Stall Margin Improvement of a Single Stage Transonic Axial Flow Compressor Using Naturally Aspirated Slots

This paper describes the method to improve the stall margin of transonic axial flow compressor by controlling the boundary layer on the suction surface of the rotor blade tip through natural aspiration. Aspiration slots in the compressor blade are intended to energize the flow by increasing its momentum on the suction surface. This phenomenon of boundary layer control can delay the flow separation and hence results in enhancement of the stall margin of the compressor stage. Flow behavior with aspiration slots and its performance are evaluated using commercially available numerical software. Steady state RANS simulations with three dimensional implicit pressure-based coupled solver and turbulence model SST k-ω are used. The effect of natural aspiration slot on the rotor blade performance is computed numerically. The main objective of the study was to identify the optimum location of the aspiration slot along the chord of the compressor on the rotor blade. The axial location chosen for the performance evaluations were 20%,40%,50%,60% and 70% of the rotor blade axial tip chord. By comparing the numerical simulation results with the steady state behavior in the absence of the aspirated slots, the optimized location of the aspiration slot that results in maximum stall improvement is identified. At the optimized location, natural aspiration slots on the rotor blade tip improved the stall margin with the minimum reduction in efficiency and stage pressure ratio when compared to base model. After critically understanding the performance with straight aspiration slots the compressor stage performance has enhanced further by orienting the aspiration slots. The numerical three dimensional results conclude an optimal improvement in the stall margin for the slots near the trailing edge of the rotor. The prediction shows that with the inclined aspiration slots at proper location it is possible to improve the stall margin of the compressor stage and also to restore the stage efficiency.

An Experimental Study of the Compressor Rotor Blade Boundary Layer

Journal of Engineering for Gas Turbines and Power ◽

10.1115/1.3239731 ◽

1985 ◽

Vol 107 (2) ◽

pp. 364-372 ◽

Cited By ~ 3

Author(s):

M. Pouagare ◽

J. M. Galmes ◽

B. Lakshminarayana

Keyword(s):

Boundary Layer ◽

Rotor Blade ◽

Radial Profile ◽

Three Dimensional ◽

Axial Flow ◽

Tip Leakage Flow ◽

Wall Boundary Layer ◽

Compressor Rotor ◽

Blade Tip ◽

The three-dimensional turbulent boundary layer developing on a rotor blade of an axial flow compressor was measured using a minature “x” configuration hot-wire probe. The measurements were carried out at nine radial locations on both surfaces of the blade at various chordwise locations. The data derived includes streamwise and radial mean velocities and turbulence intensities. The validity of conventional velocity profiles such as the “power law profile” for the streamwise profile, and Mager and Eichelbrenner’s for the radial profile, is examined. A modification to Mager’s crossflow profile is proposed. Away from the blade tip, the streamwise component of the blade boundary layer seems to be mainly influenced by the streamwise pressure gradient. Near the tip of the blade, the behavior of the blade boundary layer is affected by the tip leakage flow and the annulus wall boundary layer. The “tangential blockage” due to the blade boundary layer is derived from the data. The profile losses are found to be less than that of an equivalent cascade, except in the tip region of the blade.

Experimental Investigations of Airfoil- and Endwall Boundary Layers in a Subsonic Compressor Stage

10.1115/86-gt-143 ◽

1986 ◽

Cited By ~ 4

Author(s):

H. E. Gallus ◽

H. Hoenen

Keyword(s):

Boundary Layer ◽

Flow Field ◽

Boundary Layers ◽

Three Dimensional ◽

Axial Flow ◽

Experimental Investigations ◽

Turbulence Energy ◽

Compressor Stage ◽

Stator Blade ◽

Criteria for the maximum diffusion allowable in a blade row without reaching stall play an important part in the design of highly loaded axial-flow compressors. Most of these criteria for maximum blade loading were derived from wind tunnel measurements of 2-d-steady cascade flow. As the flow field in turbomachines is extremely unsteady and of three-dimensional nature the boundary layers are influenced by these effects. The paper deals with the results of boundary layer measurements in a stator blade channel of a subsonic axial-flow compressor stage at various operating points between unthrottled and highly throttled flow (near stall). In front of the stator, the time-averaged velocity profiles as well as the fluctuations due to the unsteady flow field downstream of the rotor were measured. The growing of the separation zones inside the channel with increased blade load is studied in detail. Photos of flow visualization in the boundary layers by dye-injection and flow patterns derived from hot-wire measurements illustrate the physics of boundary layer behavior and separation due to increasing stator blade load. The investigations include measurements of the turbulence energy and a frequency analysis of the velocity fluctuations in the boundary layers.

Effects of a Vane-Recessed Tubular-Passage Passive Stall Control Technique on a Multistage, Axial-Flow Compressor: Results of Tests on the First Stage With the Rear Stages Removed

Volume 6: Turbo Expo 2003, Parts A and B ◽

10.1115/gt2003-38301 ◽

2003 ◽

Cited By ~ 1

Author(s):

M. Akhlaghi ◽

R. L. Elder ◽

K. W. Ramsden

Keyword(s):

Rotor Blade ◽

Axial Flow ◽

Control Technique ◽

Maximum Efficiency ◽

Stall Margin ◽

Casing Treatment ◽

Axial Flow Compressor ◽

Stall Margin Improvement ◽

Blade Tip ◽

The objective of the current study was to investigate the effect of casing treatment on a multistage axial flow compressor. The main purpose of the investigation was to extend the range and operability of multistage axial compressors. The study seeks to establish whether a vane-recessed tubular-passage casing-treatment could provide beneficial stall margin improvement, without sacrificing the efficiencies of the compressor with the restricted space available for the treatment. A casing treatment that consisted of three parts: an outer casing ring, with a tubular shaped passage on the inside, a set of 120 evenly spaced curved vanes, and then a shroud or inner ring was developed from two initial designs. The casing treatment, manufactured from high quality acrylic, was positioned upstream and partly covering the tip of the first stage rotor blades. The casing treatment was tested on the first stage of a three-stage low-speed compressor with inlet guide vanes with the rear two stages removed. The rotor blade tip axial chord exposure had a significant impact on the effectiveness of the casing treatment. Seven compressor configuration incorporating casing treatments of 23.2%, 33.3%, 43.4%, 53.5%, 63.6%, 73.7% and 83.8% rotor exposure were tested. The results showed significant improvements in stall margin in all exposures and insignificant efficiency sacrifices in some exposures. Nearly 29% of stall margin improvement in terms of the corrected mass flow rate was achieved with 33.3% rotor blade tip axial chord exposure. The compressor build with 53.5% rotor exposure was the best configuration in terms of maximum efficiency gain. In terms of peak pressure rise coefficients the compressor configuration with a casing treatment of 63.6% exposure was the best design. The results also suggest that the vane-recessed tubular-passage casing treatment designed as part of this research, in most instances enabled the stall conditions in the compressor to become progressive rather than abrupt.

Three-Dimensional Flows and Loss Reduction in Axial Compressors

Journal of Turbomachinery ◽

10.1115/1.3262113 ◽

1987 ◽

Vol 109 (3) ◽

pp. 354-361 ◽

Cited By ~ 41

Author(s):

Y. Dong ◽

S. J. Gallimore ◽

H. P. Hodson

Keyword(s):

Three Dimensional ◽

Axial Compressor ◽

Axial Flow ◽

Tracer Gas ◽

Suction Surface ◽

Axial Compressors ◽

Small Clearance ◽

Oil Flow ◽

Stator Blade ◽

Measurements have been performed in a low-speed high-reaction single-stage axial compressor. Data obtained within and downstream of the rotor, when correlated with the results of other investigations, provide a link between the existence of suction surface–hub corner separations, their associated loss mechanisms, and blade loading. Within the stator, it has been shown that introducing a small clearance between the stator blade and the stationary hub increases the efficiency of the stator compared to the case with no clearance. Oil flow visualizaton indicated that the leakage reduced the extensive suction surface–hub corner separation that would otherwise exist. A tracer gas experiment showed that the large radial shifts of the surface streamlines indicated by the oil flow technique were only present close to the blade. The investigation demonstrates the possible advantages of including hub clearance in axial flow compressor stator blade rows.

Three-Dimensional Boundary Layer on a Compressor Rotor Blade at Peak Pressure Rise Coefficient

10.1115/86-gt-186 ◽

1986 ◽

Author(s):

B. Lakshminarayana ◽

P. Popovski

Keyword(s):

Boundary Layer ◽

Rotor Blade ◽

Peak Pressure ◽

Pressure Coefficient ◽

Three Dimensional ◽

Pressure Rise ◽

Layer Growth ◽

Leakage Flow ◽

Suction Surface ◽

Compressor Rotor

A comprehensive study of the three-dimensional turbulent boundary layer on a compressor rotor blade at peak pressure rise coefficient is reported in this paper. The measurements were carried out at various chordwise and radial locations on a compressor rotor blade using a rotating miniature “V” configuration hot-wire probe. The data are compared with the measurement at the design condition. Substantial changes in the blade boundary layer characteristics are observed, especially in the outer sixteen percent of the blade span. The increased chordwise pressure gradient and the leakage flow at the peak pressure coefficient have a cumulative effect in increasing the boundary layer growth on the suction surface. The leakage flow has a beneficial effect on the pressure surface. The momentum and boundary layer thicknesses increase substantially from those at the design condition, especially near the outer radii of the suction surface.

Influence of Rain Ingestion on the Endwall Treatment in an Axial Flow Compressor

Journal of Turbomachinery ◽

10.1115/1.4040550 ◽

2018 ◽

Vol 140 (8) ◽

Cited By ~ 3

Author(s):

Jichao Li ◽

Juan Du ◽

Mingzhen Li ◽

Feng Lin ◽

Hongwu Zhang ◽

...

Keyword(s):

Rotor Blade ◽

Axial Flow ◽

Pressure Rise ◽

Water Film ◽

Stall Margin ◽

Axial Flow Compressor ◽

Water Ingestion ◽

Blade Row ◽

Flow Compressor ◽

The Stability

The effects of water ingestion on the performance of an axial flow compressor are experimentally studied with and without endwall treatment. The background to the work is derived from the assessment of airworthiness for an aero-engine. The stability-enhancing effects with endwall treatments under rain ingestion are not previously known. Moreover, all the endwall treatments are designed under dry air conditions in the compressor. Water ingestion at 3% and 5% relative to the design mass flow proposed in the airworthiness standard are applied to initially investigate the effects on the performance under smooth casing (SC). Results show that the water ingestions are mainly located near the casing wall after they move through the rotor blade row. The pressure rise coefficient increases, efficiency declines, and torque increases under the proposed water ingestion. The increase of the inlet water increases the thickness of the water film downstream the rotor blade row and aggravates the adverse effects on the performances. Subsequently, three endwall treatments, namely circumferential grooves, axial slots, and hybrid slots–grooves, are tested with and without water ingestion. Compared with no water ingestion, the circumferential grooves basically have no resistance to the water ingestion. The axial slots best prevent the drop of the pressure rise coefficient induced by water ingestion, and hybrid slots–grooves are the second-best place owing to the contribution of the front axial slots. Therefore, the hybrid slots–grooves can not only extend the stall margin with less efficiency penalty compared with axial slots, but also prevent rain ingestion from worsening the compressor performance.

Wall Boundary Layer Development Near the Tip Region of an IGV of an Axial Flow Compressor

Journal of Engineering for Gas Turbines and Power ◽

10.1115/1.3239570 ◽

1984 ◽

Vol 106 (2) ◽

pp. 337-345

Author(s):

B. Lakshminarayana ◽

N. Sitaram

Keyword(s):

Boundary Layer ◽

Guide Vane ◽

Three Dimensional ◽

Axial Compressor ◽

Axial Flow ◽

Inlet Guide Vane ◽

Wall Boundary Layer ◽

Wall Boundary ◽

Boundary Layer Development ◽

The annulus wall boundary layer inside the blade passage of the inlet guide vane (IGV) passage of a low-speed axial compressor stage was measured with a miniature five-hole probe. The three-dimensional velocity and pressure fields were measured at various axial and tangential locations. Limiting streamline angles and static pressures were also measured on the casing of the IGV passage. Strong secondary vorticity was developed. The data were analyzed and correlated with the existing velocity profile correlations. The end wall losses were also derived from these data.

Three-Dimensional Flows and Loss Reduction in Axial Compressors

10.1115/86-gt-193 ◽

1986 ◽

Author(s):

Y. Dong ◽

S. J. Gallimore ◽

H. P. Hodson

Keyword(s):

Three Dimensional ◽

Axial Compressor ◽

Axial Flow ◽

Flow Visualisation ◽

Tracer Gas ◽

Suction Surface ◽

Axial Compressors ◽

Oil Flow ◽

Stator Blade ◽

Measurements have been performed in a low speed high reaction single stage axial compressor. Data obtained within and downstream of the rotor, when correlated with the results of other investigations, provide a link between the existence of suction surface-hub corner separations, their associated loss mechanisms and blade loading. Within the stator, it has been shown that introducing a small clearance between the stator blade and the stationary hub increases the efficiency of the stator compared to the case with no clearance. Oil flow visualisation indicated that the leakage reduced the extensive suction surface-hub corner separation that would otherwise exist. A tracer gas experiment showed that the large radial shifts of the surface streamlines indicated by the oil flow technique were only present close to the blade. The investigation demonstrates the possible advantages of including hub clearance in axial flow compressor stator blade rows.

Volume 1: Aircraft Engine; Marine; Turbomachinery; Microturbines and Small Turbomachinery ◽

Inner Workings of Aerodynamic Sweep

10.1115/97-gt-401 ◽

1997 ◽

Cited By ~ 9

Author(s):

A. R. Wadia ◽

P. N. Szucs ◽

D. W. Crall

Keyword(s):

Boundary Layer ◽

Flow Analysis ◽

Three Dimensional ◽

Axial Flow ◽

Tip Clearance ◽

Surface Boundary ◽

Surface Boundary Layer ◽

Stall Margin ◽

Design Speed ◽

The Impact

The recent trend in using aerodynamic sweep to improve the performance of transonic blading has been one of the more significant technological evolutions for compression components in turbomachinery. This paper reports on the experimental and analytical assessment of the pay-off derived from both aft and forward sweep technology with respect to aerodynamic performance and stability. The single stage experimental investigation includes two aft-swept rotors with varying degree and type of aerodynamic sweep and one swept forward rotor. On a back-to-back test basis, the results are compared with an unswept rotor with excellent performance and adequate stall margin. Although designed to satisfy identical design speed requirements as the unswept rotor, the experimental results reveal significant variations in efficiency and stall margin with the swept rotors. At design speed, all the swept rotors demonstrated a peak stage efficiency level that was equal to that of the unswept rotor. However, the forward-swept rotor achieved the highest rotor-alone peak efficiency. At the same time, the forward-swept rotor demonstrated a significant improvement in stall margin relative to the already satisfactory level achieved by the unswept rotor. Increasing the level of aft sweep adversely affected the stall margin. A three-dimensional viscous flow analysis was used to assist in the interpretation of the data. The reduced shock/boundary layer interaction, resulting from reduced axial flow diffusion and less accumulation of centrifuged blade surface boundary layer at the up, was identified as the prime contributor to the enhanced performance with forward sweep. The impact of tip clearance on the performance and stability for one of the aft-swept rotors was also assessed.

Effect of Axially Non-Uniform Rotor Tip Clearance on the Performance of a High Speed Axial Flow Compressor Stage

10.1115/94-gt-479 ◽

1994 ◽

Cited By ~ 3

Author(s):

K. Mohan ◽

S. A. Guruprasad

Keyword(s):

High Speed ◽

Axial Flow ◽

Separated Flow ◽

Tip Clearance ◽

Compressor Stage ◽

Tip Leakage Flow ◽

Suction Surface ◽

Wall Boundary Layer ◽

Improved Performance ◽

An axially non-uniform type of rotor tip clearance was conceived and tried on a single stage compressor. This concept is based on the advantages of a smaller tip clearance in the front portion of the blade and a larger clearance in the rear portion which allows a higher tip leakage flow to interact with the passage secondary flow, casing wall boundary layer, separated flow on the blade suction surface and the scraping vortex, which are more prominent at the rear portion of the blade. Experimental results indicated that an axially non-uniform clearance can provide improved performance of a compressor stage. Providing the tip clearance in the compressor casing instead of at the blade tip indicated certain advantages. An ‘optimum’ value of rotor tip clearance was noticed for this compressor stage, both for axially uniform and axially non-uniform clearance.