30results about How to "Reduce the Mach number" patented technology

At least part of the inner circumferential wall of the outer casing is provided with a concave surface opposing the rotor blade tips as seen in a longitudinal section. Typically, each of the rotor blades is provided with aerofoil section, and the compressor is designed as a transonic axial flow compressor. Thereby, a compressive wave is produced upstream of the shockwave so that the Mach number of the flow entering the shockwave can be reduced. As a result, the shockwave is made less severe, and the shockwave loss can be reduced. In particular, because the concave surface is provided in the casing wall as opposed to the case where the concave surface is provided in the negative pressure side of the rotor blade, the reduction in the performance owing to the change in the angle of the airflow entering the passage defined by the concave surface under a partial load condition can be avoided.

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.

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.

It is an object of the present invention to solve the problem of a drop in precision in conventional systems using a square pyramid type five-hole probe due to the drop in atmospheric pressure in high altitude ranges, and to provide a wide velocity range flight velocity vector measurement system that can prevent a drop in measurement precision. Furthermore, it is also an object of the present invention to provide a method for eliminating the effects of detection fluctuations caused by adhering water droplets, ice particles or dust in a wide velocity range flight velocity vector measurement system. The flight velocity vector measurement probe of the present invention comprises means in which a static pressure hole is formed in the tube wall surface of the probe, so that a static pressure value is obtained from the pressure detected by this static pressure hole, the Mach number M is calculated on the basis of an equation approximated by a fourth-order polynomial of the static pressure / total pressure signal and the angle of attack, and in cases where an abnormal detection value is detected, this is replaced by the preceding detection value.

The present disclosure relates to a propfan engine comprising: one or more rotor stages comprising a plurality of rotors; and an outer wall comprising an outer profile, at least a portion of the outer profile defining a substantially circular cross-section, wherein the diameter of the substantially circular cross-section increases in the direction of flow over the outer wall and downstream of a leading edge of the rotors, and the diameter increases at substantially all points defining the circumference of the substantially circular cross-section.

There is provided an axial-flow turbine comprising an exhaust chamber; a turbine including multiple stage rotor blades, said multiple stage rotor blade including terminal stage rotor blades; an annular diffuser located between the turbine and the exhaust chamber; and an annular axial-flow turbine passage defined by the turbine, the diffuser and the exhaust chamber, wherein fluid flows through the axial-flow turbine passage toward the exhaust chamber, and an annular stepped portion which inwardly projects in a radial direction is formed on the portion of an inner wall of the axial-flow turbine passage that is located on the downstream side of a trailing edge of a tip portion of the terminal stage rotor blades provided in the flow direction of the fluid. In the stepped portion, a projecting portion which inwardly projects in a radial direction may be provided.

An impeller mountable within a centrifugal compressor includes a hub having a front side and a back side, the hub being rotatable about an axis of rotation and a plurality of vanes extending outwardly from the front side of the hub such that a plurality of passages is defined between adjacent vanes. The plurality of vanes is oriented such that a flow output from the plurality of passages adjacent the back side of the impeller is arranged at an angle to the axis of rotation of less than 20 degrees.

The invention relates to a front edge design method of a transonic speed fan blade top primitive blade profile. The front edge design method comprises the steps that a reference point distant from anoriginal front edge by a certain length is obtained on a suction surface of a blade top original primitive blade profile; a linear section is made by passing the reference point, wherein a certain included angle is formed between the original suction surface and the linear section; and the linear section is connected with an original pressure surface by an arc section with a small radius, the other end of the linear section is in smooth transition with the original suction surface by a chamfer, and the axial positions of a new blade front edge and an old blade front edge are kept unchanged. Since the front edge of the blade top original primitive blade profile is subjected to sharpening treatment and the effective radius of the front edge is reduced, shock wave detachment may be delayed, and the surge-to-stall margin of a fan can be improved advantageously. In addition, with the introduction of a sharpening structure, an expansion wave at the blade front edge is split into two parts, wherein one part is an expansion wave triggered by the front edge radius and the other part is an expansion wave triggered by the transition part of the linear section and the suction surface, so thatthe mach number before a channel shock wave can be lowered advantageously, and the efficiency of the blade top original primitive blade profile is improved.

The invention discloses a pneumatic and structural feature considered three-dimensional geometric structure of a large fan connotative blade of a civil aircraft engine. The rear section of a blade profile at the root part of the large fan which is designed traditionally has a larger bent corner, a smaller acute angle is formed by a trailing edge suction surface and a hub, low-energy gas is accumulated in the acute-angle area, the flow loss is great, and the smaller acute angle between the suction surface and the hub is not beneficial to the structure modeling of the root stretching section of the blade. According to the invention, the change of an outlet metal corner of an element blade profile of the large fan connotative blade (which is the part within 20% of a blade height of the large fan blade) along the blade height is mainly optimized and a distribution curve of the outlet metal corner of a connotative blade and a projection curve of a trailing edge line in a plane vertical to the axial direction are fit. Through the verification of three-dimensional CFD (Computational Fluid Dynamics), the pneumatic and structural feature considered three-dimensional geometric structure disclosed by the invention has the advantages that the flow separation in the acute-angle area formed by the trailing edge suction surface of the fan and the hub can be effectively reduced, the fan connotative efficiency can be increased, a good entrance condition is provided for a pressure boosting grade, and the blade at the root part and the hub are relatively almost vertical, so that convenience is provided for the installation of the blade on a wheel disc.

A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

The invention provides a bump air inlet method for realizing integration of an unequal-strength wave system with a forebody, relating to the technical field of ultrasonic air inlet. The method comprises the following steps that: 1, an air inlet wave system adopts an external-compression double wave system structure based on an unequal-strength system, and the total pressure recovery coefficient sigma=sigma1*sigma2, wherein the sigma1 and the sigma2 are total pressure recovery coefficients of a conical shock wave and a normal shock wave; 2, a cone with a semi-cone angle delta c formed by the conical shock wave generates a second conical shock wave with second semi-cone angle beta in ultrasonic flow; 3, the bump height of the air inlet is h, the h / delta is 2-2.5 when the local boundary layer thickness is delta, the bump compressive plane generated in the step 2 is scaled to meet the requirement of an actual size; and 4, the lip of the air inlet adopts a conformal and sweepback lip design. The method realizes reduction of throat Mach number of the air inlet, improves the air inlet performance to ensure that the lip sweepback angle of the air inlet, the maximum turning angle of the bump compressive plane and the inlet normal shock wave angle are consistent, can increase the curve of the total pressure recovery coefficient, and reduces resistance coefficient.

A compressor rotor blade for an axial-type compressor has a blade profile having a transonic section and a profile section which extends in the transonic section and has concave and convex suction-side regions on the suction side, the convex suction-side region arranged downstream of the concave suction-side region, and has convex and concave pressure-side regions on the pressure side, the concave pressure-side region arranged downstream of the convex pressure-side region. Curvature progressions on the pressure side and on the suction side are both applied in a continuous manner over a profile chord of the profile section, the positions of the minimum values of the curvature progressions deviate from each other by not more than 10% of the length of the profile chord, and the positions of the maximum values of the curvature progressions deviate from each other by not more than 10% of the length of the profile chord.

The invention discloses a transient jet flow test device used in a pulse wind tunnel. The transient jet flow test device includes a jet flow test model, an air source supply system, a fast synchronous control system and a jet flow parameter measurement device; the jet flow test model ejects a stable jet flow; the air source supply system provides the required Stable air source, fast synchronous control system to ensure that the formation time of the stable jet flow is synchronized with the formation time of the main flow of the pulse wind tunnel; at the same time, the air source supply system ensures that the duration of the stable jet flow exceeds the effective test time of the main flow of the pulse wind tunnel, and the fast synchronous control system is in The air source supply system is automatically turned off after the set stable jet flow time is reached; the jet flow parameter measuring device is used to measure the flow field parameters of the stable jet flow from the model before the jet flow test. The transient jet flow test device can provide jet flow with accurate parameters, good stability and repeatability to evaluate the influence of jet flow disturbance on the aerodynamic characteristics of the aircraft, and can quickly complete the jet flow preparation and accurately control the injection time.

The invention discloses a method for manufacturing a valve plate of a high-efficiency and low-noise piston compressor. The compressor includes a cylinder head assembly, wherein a crescent-shaped suction port is arranged on the valve plate, and the crescent-shaped suction port is a crescent-shaped suction port. through hole. The valve plate is a powder metallurgical part; the manufacturing process of the valve plate is: mixing powder; molding with a mold; high temperature sintering; shaping; inspection; steam treatment. With the above technical solution, the pressure loss is reduced, and the energy efficiency ratio of the compressor is improved; the airflow velocity at the air inlet becomes smaller, the opening of the valve plate becomes slower, and the displacement of the valve plate becomes smaller; at the same time, the lift of the valve plate is reduced, reducing the The impact of the plate on the limit plate and the valve seat reduces the noise of the valve plate from the source; in addition, the displacement of the valve plate is reduced, the stress on the waist of the valve plate is reduced, the area of ​​the dangerous stress area of ​​the valve plate is reduced, and the valve plate is reliable. Enhanced performance; improved energy efficiency ratio, reduced airflow noise and improved compressor reliability.

A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

The invention discloses an inlet-stage blade assembly for a compressor and an axial-flow compressor containing the same. The inlet-stage blade assembly includes an inlet guide vane and a first-stage moving blade, and the inlet guide vane has the largest trailing edge metal angle. The position is the first position, and the relative blade height T at the first position satisfies 0.8≤T≤0.9; The airfoil section at the second position of T has a second chord length, and the first chord length is 3%~10% longer than the second chord length; between the second position and the tip of the first-stage moving blade, the first chord length The forward or backward sweep of the stacking axis of the first-stage rotor blade shall not exceed 1% of the absolute blade height of the first-stage rotor blade. Adopting the above structure can improve the stall margin of the blade tip of the first-stage moving blade, and make the radial deformation of the leading edge and the trailing edge of the first-stage moving blade more coordinated when the first-stage moving blade changes from a static state to a running state.

A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

A valve for controlling a fluid, in particular for controlling a gas, is described, encompassing a valve housing an actuation unit for an at least locally tubular valve armature which is guided axially displaceably and is equipped with a valve closure member by which a fluid flow between an inflow side and an outlet side is controllable and which coacts with a valve seat. The valve armature comprises a guidance collar in a region remote from the valve closure member, and is equipped with a second guidance in a region offset with respect to the guidance collar.