Flutter in Started and Unstarted Transonic Linear Cascades: Simultaneous Measurements of Unsteady Loads and Shock Dynamics

2019 ◽  
Vol 141 (12) ◽  
Author(s):  
Prahallada Jutur ◽  
Raghuraman N. Govardhan
Keyword(s):  
High Speed ◽  
Leading Edge ◽  
Suction Side ◽  
Excitation Energies ◽  
Narrow Region ◽  
Simultaneous Measurements ◽  
Shock Phase ◽  
Shock Dynamics ◽  
Order Of Magnitude ◽  
Oscillation Frequencies

AbstractIn the present work, we report on an experimental study of bending/flap mode flutter of a blade within a linear cascade at transonic conditions. Driven by the motivation to understand the contribution of shock location/dynamics to flutter characteristics, we have performed simultaneous measurements of shock dynamics using high-speed shadowgraphy combined with unsteady load measurements on an oscillating blade within the cascade. The flutter characteristics in terms of energy transfer from the fluid to the blade and shock dynamics have been mapped out over a range of blade oscillation frequencies and static pressure ratios (SPRs) across the cascade, the latter being important as they decide the mean location of the passage shocks. SPR values studied include both conditions where the shock is within the passage (started cascade) and where the passage shock is pushed ahead of the leading edge of the blades (unstarted cascade). These measurements show characteristically different flutter behavior for an unstarted cascade compared with a started cascade, the former having received very little attention in the literature. While both these cases show small excitation levels at low reduced frequencies, the unstarted cascade case exhibits an additional relative narrow region of excitation at higher reduced frequencies with approximately an order of magnitude higher excitation energies. Comparison of the shock dynamics between the two excitation regimes shows significant differences in the phase of the leading edge shock in addition to changes in the suction side shock phase indicating that the two excitation regimes are of different origin.

Cooling of Turbine Blades With Expanded Exit Holes: Computational Analyses of Leading Edge and Pressure-Side of a Turbine Blade

2017 ◽  
Vol 139 (4) ◽  
Author(s):  
Fariborz Forghan ◽  
Omid Askari ◽  
Uichiro Narusawa ◽  
Hameed Metghalchi
Keyword(s):  
High Speed ◽  
Turbine Blades ◽  
Leading Edge ◽  
Weak Shock ◽  
Suction Side ◽  
Pressure Side ◽  
Blade Surface ◽  
Cooling Effectiveness ◽  
Hot Gas ◽  
Computational Analyses

Turbine blades are cooled by a jet flow from expanded exit holes (EEH) forming a low-temperature film over the blade surface. Subsequent to our report on the suction-side (low-pressure, high-speed region), computational analyses are performed to examine the cooling effectiveness of the flow from EEH located at the leading edge as well as at the pressure-side (high-pressure, low-speed region). Unlike the case of the suction-side, the flow through EEH on the pressure-side is either subsonic or transonic with a weak shock front. The cooling effectiveness, η (defined as the temperature difference between the hot gas and the blade surface as a fraction of that between the hot gas and the cooling jet), is higher than the suction-side along the surface near the exit of EEH. However, its magnitude declines sharply with an increase in the distance from EEH. Significant effects on the magnitude of η are observed and discussed in detail of (1) the coolant mass flow rate (0.001, 0.002, and 0.004 (kg/s)), (2) EEH configurations at the leading edge (vertical EEH at the stagnation point, 50 deg into the leading-edge suction-side, and 50 deg into the leading-edge pressure-side), (3) EEH configurations in the midregion of the pressure-side (90 deg (perpendicular to the mainstream flow), 30 deg EEH tilt toward upstream, and 30 deg tilt toward downstream), and (4) the inclination angle of EEH.

Visualizations of Flow Structures in the Rotor Passage of an Axial Compressor at the Onset of Stall

2017 ◽  
Vol 139 (4) ◽  
Author(s):  
Huang Chen ◽  
Yuanchao Li ◽  
David Tan ◽  
Joseph Katz
Keyword(s):  
High Speed ◽  
Large Scale ◽  
Incidence Angle ◽  
Axial Compressor ◽  
Leading Edge ◽  
Suction Side ◽  
Stereoscopic Particle Image Velocimetry ◽  
High Speed Imaging ◽  
Vortical Structures ◽  
Flow Phenomena

Experiments preformed in the JHU refractive index matched facility examine flow phenomena developing in the rotor passage of an axial compressor at the onset of stall. High-speed imaging of cavitation performed at low pressures qualitatively visualizes vortical structures. Stereoscopic particle image velocimetry (SPIV) measurements provide detailed snapshots and ensemble statistics of the flow in a series of meridional planes. At prestall condition, the tip leakage vortex (TLV) breaks up into widely distributed intermittent vortical structures shortly after rollup. The most prominent instability involves periodic formation of large-scale backflow vortices (BFVs) that extend diagonally upstream, from the suction side (SS) of one blade at midchord to the pressure side (PS) near the leading edge of the next blade. The 3D vorticity distributions obtained from data recorded in closely spaced planes show that the BFVs originate form at the transition between the high circumferential velocity region below the TLV center and the main passage flow radially inward from it. When the BFVs penetrate to the next passage across the tip gap or by circumventing the leading edge, they trigger a similar phenomenon there, sustaining the process. Further reduction in flow rate into the stall range increases the number and size of the backflow vortices, and they regularly propagate upstream of the leading edge of the next blade, where they increase the incidence angle in the tip corner. As this process proliferates circumferentially, the BFVs rotate with the blades, indicating that there is very little through flow across the tip region.

Effect of Hole Diameter on Nozzle Endwall Film Cooling and Associated Phantom Cooling

2015 ◽  
Author(s):  
Luzeng Zhang ◽  
Juan Yin ◽  
Kevin Liu ◽  
Moon Hee-Koo
Keyword(s):  
Film Cooling ◽  
High Speed ◽  
Leading Edge ◽  
Suction Side ◽  
Hole Diameter ◽  
Two Dimensional ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Blowing Ratio

Flow fields near the turbine nozzle endwall are highly complex due to the passage vortices and endwall cross flows. Consequently, it is challenging to provide proper cooling to the endwall surfaces. An effective way to cool the endwall is to have film cooling holes forward of the leading edge, often called “inlet-film cooling”. This paper presents the results of an experimental investigation on how the film hole diameter affects the film effectiveness on nozzle endwall and associated phantom cooling effectiveness on airfoil suction side. The measurements were conducted in a high speed linear cascade, which consists of three nozzle vanes and four flow passages. Double staggered rows of film injections, which were located upstream from the nozzle leading edge, provided cooling to the contoured endwall surfaces. Film cooling effectiveness on the endwall surface and corresponding phantom cooling effectiveness on the airfoil suction side were measured separately with a Pressure Sensitive Paint (PSP) technique through the mass transfer analogy. Four different film hole diameters with the same injection angle and the same pitch to diameter ratio were studied for up to six different MFR’s (mass flow ratios). Two dimensional film effectiveness distributions on the endwall surface and two dimensional phantom cooling distributions on the airfoil suction side are presented. Film/phantom cooling effectiveness distributions are pitchwise/spanwise averaged along the axial direction and also presented. The results indicate that both the endwall film effectiveness and the suction side phantom cooling effectiveness increases with the hole diameter (as decreases in blowing ratio for a given MFR) up to a specific diameter, then starts decreasing. An optimal value of the film hole diameter (blowing ratio) for the given injection angle is also suggested based on current study.

Parametric Study of the Pressure Stability on an Oscillating Airfoil From Stable to Stalled Flow Conditions

2000 ◽  
Author(s):  
S. Svensdotter ◽  
U. Johansson ◽  
T. Fransson
Keyword(s):  
Leading Edge ◽  
Suction Side ◽  
Amplitude Dependence ◽  
Phase Lag ◽  
Suction Surface ◽  
Constant Frequency ◽  
Edge Vortex ◽  
Pressure Stability ◽  
Stall Angle ◽  
Oscillation Frequencies

This paper presents investigations on the suction side of an oscillation blade at an inlet Mach number of 0.5, with oscillation frequencies from 60 to 210Hz, mean incidence angles from 0° to 10° and with blade amplitudes up to 4.2° (for 60Hz). The time-dependent pressures show that the aerodynamic loads are not linear for the stall region. It was found that an increased mean incidence caused a higher excitation of the blade from the aerodynamic forces, while the effects of varying the blade amplitudes and frequencies were more difficult to determine. At 6° mean incidence, and below, increased blade amplitude for a constant frequency caused the blade suction surface to experience an increased excitation, while for the 7° incidence, and above, the opposite was shown. The fact that the blade amplitude dependence changes when the mean incidence increases above the steady state stall angle is explained by the increased portion of the oscillation cycle spent at stalled conditions. A shift in phase was seen for the highest investigated frequency, which might be due to a phase lag between the blade motion and the leading edge vortex, increasing with increasing frequency.

Effect of Roughness and Unsteadiness on the Performance of a New Low Pressure Turbine Blade at Low Reynolds Numbers

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Francesco Montomoli ◽  
Howard Hodson ◽  
Frank Haselbach
Keyword(s):  
Reynolds Number ◽  
Wind Tunnel ◽  
Beneficial Effect ◽  
High Speed ◽  
Leading Edge ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
High Lift ◽  
Wake Passing

This paper presents a study of the performance of a high-lift profile for low pressure turbines at Reynolds numbers lower than in previous investigations. By following the results of Coull et al. (2008, “Velocity Distributions for Low Pressure Turbines,” ASME Paper No. GT2008-50589) on the design of high-lift airfoils, the profile is forward loaded. The separate and combined effects of roughness and wake passing are compared. On a front loaded blade, the effect of incidence becomes more important and the consequences in terms of cascade losses, is evaluated. The experimental investigation was carried out in the high speed wind tunnel of Whittle Laboratory, University of Cambridge. This is a closed-circuit continuous wind tunnel where the Reynolds number and Mach number can be fixed independently. The unsteadiness caused by wake passing in front of the blades is reproduced using a wake generator with rotating bars. The results confirm that the beneficial effect of unsteadiness on losses is present even at the lowest Reynolds number examined (Re3=20,000). This beneficial effect is reduced at positive incidence. With a front loaded airfoil and positive incidence, the transition occurs on the suction side close to the leading edge and this results in higher losses. This has been found valid for the entire Reynolds range investigated (20,000≤Re3≤140,000). Roughening the surface also had a beneficial effect on the losses but this effect vanishes at the lower Reynolds numbers, i.e., (Re3≤30,000), where the surface becomes hydraulically smooth. The present study suggests that a blade with as-cast surface roughness has a lower loss than a polished one.

The Effect of an Eroded Leading Edge on the Aerodynamic Performance of a Transonic Fan Blade Cascade

2014 ◽  
Author(s):  
A. Hergt ◽  
J. Klinner ◽  
W. Steinert ◽  
S. Grund ◽  
M. Beversdorff ◽  
...  
Keyword(s):  
High Speed ◽  
Numerical Study ◽  
Separation Bubble ◽  
Leading Edge ◽  
Suction Side ◽  
Chord Length ◽  
Blade Cascade ◽  
Shock Position ◽  
Fan Blade ◽  
Blunt Leading Edge

Especially at transonic flow conditions the leading edge shape influences the performance of a fan profile. At the same time the leading edge of a fan profile is highly affected by erosion during operation. This erosion leads to a deformation of the leading edge shape and a reduction of the chord length. In the present experimental and numerical study, the aerodynamic perfomance of an original fan profile geometry is compared to an eroded fan profile with a blunt leading edge and a chord length reduced by about 1 percent. The experiments are performed at a linear fan blade cascade in the Transonic Cascade Wind Tunnel of DLR in Cologne. The inflow Mach number during the tests is 1.25 and the Reynolds number 1.5 × 106. All tests are carried out at a low inflow turbulence level of 0.8 percent. The results of the investigation show that losses are increased over the whole operating range of the cascade. At the aerodynamic design point the losses raise by 25 percent. This significant loss increase can be traced back to the increase of the shock losses at the leading edge. The change in shock structure is investigated and described in detail by means of PIV measurements and Schlieren imaging. Additionally, the unsteady fluctuation of the shock position is measured by a high speed shadowgraphy. Then the frequency range of the fluctuation is obtained by a Fourier analysis of the time resolved shock position. Furthermore, liquid crystal measurements are performed in order to analyze the influence of the leading edge shape on the development of the suction side boundary layer. The results show that for the original fan blade the transition occurs at the shock position on the blade suction side by a separation bubble whereas the transition onset is shifted upstream for the fan blade with the blunt leading edge.

The Effect of an Eroded Leading Edge on the Aerodynamic Performance of a Transonic Fan Blade Cascade

2014 ◽  
Vol 137 (2) ◽  
Author(s):  
Alexander Hergt ◽  
J. Klinner ◽  
W. Steinert ◽  
S. Grund ◽  
M. Beversdorff ◽  
...  
Keyword(s):  
High Speed ◽  
Numerical Study ◽  
Separation Bubble ◽  
Leading Edge ◽  
Suction Side ◽  
Aerodynamic Performance ◽  
Chord Length ◽  
Blade Cascade ◽  
Shock Position ◽  
Fan Blade

Especially at transonic flow conditions the leading edge shape influences the performance of a fan profile. At the same time the leading edge of a fan profile is highly affected by erosion during operation. This erosion leads to a deformation of the leading edge shape and a reduction of the chord length. In the present experimental and numerical study, the aerodynamic performance of an original fan profile geometry is compared to an eroded fan profile with a blunt leading edge (BLE) and a chord length reduced by about 1%. The experiments are performed at a linear fan blade cascade in the Transonic Cascade Wind Tunnel of DLR in Cologne. The inflow Mach number during the tests is 1.25 and the Reynolds number 1.5 × 106. All tests are carried out at a low inflow turbulence level of 0.8%. The results of the investigation show that losses are increased over the whole operating range of the cascade. At the aerodynamic design point (ADP) the losses raise by 25%. This significant loss increase can be traced back to the increase of the shock losses at the leading edge. The change in shock structure is investigated and described in detail by means of particle image velocimetry (PIV) measurements and Schlieren imaging. Additionally, the unsteady fluctuation of the shock position is measured by a high-speed shadowgraphy. Then the frequency range of the fluctuation is obtained by a Fourier analysis of the time resolved shock position. Furthermore, liquid crystal measurements are performed in order to analyze the influence of the leading edge shape on the development of the suction side boundary layer. The results show that for the original fan blade the transition occurs at the shock position on the blade suction side by a separation bubble whereas the transition onset is shifted upstream for the fan blade with the BLE.

Visualizations of Flow Structures in the Rotor Passage of an Axial Compressor at the Onset of Stall

2016 ◽  
Author(s):  
Huang Chen ◽  
Yuanchao Li ◽  
David Tan ◽  
Joseph Katz
Keyword(s):  
High Speed ◽  
Large Scale ◽  
Axial Compressor ◽  
Leading Edge ◽  
Suction Side ◽  
Flow Structures ◽  
Tip Leakage Flow ◽  
High Speed Imaging ◽  
Vortical Structures ◽  
Tip Leakage

Flow visualizations and stereoscopic PIV (SPIV) measurements are carried out to study the flow phenomena developing in the rotor passage of an axial compressor at the onset of stall. Experiments have been performed in the JHU optically index-matched facility, using acrylic blades and liquid that have the same optical refractive index. The blade geometries are based on the first one and a half stages of the Low Speed Axial Compressor (LSAC) facility at NASA Glenn. The SPIV measurements provide detailed snapshots and ensemble statistics on the flow in a series of meridional planes. Data recorded in closely spaced planes enable us to obtain ensemble averaged 3D vorticity distributions. High speed imaging of cavitation, performed at low pressure, is used to qualitatively visualize the vortical structures within the rotor passage. The observations are performed just above and at stall conditions. At pre-stall condition, shortly after it rolled up, the tip leakage vortex (TLV) breaks up into widely distributed intermittent vortical structures. In particular, interaction of the backward tip leakage flow with the nearly opposite direction main passage flow under (radially inward) it results in periodic generation of large scale vortices that extend upstream, from the suction side (SS) of one blade to the pressure side (PS) or even near the leading edge of the next blade. When these structures penetrate to the next passage, they trigger formation of a similar phenomenon there, initiating a process that sustains itself. Once they form, these vortices rotate with the blade, indicating little through flow in the tip region. The 3D velocity and vorticity distributions confirm the presence of these large flow structures at the transition between the high circumferential velocity region below the TLV center and the main flow deeper in the passage. Further reduction in flow rate into the stall range caused a rapid increase in the number and scale of these vortices, demonstrating that their formation and proliferation plays a key role in the onset of stall.

scholarly journals Investigation of Separation Phenomena in a Radial Pump at Reduced Flow Rate by Large-Eddy Simulation

2016 ◽  
Vol 138 (12) ◽  
Author(s):  
Antonio Posa ◽  
Antonio Lippolis ◽  
Elias Balaras
Keyword(s):  
Reverse Flow ◽  
Leading Edge ◽  
Suction Side ◽  
Secondary Flows ◽  
Vaned Diffuser ◽  
Flow Rates ◽  
Turbulent Statistics ◽  
Large Eddy ◽  
Radial Pump ◽  
Reduced Flow

Turbopumps operating at reduced flow rates experience significant separation and backflow phenomena. Although Reynolds-Averaged Navier–Stokes (RANS) approaches proved to be usually able to capture the main flow features at design working conditions, previous numerical studies in the literature verified that eddy-resolving techniques are required in order to simulate the strong secondary flows generated at reduced loads. Here, highly resolved large-eddy simulations (LES) of a radial pump with a vaned diffuser are reported. The results are compared to particle image velocimetry (PIV) experiments in the literature. The main focus of the present work is to investigate the separation and backflow phenomena occurring at reduced flow rates. Our results indicate that the effect of these phenomena extends up to the impeller inflow: they involve the outer radii of the impeller vanes, influencing significantly the turbulent statistics of the flow. Also in the diffuser vanes, a strong spanwise evolution of the flow has been observed at the reduced load, with reverse flow, located mainly on the shroud side and on the suction side (SS) of the stationary channels, especially near the leading edge of the diffuser blades.

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