Mach Number Effect on Symmetric and Antisymmetric Modes of Base Pressure Fluctuations

2018 ◽  
Vol 141 (2) ◽  
Author(s):  
N. S. Vikramaditya ◽  
M. Viji
Keyword(s):  
Mach Number ◽  
Vortex Shedding ◽  
Pressure Fluctuations ◽  
Base Pressure ◽  
Antisymmetric Mode ◽  
Recirculation Bubble ◽  
Pressure Drops ◽  
Wide Range ◽  
Convective Motions ◽  
Mach Numbers

An experimental study aimed at evaluating the influence of Mach number on the base pressure fluctuations of a cylindrical afterbody was performed over a wide range of Mach numbers from subsonic to supersonic speeds. Time-averaged results indicate that the coefficient of base pressure drops with the increase in the freestream Mach number at subsonic speeds and increases at supersonic Mach numbers. The coefficient of root-mean-square of the pressure fluctuations follows a decreasing trend with the increase in the Mach number. Examination of the spectra reveals different mechanisms dominate the pressure fluctuations from the center to the periphery of the base as well as with the change in the Mach number. Analysis of the azimuthal coherence indicates that all the dominant tones in the spectra can be classified either into a symmetric or an antisymmetric mode at subsonic Mach numbers. However, at supersonic Mach numbers, all the dominant tones in the spectra are symmetric in nature. The results from the cross-correlation suggest that two possible mechanisms of recirculation bubble pulsing and convective motions/vortex shedding are driving the dynamics on the base at subsonic Mach numbers. However, at supersonic Mach numbers, only single mechanism of the recirculation bubble pulsing dominates. Moreover, it indicates that the symmetric mode is associated with the dynamics of the recirculation bubble and the antisymmetric mode is related to the convective motions/vortex shedding.

Mitigation of Antisymmetric Mode and Drag with Base Cavities

2021 ◽  
Author(s):  
Viji M ◽  
Srinivas Nistala Vikramaditya
Keyword(s):  
Cavity Length ◽  
Pressure Fluctuations ◽  
Substantial Reduction ◽  
Base Pressure ◽  
Antisymmetric Mode ◽  
Azimuthal Mode ◽  
Total Drag ◽  
Base Cavity ◽  
Mach Numbers ◽  
The Impact

Abstract Experiments were carried out to examine the impact of base cavities on the base pressure fluctuations and total drag of a cylindrical afterbody for freestream Mach numbers 0.6-1.5. Significant improvement in the base pressure and a substantial reduction in the afterbody drag was noticed in the presence of a base cavity at subsonic Mach numbers. However, on increasing the cavity length beyond a certain value its performance deteriorates. At supersonic Mach numbers their effectiveness drops drastically. Tones in the spectra can be classified into two types depending on the dominant azimuthal mode which is either 0 or 1 and are referred to as symmetric and an antisymmetric mode, respectively. Spectra at subsonic Mach numbers exhibit tones which are related either to mode 0 or 1. However, at supersonic Mach numbers only tones related to mode 0 exist. The base cavity either, effectively suppress the antisymmetric mode or modify it into a symmetric mode resulting in mitigation of the tones related to antisymmetric mode.

Heat Transfer in Rotating Serpentine Coolant Passage With Ribbed Walls at Low Mach Numbers

2015 ◽  
Vol 7 (1) ◽  
Author(s):  
Shang-Feng Yang ◽  
Je-Chin Han ◽  
Salam Azad ◽  
Ching-Pang Lee
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Reynolds Numbers ◽  
Working Fluid ◽  
Transfer Coefficients ◽  
Low Mach Number ◽  
Heat Transfer Coefficients ◽  
Rotation Numbers ◽  
Wide Range ◽  
Mach Numbers

This paper experimentally investigates the effect of rotation on heat transfer in typical turbine blade serpentine coolant passage with ribbed walls at low Mach numbers. To achieve the low Mach number (around 0.01) condition, pressurized Freon R-134a vapor is utilized as the working fluid. The flow in the first passage is radial outward, after the 180 deg tip turn the flow is radial inward to the second passage, and after the 180 deg hub turn the flow is radial outward to the third passage. The effects of rotation on the heat transfer coefficients were investigated at rotation numbers up to 0.6 and Reynolds numbers from 30,000 to 70,000. Heat transfer coefficients were measured using the thermocouples-copper-plate-heater regional average method. Heat transfer results are obtained over a wide range of Reynolds numbers and rotation numbers. An increase in heat transfer rates due to rotation is observed in radially outward passes; a reduction in heat transfer rate is observed in the radially inward pass. Regional heat transfer coefficients are correlated with Reynolds numbers for nonrotation and with rotation numbers for rotating condition, respectively. The results can be useful for understanding real rotor blade coolant passage heat transfer under low Mach number, medium–high Reynolds number, and high rotation number conditions.

Cellular vortex shedding in the wake of a tapered plate

2008 ◽  
Vol 617 ◽  
pp. 355-379 ◽  
Author(s):  
VAGESH D. NARASIMHAMURTHY ◽  
HELGE I. ANDERSSON ◽  
BJØRNAR PETTERSEN
Keyword(s):  
Experimental Data ◽  
Reynolds Number ◽  
Vortex Shedding ◽  
High Reynolds Number ◽  
Three Dimensional ◽  
Base Pressure ◽  
Recirculation Bubble ◽  
Apparent Lack ◽  
Secondary Motion ◽  
High Reynolds

Direct numerical simulation (DNS) of vortex shedding behind a tapered plate with the taper ratio 20 placed normal to the inflow has been performed. The Reynolds numbers based on the uniform inflow velocity and the width of the plate at the wide and narrow ends were 1000 and 250, respectively. For the first time ever cellular vortex shedding was observed behind a tapered plate in a numerical experiment (DNS). Multiple cells of constant shedding frequency were found along the span of the plate. This is in contrast to apparent lack of cellular vortex shedding found in the high-Reynolds-number experiments by Gaster & Ponsford (Aero. J., vol. 88, 1984, p. 206). However, the present DNS data is in good qualitative agreement with similar high-Reynolds-number experimental data produced by Castro & Watson (Exp. Fluids, vol. 37, 2004, p. 159). It was observed that a tapered plate creates longer formation length coupled with higher base pressure as compared to non-tapered (i.e. uniform) plates. The three-dimensional recirculation bubble was nearly conical in shape. A significant base pressure reduction towards the narrow end of the plate, which results in a corresponding increase in Strouhal number, was noticed. This observation is consistent with the experimental data of Castro & Rogers (Exp. Fluids, vol. 33, 2002, p. 66). Pressure-driven spanwise secondary motion was observed, both in the front stagnation zone and also in the wake, thereby reflecting the three-dimensionality induced by the tapering.

The Influence of Reynolds Number, Mach Number and Incidence Effects on Loss Production in Low Pressure Turbine Airfoils

2006 ◽  
Author(s):  
Rau´l Va´zquez ◽  
Antonio Antoranz ◽  
David Cadrecha ◽  
Leyre Arman˜anzas
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Low Pressure ◽  
Low Pressure Turbine ◽  
Wide Range ◽  
Flow Features ◽  
Mach Numbers ◽  
Turbine Airfoils ◽  
Annular Cascade ◽  
The Impact

This paper presents an experimental study of the flow field in an annular cascade of Low Pressure Turbine airfoils. The influence of Reynolds number, Mach number and incidence on profile and end wall losses have been investigated. The annular cascade consisted of 100 high lift, high aspect ratio, high turning blades that are characteristic of modern LP Turbines. The investigation was carried out for a wide range of Reynolds numbers, extending from 120k to 315k, exit Mach numbers, from 0.5 to 0.9, and incidences from −20 to +14 degrees. Results clearly indicate a significant effect of incidence and Mach number in secondary loss production; however, the Reynolds number shows it much weaker impact. It has also been found that the profile loss production is strongly influenced by both Reynolds and Mach numbers, being the impact of the incidence weaker. Finally, measured data suggest that, in order to properly reproduce the performance of these types of airfoils, annular cascades can be required as far as linear cascades may miss some essential flow features.

scholarly journals Unsteady Pressures in a Transonic Fan Cascade Due to a Single Oscillating Airfoil

2002 ◽  
Author(s):  
J. Lepicovsky ◽  
E. R. McFarland ◽  
V. R. Capece ◽  
J. Hayden
Keyword(s):  
Computer Simulation ◽  
Mach Number ◽  
Pressure Fluctuations ◽  
Experimental Program ◽  
Influence Coefficient ◽  
Pressure Data ◽  
Data Set ◽  
Unsteady Pressure ◽  
Mach Numbers ◽  
Transonic Fan

An extensive set of unsteady pressure data was acquired along the midspan of a modern transonic fan blade for simulated flutter conditions. The data set was acquired in a nine-blade linear cascade with an oscillating middle blade to provide a database for the influence coefficient method to calculate instantaneous blade loadings. The cascade was set for an incidence of 10 dg. The data were acquired on three stationary blades on each side of the middle blade that was oscillated at an amplitude of 0.6 dg. The matrix of test conditions covered inlet Mach numbers of 0.5, 0.8, and 1.1 and the oscillation frequencies of 200, 300, 400, and 500 Hz. A simple quasi-unsteady two-dimensional computer simulation was developed to aid in the running of the experimental program. For high Mach number subsonic inlet flows the blade pressures exhibit very strong, low-frequency, self-induced oscillations even without forced blade oscillations, while for low subsonic and supersonic inlet Mach numbers the blade pressure unsteadiness is quite low. The amplitude of forced pressure fluctuations on neighboring stationary blades strongly depends on the inlet Mach number and forcing frequency. The flowfield behavior is believed to be governed by strong nonlinear effects, due to a combination of viscosity, compressibility, and unsteadiness. Therefore, the validity of the quasi-unsteady simplified computer simulation is limited to conditions when the flowfield is behaving in a linear, steady manner. Finally, an extensive set of unsteady pressure data was acquired to help development and verification of computer codes for blade flutter effects.

Aerodynamic Design and Testing of Three Low Solidity Steam Turbine Nozzle Cascades

2004 ◽  
Author(s):  
Bo Song ◽  
Wing F. Ng ◽  
Joseph A. Cotroneo ◽  
Douglas C. Hofer ◽  
Gunnar Siden
Keyword(s):  
Mach Number ◽  
Steam Turbine ◽  
Boundary Layer Transition ◽  
Experimental Results ◽  
Blade Surface ◽  
Layer Transition ◽  
Number Range ◽  
Baseline Design ◽  
Wide Range ◽  
Mach Numbers

Three sets of low solidity steam turbine nozzle cascades were designed and tested. The objective was to reduce cost through a reduction in parts count while maintaining or improving performance. The primary application is for steam turbine high pressure sections where Mach numbers are subsonic and high levels of unguided turning can be tolerated. The baseline Design A has a ratio of pitch to axial chord of 1.2. This is the pitch diameter section of a 50% reaction stage that has been verified by multistage testing on steam to have a high level of efficiency. Designs B and C have ratios of pitch to axial chord of 1.5 and 1.8 respectively. All three designs satisfy the same inlet and exit vector diagrams. Analytical surface Mach number distributions and boundary layer transition predictions are presented. Extensive cascade test measurements were carried out for a broad incidence range from −60 to +35 degrees. At each incidence, four outlet Mach numbers were tested, ranging from 0.2 to 0.8, with the corresponding Reynolds number variation from 1.8×105 to 9.0×105. Experimental results of loss coefficient and blade surface Mach number are presented and compared for the three cascades. The experimental results have demonstrated low losses over the tested Mach number range for a wide range of incidence from −45 to 15 degrees. Designs B and C have lower profile losses than Design A. The associated flow physics is interpreted using the results of wake profile, blade surface Mach number distribution and blade surface oil flow visualization, with the emphasis placed on the loss mechanisms for different flow conditions and the loss reduction mechanism with lower solidity. The effect of the higher profile loading of the lower solidity designs on increased end wall losses induced by increased secondary flow, especially on low aspect ratio designs, is the subject of ongoing studies.

Mach number effects on vortex shedding of a square cylinder and thick symmetrical airfoil arranged in tandem

1986 ◽  
Vol 407 (1833) ◽  
pp. 283-297 ◽  
Keyword(s):  
Mach Number ◽  
Shock Waves ◽  
Vortex Shedding ◽  
Pressure Fluctuations ◽  
Shear Layers ◽  
Side Length ◽  
Square Cylinder ◽  
Formation Region ◽  
Spacing Ratio ◽  
Vortex Shedding Frequency

An experimental study was done to elucidate the Mach number effects on vortex shedding of a square cylinder (side length D = 20 mm) and thick symmetrical airfoil (NACA 0018, chord length 20 mm) arranged in tandem , at free stream Mach numbers between 0.1526 and 0.9081, and at free stream Reynolds numbers (based on the side length D ) between 0.702 x 10 5 and 4.188 x 10 5 . The spacing ratio of the central distance, L , between the square cylinder and the airfoil to the side length, D , of the square cylinder was varied from 1.125 to 5.5. It was found that the regular vortex shedding is not suppressed by steady shock waves in the local supersonic flow regions; the periodic vortex shedding is still present, irrespective of the appearance of the shock waves. When the spacing ratio is fixed, the Strouhal number behind the square cylinder is almost constant up to the critical Mach number of about 0.70, but it increases rapidly with further increase of the Mach number. However, once the shock waves are formed on both sides of the vortex formation region, various frequency components, other than the vortex shedding frequency appear; the spectral peaks lower than those of the vortex shedding frequency were identified as frequencies of an acoustic-feedback oscillation and the resonance of the wind tunnel structural system. With increasing the Mach number, the formation region becomes small and asymmetric, and the separating shear layers become wavy. These changes result in an increase of the scale and strength of the vortices and thus enhance the vortex shedding process. However, when the Mach number exceeds the critical value, the streamwise length of the formation region increases suddenly and becomes long enough to enclose the downstream airfoil. Under this circumstance, the formation region is almost symmetrical with respect to the wake axis, and shock waves are formed on the upper and lower separating shear layers. The shock waves are almost normal to the wake axis at M = 0.7512 and 0.8215, but incline to the downstream direction at M = 0.9081. Acoustic waves travelling upstream have been observed most clearly when the vortex shed from the square cylinder hits the leading edge of the airfoil at a Mach number of about 0.63, which is close to, but slightly smaller than the critical value. The mean pressure and the amplitude of the pressure fluctuations in the test section decreases and increases, respectively, with increasing the Mach number. However, the amplitude of the pressure fluctuations decreases suddenly when the steady shock waves are formed on the upper and lower separating shear layers.

Heat Transfer in Rotating Serpentine Coolant Passage With Ribbed Walls at Low Mach Numbers

2013 ◽  
Author(s):  
Shang-Feng Yang ◽  
Je-Chin Han ◽  
Salam Azad ◽  
Ching-Pang Lee
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Reynolds Numbers ◽  
Working Fluid ◽  
Transfer Coefficients ◽  
Low Mach Number ◽  
Heat Transfer Coefficients ◽  
Rotation Numbers ◽  
Wide Range ◽  
Mach Numbers

This paper experimentally investigates the effect of rotation on heat transfer in typical turbine blade serpentine coolant passage with ribbed walls at low Mach numbers. To achieve the low Mach number (around 0.01) condition, pressurized Freon R-134a vapor is utilized as the working fluid. The flow in the first passage is radial outward, after the 180° tip turn the flow is radial inward to the second passage, and after the 180° hub turn the flow is radial outward to the third passage. The effects of rotation on the heat transfer coefficients were investigated at rotation numbers up to 0.6 and Reynolds numbers from 30,000 to 70,000. Heat transfer coefficients were measured using the thermocouples-copper-plate-heater regional average method. Heat transfer results are obtained over a wide range of Reynolds numbers and rotation numbers. An increase in heat transfer rates due to rotation is observed in radially outward passes; a reduction in heat transfer rate is observed in the radially inward pass. Regional heat transfer coefficients are correlated with Reynolds numbers for non-rotation and with rotation numbers for rotating condition, respectively. The results can be useful for understanding real rotor blade coolant passage heat transfer under low Mach number, medium-high Reynolds number and high rotation number conditions.

Aerodynamic Design and Testing of Three Low Solidity Steam Turbine Nozzle Cascades

2004 ◽  
Vol 129 (1) ◽  
pp. 62-71 ◽  
Author(s):  
Bo Song ◽  
Wing F. Ng ◽  
Joseph A. Cotroneo ◽  
Douglas C. Hofer ◽  
Gunnar Siden
Keyword(s):  
Mach Number ◽  
Steam Turbine ◽  
Boundary Layer Transition ◽  
Experimental Results ◽  
Base Line ◽  
Blade Surface ◽  
Number Range ◽  
Wide Range ◽  
Mach Numbers ◽  
Line Design

Three sets of low solidity steam turbine nozzle cascades were designed and tested. The objective was to reduce cost through a reduction in parts count while maintaining or improving performance. The primary application is for steam turbine high pressure sections where Mach numbers are subsonic and high levels of unguided turning can be tolerated. The base line design A has a ratio of pitch to axial chord of 1.2. This is the pitch diameter section of a 50% reaction stage that has been verified by multistage testing on steam to have a high level of efficiency. Designs B and C have ratios of pitch to axial chord of 1.5 and 1.8, respectively. All three designs satisfy the same inlet and exit vector diagrams. Analytical surface Mach number distributions and boundary layer transition predictions are presented. Extensive cascade test measurements were carried out for a broad incidence range from −60to+35deg. At each incidence, four outlet Mach numbers were tested, ranging from 0.2 to 0.8, with the corresponding Reynolds number variation from 1.8×105 to 9.0×105. Experimental results of loss coefficient and blade surface Mach number are presented and compared for the three cascades. The experimental results have demonstrated low losses over the tested Mach number range for a wide range of incidence from −45to15deg. Designs B and C have lower profile losses than design A. The associated flow physics is interpreted using the results of wake profile, blade surface Mach number distribution, and blade surface oil flow visualization, with the emphasis placed on the loss mechanisms for different flow conditions and the loss reduction mechanism with lower solidity. The effect of the higher profile loading of the lower solidity designs on increased end wall losses induced by increased secondary flow, especially on low aspect ratio designs, is the subject of ongoing studies.

Compressibility Effects in Turbulent Subsonic Jets

2003 ◽  
Author(s):  
Zhexuan Wang ◽  
Yiannis Andreopoulos
Keyword(s):  
Mach Number ◽  
Pressure Fluctuations ◽  
Turbulent Jets ◽  
Spreading Rate ◽  
Jet Flows ◽  
Mean Velocity ◽  
Pressure Transducers ◽  
High Frequency Response ◽  
Mach Numbers ◽  
Response Pressure

The behavior of compressible turbulent jets issuing in still air in the absence of shock waves has been investigated at three different subsonic Mach numbers, 0.3, 0.6 and 0.9. Helium, nitrogen and krypton gases were used to generate the jet flows and investigate the density effects on the structure of turbulence. Particle Image Velocimetry and high-frequency response pressure transducers were used to obtain velocity, Mach number inside the flow field. The decay of the Mach number at the centerline of the axisymmetric jets increases with increasing the initial Mach number at the exit of the flow for all jets. The decay of mean velocity at the centerline of the jets is also higher at higher exit Mach numbers. However, the velocity non-dimensionalized by the exit velocity seems to decrease faster at low exit Mach numbers suggesting a reduced mixing with increasing the exit flow Mach numbers. Helium jets were found to have the largest spreading rate among the three different gas jets used in the present investigation, while krypton had the lowest spreading rate. Total pressure fluctuations appear to decrease with increasing exit flow Mach numbers. Unusually high turbulence intensities were measured in helium jets issuing in still air.

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