Stator Performance and Unsteady Loading in Transonic Compressor Stages

1998 ◽  
Vol 120 (2) ◽  
pp. 224-232 ◽  
Author(s):  
M. Durali ◽  
J. L. Kerrebrock
Keyword(s):  
Mach Number ◽  
Pressure Distribution ◽  
Pressure Ratio ◽  
Pressure Fluctuations ◽  
Leading Edge ◽  
Large Degree ◽  
Angle Of Incidence ◽  
Mean Flow ◽  
Flow Angle ◽  
Transonic Compressor

The structure and behavior of wakes from a transonic compressor rotor and their effect on the loading and performance of the downstream stator have been investigated experimentally. The rotor was 23.25 inches in diameter with a measured tip Mach number of 1.23 and a pressure ratio of 1.66. Time and space-resolved measurements have been completed of the rotor and stator outflow, as well of the pressure distribution on the surface of the stator blades. It was found that the wakes from this rotor have large flow angle and flow Mach number variations from the mean flow, significant pressure fluctuations, and a large degree of variation from hub to tip. There was a significant total pressure defect and practically no static pressure variation associated with the stator wakes. Wakes from the rotor exist nearly undiminished in the exit flow of the stator and decay in the annular duct behind the stator. The pressure at all points along the chord over each of the stator blades’ surfaces fluctuated nearly in phase in response to the rotor wakes, that is the unsteady chordwise pressure distribution is determined mainly by the change in angle of incidence to the blade and not by the local velocity fluctuations within the passage. The unsteady forces on the stator blades, induced by the rotor wakes, were as high as 25 percent of the steady forces, and lagged the incidence of the wakes on the leading edge by approximately 180 deg at most radii.

Boundary Layer Suction via a Slot in a Transonic Compressor: Numerical Parameter Study and First Experiments

2004 ◽  
Author(s):  
K. Hubrich ◽  
A. Bo¨lcs ◽  
P. Ott
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Parameter Study ◽  
Flow Angle ◽  
Boundary Layer Suction ◽  
Transonic Compressor ◽  
Shock Mach Number

In the present paper a numerical and experimental study aiming at the enhancement of the working range of a transonic compressor via boundary layer suction (BLS) is presented. The main objective of the investigation is to study the influence of BLS on the interference between shock wave and boundary layer and to identify the possible benefit of BLS on the compressor working characteristics. An extensive numerical study has been carried out for the DATUM blade and for 2 different suction location configurations for one speed line and varying back-pressure levels, ranging from choked conditions to stall. It was found that the working range of the transonic compressor with a nominal inlet Mach number of 1.2 and a nominal pre-shock Mach number of 1.35 could be increased by sucking 2% of flow on the SS away, in such a way that the maximum pressure ratio and maximum diffusion could both be increased by 10%, when compared to the DATUM case. For smaller pressure ratios with respect to the design pressure ratio, the BLS is located in a supersonic flow region and thus creates additional losses due to a more divergent flow channel, which additionally accelerates the flow and results in a higher pre-shock Mach number creating higher losses. First measurements carried out in LTTs annular cascade, do show reasonable agreement with the computations in terms of inlet Mach number, flow angle, main shock location and stall limit. The most pronounced difference between measurements and computations is the occurrence of a terminal normal channel shock behind a bowed detached shock wave and a separation on the SS of the blade, which were both not predicted by the CFD.

scholarly journals The Role of Impeller Outflow Conditions on the Performance of Vaned Diffusers

2016 ◽  
Author(s):  
Jonathan Everitt ◽  
Zoltán Spakovszky ◽  
Daniel Rusch ◽  
Jürg Schiffman
Keyword(s):  
Mach Number ◽  
Swirling Flow ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Pressure Recovery ◽  
Strong Mixing ◽  
Mixing Processes ◽  
Vaned Diffuser ◽  
Flow Angle ◽  
High Loss

Highly-loaded impellers, typically used in turbocharger and gas turbine applications, exhaust an unsteady, transonic flow that is non-uniform across the span and pitch and swirling at angles approaching tangential. With the exception of the flow angle, conflicting data exist regarding whether these attributes have substantial influence on the performance of the downstream diffuser. This paper quantifies the relative importance of the flow angle, Mach number, non-uniformity and unsteadiness on diffuser performance, through diffuser experiments in a compressor stage and in a rotating swirling flow test rig. This is combined with steady and unsteady Reynolds-Averaged Navier Stokes computations. The test article is a pressure ratio 5 turbocharger compressor with an airfoil vaned diffuser. The swirling flow rig is able to generate rotor outflow conditions representative of the compressor except for the periodic pitchwise unsteadiness, and fits a 0.86 scale diffuser and volute. In both rigs, the time-mean impeller outflow is mapped across a diffuser pitch using miniaturized traversing probes developed for the purpose. Across approximately two-thirds of the stage operating range, diffuser performance is well correlated to the average impeller outflow angle when the metric used is effectiveness, which describes the pressure recovery obtained relative to the maximum possible given the average inflow angle and Mach number and the vane exit metal angle. Utilizing effectiveness captures density changes through the diffuser at higher Mach numbers; a 10% increase in pressure recovery is observed as the inlet Mach number is increased from 0.5 to 1. Further, effectiveness is shown to be largely independent of the time-averaged spanwise and unsteady pitchwise non-uniformity from the rotor; this independence is reflective of the strong mixing processes that occur in the diffuser inlet region. The observed exception is for operating points with high time-averaged vane incidence. Here, it is hypothesized that temporary excursions into high-loss flow regimes cause a nonlinear increase in loss as large unsteady angle variations pass by from the rotor. Given that straight-channel diffuser design charts typically used in preliminary radial vaned diffuser design capture neither streamtube area changes from impeller exit to the diffuser throat nor vane incidence effects, their utility is limited. An alternative approach, utilizing effectiveness and vane leading edge incidence, is proposed.

scholarly journals The Role of Impeller Outflow Conditions on the Performance of Vaned Diffusers

2017 ◽  
Vol 139 (4) ◽  
Author(s):  
Jonathan N. Everitt ◽  
Zoltán S. Spakovszky ◽  
Daniel Rusch ◽  
Jürg Schiffmann
Keyword(s):  
Mach Number ◽  
Swirling Flow ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Pressure Recovery ◽  
Strong Mixing ◽  
Mixing Processes ◽  
Vaned Diffuser ◽  
Flow Angle ◽  
High Loss

Highly loaded impellers, typically used in turbocharger and gas turbine applications, exhaust an unsteady, transonic flow that is nonuniform across the span and pitch and swirling at angles approaching tangential. With the exception of the flow angle, conflicting data exist regarding whether these attributes have substantial influence on the performance of the downstream diffuser. This paper quantifies the relative importance of the flow angle, Mach number, nonuniformity, and unsteadiness on diffuser performance, through diffuser experiments in a compressor stage and in a rotating swirling flow test rig. This is combined with steady and unsteady Reynolds-averaged Navier–Stokes (RANS) computations. The test article is a pressure ratio 5 turbocharger compressor with an airfoil vaned diffuser. The swirling flow rig is able to generate rotor outflow conditions representative of the compressor except for the periodic pitchwise unsteadiness and fits a 0.86 scale diffuser and volute. In both rigs, the time-mean impeller outflow is mapped across a diffuser pitch using miniaturized traversing probes developed for the purpose. Across approximately two-thirds of the stage operating range, diffuser performance is well correlated to the average impeller outflow angle when the metric used is effectiveness, which describes the pressure recovery obtained relative to the maximum possible given the average inflow angle and Mach number and the vane exit metal angle. Utilizing effectiveness captures density changes through the diffuser at higher Mach numbers; a 10% increase in pressure recovery is observed as the inlet Mach number is increased from 0.5 to 1. Further, effectiveness is shown to be largely independent of the time-averaged spanwise and unsteady pitchwise nonuniformity from the rotor; this independence is reflective of the strong mixing processes that occur in the diffuser inlet region. The observed exception is for operating points with high time-averaged vane incidence. Here, it is hypothesized that temporary excursions into high-loss flow regimes cause a nonlinear increase in loss as large unsteady angle variations pass by from the rotor. Given that straight-channel diffuser design charts typically used in preliminary radial vaned diffuser design capture neither streamtube area changes from impeller exit to the diffuser throat nor vane incidence effects, their utility is limited. An alternative approach, utilizing effectiveness and vane leading edge incidence, is proposed.

Effect of Airfoil Vortex Generator on the Performance and Stability of a Transonic Axial Compressor

2021 ◽  
Author(s):  
Subbaramu Shivaramaiah ◽  
Mahesh K. Varpe
Keyword(s):  
Pressure Ratio ◽  
Research Work ◽  
Vortex Generator ◽  
Leading Edge ◽  
Leakage Flow ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Tip Leakage ◽  
Transonic Compressor

Abstract In the present research work, effect of airfoil vortex generator on performance and stability of transonic compressor stage is investigated through CFD simulations. In turbomachines vortex generators are used to energize boundary and generated vortex is made to interact with tip leakage flow and secondary flow vortices formed in rotor and stator blade passage. In the present numerical investigation symmetrical airfoil vortex generator is placed on rotor casing surface close to leading edge, anticipating that vortex generated will be able to disturb tip leakage flow and its interaction with rotor passage core flow. Six different vortex generator configuration are investigated by varying distance between vortex generator trailing edge and rotor leading edge. Particular vortex generator configuration shows maximum improvement of stall margin and operating range by 5.5% and 76.75% respectively. Presence of vortex generator alters flow blockage by modifying flow field in rotor tip region and hence contributes to enhancement of stall margin. As a negative effect, interaction of vortex generator vortices and casing causes surface friction and high entropy generation. As a result compressor stage pressure ratio and efficiency decreases.

Transonic Compressor Rotor Cascade With Boundary-Layer Separation: Experimental and Theoretical Results

1993 ◽  
Author(s):  
R. Fuchs ◽  
W. Steinert ◽  
H. Starken
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Computational Design ◽  
Boundary Layer Separation ◽  
Ratio Test ◽  
Integral Boundary ◽  
Flow Angle ◽  
Laminar Separation ◽  
Transonic Compressor ◽  
Compressor Rotor

A transonic compressor rotor cascade designed for an inlet Mach number of 1.09 and 14 degrees of flow turning has been redesigned for higher loading by an increased pitch-to-chord ratio. Test results, showing the influence of inlet Mach number and flow angle on cascade performance are presented and compared to data of the basic design. Loss-levels of both, the original and the redesigned higher loaded blade were identical at design condition, but the new design achieved even lower losses at lower inlet Mach numbers. The computational design and analysis has been performed by a fast inviscid time-dependent code coupled to a viscous direct/inverse integral boundary-layer code. Good agreement was achieved between measured and predicted surface Mach number distributions as well as exit-flow angles. A boundary-layer visualization method has been used to detect laminar separation bubbles and turbulent separation regions. Quantitative results of measured bubble positions are presented and compared to calculated results.

Mixed Flow Turbines: Inlet and Exit Flow Under Steady and Pulsating Conditions

2000 ◽  
Author(s):  
N. Karamanis ◽  
R. F. Martinez-Botas ◽  
C. C. Su
Keyword(s):  
High Speed ◽  
Pulse Propagation ◽  
Combustion Engine ◽  
Pressure Ratio ◽  
Flow Characteristics ◽  
Leading Edge ◽  
Mixed Flow ◽  
Flow Measurements ◽  
Flow Angle ◽  
Isentropic Efficiency

The performance and detailed flow characteristics of a high pressure ratio mixed flow turbine has been investigated under steady and pulsating flow conditions. The rotor has been designed to have a nominal constant incidence (based on free vortex flow in the volute) and it is for use in an automotive high speed diesel turbocharger. The results indicated a departure from the quasi-steady analysis commonly used in turbocharger turbine design. The pulsations from the engine have been followed through the inlet pipe and around the volute; the pulse has been shown to propagate close to the speed of sound and not according to the bulk flow velocity as stated by some researchers. The flow entering and exiting the blades has been quantified by a laser Doppler velocimetry system. The measurements were performed at a plane 3.0 mm ahead of the rotor leading edge and 9.5 mm behind the rotor trailing edge. The turbine test conditions corresponded to the peak efficiency point at 29,400 and 41,300 rpm. The results were resolved in a blade-to-blade sense to examine in greater detail the nature of the flow at turbocharger representative conditions. A correlation between the combined effects of incidence and exit flow angle with the isentropic efficiency has been shown. The unsteady flow characteristics have been investigated at two flow pulse frequencies, corresponding to internal combustion engine speeds of 1600 and 2400 rpm. Four measurement planes have been investigated: one in the pipe feeding the volute, two in the volute (40° and 130° downstream of the tongue) and one at the exit of the turbine. The pulse propagation at these planes has been investigated; the effect of the different planes on the evaluation of the unsteady isentropic efficiency is shown to be significant. Overall, the unsteady performance efficiency results indicated a significant departure from the corresponding steady performance, in accordance with the inlet and exit flow measurements.

Effect of airfoil leading edge waviness on flow structures and noise

2016 ◽  
Vol 26 (6) ◽  
pp. 1821-1842 ◽  
Author(s):  
Man Zhang ◽  
Abdelkader Frendi
Keyword(s):  
Mach Number ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Space Time ◽  
Flow Structures ◽  
Mean Flow ◽  
Content Type ◽  
Acoustic Signature ◽  
Depth Analysis ◽  
Time Correlations

Purpose – The tubercles at the leading edge of Humpback Whale flippers have been shown to increase aerodynamic efficiency. The purpose of this paper is to compute the flow structures and noise signature of a NACA0012 airfoil with and without leading edge waviness, and located in the wake of a cylinder using the hybrid RANS-LES method. Design/methodology/approach – The mean flow Mach number is 0.2 and the angle of attack used is 2°. After benchmarking the method using existing experimental results, unsteady computations were then carried-out on both airfoil geometries and for a 2° angle of attack. Findings – Results from these computations confirmed the aerodynamic benefits of the leading edge waviness. Moreover, the wavy leading edge airfoil was found to be at least 4 dB quieter than its non-wavy counterpart. In-depth analysis of the computational results revealed that the wavy leading edge airfoil breaks up the large coherent structures which are then convected at higher speeds down the trough region of the waviness in agreement with previous experimental observations. This result is supported by both the two-point and space-time correlations of the wall pressure. Research limitations/implications – The limitations of the current findings reside in the fact that both the Reynolds number and the flow Mach number are low, therefore not applicable to aircrafts. In order to extend the study to practical aircrafts one needs huge grids and large computational resources. Practical implications – The results obtained here could have a huge implications on the design of future aircrafts and spacecrafts. More specifically, the biggest benefit from such redesign is the reduction of acoustic signature as well as increased efficiency in fuel consumption. Social implications – Reducing acoustic signature from aircrafts has been a major research thrust for NASA and Federal Aviation Administration. The social impact of such reduction would be improved quality of life in airport communities. For military aircrafts, this could results in reduced detectability and hence saving lives. Originality/value – Humpback Whales have been studied by various researchers to understand the effects of leading edge “tubercles” on flow structures. What is new in this study is the numerical confirmation of the effects of the tubercles on the flow structures and the resulting noise radiations. It is shown through the use of two-point correlations and space-time correlations that the flow structures in the trough area are indeed vortex tubes.

Blade Excitation by Broad-Band Pressure Fluctuations in a Centrifugal Compressor

1988 ◽  
Vol 110 (1) ◽  
pp. 129-137 ◽  
Author(s):  
U. Haupt ◽  
M. Rautenberg ◽  
A. N. Abdel-Hamid
Keyword(s):  
Centrifugal Compressor ◽  
Reverse Flow ◽  
Pressure Ratio ◽  
Pressure Fluctuations ◽  
Broad Band ◽  
Rotating Stall ◽  
High Mass ◽  
Blade Vibration ◽  
Flow Angle ◽  
Unsteady Pressure

The mechanism of blade excitation during the operation of a high-mass-flow, high-pressure-ratio centrifugal compressor has been investigated. This was carried out in the compressor operating range below 60 percent of design speed and in the zone of unsteady flow occurrence, where considerable blade vibration has been measured but no periodic unsteady pressure pattern such as rotating stall could be identified. Experiments conducted to study the mechanism of interactions between flow and blades were accomplished using several measuring methods simultaneously, such as measurements of blade vibration, flow angle at impeller inlet, unsteady pressure at different meridional and peripheral locations, as well as flow visualization by means of oil pattern. Analysis of the measurements showed typical broad-band characteristics of the unsteady pressure field and also for the blade vibration behavior. Results of flow angle investigations at the impeller inlet together with the analysis of oil pattern show that the broad-band pressure fluctuations and blade excitation can be attributed to a strong reverse flow near the suction side of the radial blade in the shroud zone. This reverse flow has its source downstream of the impeller and is extending back up to a location ahead of the impeller inlet. Similar results were obtained when the compressor was operated with vaneless and vaned diffuser configurations.

Experimental Evaluation of the Effects of a Blunt Leading Edge on the Performance of a Transonic Rotor

1973 ◽  
Vol 95 (3) ◽  
pp. 199-204 ◽  
Author(s):  
L. Reid ◽  
D. C. Urasek
Keyword(s):  
Mach Number ◽  
Experimental Evaluation ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Major Portion ◽  
Peak Efficiency ◽  
Design Speed ◽  
Blunt Leading Edge ◽  
Transonic Rotor

A rotor designed for a tip relative Mach number of 1.4 and a pressure ratio of 1.8 was tested with two different leading edge configurations, to evaluate experimentally the effects of leading edge thickness on the performance of a transonic rotor. The rotor was first tested with a leading edge thickness that resulted in a normal blade gap blockage of approximately 3 percent at the rotor tip and increased to about 4 percent at the rotor hub. The rotor leading edge was then cut back to produce a leading edge thickness which resulted in a normal blade gap blockage of about 6 percent at the tip and the hub remained at about 4 percent. The increased leading edge thickness resulted in a decrease in the rotor overall peak efficiency of 3.5 points at design speed. The major portion of this loss in efficiency occurred in the outer 30 percent of the blade span. At 90 and 70 percent of design speed, the increased leading edge thickness had a relatively small effect on rotor efficiency.

A Comparison between Measured and Computed Flow Fields in a Transonic Compressor Rotor

1980 ◽  
Vol 102 (4) ◽  
pp. 883-889 ◽  
Author(s):  
P. W. McDonald ◽  
C. R. Bolt ◽  
R. J. Dunker ◽  
H. B. Weyer
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Computational Procedure ◽  
Operating Conditions ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Pressure Profiles

The flow field within the rotor of a transonic axial compressor has been computed and compared to measurements obtained with an advanced laser velocimeter. The compressor was designed for a total pressure ratio of 1.51 at a relative tip Mach number of 1.4. The comparisons are made at 100 percent design speed (20,260 RPM) with pressure ratios corresponding to peak efficiency, near surge, and wide open discharge operating conditions. The computational procedure iterates between a blade-to-blade calculation and an intrablade through flow calculation. Calculated Mach number contours, surface pressure distributions, and exit total pressure profiles are in agreement with the experimental data demonstrating the usefulness of quasi three-dimensional calculations in compressor design.

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