Effects of increased leading-edge thickness on performance of a transonic rotor blade

A single-stage transonic compressor was tested with two rotor blade leading-edge configurations to investigate the effect of increased leading-edge thickness on the performance of a transonic blade row. The original rotor blade configuration was modified by cutting back the leading edge sufficiently to double the blade leading-edge thickness and thus the blade gap blockage in the tip region. At design speed this modification resulted in a decrease in rotor overall peak efficiency of four points. The major portion of this decrement in rotor overall peak efficienty was attributed to the flow conditions in the outer 30 percent of the blade span. At 70 and 90 percent of design speed, the modification had very little effect on rotor overall performance

Effect of casing treatment on performance of a two-stage high-pressure-ratio fan

A two-stage fan, previously tested with a solid casing, was tested with a casing with circumferential grooves over the tips of both rotors (casing treatment). Tests were conducted at 80 and 100 percent of design speed with uniform flow. The casing treatment improved the flow range and stall margin significantly without changing the characteristics overall performance curves of total-pressure and efficiency as functions of weight flow, other than extending them to lower weight flows

Performance of transonic fan stage with weight flow per unit annulus area of 208 kilograms per second per square meter (42.6 (lb/sec)/sq ft)

Performance was obtained for a 50-cm-diameter compressor designed for a high weight flow per unit annulus area of 208 (kg/sec)/sq m. Peak efficiency values of 0.83 and 0.79 were obtained for the rotor and stage, respectively. The stall margin for the stage was 23 percent, based on equivalent weight flow and total-pressure ratio at peak efficiency and stall

Space Chemical Propulsion Test Facilities at NASA Lewis Research Center

The NASA Lewis Research Center, located in Cleveland, Ohio, has a number of space chemical propulsion test facilities which constitute a significant national space testing resource. The purpose of this paper is to make more users aware of these test facilities and to encourage their use through cooperative agreements between the government, industry, and universities. Research which is of interest to the government is especially encouraged and often can be done in a cooperative manner that best uses the resources of all parties. An overview of the Lewis test facilities is presented

Performance of transonic fan stage with weight flow per unit annulus area of 178 kilograms per second per square meter (6.5(lb/sec)/(sq ft))

The overall and blade-element performances are presented over the stable flow operating range from 50 to 100 percent of design speed. Stage peak efficiency of 0.834 was obtained at a weight flow of 26.4 kg/sec (58.3 lb/sec) and a pressure ratio of 1.581. The stall margin for the stage was 7.5 percent based on weight flow and pressure ratio at stall and peak efficiency conditions. The rotor minimum losses were approximately equal to design except in the blade vibration damper region. Stator minimum losses were less than design except in the tip and damper regions

Performance of a single-stage transonic compressor with a blade-tip solidity of 1.5 and comparison with 1.3 and 1.7 solidity stages

The overall and blade-element performance of a transonic compressor stage with a tip solidity of 1.5 is presented over the stable operating range at rotative speeds from 50 to 100 percent of design speed. State peak efficiency of 0.82 was obtained at a weight flow of 29.4 kg.sec (200.4 (kg/sec)/m2 of annulus area) and a pressure ratio of 1.71. Stall margin at design speed was 14 percent. A comparison of three stages in a solidity study showed that the performance of the 1.5 solidity stage and the 1.3 solidity stage were nearly identical but that the performance of the 1.7 solidity stage was significantly lower

Performance of two-stage fan with larger dampers on first-stage rotor

The performance of a two stage, high pressure-ratio fan, having large, part-span vibration dampers on the first stage rotor is presented and compared with an identical aerodynamically designed fan having smaller dampers. Comparisons of the data for the two damper configurations show that with increased damper size: (1) very high losses in the damper region reduced overall efficiency of first stage rotor by approximately 3 points, (2) the overall performance of each blade row, downstream of the damper was not significantly altered, although appreciable differences in the radial distributions of various performance parameters were noted, and (3) the lower performance of the first stage rotor decreased the overall fan efficiency more than 1 percentage point

Design and performance of a 427-meter-per-second-tip-speed two-stage fan having a 2.40 pressure ratio

The aerodynamic design and the overall and blade-element performances are presented of a 427-meter-per-second-tip-speed two-stage fan designed with axially spaced blade rows to reduce noise transmitted upstream of the fan. At design speed the highest recorded adiabatic efficiency was 0.796 at a pressure of 2.30. Peak efficiency was not established at design speed because of a damper failure which terminated testing prematurely. The overall efficiencies, at 60 and 80 percent of design speed, peaked at approximately 0.83

Effect of casing treatment on performance of an inlet stage for a transonic multistage compressor

An inlet stage of a transonic compressor was tested with three rotor tip casing treatment configurations: blade angle slots, circumferential grooves, and axial skewed slots. Significant increases in both rotor and stage total pressure ratio, total temperature ratio, efficiency, flow range, and very large improvements in stall margin were obtained with all three casing treatment configurations. The greatest improvement in performance was achieved with axial skewed slots

Stalled and stall-free performance of axial-flow compressor stage with three inlet-guide-vane and stator-blade settings

The performance of the first stage of a transonic, multistage compressor was mapped over a range of inlet-guide-vane and stator-blade settings. Both stall-free and deep-stall performance data were obtained. For the settings tested, as stall was encountered and flow was further reduced, a relatively sharp drop in pressure ratio occurred and was followed by a continuing but more gradual reduction in pressure ratio with reduced flow. The position of the stall line on the map of pressure ratio against equivalent weight flow was essentially unaffected over the range of inlet-guide-vane and stator-blade settings