Experimental Investigation of Pressure Side Flow Separation on the T106C Airfoil at High Suction Side Incidence Flow

2017 ◽  
Vol 139 (5) ◽  
Author(s):  
Stephan Stotz ◽  
Yavuz Guendogdu ◽  
Reinhard Niehuis
Keyword(s):  
Mach Number ◽  
Wind Tunnel ◽  
Separation Bubble ◽  
Suction Side ◽  
Pressure Side ◽  
Blade Passage ◽  
Wide Range ◽  
Profile Losses ◽  
Exit Mach Number ◽  
Side Flow

The objective of this work is to study the influence of a pressure side separation bubble on the profile losses and the development of the bubble in the blade passage. For the experimental investigations, the T106 profile is used, with an increased loading due to an enlarged pitch to chord ratio from 0.799 to 0.95 (T106C). The experiments were performed at the high-speed cascade wind tunnel of the Institute of Jet Propulsion at the University of the Federal Armed Forces Munich. The main feature of the wind tunnel is to vary Reynolds and Mach number independently to achieve realistic turbomachinery conditions. The focus of this work is to determine the influence of a pressure side separation on the profile losses and hence the robustness to suction side incidence flow. The cascade is tested at four incidence angles from 0 deg to −22.7 deg to create separation bubbles of different sizes. The influence of the Reynolds number is investigated for a wide range at constant exit Mach number. Therefore, a typical exit Mach number for low pressure turbines in the range of 0.5–0.8 is chosen in order to consider compressible effects. Furthermore, two inlet turbulence levels of about 3% and 7.5% have been considered. The characteristics of the separation bubble are identified by using the profile pressure distributions, whereas wake traverses with a five hole probe are used to determine the influence of the pressure side separation on the profile losses. Further, time-resolved pressure measurements near the trailing edge as well as single hot wire measurements in the blade passage are conducted to investigate the unsteady behavior of the pressure side separation process itself and also its influence on the midspan passage flow.

Experimental Investigation of Pressure Side Flow Separation on the T106C Airfoil at High Suction Side Incidence Flow

2016 ◽  
Author(s):  
Stephan Stotz ◽  
Reinhard Niehuis ◽  
Yavuz Guendogdu
Keyword(s):  
Mach Number ◽  
Wind Tunnel ◽  
Separation Bubble ◽  
Suction Side ◽  
Pressure Side ◽  
Blade Passage ◽  
Wide Range ◽  
Profile Losses ◽  
Exit Mach Number ◽  
Side Flow

The objective of this work is to study the influence of a pressure side separation bubble on the profile losses and the development of the bubble in the blade passage. For the experimental investigations the T106 profile is used, with an increased loading due to an enlarged pitch to chord ratio from 0.799 to 0.95 (T106C). The experiments were performed at the high-speed cascade wind tunnel of the Institute of Jet Propulsion at the University of the Federal Armed Forces Munich. The main feature of the wind tunnel is to vary Reynolds and Mach number independently to achieve realistic turbomachinery conditions. The focus of this work is to determine the influence of a pressure side separation on the profile losses and hence the robustness to suction side incidence flow. The cascade is tested at four incidence angles from 0° to −22.7° to create separation bubbles of different sizes. The influence of the Reynolds number is investigated for a wide range at constant exit Mach number. Therefore a typical exit Mach number for low pressure turbines in the range of 0.5–0.8 is chosen in order to consider compressible effects. Furthermore, two inlet turbulence levels of about 3% and 7.5% have been considered. The characteristics of the separation bubble are identified by using the profile pressure distributions, whereas wake traverses with a five hole probe are used to determine the influence of the pressure side separation on the profile losses. Further, time-resolved pressure measurements near the trailing edge as well as single hot wire measurements in the blade passage are conducted to investigate the unsteady behavior of the pressure side separation process itself and also its influence on the midspan passage flow.

scholarly journals Effects of Tip Clearance Gap and Exit Mach Number on Turbine Blade Tip and Near-Tip Heat Transfer

2013 ◽  
Author(s):  
K. Anto ◽  
S. Xue ◽  
W. F. Ng ◽  
L. J. Zhang ◽  
H. K. Moon
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Turbine Blade ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Side ◽  
Engine Blade ◽  
Blade Tip ◽  
Exit Mach Number

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet. Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.

The Effects of Inlet Reynolds Number, Exit Mach Number and Incidence Angle on Leading Edge Film Cooling Effectiveness of a Turbine Blade in a Linear Transonic Cascade

2015 ◽  
Author(s):  
Cong Liu ◽  
Hui-ren Zhu ◽  
Zhong-yi Fu ◽  
Run-hong Xu
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Film Cooling ◽  
Incidence Angle ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Exit Mach Number

This paper experimentally investigates the film cooling performance of a leading edge with three rows of film holes on an enlarged turbine blade in a linear cascade. The effects of blowing ratio, inlet Reynolds number, isentropic exit Mach number and off-design incidence angle (i<0°) are considered. Experiments were conducted in a short-duration transonic wind tunnel which can model realistic engine aerodynamic conditions and adjust inlet Reynolds number and exit Mach number independently. The surface film cooling measurements were made at the midspan of the blade using thermocouples based on transient heat transfer measurement method. The changing of blowing ratio from 1.7 to 3.3 leads to film cooling effectiveness increasing on both pressure side and suction side. The Mach number or Reynolds number has no effect on the film cooling effectiveness on pressure side nearly, while increasing these two factors has opposite effect on film cooling performance on suction side. The increasing Mach number decreases the film cooling effectiveness at the rear region mainly, while at higher Reynolds number condition, the whole suction surface has significantly higher film cooling effectiveness because of the increasing cooling air mass flow rate. When changing the incidence angle from −15° to 0°, the film cooling effectiveness of pressure side decreases, and it presents the opposite trend on suction side. At off-design incidence of −15° and −10°, there is a low peak following the leading edge on the pressure side caused by the separation bubble, but it disappears with the incidence and blowing ratio increased.

The Effect of Pulsed Blowing on the Boundary Layer of a Highly Loaded Low Pressure Turbine Blade

2013 ◽  
Author(s):  
Marion Mack ◽  
Roland Brachmanski ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Mach Number ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Trailing Edge ◽  
Low Pressure Turbine ◽  
Wide Range ◽  
Highly Loaded

The performance of the low pressure turbine (LPT) can vary appreciably, because this component operates under a wide range of Reynolds numbers. At higher Reynolds numbers, mid and aft loaded profiles have the advantage that transition of suction side boundary layer happens further downstream than at front loaded profiles, resulting in lower profile loss. At lower Reynolds numbers, aft loading of the blade can mean that if a suction side separation exists, it may remain open up to the trailing edge. This is especially the case when blade lift is increased via increased pitch to chord ratio. There is a trend in research towards exploring the effect of coupling boundary layer control with highly loaded turbine blades, in order to maximize performance over the full relevant Reynolds number range. In an earlier work, pulsed blowing with fluidic oscillators was shown to be effective in reducing the extent of the separated flow region and to significantly decrease the profile losses caused by separation over a wide range of Reynolds numbers. These experiments were carried out in the High-Speed Cascade Wind Tunnel of the German Federal Armed Forces University Munich, Germany, which allows to capture the effects of pulsed blowing at engine relevant conditions. The assumed control mechanism was the triggering of boundary layer transition by excitation of the Tollmien-Schlichting waves. The current work aims to gain further insight into the effects of pulsed blowing. It investigates the effect of a highly efficient configuration of pulsed blowing at a frequency of 9.5 kHz on the boundary layer at a Reynolds number of 70000 and exit Mach number of 0.6. The boundary layer profiles were measured at five positions between peak Mach number and the trailing edge with hot wire anemometry and pneumatic probes. Experiments were conducted with and without actuation under steady as well as periodically unsteady inflow conditions. The results show the development of the boundary layer and its interaction with incoming wakes. It is shown that pulsed blowing accelerates transition over the separation bubble and drastically reduces the boundary layer thickness.

scholarly journals LDA Study of the Flow Development Through an Orthogonally Rotating U-Bend of Strong Curvature and Rib Roughened Walls

1996 ◽  
Author(s):  
H. Iacovides ◽  
D. C. Jackson ◽  
H. Ji ◽  
G. Kelemenis ◽  
B. E. Launder ◽  
...  
Keyword(s):  
Laser Doppler Anemometry ◽  
Separation Bubble ◽  
Turbine Blades ◽  
Symmetry Plane ◽  
Suction Side ◽  
Outer Wall ◽  
Outer Side ◽  
Pressure Side ◽  
Mean Flow ◽  
Flow Development

This paper reports laser Doppler anemometry (LDA) and wall pressure measurements of turbulent flow in a square-sectioned, rotating U-bend typical of coolant passages employed in modern gas turbine blades. In the upstream and downstream tangents, the pressure and suction (inner and outer) surfaces are roughened with discrete square-sectioned ribs in a staggered arrangement for a rib-height to duct-diameter ratio of 0.1. Three cases have been examined at a passage Reynolds number of 105: a stationary case; a case of positive rotation (the pressure side coinciding with the outer side of the U-bend) at a rotation number (Ro=ΩD/Um) of 0.2; and a case of negative rotation at Ro=−0.2. Measurements have been obtained along the symmetry plane of the duct. In the upstream section, the separation bubble behind each rib is about 2.5 rib-heights long. Rotation displaces the high momentum fluid towards the pressure side, enhances turbulence along the pressure side and suppresses turbulence along the suction side. The introduction of ribs in the straight sections reduces the size of the separation bubble along the inner wall of the U-bend, by raising turbulence levels at the bend entry; it also causes the formation of an additional separation bubble over the first rib interval along the outer wall, downstream of the bend exit. Rotation also modifies the mean flow development within the U-bend, with negative rotation speeding up the flow along the inner wall and causing a wider inner-wall separation bubble at exit. Turbulence levels within the bend are generally increased by rotation and, over the first two diameters downstream of the bend, negative rotation increases turbulence while positive rotation on the whole has the opposite effect.

Comparative Investigation of Three Highly Loaded LP Turbine Airfoils: Part II — Measured Profile and Secondary Losses at Off-Design Incidence

2007 ◽  
Author(s):  
T. Zoric ◽  
I. Popovic ◽  
S. A. Sjolander ◽  
T. Praisner ◽  
E. Grover
Keyword(s):  
Separation Bubble ◽  
Suction Side ◽  
Comparative Investigation ◽  
Secondary Flows ◽  
The Other ◽  
Pressure Probe ◽  
Low Pressure Turbine ◽  
Profile Losses ◽  
Lp Turbine ◽  
Turbine Airfoils

The first part of the paper compared the midspan aerodynamics and the secondary flows for a family of three low-pressure turbine (LPT) airfoils at design conditions. However, since a typical engine spends much of its time operating at off-design conditions, good tolerance of LPT airfoils to off-design operation is desired. The sensitivity of the midspan flow to Reynolds number was examined for the three airfoils in a paper presented at the 2006 ASME-IGTI Turbo-Expo. The present paper examines the performance of the airfoils for three values of incidence: −5, 0, and +5 degrees relative to design. Both the profile and secondary losses are considered. Detailed loading distributions measured at midspan are used to explain the behaviour of the profile flow and the resulting change in losses as the incidence was varied. The secondary flow behaviour is determined as at the design incidence from detailed flowfield measurements made downstream of the trailing edge using a seven-hole pressure probe. The results show that in terms of profile losses the baseline airfoil (which has a Zweifel coefficient Z = 1.08) and the front-loaded one with Z = 1.37 have comparable losses over the range of incidences examined. However, the aft-loaded airfoil with Z = 1.37 had noticeably higher profile losses than the other two. On the other hand, the front-loaded one has higher secondary losses than its aft-loaded counterpart at all conditions examined. This obviously poses a dilemma for the designer in terms of the choice of loading distribution. It was also noted that the distribution of loading seems to affect the secondary losses more than the loading level (Zweifel coefficient). An interaction of the secondary flows with the suction side separation bubble might be responsible in part for this finding.

Profile Loss of Ultra-Highly Loaded Turbine Cascade at Transonic Flow Condition

2019 ◽  
Author(s):  
Hoshio Tsujita ◽  
Masanao Kaneko
Keyword(s):  
Mach Number ◽  
Transonic Flow ◽  
Compressible Flows ◽  
Detailed Examination ◽  
Aerodynamic Performance ◽  
Turbine Cascade ◽  
Wide Range ◽  
Exit Mach Number ◽  
Profile Loss ◽  
Highly Loaded

Abstract The aerodynamic performance of turbine components constituting the gas turbine engine is seriously required to be improved in order to reduce environmental load. The energy recovery efficiency in turbine component can be enhanced by the increase of turbine blade loading. In this study, as the first stage to investigate the aerodynamic performance of an ultra-highly loaded turbine cascade (UHLTC) with a turning angle of 160 degrees at transonic flow regime, two-dimensional steady compressible flows in UHLTC were analyzed numerically by using a commercial CFD code to focus on the profile loss. In the computations, the isentropic exit Mach number was varied in the wide range from 0.3 to 1.8 in order to examine the effects of exit Mach number on the shock wave formation and the associated profile loss generation. The computed results were examined in detail by comparing with those for a typical transonic turbine cascade. The detailed examination for the present computed results clarified the variation of shock pattern with the increase of exit Mach number and the loss “plateau” behavior in the present UHLTC.

Aerodynamic Losses of a Cambered Turbine Vane: Influences of Surface Roughness and Freestream Turbulence Intensity

2005 ◽  
Author(s):  
Qiang Zhang ◽  
Phillip M. Ligrani
Keyword(s):  
Surface Roughness ◽  
Mach Number ◽  
Turbulence Intensity ◽  
Suction Side ◽  
Freestream Turbulence ◽  
Smooth Test ◽  
Profile Losses ◽  
Turbine Vane ◽  
Mach Number Distribution ◽  
Aerodynamic Losses

The effects of surface roughness and freestream turbulence level on the aerodynamic performance of a turbine vane are experimentally investigated. Wake profiles are measured with three different freestream turbulence intensity levels (1.1%, 5.4% and 7.7%) at two different locations downstream of the test vane trailing edge (one and 0.25 axial chord lengths). Chord Reynolds number based on exit flow conditions is 0.9 × 106. The Mach number distribution and the test vane configuration both match arrangements employed in an industrial application. Four cambered vanes with different surface roughness levels are employed in this study. Effects of surface roughness on the vane pressure side on the profile losses are relatively small compared with suction side roughness. Overall effects of turbulence on local wake deficits of total pressure, Mach number, and kinetic energy are almost negligible in most parts of the wake produced by the smooth test vane, except that higher freestream losses are present at higher turbulence intensity levels. Profiles produced by test vanes with rough surfaces show apparent lower peak values in the center of the wake. Integrated Aerodynamic Losses (IAL) and area-averaged loss coefficient YA are also presented and compared with results from other research groups.

Aerodynamic Performance of Suction-Side Gill Region Film Cooling

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas ◽  
Edward Vlasic
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Exit Mach Number ◽  
Total Pressure Loss Coefficient

The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77–1.99 similar to values present in operating gas turbine engines. Presented are the local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated aerodynamic losses (IAL) increase anywhere from 4% to 45% compared with a smooth blade with no film injection. The performance of each hole type depends on the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL are generally present at any particular wake location for the RR or SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

Numerical investigation of the tone noise mechanism over laminar airfoils

2007 ◽  
Vol 591 ◽  
pp. 155-182 ◽  
Author(s):  
G. DESQUESNES ◽  
M. TERRACOL ◽  
P. SAGAUT
Keyword(s):  
Numerical Simulation ◽  
Boundary Layer ◽  
Numerical Investigation ◽  
Separation Bubble ◽  
Experimental Studies ◽  
Suction Side ◽  
The Other ◽  
Pressure Side ◽  
Tone Frequency ◽  
Noise Phenomenon

This paper presents the first numerical investigation via direct numerical simulation of the tone noise phenomenon occurring in the flow past laminar airfoils. This phenomenon corresponds to the radiation of discrete acoustic tones in some specific flow conditions, and has received much attention since the 1970s, and several experimental studies have been carried out to identify and understand the underlying physical mechanisms. However, several points remain to be clarified in order to provide a complete explanation of its origin. The flow around a two-dimensional NACA0012 airfoil is considered in order to have a deeper understanding of the tone noise phenomenon. Consistently with previous experimental studies, it is shown that depending on the Reynolds number and angle of attack, two different types of acoustic spectrum are observed: one which exhibits a broadband contribution with a dominant frequency together with a sequence of regularly spaced discrete frequencies, while the other one is only characterized by a simple broadband contribution. The first configuration is typical of the tone noise phenomenon. The present work shows that in this case, the mean flow on the pressure side of the airfoil exhibits a separation bubble near the trailing edge and the main tone frequency is close to the most amplified frequency of the boundary layer. The mechanism proposed in previous works for the main tone generation – which implies the existence of a separation bubble at the pressure side – is therefore validated by numerical simulation. On the other hand, the analysis of the suction side boundary layer reveals that there is no separation and that the most amplified frequency is different from the main tonal one. However, the suction side boundary layer is highly receptive to the tone frequency. Finally, an original explanation for the existence of the secondary discrete frequencies observed in the radiated pressure spectrum is given. They are associated to a bifurcation of the airfoil wake from a symmetric to a non-symmetric vortex pattern. A possible explanation for the existence of this bifurcation is the interaction between the disturbances which are the most amplified by the suction side boundary layer and those originating in the forcing of the suction side flow by the main tone noise mechanism.

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