Design and two dimensional cascade test of a jet-flap turbine stator blade with ratio of axial chord to spacing of 0.5

A jet-flap blade was designed for a velocity diagram typical of the first-stage stator of a jet engine turbine and was tested in a simple two-dimensional cascade of six blades. The principal measurements were blade surface static pressure and cross-channel surveys of exit total pressure, static pressure, and flow angle. The results of the experimental investigation include blade loading, exit angle, flow, and loss data for a range of exit critical velocity ratios and three jet flow conditions

Incidence loss for a core turbine rotor blade in a two-dimensional cascade

The effect of incidence angle on the aerodynamic performance of an uncooled core turbine rotor blade was investigated experimentally in a two-dimensional cascade. The cascade test covered a range of incidence angles from minus 15 deg to 15 deg in 5-degree increments and a range of pressure ratios corresponding to ideal exit critical velocity ratios of 0.6 to 0.95. The principal measurements were blade-surface static pressures and cross-channel surveys of exit total pressure, static pressure, and flow angle. The results of the investigation include blade-surface velocity distribution and overall performance in terms of weight flow and loss for the range of incidence angles and exit velocity ratios investigated. The measured losses are also compared with two common methods of predicting incidence loss

Aerodynamic performance of a fully film cooled core turbine vane tested with cold air in a two-dimensional cascade

The aerodynamic performance of a fully film cooled core turbine vane was investigated experimentally in a two-dimensional cascade of 10 vanes. Three of the 10 vanes were cooled; the others were solid (uncooled) vanes. Cold air was used for both the primary and coolant flows. The cascade test covered a range of pressure ratios corresponding to ideal exit critical velocity ratios of 0.6 to 0.95 and a range of coolant flow rates to 7.5 percent of the primary flow. The coolant flow was varied by changing the coolant supply pressure. The principal measurements were cross-channel surveys of exit total pressure, static pressure, and flow angle. The results presented include exit survey data and overall performance in terms of loss, flow angle, and weight flow for the range of exit velocity ratios and coolant flows investigated. The performance of the cooled vane is compared with the performance of an uncooled vane of the same profile and also with the performance obtained with a single cooled vane in the 10-vane cascade

Performance of a high-work, low-aspect-ratio turbine stator tested with a realistic inlet radial temperature gradient

A 0.767-scale model of a turbine stator designed for the core of a high-bypass-ratio aircraft engine was tested with uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The principal measurements were radial and circumferential surveys of stator-exit total temperature, total pressure, and flow angle. The stator-exit flow field was also computed by using a three-dimensional Navier-Stokes solver. Other than temperature, there were no apparent differences in performance due to the inlet conditions. The computed results compared quite well with the experimental results

Cold-air performance evaluation of scale model oxidizer pump-drive turbine for the M-1 hydrogen-oxygen rocket engine. 3 - Performance of first stage with inlet-feedpipe-manifold assembly

Cold air performance test of scale model oxidizer pump drive turbine for M-1 engine - performance of first stage with inlet feedpipe manifold assembl

Description of the warm core turbine facility recently installed at NASA Lewis Research Center

The two net facilities were installed and operated at their design, or rated conditions. The important feature of both of these facilities is that the ratio of turbine inlet temperature to coolant temperature encountered in high temperature engines can be duplicated at moderate turbine inlet temperature. The limits of the facilities with regard to maximum temperature, maximum pressure, maximum mass flow rate, turbine size, and dynamometer torque-speed characteristics are discussed

Performance of a high-work low aspect ration turbine tested with a realistic inlet radial temperature profile

Experimental results are presented for a 0.767 scale model of the first stage of a two-stage turbine designed for a high by-pass ratio engine. The turbine was tested with both uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The inlet temperature profile was essentially mixed-out in the rotor. There was also substantial underturning of the exit flow at the mean diameter. Both of these effects were attributed to strong secondary flows in the rotor blading. There were no significant differences in the stage performance with either inlet condition when differences in tip clearance were considered. Performance was very close to design intent in both cases

Effect of cooling-hole geometry on aerodynamic performance of a film-cooled turbine vane tested with cold air in a two-dimensional cascade

The effect of the orientation and cooling-hole size on turbine-vane aerodynamic losses was evaluated. The contribution of individual vane regions to the overall effect was also investigated. Test configurations were based upon a representative configuration having 45 spanwise rows of holes spaced about the entire vane profile. Nominal hole diameters of 0.0254 and 0.0356 cm and nominal hole orientations of 35 deg, 45 deg, and 55 deg from the local vane surface and 0 deg, 45 deg, and 90 deg from the main-stream flow direction were investigated. Flow conditions and aerodynamic losses were determined by vane-exit surveys of total pressure, static pressure, and flow angle

Incidence loss for fan turbine rotor blade in two-dimensional cascade

The effect of incidence angle on the aerodynamic performance of a fan turbine rotor blade was investigated experimentally in a two dimensional cascade. The test covered a range of incidence angles from -15 deg to 10 deg and exit ideal critical velocity ratios from 0.75 to 0.95. The principal measurements were blade-surface static pressures and cross-channel survey of exit total pressure, static pressure, and flow angle. Flow adjacent to surfaces was examined using a visualization technique. The results of the investigation include blade-surface velocity distribution and overall kinetic energy loss coefficients for the incidence angles and exit velocity ratios tested. The measured losses are compared with those from a reference core turbine rotor blade and also with two common analytical methods of predicting incidence loss