The paper describes effects of the incidence on cascade three-dimensional flows and on the associated loss mechanisms occurring in a low-speed linear turbine rotor cascade. For each of five nominal incidences ranging from 7.2 to −53.3 deg, the cascade flow was surveyed at 15 or 16 planes located axially throughout the cascade. Blade-to-blade flows at the cascade midspan and near the endwalls, meridional flows along the blade surfaces, and static and total pressures within the cascade were analyzed in detail. The results were represented by two- or three-dimensional tomograms, on the surfaces of which the cascade flows were drawn by vectors, scalar contours, and streaklines, and from which one can easily understand the extraordinarily complicated cascade flows and the loss generation mechanisms. The present study will not only give new insight into the incidence effects, but will also contribute many solid experimental facts of a quantitative nature to our current knowledge of turbine cascade flows.
Hypersonic air-breathing engines will make the Earth-to-orbit vehicle completely different from the present one powered by rocket engines. The space plane propelled by a certain hypersonic air-breathing propulsion system is expected to appear in the next century. The turbomachinery-based engine (turboengine) is a candidate for the space plane propulsion system and will be combined with scramjet and rocket engines. Turboengines, including turboramjet, air-turboramjet, and their modifications, may be applied as the accelerators to the space plane having a high specific impulse at a rather low supersonic Mach number. Here, a conceptual study of these turboengines with preliminary system design, performance calculations, and consideration of relative merits of the engine concepts is performed for the configuration, performance, weight, and size. An engine evaluation with mission capability of the space plane for assumed requirements is made. As a result, engine performance depends on the liquid oxygen utilization, and weight and size of the engine are important factors for application to the space plane. Thus a certain optimization of the engine system itself and of a combination of the engines would be necessary.
In order to investigate several problems associated with the turbine cooling, an air-cooled two-stage axial flow turbine for an aircraft engine application was designed. Aerodynamic characteristics of the two-stage turbine without coolants were obtained first from the cold air turbine tests, and predictions of the turbine performance with supplying of coolants were made using the test results. Following these experiments, cooling tests of the first stage turbine were conducted in the range of turbine inlet gas temperatures lower than 1360 K by the another test apparatus. The descriptions of the turbine and the two test apparatus and the experimental results of the two test turbines are presented. The performance prediction, coolant effects and Reynolds number effect on the turbine performance are also described.
Two-dimensional turbine cascade tests have been carried out in order to make sure of aerodynamic and cooling performances and study the influence of the coolant ejection from the blade surface of the air-cooled turbine cascade in advance of their application to high temperature turbines. Several kinds of air-cooled nozzle and rotor blades have been tested in two cascade wind tunnels of different types. Chordwise distributions of the cooling effectiveness, mean cooling effectivenesses, iso-thermal contours on the blade surface by infrared thermal cameras, and the effects of coolant ejection on the aerodynamic performance are presented.
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