52results about How to "Increase the total pressure recovery coefficient" patented technology

The invention relates to a supersonic flow channel design method and apparatus. The method comprises the following steps: step S1, generating a reference flow field molded surface; step S2, calculating the reference flow field obtained in the step S1; step S3, generating a three-dimensional internal steering air inlet channel by using a streamline tracing method; step S4, calculating the flow field of the three-dimensional internal steering air inlet channel obtained in the step S3; and step S5, repeatedly iterating step S1 to step S5 by using an optimization algorithm to obtain a three-dimensional internal steering air inlet channel meeting traffic capture performance, geometric parameters and aerodynamic performance. Compared with the prior art, the method has the advantages of being able to solve the problems that the traditional air inlet channel design does not directly aim at the aerodynamic performance of the air inlet channel and starts from the reference flow field performance, such that it is difficult to ensure the same performance of the air inlet channel, the experience is too high to ensure the optimal value of the performance of the air inlet channel and that the total shrinkage ratio and the internal shrinkage ratio of the internal steering air inlet channel are uncontrollable.

The invention discloses a half split hierarchical S-shaped air inlet channel design method. First, by use of two stationary blade row mean camber line design for reference, two-dimensional half split design idea is used, dimensionless parameter ''load coefficient'' ''expansion coefficient'' control of air inlet channel normal and axial pressure distribution is proposed, a two-dimensional wall surface molded line of an S-shaped air inlet channel is designed, and a two-dimensional wall surface geometric parameterized model is established; total pressure recovery coefficient is used as an optimization objective, and by combination of genetic algorithm and an S-shaped air inlet channel modelling program, the two-dimensional parameterized model is optimized; by use of three-dimensional hierarchy design idea and by introduction of dimensionless parameter ''distribution coefficient'', an S-shaped air inlet channel three-dimensional geometric model is further obtained by determination of center line along the distance; and finally, the total pressure recovery coefficient and outlet distortion are used as an optimization objective, the genetic algorithm is used for multi-objective optimization of the model, and by discuss of physical relationship of the parameters and the air inlet total channel pressure loss and the outlet distortion, an S-shaped air inlet channel with better aerodynamic performance can be obtained.

The invention discloses an embedded type air inlet channel using a combined opening surface vortex control method. The embedded type air inlet channel comprises an air inlet channel inlet, a flow guide surface, a side wall, a rear lip opening and multiple micro vortex generators, wherein the micro vortex generators are arranged on the flow guide surface; the air inlet channel inlet comprises a main inlet and a pair of overflow grooves; and the main inlet is arranged at the front part of the air inlet channel inlet and the overflow grooves extend backwards from two sides of the rear of the main inlet. The air intake strengthening function of the opening surface main vortex of the embedded type air inlet channel is used; directly flowed secondary vortexes, which continuously expand to two sides of the air inlet channel inlet, are generated by the vortex generators, and main vortexes at two sides of the air inlet channel inlet are thrust to the outside of the inlet, so that most of vortex air flow is exhausted to the outside of the air inlet channel through the pair of overflow grooves at the rear end of the inlet; and besides, at the time, the advantage of the embedded type air inlet channel is realized, the adverse pneumatic impact caused by the main vortex is ingeniously avoided, so that the purposes of increasing the total pressure recovery coefficient of the embedded air inlet channel, reducing the distortion index and improving air intake capability are achieved.

The invention provides a contra-rotating rotary-jetting engine which can remarkably raise heat efficiency and reduce pollutant emission, manufacturing cost as well as maintenance cost. The engine works by utilizing a principle that a rotary jet turbine of the engine ejects fuel gas at high speed from a tangential spout to allow the self of the rotary jet turbine to rotate reversely so as to convert fuel gas internal energy into mechanical energy; meanwhile, the engine also utilizes a tangential impact turbine to recycle the kinetic energy of the fuel gas ejected from the rotary jet turbine, utilizes a heat regenerator to realize regenerative cycle, adopts a contra-rotating axial flow and contra-rotating centrifugal pressure gas technique to obtain high total pressure ratio, adopts a rotary flame tube technique to obtain high fuel gas temperature in front of the turbine, adopts the novel structural design to greatly reduce the number of components and the consumption of precious materials, and adopts a power confluence mechanism of a cone gear and an overrunning clutch to realize the power confluent output of the mechanical energy of the rotary jet turbine and the tangential impact turbine. Therefore, the engine has high energy conversion efficiency, low manufacturing cost and pollutant emission.

The invention belongs to the technical field of aero-engines, and particularly relates to a ramjet combustor rectification grille. The ramjet combustor rectification grille comprises a front grille body and a rear grille body, the front grille body is arranged at a diffuser inlet of a ramjet combustor and used for buffering high-speed airflow of the diffuser inlet and guiding part of the bufferedhigh-speed airflow to a center area of a diffuser, and the rear grille body is arranged at a diffuser outlet of the ramjet combustor and used for buffering the air flow in the center area of the diffuser and evenly diffusing the buffered air flow in the center area of the diffuser outwards along the center area. According to the ramjet combustor rectification grille, through the design scheme thatthe front and rear rectification grille bodies are arranged at the inlet and outlet of the diffuser for combined rectification, the defects of an existing ramjet combustor rectification grille are overcome; and compared with a traditional single-layer grille, the rectification capacity of the ramjet combustor rectification grille can be obviously improved, the unevenness of distorted airflow is reduced, and the total pressure loss of the grille can be effectively reduced.

The invention relates to a fluid-solid coupled gas shock wave adjusting bearing. At present, a widely used dynamic pressure foil-bearing is limited in carrying capacity and lacks of stability and the coating process requirements on the surface of a foil are high, and a static pressure-dynamic pressure mixed bearing is insufficient in carrying capacity at a slow rotating speed while is pneumatically unstable at a high rotating speed. The bearing provided by the invention comprises a bearing bush (1) which is provided with air supply throttling orifices (3), damping cavities (4) are arranged in the outlet positions of the air supply throttling orifices, one or more than one row of air supply throttling orifices and the damping cavities is uniformly distributed circumferentially on the inner surface of the bearing bush, and foil structures (7) are arranged between the outlets of the damping cavities and a rotor (2); the foil structures comprises bump foils (6) and flat foils (5). By changing the air supply pressure of the bearing, the throttling orifice structures and the damping cavity structures, the fluid-solid coupled gas shock wave adjusting bearing can be matched with a large-gap high-load bearing working condition. The invention is the fluid-solid coupled gas shock wave adjusting bearing.

The invention discloses a low-resistance efficient subsonic speed air inlet channel which comprises an air inlet channel opening face located at the front end of the air inlet channel and an air inlet channel inner channel extending backwards from the air inlet channel opening face, an air inlet channel lip cover is arranged outside the air inlet channel opening face in a surrounding mode, and the air inlet channel opening face and the air inlet channel lip cover are connected with a machine body. The air inlet channel structure is different from a conventional subsonic air inlet channel, no boundary layer separation channel and no protruding bulge exist in front of the opening of the air inlet channel, through integrated fusion of the sweepback sharp lip and the fuselage, the windward area of an aircraft can be reduced while the high total pressure recovery coefficient and the low flow field distortion index of the air inlet channel are kept, and the air inlet channel is more compact in structure. Therefore, resistance is reduced.

The invention provides a low-external-resistance supersonic / hypersonic speed inlet. A two-dimensional bump is arranged at an inlet of the supersonic / hypersonic speed inlet, and partial air discharge seam / hole arrays are arranged at the inward concave positions of the windward side and the leeward side; and meanwhile, an air discharge cavity below the bump is divided into two sub-cavities, and electromagnets and a metal plate with a folded corner are adopted for achieving control over opening and closing of the two air discharge sub-cavities. By means of control over opening and closing of the two air discharge sub-cavities, the low-external-resistance supersonic / hypersonic speed inlet can adapt to shock wave intensity and incidence point position changes brought by inlet working mach number and attitude angle changes, the optimal control ability on boundary layer separation caused by shock wave / boundary layer disturbance of a lip cover is obtained, and the remarkable effects of widening the working range of the low-external-resistance supersonic / hypersonic speed inlet and increasing the total pressure recovery coefficient are achieved. Due to the high flowing control ability of the low-external-resistance supersonic / hypersonic speed inlet, the inlet is allowed to adopt the design of the horizontal lip cover, and therefore the external pneumatic resistance of an aircraft can be effectively reduced.

The invention discloses a method for designing a fixed-geometry two-dimensional mixed-compression type supersonic velocity air inlet channel. A basic structure of the adopted two-dimensional mixed-compression type air inlet channel comprises an outer pressing section, an inner pressing section, a throat section, an expanding section and a lip. The lip is of a sharp plate structure with a certain wedge angle. When the air inlet channel is designed, designed starting Mach number is improved appropriately, the lip shape is changed, and the starting mach number is effectively reduced to be lower than a requested value through a certain flow on the premise that the flow requirement of the air inlet channel is met. By means of the method, the requirement for starting performance can be met, contraction ratio of the air inlet channel can be increased in design, and accordingly high back pressure resisting capability and low outlet Mach number are obtained in a wide working range to facilitate efficiently combusted organizations in a combustion chamber. Furthermore, small total pressure loss in the working process of the air inlet channel can be achieved through the change of the lip shape. The method for designing the fixed-geometry two-dimensional mixed-compression type supersonic velocity air inlet channel has good actual application value.

The invention relates to a design method of a hypersonic velocity air inlet channel. The design method comprises the steps of creating a reference flow field profile based on a functional relation; optimally designing axial symmetric reference flow fields of the air inlet channel; and then obtaining a three-dimensional profile of the air inlet channel through a streamline tracking method. Comparedwith the prior art, the method has the advantage that the problem that the design of a traditional air inlet channel is performed by designing inner flow field condition firstly, and the total pressure recovery coefficients of the air inlet channel cannot reach the maximum value while the shock wave sealing of a port is realized, can be solved.

The invention discloses a plasma vortex generator-based inward-rotating air inlet component and a control method thereof. The assembly includes a machine body plane and an inward turning air inlet body, wherein one or more sets of arc spark discharge plasma vortex generators are arranged at a certain distance upstream along the compression surface below the machine body plane. This method uses the flow vortex formed by the interaction between the plasma jet and the free flow to promote the momentum exchange between the main flow and the low-energy flow in the boundary layer, enhance the ability of the boundary layer to resist the reverse pressure gradient, and reduce the shock wave / boundary layer induced by the downstream internal compression surface. The interference separation zone has a certain delay or inhibition effect, thereby effectively improving the performance of the intake port. Compared with the traditional method, the control method proposed by the present invention does not require additional air source, no flow loss, no additional resistance, and the control method is accurate and efficient, which has great advantages for the wide speed range and integrated design of the internal rotation air inlet. important promotion.

The invention provides a diverterless subsonic air inlet passage integrated with an aircraft. The diverterless subsonic air inlet passage comprises a lip cover on the basis of electromagnetic stealth and pneumatic consideration and a fully shielded inner pipeline. The shape of an air inlet passage opening surface and a fuselage are in integrated combination designs, and boundary layers at the fronts of inlets can be eliminated by pressure gradients formed by the lip cover and the surface of the fuselage. The lip cover is in sweepback designs in the overlooking and side-looking directions, and sweepback angles are consistent with preset stealth angles of the aircraft. The inner pipeline with an air inlet passage is provided with a selected fully shielded double-S bending type air inlet passage, and the cross sections of outlets of the air inlet passage cannot be seen from optional angles of air inlets. The diverterless subsonic air inlet passage has the advantages that specific bumps, which are protruded on the surfaces of existing fuselages, of conventional diverterless air inlet passages and boundary layer diverters on the existing fuselages can be omitted on the premise that requirements on total pressure recovery coefficients and flow field distortion indexes of the air inlet passage are met, and accordingly pneumatic resistance of the aircraft can be lowered; the radar stealth performance of the air inlet passage can be enhanced under the joint effect of the aircraft, the sweepback lip cover and the fully shielded inner pipeline.

The invention provides an air inlet wave system configuration method for independent convergence of shock waves and isentropic compression waves. According to the method, shock waves of a compression surface of an air inlet are all converged at a lip position, and a virtual convergent point of isentropic compression waves behind the shock waves is arranged at the right upper part of the lip of the air inlet, so that the isentropic compression waves and other shock waves of the compression surface of the air inlet are independently converged, and the isentropic compression waves and lip cover shock waves generated by the lip cover of the air inlet form heterogeneous shock wave-isentropic compression wave interference. Compared with a conventional wave distribution mode that all shock waves and isentropic compression waves are converged at the lip position, the method has the advantages that the total pressure recovery coefficient of the throat of the air inlet can be greatly improved, the strength of the shock waves of the lip cover of the air inlet is weakened, the scale of a separation area formed by interference of the shock waves / boundary layers of the lip cover is reduced, and the low Mach number flow coefficient of the air inlet is improved. Especially for an air inlet with strong lip cover shock waves, the technical advantages of the invention are more obvious.