Anthony J. Strazisar

1997 Best Paper Award—Controls and Diagnostics Committee: Active Stabilization of Rotating Stall and Surge in a Transonic Single-Stage Axial Compressor

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 11 and 3.5 percent reductions in stalling mass flow, respectively, with injection adding 3.6 percent of the design compressor mass flow. The aerodynamic effects of the injection have also been examined. At a tip Mach number, M tip , 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 prestall perturbations using constant gain feedback. At a M tip of 1.5 (design rotor speed), the prestall 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 ten 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 percent 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 a this speed. A dynamic, model-based robust controller was therefore us to stabilize the multiple modes that co prise the first three harmonic perturbations in this transonic region of operation.

Recommendations for Achieving Accurate Numerical Simulation of Tip Clearance Flows in Transonic Compressor Rotors

The tip clearance flows of transonic compressor rotors are important because they have a significant impact on rotor and stage performance. A wall-bounded shear layer formed by the relative motion between the overtip leakage flow and the shroud wall is found to have a major influence on the development of the tip clearance flow field. This shear layer which has not been recognized by earlier investigators, impacts the stable operating range of the rotor. Simulation accuracy is dependent on the ability of the numerical code to resolve this layer. While numerical simulations of these flows are quite sophisticated, they are seldom verified through rigorous comparisons of numerical and measured data because these kinds of measurements are rare in the detail necessary to be useful in high-speed machines. In this paper we compare measured tip-clearance flow details (e.g., trajectory and radial extent) with corresponding data obtained from a numerical simulation. Laser-Doppler Velocimeter (LDV) measurements acquired in a transonic compressor rotor, NASA Rotor 35, are used. The tip clearance flow field of this transonic rotor is simulated using a Navier-Stokes turbomachinery solver that incorporates an advanced k-e turbulence model derived for flows that are not in local equilibrium. A simple method is presented for determining when the wall-bounded shear layer is an important component of the tip clearance flow field.

Compressor Stability Enhancement Using Discrete Tip Injection

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-average 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% speed the stalling flow coefficient can be reduced by 30% using an injected massflow equivalent to 1% of the annulus flow. At design speed, the stalling flow coefficient was reduced by 6% using an injected massflow equivalent to 2% 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.Copyright

Active Flow Separation Control of a Stator Vane Using Embedded Injection in a Multistage Compressor Experiment

Active flow control has been applied to the suction surface of stator vanes in a low speed axial compressor. Injection from the suction surface has been shown to reduce separation on vanes that were induced to separate by increasing the vane stagger angle by 3° deg. Various configurations were investigated including injector geometry (slots versus holes) and type of injection (steady versus unsteady). Unsteady injection was realized using two different approaches; external actuation through a high frequency valve and embedded actuation using a fluidic device internal to the vane. Using total pressure loss through the vane passage as a metric, reductions in area-averaged loss of 25% were achieved using injected mass flow rates on the order of 1% of compressor throughflow. The development of a tracking control algorithm was also explored for the purpose of closed-loop control. A reliable method of detecting surface separation was implemented using unsteady pressure measurements on the compressor casing near the vane suction surface.

Compressor Stall Control Through Endwall Recirculation

Experiments that demonstrate the use of endwall recirculation to control the stall of transonic compressor stages are described. Endwall recirculation of a compressor stage is implemented by bleeding air from the casing downstream of a stator blade row and injecting the air as a wall jet upstream of a preceding rotor blade row. The bleed ports, injection ports, and recirculation channels are circumferentially discrete, and occupy only 20–30% of the circumference. The development of compact wall-jet injectors is described first. Next, the results of proof-of-concept steady recirculation tests on a single-stage transonic compressor are presented. Finally, the potential for using endwall recirculation to increase the stability of transonic highly-loaded multistage compressors is demonstrated through results from a rig test of simulated recirculation driving both a steady injected flow and an unsteady injected flow commanded by closed-loop active control during compressor operation at 78–100% of design speed. In this test air from an external source was injected upstream of several rotor blade rows while compressor bleed was increased by an amount equivalent to the injected massflow. During closed loop control, wall static pressure fluctuations were monitored and the injected flow rate was controlled to reduce the stalling mass flow. The use of wall jet injection to study the dynamics of transonic compressor stages is also discussed.Copyright

An Improved Aspirating Probe for Total-Temperature and Total-Pressure Measurements in Compressor Flows

The aspirating probe originally designed by Epstein and Ng at MIT was modified by replacing the two platinum-coated tungsten hot wires normally used with platinum-iridium alloy wires. The resulting improved unsteady total pressure and total temperature resolution of the modified probe is demonstrated. Flowfield measurements were made downstream of NASA Rotor 37 for a part-speed operating condition to test the performance of the probe. Time-resolved blade-to-blade total temperature and total pressure as calculated from the two platinum-iridium hot-wire voltages are shown. The flowfield measurements are compared with independent measurements of total pressure with high response transducers and total temperature calculated from laser anemometer measurements. Limitations ofa more often used unsteady temperature data reduction method, which involves only one aspirating probe hot-wire voltage and a high-response pressure measurement, are discussed.

Active Flow Separation Control of a Stator Vane Using Surface Injection in a Multistage Compressor Experiment

Micro-flow control actuation embedded in a stator vane was used to successfully control separation and improve near stall performance in a multistage compressor rig at NASA Glenn. Using specially designed stator vanes configured with internal actuation to deliver pulsating air through slots along the suction surface, a research study was performed to identify performance benefits using this microflow control approach. Pressure profiles and unsteady pressure measurements along the blade surface and at the shroud provided a dynamic look at the compressor during microflow air injection. These pressure measurements lead to a tracking algorithm to identify the onset of separation. The testing included steady air injection at various slot locations along the vane. The research also examined the benefit of pulsed injection and actively controlled air injection along the stator vane. Two types of actuation schemes were studied, including an embedded actuator for on-blade control. Successful application of an online detection and flow control scheme will be discussed. Testing showed dramatic performance benefit for flow reattachment and subsequent improvement in diffusion through the use of pulsed controlled injection. The paper will discuss the experimental setup, the blade configurations, and preliminary CFD results which guided the slot location along the blade. The paper will also show the pressure profiles and unsteady pressure measurements used to track flow control enhancement, and will conclude with the tracking algorithm for adjusting the control.

The Effect of Variable Chord Length on Transonic Axial Rotor Performance

Aircraft fan and compressor blade leading edges suffer from atmospheric particulate erosion that reduces aerodynamic performance. Recontouring the blade leading edge region can restore blade performance. This process typically results in blades of varying chord length. The question therefore arises as to whether performance of refurbished fans and compressors could be further improved if blades of varying chord length are installed into the disk in a certain order. To investigate this issue the aerodynamic performance of a transonic compressor rotor operating with blades of varying chord length was measured in back-to-back compressor test rig entries. One half of the rotor blades were the full nominal chord length while the remaining half of the blades were cut back at the leading edge to 95% of chord length and recontoured. The rotor aerodynamic performance was measured at 100, 80, and 60% of design speed for three blade installation configurations: nominal-chord blades in half of the disk and short-chord blades in half of the disk; four alternating quadrants of nominal-chord and short-chord blades; nominal-chord and short-chord blades alternating around the disk. No significant difference in performance was found between configurations, indicating that blade chord variation is not important to aerodynamic performance above the stall chord limit if leading edges have the same shape. The stall chord limit for most civil aviation turbofan engines is between 94-96% of nominal (new) blade chord.

High Resolution RANS Nonlinear Harmonic Study of Stage 67 Tip Injection Physics

Numerical prediction of the Stage 67 transonic fan stage employing wall jet tip injection flow control and study of the physical mechanisms leading to stall suppression and stability enhancement afforded by endwall recirculation/injection is the focus of this paper. Reynolds averaged Navier–Stokes (RANS) computations were used to perform detailed analysis of the Stage 67 configuration experimentally tested at NASAs Glenn Research Center in 2004. Time varying predictions of the stage plus recirculation and injection flowpath were executed utilizing the nonlinear harmonic (NLH) approach. Significantly higher grid resolution per passage was achieved than what has been generally employed in prior reported numerical studies of spike stall phenomena in transonic compressors. This paper focuses on characterizing the physics of spike stall embryonic stage phenomena and the influence of tip injection, resulting in experimentally and numerically demonstrated stall suppression.

The Effect of Adding Roughness and Thickness to a Transonic Axial Compressor Rotor

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 microns (100-125 RMS microinches) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10% at the hub and 20% at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9% reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effect 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 microns (10-20 RMS microinches), compared to the bare metal blade surface finish of 0.508 RMS microns (20 RMS microinches). 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% of design speed. The results indicate that thickness/roughness over the first 10% of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70% 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-3D 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 coating the blade 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.