Chunill Hah

An Experimental and Numerical Investigation into the Mechanisms of Rotating Instability

This paper reports on an experimental and numerical investigation aimed at understanding the mechanisms of rotating instabilities in a low speed axial flow compressor. The phenomena of rotating instabilities in the current compressor were first identified with an experimental study. Then, an unsteady numerical method was applied to confirm the phenomena and to interrogate the physical mechanisms behind them. The experimental study was conducted with high-resolution pressure measurements at different clearances, employing a double phase-averaging technique. The numerical investigation was performed with an unsteady 3-D Navier-Stokes method that solves for the entire blade row. The current study reveals that a vortex structure forms near the leading edge plane. This vortex is the result of interactions among the classical tip-clearance flow, axially reversed endwall flow, and the incoming flow. The vortex travels from the suction side to the pressure side of the passage at roughly half of the rotor speed. The formation and movement of this vortex seem to be the main causes of unsteadiness when rotating instability develops. Due to the nature of this vortex, the classical tip clearance flow does not spill over into the following blade passage. This behavior of the tip clearance flow is why the compressor operates in a stable mode even with the rotating instability, unlike traditional rotating stall phenomena.

Development of Hub Corner Stall and Its Influence on the Performance of Axial Compressor Blade Rows

A detailed investigation has been performed to study hub corner stall phenomena in compressor blade rows. Three-dimensional flows in a subsonic annular compressor stator and in a transonic compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier-Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of compressor hub corner stall. Both the measurements and the numerical solutions for the stator indicate that a strong twisterlike vortex is formed near the rear part of the blade suction surface. Low-momentum fluid inside the hub boundary layer is transported toward the suction side of the blade by this vortex. On the blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic compressor rotor studied in this paper, the numerical solution indicates that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced blade loading, which occurs at 60 percent rotor design speed. The present study demonstrates that hub corner stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the blade row.

SHORT LENGTH-SCALE ROTATING STALL INCEPTION IN A TRANSONIC AXIAL COMPRESSOR - CRITERIA AND MECHANISMS

The current paper reports on investigations aimed at advancing the understanding of the flow mechanism that leads to the onset of short-length scale rotating stall in a transonic axial compressor. Experimental data show large oscillation of the tip clearance vortex as the rotor operates near the stall condition. Inception of spike-type rotating stall is also measured in the current transonic compressor with high response pressure transducers. Computational studies of a single passage and the full annulus were carried out to identify flow mechanisms behind the spike-type stall inception in the current transonic compressor rotor. Steady and unsteady single passage flow simulations were performed, first to get insight into the interaction between the tip clearance vortex and the passage shock. The conventional Reynolds-averaged Navier-Stokes method with a standard turbulence closure scheme does not accurately reproduce tip clearance vortex oscillation and the measured unsteady pressure field. Consequently, a Large Eddy Simulation (LES) was carried out to capture more relevant physics in the computational simulation of the rotating stall inception. The unsteady random behavior of the tip clearance vortex and it’s interaction with the passage shock seem to be critical ingredients in the development of spike-type rotating stall in a transonic compressor. The Large Eddy Simulation was further extended to the full annulus to identify flow mechanisms behind the measured spike-type rotating stall inception. The current study shows that the spike-type rotating stall develops after the passage shock is fully detached from the blade passages. Interaction between the tip clearance vortex and the passage shock creates a low momentum area near the pressure side of the blade. As the mass flow rate decreases, this low momentum area moves further upstream and reversed tip clearance flow is initiated at the trailing edge plane. Eventually, the low momentum area near the pressure side reaches the leading edge and forward spillage of the tip clearance flow occurs. The flows in the affected blade passage or passages then stall. As the stalled blade passages are formed behind the passage shock, the stalled area rotates counter to the blade rotation just like the classical Emmon’s type rotating stall. Both the measurements and the computations show that the rotating stall cell covers one to two blade passage lengths and rotates at roughly 50% of the rotor speed.Copyright

Role of Tip-Leakage Vortices and Passage Shock in Stall Inception in a Swept Transonic Compressor Rotor

The current paper reports on investigations aimed at advancing the understanding of the flow field near the casing of a forward-swept transonic compressor rotor. The role of tip clearance flow and its interaction with the passage shock on stall inception are analyzed in detail. Steady and unsteady three-dimensional viscous flow calculations are applied to obtain flow fields at various operating conditions. The numerical results are first compared with available measured data. Then, the numerically obtained flow fields are interrogated to identify the roles of flow interactions between the tip clearance flow, the passage shock, and the blade/endwall boundary layers. In addition to the flow field with nominal tip clearance, two more flow fields are analyzed in order to identify the mechanisms of blockage generation: one with zero tip clearance, and one with nominal tip clearance on the forward portion of the blade and zero clearance on the aft portion. The current study shows that the tip clearance vortex does not break down, even when the rotor operates in a stalled condition. Interaction between the shock and the suction surface boundary layer causes the shock, and therefore the tip clearance vortex, to oscillate. However, for the currently investigated transonic compressor rotor, so-called breakdown of the tip clearance vortex does not occur during stall inception. The tip clearance vortex originates near the leading edge tip, but moves downward in the spanwise direction inside the blade passage. A low momentum region develops above the tip clearance vortex from flow originating from the casing boundary layer. The low momentum area builds up immediately downstream of the passage shock and above the core vortex. This area migrates toward the pressure side of the blade passage as the flow rate is decreased. The low momentum area prevents incoming flow from passing through the pressure side of the passage and initiates stall inception. It is well known that inviscid effects dominate tip clearance flow. However, complex viscous flow structures develop inside the casing boundary layer at operating conditions near stall.Copyright

A Viscous Flow Study of Shock-Boundary Layer Interaction, Radial Transport, and Wake Development in a Transonic Compressor

A numerical study based on the three-dimensional Reynolds-averaged Navier–Stokes equation has been conducted to investigate the detailed flow physics inside a transonic compressor. Three-dimensional shock structure, shock-boundary layer interaction, flow separation, radial mixing, and wake development are all investigated at design and off-design conditions. Experimental data based on laser anemometer measurements are used to assess the overall quality of the numerical solution. An additional experimental study to investigate end-wall flow with a hot film was conducted, and these results are compared with the numerical results. Detailed comparison with experimental data indicates that the overall features of the three-dimensional shock structure, the shock-boundary layer interaction, and the wake development are all calculated very well in the numerical solution. The numerical results are further analyzed to examine the radial mixing phenomena in the transonic compressor. A thin sheet of particles is injected in the numerical solution upstream of the compressor. The movement of particles is traced with a three-dimensional plotting package. This numerical survey of tracer concentration reveals the fundamental mechanisms of radial transport in this transonic compressor. Strong radially outward flow is observed inside a separated flow region and this outward flow accounts for about 80 percent of the total radial transport. The radially inward flow is mainly due to the traditional secondary flow.

Effect of Circumferential Grooves on the Aerodynamic Performance of an Axial Single-Stage Transonic Compressor

This paper reports on experimental and numerical investigations on circumferential grooves in an axial single-stage transonic compressor. Total pressure ratio and efficiency speedlines were taken at design speed and three off-design conditions. The experiments comprise four different configurations with deep and shallow grooves and variable coverage of the projected rotor axial chord. All casing treatments proved to have a beneficial effect on stall range while maintaining high levels of efficiency, even at off-design operation. Deep grooves extending almost to the trailing edge showed the biggest potential: the mass flow at stall inception for design speed could be strongly reduced, and the operating range could be enlarged by 56.1%. When three shallow grooves were applied to the compressor, the stage efficiency at design speed was shifted to slightly higher values. A possible explanation could be a favorable change in stator aerodynamics due to the reduction of corner separation. For a closer look into the physical effects of grooves on the tip leakage flow, a rotor-only CFD analysis has been carried out using a steady state calculation. A multi-block grid with approximately 1.2 million nodes was used. The numerical simulations reveal strong effects of circumferential grooves on the rotor flow field at tip. Mach-number contours, axial velocity distributions and particle traces for the smooth casing and six deep grooves are presented at stall mass flow. Compared to the smooth wall case, the treated casing significantly reduces blockage in the tip area and weakens the roll-up of the core vortex. These mechanisms prevent an early spillage of low momentum fluid into the adjacent blade passage and delay the onset of rotating stall.Copyright

The Application of Pressure- and Temperature-Sensitive Paints to an Advanced Compressor

An innovative pressure-measurement technique that employs the tools of molecular spectroscopy has been widely investigated by the aerospace community. Measurements are made via oxygen-sensitive molecules attached to the surface of interest as a coating, or paint. The pressure-sensitive-paint (PSP) technique is now commonly used in stationary wind-tunnel tests; this paper presents the use of this technique in advanced turbomachinery applications. New pressure- and temperature-sensitive paints (P/TSPs) have been developed for application to a state-of-the-art transonic compressor where pressures up to 1.4 atm and surface temperatures to 90°C are expected for the suction surface of the first-stage rotor. PSP and TSP data images have been acquired from the suction surface of the first-stage rotor at 85 percent of the corrected design speed for the compressor near-stall condition. A comparison of experimental results with CFD calculations is discussed.

Tip Clearance Vortex Oscillation, Vortex Shedding and Rotating Instabilities in an Axial Transonic Compressor Rotor

Unsteady flow characteristics in a modern transonic axial compressor operating near stall are studied in detail. Measured data from high-response pressure probes show that the tip clearance vortex oscillates substantially near stall. Instantaneous flow structure varies substantially among different blade passages even with uniform inlet flow. Fast Fourier transformation of measured wall pressure shows a dominant frequency component that is between 30% and 40% of the rotor speed. To identify and analyze this phenomenon, computational studies based on a single passage and full annulus were carried out. The flow field in a transonic compressor near stall is heavily influenced by the unsteady motion of tip clearance vortices. Therefore, a Large Eddy Simulation (LES) was carried out to capture transient characteristics of the tip clearance vortex more realistically. The wall pressure spectrum from the current full annulus analysis also shows a dominant frequency when the rotor operates near stall. The calculated peak frequency is about 30% of the rotor frequency. The dominant frequency, which is non-synchronous with the rotor blade, is due to rotating flow instabilities. Flow interactions across blade passages due to synchronized tip clearance vortex oscillation seem to be the main cause.Copyright

Suppression of Cavitation Instabilities in an Inducer by Circumferential Groove and Explanation of Higher Frequency Components

The purpose of the present research is to suppress cavitation instabilities by using a circumferential groove. The circumferential groove was designed based on CFD so that the tip leakage vortex is trapped by the groove and does not interact with the next blade. Experimental results show that the groove can suppress rotating cavitation, asymmetric cavitation and cavitation surge. However, weak instabilities with higher frequency could not be suppressed by the groove. From the analysis of pressure pattern similar to that for rotor-stator interaction, it was found that the higher frequency components are caused by the interaction of backflow vortices with the inducer blades.

Development of Secondary Flow Field in a Low Solidity Diffuser in a Transonic Centrifugal Compressor Stage

In the present paper, the flow structure inside a low-solidity diffuser of a transonic compressor was investigated in detail. Steady computations were carried out and compared to experimental data. The secondary flow development inside the diffuser was analyzed and the reason for the stall inception was detected. Unsteady calculations were performed for two operating points, one close to the choke and the other one close to the stall of the compressor, in order to assess the effect of the unsteadiness in the diffuser secondary flow development.Copyright