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

The invention disclosed herein provides a subsonic wind tunnel capable for accurately maintaining a desired Mach number, pressure and temperature for use in aerothermal testing of materials. The invention allows the Mach number to be precisely controlled in the test section of the wind tunnel by employing a restricted outlet acting as a sonic throat for the wind tunnel. In the preferred embodiment, the restricted outlet is constructed to be variable in cross sectional area allowing a range of Mach numbers to be tested. The variable outlet is varied during the operation of the wind tunnel so that an actual trajectory with changing Mach numbers, pressure, and temperature is simulated.

Airfoils are provided having non-monotonic meanline angle distributions and local negative camber along a length of the meanline between a point on a leading edge of the airfoil and a point on the trailing edge of the airfoil. The improved airfoil shape decreases the peak of a Mach number of a shock wave that develops in a passage between adjacent airfoils and attenuates or eliminates the shock wave at the suction surface of the airfoil within the passage. A blade constructed by this type of airfoil provides improved fluid flow in the passages between the airfoils, increased efficiency and an improved stall margin, among other benefits.

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.

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.

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.

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.

A hybrid air intake system and method for air-breathing missiles and aircraft comprising providing to an airborne vehicle fuselage an air intake opening comprising at least a pair of side walls and at least an upper wall and a lower wall extending between the side walls, wherein the lower wall extends below a plane defined by the vehicle fuselage, and not employing a boundary layer diverter between the fuselage and the upper wall.

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 discloses an inlet stage blade assembly for a gas compressor and an axial flow gas compressor comprising the same. The inlet stage blade assembly comprises an inlet guide blade and a first stage moving blade, wherein the position of the inlet guide blade with the largest trailing edge metal angle is a first position, and the relative blade height T at the first position is larger than or equal to 0.8 and smaller than or equal to 0.9; the blade-shaped section of the first-stage moving blade at the blade tip has a first chord length, the blade-shaped section of the first-stage moving blade at the second position with the relative blade height being T has a second chord length, and the first chord length is 3%-10% larger than the second chord length; and between the second position and the blade tip of the first-stage moving blade, the sweepforward or sweepback degree of the stacking shaft of the first-stage moving blade does not exceed 1% of the absolute blade height of the first-stage moving blade. By the adoption of the structure, the stall margin of the blade tip of the first-stage movable blade can be improved, and radial deformation of the front edge and the tail edge of the first-stage movable blade from the static state to the running state is more coordinated.

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 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 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 wide-speed-range air inlet channel design method based on double-incidence bending shock waves. The method comprises the following steps of designing a double-incidence bending shock wave reference flow field based on the bending shock wave theory; designing an air inlet channel initial throat section molded line, and performing reverse flow streamline tracking in the reference flow field to obtain an initial air inlet channel compression molded surface; designing an air inlet channel isolation section outlet section, expanding an air inlet channel initial throat section in a double-S-bend form, and stretching in an equal-straight manner to obtain a three-dimensional inward-turning wide-speed-range air inlet channel isolation section; carrying out wall surface fairing and viscosity correction; taking the tail end of a first-stage compression molded surface of an air inlet channel as the rotating shaft position of a first-stage splitter plate, and taking the length of a second-stage compression molded surface as the length of the first-stage splitter plate; taking the position where the second derivative of an upper wall curve of the air inlet isolation section is zero, taking the position as the position of a second-stage splitter plate rotating shaft, and taking the length from the second-stage splitter plate rotating shaft to the air inlet channel initial throat section as the length of the second-stage splitter plate; and designing a discharge port area under the Mach number and a throat area under the Mach number. Expansion waves at the position of the rotating shaft can be eliminated; and performance requirements are met.

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.

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.

The invention discloses a manufacturing method of a high-efficiency low-noise piston compressor valve plate. A compressor comprises an air cylinder head assembly, wherein a crescent-shaped air suctionport is formed in the valve plate and is a crescent-shaped through hole; and the valve plate is a powder metallurgical part. The technological process of the valve plate manufacturing method comprises the steps of powder mixing, compression molding through a die, high temperature sintering, shaping, inspection and steam treatment. By adopting the above technical scheme, pressure losses are reduced, and the compressor energy efficiency ratio is increased; the airflow velocity at an air inlet is decreased, a valve block is opened in a slow-down mode, and valve block displacement is reduced; meanwhile, the valve block lift is reduced, impact to a limiting plate and a valve seat is reduced, and noise of the valve block is reduced from the source; in addition, valve block displacement is reduced, stress to the waist of the valve block is reduced, thus, the critical stress zone area of the valve block is decreased, and the reliability of the valve block is improved; and the energy efficiency ratio is increased, airflow noise is reduced, and the reliability of the compressor is improved.

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.

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.