycle of 2tau. Compared with a conventional permanent magnet synchronous linear motor, the continuous staggered pole permanent magnet synchronous linear motor has the advantages of less permanent magnets, a high utilization ratio of the permanent magnets, good thrust performance, high economy and practicability and can be applied in industrial occasions with low speed movement.

The invention discloses a stationary detonation engine based on the variable wedge angle. The stationary detonation engine based on the variable wedge angle comprises a gas inlet channel, an oblique detonation combustion chamber, a tail jet pipe, a fuel oil jetting and atomizing system and a wedge angle control system. The gas inlet channel enables an incoming flow to generate an oblique shock wave, and pressurizing for temperature increasing is conducted. The oblique detonation combustion chamber contains the incoming flow to mix the incoming flow and fuel, and through the wedge face on the rear portion of the oblique detonation combustion chamber, an oblique shock wave is induced, and thus mixed gas is ignited to generate the oblique detonation shock wave. Combustion products generated in the combustion chamber are further expanded and accelerated in a tail jetting pipe flow channel. The fuel oil jetting and atomizing system jets the fuel in the front of the combustion chamber to promote mixing of the fuel and the incoming flow, and meanwhile pre-combustion of the incoming flow is prevented. According to the aerodynamic condition of the mixed gas of the combustion chamber, the wedge angle control system adjusts the forms such as the wedge face angle in real time, and thus the oblique detonation wave is at an inlet of the tail jet pipe in a stationary mode. According to the stationary detonation engine based on the variable wedge angle, through a wedge face control device on the rear portion of the combustion chamber, combustion of the mixed gas is controlled at stable theoblique detonation form, thus combustion is basically at an optimum state so as to optimize the thrust performance of the engine, and the effect that the engine can continuously work under the variable work conditions is achieved.

The invention relates to an outside-sealed displacement adjusting mechanism for a parallel-connection spray pipe, and relates to a combined engine tail spray pipe. The outside-sealed displacement adjusting mechanism for the parallel-connection spray pipe is provided with an actuator cylinder, a connecting rod, a rocker arm, a fixed hinge, a long sliding block, a long sliding block cam shaft, a spray pipe side plate, a spray pipe intermediate diaphragm, a spray pipe upper wall surface, a splitter plate and a spray pipe lower wall surface. The actuator cylinder can move left and right; the rocker arm rotates around the fixed hinge; a through hole is formed in the lower end of the rocker arm; the long sliding block cam shaft can move in the hole; a side plate slot is formed in a spray pipe side wall; a through hole is formed in the center of the slide plate slot; the long slide block can slide in the side plate slot; the length of the long sliding block is larger than the length of the through hole of the side plate slot, so that during a sliding process of the long sliding block, the sliding block can always cover the through hole of the side plate slot, and a sealing effect is achieved; the long sliding block cam shaft is arranged at positions, on two sides of the side plate, on the long sliding block; the long sliding block cam shaft on the outer side of the side plate is connected with the rocker arm through the through hole; and the long sliding block cam shaft on the inner side of the side plate is connected with the splitter plate through a splitter plate sliding chute.An air flow state of a spray pipe outlet can be improved, and better thrust performance is obtained.

The invention provides a boundary pumping control method for supersonic velocity knocking stabilizing and self-maintaining. A supersonic combustion ramjet engine in the method comprises a gas input channel, an isolating section, a combustion chamber and a supersonic velocity spraying pipe. A heat jet flow hole for heat jet flow in a heat jet flow pipe to be sprayed out is arranged in the combustion chamber of the engine. A plurality of pumping holes are formed in the inner walls, before and behind the heat jet flow hole, of the combustion chamber. The pumping holes communicate with a pumping pump through pipelines. When the heat jet flow is adopted for carrying out knocking detonating in supersonic velocity airflow, the pumping pump and the pumping holes pump the airflow of an inner boundary layer of the combustion chamber in a pumping manner, and meanwhile, the heat jet flow is sprayed in to achieve knocking detonating. According to the method, in the knocking detonating process, thepumping pump is started, the pumping holes pump the airflow of the low-speed area of the inner boundary layer of the combustion chamber, the mutual effect between detonating knocking waves and the boundary layer is restrained to prevent forward passing of combustion waves, the influence of the boundary layer to detonating of the knocking waves is eliminated, and therefore the knocking waves can betransferred in the combustion chamber dynamically and stably.

The invention discloses a side-feed supply device for a pulsed plasma electric thruster. The side-feed supply device comprises two sets of supply assemblies of symmetric structures, the supply assemblies are fixedly connected to a bottom plate of the electric thruster, the supply assemblies comprise a first supply bracket, a first tension spring, a first push rod, a first supply blocking sheet, afirst solid propellant, a second supply bracket, a second tension spring, a second push rod, a second supply blocking sheet and a second solid propellant, the opposite ends of the first solid propellant and the second solid propellant are of bevel structures which constitute a V-shaped discharge groove, an opening of the V-shaped discharge groove faces the tail end of an electrode plate, and a cone limiting table is arranged at the position, at the included angle of the V-shaped discharge groove, of the inner side of an anode plate. According to the side-feed supply device, the two solid propellants can be fed, a V shape is constituted by the ablation end faces of the propellants in a discharge room, thus the area swept by an electrical discharge arc can be increased, and the ablation quality of single pulse under the same condition is also improved.

The invention relates to a mute barbell roller bearing which has a structure support technology that is characterized in that the external dimension standard of the mute barbell roller bearing is in line with the technical standard of an international general roller bearing of international standardization organization (ISO). The mute barbell roller bearing is formed by correspondingly matching an inner sleeve ring formed by two separated rings 1, a row of barbell rollers 3, an outer sleeve ring 2 and an invisible retainer 4. The invention aims at solving the technical problem that the traditional N-type series cylindrical roller bearing is not capable of thrusting. The mute barbell roller bearing has the most remarkable selling points that (1) the dynamic flexibility is improved by 20%; (2) the jump noise is reduced by 5-14 decibels (dB); (3) the rotating speed is accelerated by 12-30%; (4) the loading capacity is increased by 30-50%; (5) the safety factor is increased by 2%; (6) a sealing cover can be arranged on the mute barbell roller bearing; (7) axial thrust capability is improved by 70%; and (8) the servo life is prolonged by 14-30%, therefore, the mute barbell roller bearing is a three generation leading edge innovative technology product which is economical and environment-friendly.

The invention discloses a rocket ramjet wide-speed-range engine based on annular pressurization center body tail confluence. The engine comprises a center body, an air inlet channel, a combustion chamber with a concave cavity and a spray pipe. An embedded rocket in a conventional rocket ramjet engine is arranged on the outer side of the center body, the rocket gas and air are mixed in an annular mixing channel, and the contact area of rocket jet flow and the air is increased, so that the mixing distance of two air flows can be effectively shortened. Meanwhile, annular airflow is converted into central pipe flow through a central body contraction section at the spray pipe, the outlet area of the engine spray pipe can be reduced, and therefore the structural size of the engine is reduced, and the overall thrust performance of the rocket ramjet engine can be effectively improved by adopting the scheme.

The invention discloses a throat excursion type pneumatic vector nozzle with asymmetric rear body profiles. The throat excursion type pneumatic vector nozzle comprises a nozzle body, and the rear bodyprofiles of the nozzle body are arranged to be of asymmetrical structures. Under general conditions, the upper and lower rear body profiles of the outside of an outlet of the nozzle are asymmetrical,main flow flowing out of the outlet of the nozzle acts with an asymmetric nozzle rear body to generate a low pressure zone, the main flow is promoted to further be close to the wall surface of the nozzle rear body, and the nozzle vector angle is increased. Meanwhile, the wall face of the wall surface of the asymmetric nozzle rear body and the main flow of the nozzle generate interaction with different strengths and different ranges, and the nozzle pitch vector angles are not equal. Through the design of the asymmetric nozzle rear body profiles, the vector angle of the throat offset pneumaticvector nozzle is further increased, the pitch vector angles are not equal, the infrared stealth performance of the nozzle is improved, and the actual needs of the aircraft are met.

The invention provides a musical whistle for physical education. The musical whistle includes a whistle mouth, and the whistle mouth is connected with a resonant cavity with a circular vertical section; a sharp separation block is fixed at the connection position of the whistle mouth and the resonant cavity, and a vent hole is formed in the portion, at the rear upper portion of the sharp separation block, of the resonant cavity. A movable and small ball is provided in the resonant cavity, the inside of the movable and small ball is a hollow cavity, and air inlet hole is formed in the movable and small ball. The sound generated by the musical whistle for physical education has a higher degree of roundness and the whistle is more tuneful than a traditional whistle.

The invention relates to a grouting sleeve with multi-point-distributed inner cavity bulges and a manufacturing method thereof. The grouting sleeve comprises a metal pipe, wherein the two ends of an inner cavity of the metal pipe are provided with the plurality of bulges correspondingly, the bulges are arranged towards the central axis of the metal pipe, the bulges can restrain grouting materialspoured in the metal pipe, the bulges are uniformly distributed in the axial direction and the circumferential direction of the metal pipe, stress concentration of the metal pipe at the bulges is reduced, each bulge comprises a vertical face and an inclined face, the vertical faces are perpendicular to the central axis of the metal pipe, the vertical faces are arranged on one sides, close to the center of the metal pipe, of the inclined faces correspondingly, the plurality of columns of bulges are arranged in the axial direction of the metal pipe, a corresponding thrust-stopping inclined tableis formed between the every two adjacent bulges in the same column, and the heights of one ends, close to the center of the metal pipe, of the thrust-stopping inclined tables are smaller than the heights of one ends, away from the center of the metal pipe, of the thrust-stopping inclined tables.

The invention relates to a supersonic combustion chamber fuel injection design method based on air inlet channel non-uniform airflow, and the method comprises the steps: obtaining spray hole parameters of a combustion chamber according to supersonic air inlet channel parameters and engine parameters, wherein the spray hole parameters comprise the number of spray holes, spray hole groups and spray hole aperture ranges; obtaining the set distance range of the spray hole according to the fuel diffusion characteristic and the ignition delay characteristic; obtaining the set angle range of the spray holes according to the distribution angle of the low-speed airflow area in the combustion chamber, and obtaining the global equivalent ratio of each group of spray holes according to the angle proportion of the distribution angle in the airflow flow direction section. According to the method, the arrangement mode of the spray holes is determined based on the circumferential distribution characteristics of the airflow velocity of the fuel on the flow direction section of the combustion chamber, the circumferential and radial blending combustion efficiency of the fuel on the circular section of the combustion chamber can be improved, and then the engine thrust performance is improved.

The invention relates to a square voice coil motor, consisting of a rotor and a stator. The stator of the motor at least comprises two clearance magnetic fields with rectangular sections, and each clearance magnetic field is formed under the action of an upper magnetic block and a lower magnetic block. The clearance of the magnetic field is increased by magnets with the same volume while the intensity of the magnetic field is not weakened, and the thrust force of the motor can be modified by adding the coils; in addition, the effective travel of the motor is increased under the condition of the same volume, and the cost is saved.

The invention discloses a confluence rocket ramjet wide-speed-range engine based on an annular pressurization center body mixing section. The confluence rocket ramjet wide-speed-range engine based on the annular pressurization center body mixing section comprises a center body, an air inlet channel, a combustion chamber with a concave cavity and a spray pipe. An embedded rocket in a conventional rocket ramjet engine is arranged on the outer side of the center body in a front-arranged mode, mixed pressurization of the rocket gas and air is conducted in an annular mixing channel, and then central pipe flow is formed through confluence. Due to the fact that the contact area of rocket jet flow and air is increased, the mixing distance of two air flows can be effectively shortened. By the adoption of the scheme, the overall thrust performance of the rocket ramjet engine can be effectively improved.

A contrast similar expansion method of product function is characterized by provided with a known product knowledge base storing a function tree and bounded relation information of known product and building an original function tree and bounded relation information of novel product; on the base of combining a similar theory and an extension theory and according to the specific needs of product designers, any function node of the original function tree of novel product is compared with the product functions in the known knowledge base; according to a certain rule, the novel functions are appended to the product function tree to generate a novel product function tree. Compared with the original product function tree, the novel product function tree is richer; therefore, the possibility of obtaining an innovation function is greatly increased. The invention gives full play to the guiding function of designers in the process of product design and makes designers release from the complicated and mechanical work of data comparison, thereby providing a design method which has high freedom, strong deduction capacity and aims at innovation of product function.