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.

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.

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.

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.

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.