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Noncavitating operation data from different axial flow pump rotor configurations and correlation of flow pump deviation angles
The experimental performance of a 20-inch-diameter axial-flow transonic compressor rotor with small dampers is presented. The compressor rotor was tested earlier with large dampers which were twice in size, and comparisons of overall performance and radial distributions of selected flow and performance parameters are made. The rotor with small dampers experienced lower losses in the damper region which resulted in locally higher values of temperature rise efficiency and total pressure ratio. However, there was no appreciable effect on overall efficiency and pressure ratio. A greater stall margin was measured for the rotor with small dampers at design speed, but at 70 and 90 percent of design speed the rotor with large dampers had somewhat greater flow range.
A multi-task research program on aerodynamic problems in advanced axial-flow turbomachine configurations was carried out at Iowa State University. The elements of this program were intended to contribute directly to the improvement of compressor, fan, and turbine design methods. Experimental efforts in intra-passage flow pattern measurements, unsteady blade row interaction, and control of secondary flow are included, along with computational work on inviscid-viscous interaction blade passage flow techniques. This final report summarizes the results of this program and indicates directions which might be taken in following up these results in future work. In a separate task a study was made of existing turbomachinery research programs and facilities in universities located in the United States. Some potentially significant research topics are discussed which might be successfully attacked in the university atmosphere.
The compressor case of a 13-stage axial-flow compressor was modified to incorporate air-bleed systems over the fifth and tenth rotor stages. Rotating stall-exited rotor-blade vibrations were measured in the first and second stages at 60 and 68 percent of rated speed, respectively. Either the fifth- or the tenth-stage bleeds satisfactorily eliminated the peak vibratory stresses.
Several modifications have been made to the axial-flow turbine preliminary sizing code TURBAN. Turbine cooling has been added to the analysis. New alternative input options allow direct specification of stage reaction, stage work factor, and stage work split. The Reynolds number loss dependency was modified and an internal calculation of air viscosity was added. A complete description of input and output along with sample cases are included.
Efficiency, mass flow rate, torque, and exit diffuser effectiveness of 5-inch axial flow turbine
Arc model describing symptomatic behavior of fluctuating mode when axial gas flow is superimposed on electric arc
Containment tests of vortexes with radial outflow in basic vortex tube, and in axial-flow vortex tube
Computer program determines the alternative geometries and associated design point performance of axial-flow turbines capable of satisfying specified design requirements. The program solves the flow field within the turbine without making the simplifying assumptions that result in restrictive designs.
Effect of rotor tip clearance on performance of 5-inch single-stage axial flow turbine
Performance and cavitation damage to axial flow pump rotor blades in high temperature liquid sodium
Cavitation damage and effect of fluid temperature on performance of axial flow pumps in liquid sodium
A 22.9-centimeter diameter axial flow rotor with a 0.8 hub-tip radius ratio, a design flow coefficient of 0.466, and a blade tip design diffusion factor of 0.55 was tested in cold water under both cavitating and noncavitating conditions. Radial surveys of the flow conditions at the rotor inlet and outlet were made. At design flow, the rotor produced an overall headrise coefficient of 0.360 with an overall efficiency of 95.0 percent. The efficiency remained greater than 88 percent over the entire flow coefficient range which varied from 0.350 to 0.615.
Rotor blade aerodynamic damping is experimentally determined in a three-stage transonic axial flow compressor having design aerodynamic performance goals of 4.5:1 pressure ratio and 65.5 lbm/sec weight flow. The combined damping associated with each mode is determined by a least squares fit of a single degree of freedom system transfer function to the nonsynchronous portion of the rotor blade strain gage output power spectra. The combined damping consists of the aerodynamic damping and the structural and mechanical damping. The aerodynamic damping varies linearly with the inlet total pressure for a given corrected speed, weight flow, and pressure ratio while the structural and mechanical damping is assumed to remain constant. The combined damping is determined at three inlet total pressure levels to obtain the aerodynamic damping. The third-stage rotor blade aerodynamic damping is presented and discussed for the design equivalent speed with the stator blades reset for maximum efficiency. The compressor overall preformance and experimental Campbell diagrams for the third-stage rotor blade row are also presented.
Rotor blade aerodynamic damping is experimentally determined in a three-stage transonic axial flow compressor having design aerodynamic performance goals of 4.5:1 pressure ratio and 65.5 lbm/sec weight flow. The combined damping associated with each mode is determined by a least squares fit of a single degree of freedom system transfer function to the nonsynchronous portion of the rotor blade strain gauge output power spectra. The combined damping consists of aerodynamic and structural and mechanical damping. The aerodynamic damping varies linearly with the inlet total pressure for a given equivalent speed, equivalent mass flow, and pressure ratio while structural and mechanical damping are assumed to be constant. The combined damping is determined at three inlet total pressure levels to obtain the aerodynamic damping. The third stage rotor blade aerodynamic damping is presented and discussed for 70, 80, 90, and 100 percent design equivalent speed. The compressor overall performance and experimental Campbell diagrams for the third stage rotor blade row are also presented.
Rotor blade aerodynamic damping is experimentally determined in a three-stage transonic axial flow compressor having design aerodynamic performance goals of 4.5:1 pressure ratio and 65.5 lbm/sec weight flow. The combined damping associated with each mode is determined by a least squares fit of a single degree of freedom system transfer function to the nonsynchronous portion of the rotor blade strain gage output power spectra. The combined damping consists of the aerodynanmic damping and the structural and mechanical damping. The aerodynamic damping varies linearly with the inlet total pressure for a given corrected speed, weight flow, and pressure ratio while the structural and mechanical damping is assumed to remain constant. The combined damping is determined at three inlet total pressure levels to obtain the aerodynamic damping. The third-stage rotor blade aerodynamic damping is presented and discussed for the design equivalent speed with the stator blades reset for maximum efficiency. The compressor overall performance and experimental Campbell diagrams for the third-stage rotor blade row are also presented.
A preliminary investigation of an axial-flow gas turbine-propeller engine was conduxted. Performance data were obtained for engine speeds from 8000 to 13,000 rpm and altitudes from 5000 to 35,000 feet and compressor inlet ram pressure ratios from 1.00 to 1.17.
A set of modified NACA 5509-34 rotor and stator blades was investigated with rough-machine, hand-filed, and highly polished surface finishes over a range of weight flows at six equivalent tip speeds from 672 to 1092 feet per second to determine the effect of blade-surface finish on the performance of a single-stage axial-flow compressor. Surface-finish effects decreased with increasing compressor speed and with decreasing flow at a given speed. In general, finishing blade surfaces below the roughness that may be considered aerodynamically smooth on the basis of an admissible-roughness formula will have no effect on compressor performance.