A l a s e r anomometer system employing an e f f i c i e n t data acquisition technique has been used to make measurements upstream, within, and downstream of the c o w pressor rotor. A fluorescent dye technique allowed measurements within endwali boundary layers. Adjustable l a s e r beau orientation minimized shadowed regions and enabled r a d i a l velocity measurements outside of the blade row. The flow phenomena investigated include flow variations from passage t o passage, the rotor shocksystem, three-dimensional flows i n t h e blade wake, and the development of the outer endwall beundaty layer. Laser anemometer measurements a r e compared t o a numerical solution of the streamfunction equations and t o measurements made with conventional instmuentation.

Rotating stall and surge have been stabilized in a transonic single-stage axial compressor using active feedback control. The control strategy is to sense upstream wall static pressure patterns and feed back the signal to an annular array of twelve separately modulated air injectors. At tip relative Mach numbers of 1.0 and 1.5 the control achieved a 11% and 3.5% reduction in stalling mass flow respectively, with injection adding 3.6% of the design compressor mass flow. The aerodynamic effects of the injection have also been examined. At a tip Mach number, Mtip, of 1.0, the stall inception dynamics and effective active control strategies are similar to results for low-speed axial compressors. The range extension was achieved by individually damping the first and second spatial harmonics of the pre-stall perturbations using constant gain feedback. At a Mtip of 1.5 (design rotor speed), the pre-stall dynamics are different than at the lower speed. Both one-dimensional (surge) and two-dimensional (rotating stall) perturbations needed to be stabilized to increase the compressor operating range. At design speed, the instability was initiated by approximately 10 rotor revolutions of rotating stall followed by classic surge cycles. In accord with the results from a compressible stall inception analysis, the zeroth, first, and second spatial harmonics each include more than one lightly damped mode which can grow into the large amplitude instability. Forced response testing identified several modes traveling up to 150% of rotor speed for the first three spatial harmonics; simple constant gain control cannot damp all of these modes and thus cannot stabilize the compressor at this speed. A dynamic, model-based robust controller was therefore used to stabilize the multiple modes which comprise the first three harmonic perturbations in this transonic region of operation.

Mass injection upstream of the tip of a high-speed axial compressor rotor is a stability enhancement approach known to be effective in suppressing stall in tip-critical rotors. This process is examined in a transonic axial compressor rotor through experiments and time-averaged Navier-Stokes CFD simulations. Measurements and simulations for discrete injection are presented for a range of injection rates and distributions of injectors around the annulus. The simulations indicate that tip injection increases stability by unloading the rotor tip and that increasing injection velocity improves the effectiveness of tip injection. For the tested rotor, experimental results demonstrate that at 70 percent speed the stalling flow coefficient can be reduced by 30 percent using an injected massflow equivalent to 1 percent of the annulus flow. At design speed, the stalling flow coefficient was reduced by 6 percent using an injected massflow equivalent to 2 percent of the annulus flow. The experiments show that stability enhancement is related to the mass-averaged axial velocity at the tip. For a given injected massflow, the mass-averaged axial velocity at the tip is increased by injecting flow over discrete portions of the circumference as opposed to full-annular injection. The implications of these results on the design of recirculating casing treatments and other methods to enhance stability will be discussed.

The performance deterioration of a high-speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54–3.18 rms μm (100–125 rms μin.) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10 percent at the hub and 20 percent at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9 percent reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effects of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254–0.508 rms μm (10–20 rms μin.), compared to the bare metal blade surface finish of 0.508 rms pm (20 rms μin.). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60, 80, and 100 percent of design speed. The results indicate that thickness/roughness over the first 2 percent of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70 percent of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-three-dimensional Navier–Stokes flow solver, which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that adding roughness at the blade leading edge causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage, which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

A laser-doppler anemometer was used to obtain flow-field velocity measurements in a 4:1 pressure ratio, 4.54 kg/s (10 lbm/s), centrifugal impeller, with splitter blades and backsweep, which was configured with a vaneless diffuser. Measured through-flow velocities are reported for ten quasi-orthogonal survey planes at locations ranging from 1% to 99% of main blade chord. Measured through-flow velocities are compared to those predicted by a 3-D viscous steady flow analysis (Dawes) code.The measurements show the development and progression through the impeller and vaneless diffuser of a through-flow velocity deficit which results from the tip clearance flow and accumulation of low momentum fluid centrifuged from the blade and hub surfaces. Flow traces from the CFD analysis show the origin of this deficit which begins to grow in the inlet region of the impeller where it is first detected near the suction surface side of the passage. It then moves toward the pressure side of the channel, due to the movement of tip clearance flow across the impeller passage, where it is cut by the splitter blade leading edge. As blade loading increases toward the rear of the channel the deficit region is driven back toward the suction surface by the cross-passage pressure gradient. There is no evidence of a large wake region that might result from flow separation and the impeller efficiency is relatively high. The flow field in this impeller is quite similar to that documented previously by NASA Lewis in a large low-speed backswept impeller.