154 results about "Transonic" patented technology

The invention discloses a three-hole transonic speed pressure probe. As a pneumatic technology develops, a transonic speed air flow appears more and more frequently, for instance, many air flows in an aeroengine reach a transonic speed state, so these transonic speed flow fields need to be measured and analyzed during development of engine components. Conventionally, the transonic speed flow fields are measured respectively by use of a total pressure probe and a direction probe so that total pressure and air flow directions can be obtained. When a flow speed reaches a transonic speed, the probes are under great resistance, and the probes are blown bent easily. When the flow speed reaches an ultrasonic speed, due to the existence of shock waves, there is large total pressure loss, and the influences exerted by the probes on the flow fields are quite large. Though these conventional probes can be used in an ultrasonic air flow, the influences exerted by the probes on the flow fields cannot be neglected, and the requirement for the strength of the probes is quite high. The novel three-hole transonic speed pressure probe provided by the invention is used for measuring the transonic speed flow fields. By use of the probe provided by the invention, the intensity of the shock waves in front of the probe can be weakened during measurement, the influences exerted by the probe on the flow fields are reduced, air-flow total pressure, static pressure, a Mach number and an air-flow deflection angle can be measured simultaneously, the resistance acting on the probe can be effectively reduced, and the firmness and reliability of the probe are guaranteed.

A transonic blade is provided that operates in a flow field where flow has a transonic speed or higher in an axial-flow rotating machine and that concurrently achieves a reduction in shock loss and in the local stress of the blade.The transonic blade includes a hub cross-sectional surface joined to a rotating shaft or an outer circumferential side casing of a rotating machine; a tip cross-sectional surface located furthest from the hub cross-sectional surface in a spanwise direction which is a vertical direction of the rotating shaft; a leading edge located on an upstream side; and a trailing edge located on a downstream side. At least a part of a passing working fluid flow has a transonic speed or higher. A portion of a stacking line which is a line connecting together respective gravity centers of cross-sectional surfaces located from the hub cross-sectional surface to the tip cross-sectional surface is located on a downstream side of a stacking center in a flow direction of a working fluid main flow.

A turbine blade system including a blade airfoil having an airfoil shape with the blade airfoil having a nominal profile substantially in accordance with normalized Cartesian coordinate values Z set forth in a table 1 below and which values are dimensionless values that are convertible to corresponding absolute distance values that define nominal airfoil profile sections and which, when joined smoothly with adjacent ones thereof, form a complete nominal airfoil shape that is substantially matched by the airfoil shape of the blade airfoil. This blade airfoil can be supported on a ring platform having a support surface with a support surface shape in the vicinity of the location at which that blade airfoil is supported thereon formed in a similar manner.

A turbine blade system including a blade airfoil having an airfoil shape with the blade airfoil having a nominal profile substantially in accordance with normalized Cartesian coordinate values Z set forth in a table 1 below and which values are dimensionless values that are convertible to corresponding absolute distance values that define nominal airfoil profile sections and which, when joined smoothly with adjacent ones thereof, form a complete nominal airfoil shape that is substantially matched by the airfoil shape of the blade airfoil. This blade airfoil can be supported on a ring platform having a support surface with a support surface shape in the vicinity of the location at which that blade airfoil is supported thereon formed in a similar manner.

An apparatus and method is provided for improving efficiency of a transonic gas turbine engine compressor by bleeding off a shockwave-induced boundary layer from the gas flow passage of the compressor using an array of bleed holes having a downstream edge aligned with a foot of an oblique shock wave which originates on the leading edge of an adjacent transonic rotor blade tip.

The Longitudinal Flying Wing aircraft idea provides for design of large cargo and passenger aircraft in range from low to high subsonic and transonic speed. Such aircraft would have up to twice lower fuel consumption per unit of payload, higher lift capacity, and a significantly longer range, while having a significantly lower level of noise inside passenger cabin and cockpit relative to classical concept aircraft. This idea is further providing for efficient, reliable, and simple flight controls, hence it may be successfully applied for design of all-size, long range, high-lift-capacity unmanned aircraft throughout the entire range of subsonic speeds.

The present invention relates to the aerodynamic design of moving blades, pertaining to later stages of axial steam turbines where the inlet flow is non-uniform over the blade height. The claim made herein is a set of six invented transonic blade profiles which can be used to develop various type of 3D twisted blades for axial steam turbine. The aerodynamic characteristics of these 6 base profiles are evaluated herein as a function of stagger angle and pitch / chord ratios.

The invention discloses a supersonic / hypersonic aerocraft engine overexpansion nozzle bypass type device. The device comprises a nozzle contracting segment, a nozzle expanding segment, an air-bleeding bypass and a control device for controlling the air-bleeding bypass to be opened and closed, the air-bleeding bypass is an internal flowing passage communicating the nozzle contracting segment and the nozzle expanding segment, and an internal flow field is established by the aid of natural pressure difference between an air inlet of the nozzle contracting segment and an air outlet of the nozzle expanding segment. When the aerocraft works in the transonic and over-expanding state, the bypass is opened, high-pressure air of the nozzle contracting segment is bled to the nozzle expanding segment to form secondary flow injecting effect, and thrust coefficient of the nozzle is increased; and when the aerocraft in the normal working state, the air-bleeding bypass passage is closed, and nozzle design point performances cannot be affected. The supersonic / hypersonic aerocraft engine overexpansion nozzle bypass type device is simple in structure, needs no extra air bleeding sources and is free of additional weight, and thrust performances of the nozzle are improved.

A transonic blade is provided that operates in a flow field where flow has a transonic speed or higher in an axial-flow rotating machine and that concurrently achieves a reduction in shock loss and in the local stress of the blade. The transonic blade includes a hub cross-sectional surface joined to a rotating shaft or an outer circumferential side casing of a rotating machine; a tip cross-sectional surface located furthest from the hub cross-sectional surface in a spanwise direction which is a vertical direction of the rotating shaft; a leading edge located on an upstream side; and a trailing edge located on a downstream side. At least a part of a passing working fluid flow has a transonic speed or higher. A portion of a stacking line which is a line connecting together respective gravity centers of cross-sectional surfaces located from the hub cross-sectional surface to the tip cross-sectional surface is located on a downstream side of a stacking center in a flow direction of a working fluid main flow.

Sectional profiles close to a tip 124 and a part between a midportion 125 and a hub 123 are shifted to the upstream of an operating fluid flow in a sweep direction. Accordingly, an S shape is formed in which the tip 124 and the part between the midportion 125 and the hub 123 protrude. As a result, it is possible reduce various losses due to shook, waves, thereby forming a transonic airfoil having an excellent aerodynamic characteristic.

An air vehicle having a fuselage and two wings extending laterally therefrom and having a first surface between leading and trailing edges and an opposite second surface. The vehicle includes adjacent upstream and downstream orifices positioned on at least one first surface. Each upstream orifice is closer to the leading edge than the downstream orifice. Each second surface is substantially free of orifices. The vehicle includes an actuator within each wing having orifices. Each actuator is connected to corresponding upstream and downstream orifices for creating a negative pressure differential at the upstream orifice and a positive pressure differential at the downstream orifice so air is drawn into the upstream orifice and air is pushed away form the downstream orifice. The orifice is configured so air is drawn into and directed out of the upstream and downstream orifices, respectively, at an angle of about 90% with respect to the first surface.

A process of designing a transonic airfoil that results in an exit deviation angle that is optimized to minimize downstream shock induced flowfield disturbances. Specifically, a design process of a transonic airfoil minimizes pitchwise variation in downstream static pressure that includes determining an airfoil exit deviation angle and downstream turning for reduced shock from a trailing edge of the transonic airfoil.