scholarly journals Unsteady pressure measurement and numerical simulations in an end-wall region of a linear blade cascade

2021 ◽  
Vol 3 (8) ◽  
Author(s):  
Erik Flídr ◽  
Petr Straka ◽  
Milan Kladrubský ◽  
Tomáš Jelínek
Keyword(s):  
Numerical Simulations ◽  
Pressure Measurement ◽  
Power Spectra ◽  
Surface Flow ◽  
Wall Region ◽  
Suction Surface ◽  
Flow Angle ◽  
Unsteady Pressure ◽  
Flow Visualizations ◽  
End Wall

AbstractThis contribution describes experimental and numerical research of an unsteady behaviour of a flow in an end-wall region of a linear nozzle cascade. Effects of compressibility ($$M_\mathrm {2,is}$$ M 2 , is ) and inlet flow angle ($$\alpha _1$$ α 1 ) were investigated. Reynolds number ($$Re_\mathrm {2,is}$$ R e 2 , is $$=8.5\times 10^5$$ = 8.5 × 10 5 ) was held constant for all tested cases. Unsteady pressure measurement was performed at the blade mid-span in the identical position $${\mathfrak {s}}$$ s to obtain reference data. Surface flow visualizations were performed as well as the steady pressure measurement to support conclusions obtained from the unsteady measurements. Comparison of the surface Mach number distributions obtained from the experiments and from the numerical simulations are presented. Flow visualizations are then compared with calculated limiting streamlines on the blade suction surface. It was shown, that the flow structures in the end-wall region were not affected by the primary flow at the blade mid-span, even when the shock wave formed. This conclusion was made from the experimental, numerical, steady as well as unsteady points of view. Three significant frequencies in the power spectra suggested that there was a periodical interaction between the vortex structures in the end-wall region. Based on the data analyses, anisotropic turbulence was observed in the cascade.

A CFD-Based Investigation of Boundary Layer Skew in the Hub Region of a Low Speed Axial Compressor and its Influence on Performance and Losses

2009 ◽  
Author(s):  
M. Hilgert ◽  
M. Bo¨hle
Keyword(s):  
Boundary Layer ◽  
Axial Compressor ◽  
Surface Flow ◽  
Flow Coefficient ◽  
Wall Region ◽  
Suction Surface ◽  
Local Flow ◽  
Low Speed ◽  
Wall Area ◽  
End Wall

The flow in the hub end-wall region has a substantial influence on the aerodynamic performance of axial-flow compressors. A numerical investigation of this three-dimensional flow phenomenon, often referred to as boundary layer skew, that contributes to the interaction between the end-wall and blade suction surface flow, thus determining the losses in this area, is the topic of this report. A single-staged model compressor (flow coefficient: 0.5, work coefficient: 0.34) with a 7-blade rotor row and a highly loaded 11-blade stator row was designed for the simulation. To account for the complex time-dependent flow patterns in the end-wall area, a transient calculation with the rotor and stator rows fully modeled (360°) is carried out. The flow is assumed to be incompressible as the velocities of the low-speed axial compressor do not exceed a Mach-number value of 0.3. The calculations show that the boundary layer in the end-wall region is highly skewed and transient. There is a direct connection to local flow phenomena such as separation as well as to global pressure loss coefficient distributions in pitchwise and streamwise direction. Different levels of flow overturning in the end-wall area are observed, depending on the boundary layer skewness varying by simulating the compressor at different operational points.

Unsteady Pressure Investigations of Corner Separated Flow in a Linear Compressor Cascade

2015 ◽  
Author(s):  
Gherardo Zambonini ◽  
Xavier Ottavy
Keyword(s):  
Transfer Functions ◽  
Leading Edge ◽  
Positive Pressure ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Linear Compressor ◽  
Corner Separation ◽  
Unsteady Pressure ◽  
Linear Compressor Cascade ◽  
End Wall

The aim of this work is to present detailed unsteady pressure measurements of three-dimensional flow field in a NACA 65 linear compressor cascade. Chord-based Reynolds number of 382000 and incidence angle of 4 degrees were chosen as target configuration of the rig, which clearly presents the corner separation phenomenon at the juncture of the blade suction side and the end-wall. Concerning the experiments, a characterization of the mean and fluctuating component of wall static pressure on the surface of a specially developed blade is achieved at first. This fluctuating component is investigated utilizing nineteen high sensitivity condenser microphones plugged into blade cavities which have been carefully calibrated. Transfer functions obtained by calibration are exploited to reconstruct the time-dependent pressure signal and finally statistics, conditional ensemble averages, coherence and spectra analyses of fluctuations are presented in order to investigate the unsteady characteristics of the corner separation. High values of root mean square are individuated near the leading edge and in the separation region on the suction surface of the blade. Skewness and kurtosis show an intermittent behavior of the separation onset, which moves upstream and downstream on the suction surface. This intermittency of the separation line is probably linked with the existence of a bimodal behavior of the size of the corner separation. The analyses of coherence and conditional ensemble average between the signals at the leading edge and at the onset of the separation suggest a critical influence of angle and velocity of the incoming end-wall boundary layer on the positive pressure signatures of the shear layer, which characterize the inception of the separation.

scholarly journals Effect of Reynolds Number on Separation Bubbles on Controlled-Diffusion Compressor Blades in Cascade

1998 ◽  
Author(s):  
Garth V. Hobson ◽  
Denis J. Hansen ◽  
David G. Schnorenberg ◽  
Darren V. Grove
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Separation Bubble ◽  
Boundary Layer Separation ◽  
Surface Flow ◽  
Flow Reversal ◽  
Surface Boundary ◽  
Suction Surface ◽  
Flow Angle ◽  
Controlled Diffusion

A detailed experimental investigation of second-generation, controlled-diffusion, compressor stator blades at an off-design inlet-flow angle was performed in a low-speed cascade wind tunnel primarily using laser-Doppler velocimetry (LDV). The object of the study was to characterize the off-design flowfield and to obtain LDV measurements of the suction surface boundary layer separation which occurred near mid chord. The effect of Reynolds number on the flow separation in the regime of 210,000 to 640,000 was investigated. Surface flow visualization showed that at the low Re. no. the mid-chord separation bubble started laminar and reattached turbulent within 20% chord on the suction side of the blade. The extent of the bubble compared very well with the measured blade surface pressure distribution which showed a classical plateau and then diffusion in the turbulent region. LDV measurements of the flow reversal in the bubble were performed. At the intermediate Re. no. the boundary layer was transitional before the bubble which had decreased significantly in size (down to 10% chord). At the highest Re. no. the flow was turbulent from close to the leading edge, and three-dimensional flow reversal as a result of endwall effects appeared at approximately 80% chord which did not reattach.

The Application of Non-Axisymmetric Profiled End-Walls for Axial Flow Turbines in the Embedded Stage Environment

2013 ◽  
Author(s):  
Andreas Lintz ◽  
Liping Xu ◽  
Marios Karakasis
Keyword(s):  
Axial Flow ◽  
Leading Edge ◽  
Surface Flow ◽  
Secondary Flows ◽  
Experimental Results ◽  
Front Part ◽  
Flow Conditions ◽  
Suction Surface ◽  
Inlet Conditions ◽  
End Wall

In this paper, an assessment of the effectiveness of non-axisymmetric profiled end-walls in the embedded stage environment at varying inlet conditions is presented. Both numerical and experimental results were obtained in a three-stage model turbine which offers flow conditions representative of embedded blade rows in a typical high pressure steam turbine. The end-wall profile design was carried out using automatic optimization in conjunction with 3D RANS CFD. The design target is to reduce the end-wall losses by reducing the loading in the front part of the passage, which resulted in a single trough close to the blade suction surface in the leading edge region. 5-hole probe traverses and surface flow visualization show that the intensity of the secondary flows is reduced by about 10%, but overall loss is only reduced slightly. Experimental results have been obtained for the cylindrical end-wall and three different trough depths. With increasing depth, transitional effects at the end-walls might come into play, increasing the total pressure loss in the boundary layer region. The effects of the end-wall design is similar at positive and negative incidence, despite the reduced loading in the front part of the passage at negative incidence. At very high negative incidence angles, such as those occurring at the stator tip with rotor shroud leakage flows, the mechanism of secondary flow generation changes, so that a design under nominal inlet flow conditions shows no effect on the exit flow field.

Incidence Effect on the Aero-Thermal Performance of a Film Cooled Nozzle Vane Cascade

2018 ◽  
Author(s):  
H. Abdeh ◽  
G. Barigozzi ◽  
A. Perdichizzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Thermal Performance ◽  
Cooling System ◽  
Negative Impact ◽  
Surface Flow ◽  
Intensity Level ◽  
Inlet Flow ◽  
Flow Angle ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane ◽  
End Wall

In the present paper, the influence of inlet flow incidence on the aerodynamic and thermal performance of a film cooled linear nozzle vane cascade is fully assessed. Tests have been carried out on a solid and a cooled cascade. In the cooled cascade, coolant is ejected at the end wall through a slot located upstream of the leading edge plane. Moreover, a vane showerhead cooling system is also realized through 4 rows of cylindrical holes. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at a constant inlet Mach number of 0.12 and nominal cooling condition, varying the inlet flow angle in the range ±20°. The aero-thermal characterization of vane platform was obtained through 5-hole probe and end wall adiabatic film cooling effectiveness measurements. Vane load distributions and surface flow visualizations supported the discussion of the results. A relevant negative impact of positive inlet flow incidence on the cooled cascade aerodynamic and thermal performance was detected. A negligible influence was instead observed at negative incidence, even at the lowest tested value of −20°.

Numerical Investigations on Application of Cantilever Stator on Aerodynamic Performance of Tandem Bladed Axial-Flow Compressor

2021 ◽  
Author(s):  
Bhanu Pratap Singh Tanwar ◽  
Ajey Singh ◽  
Chetan S. Mistry
Keyword(s):  
Secondary Flow ◽  
Operating Conditions ◽  
Design Flow ◽  
Leakage Flow ◽  
Pressure Side ◽  
Tip Leakage Flow ◽  
Suction Surface ◽  
Flow Angle ◽  
Bladed Rotor ◽  
End Wall

Abstract Adoption of a tandem bladed rotor configuration brings special flow features at the exit compared to the conventional rotor. For tandem bladed rotor, there is the presence of strong dual-tip leakage flow, atypical exit flow angle distributions, corner blade separations leading to thicker dual wakes at the exit of the rotor to name a few. This makes the aerodynamic design of downstream stator more challenging in terms of overall performance as well as operational stability. The modern compressor requisite of being lighter and cost-efficient needs to be taken care of both aerodynamic and mechanical requirements. To overcome all these challenges, the cantilever type stator (without hub rotation) has been chosen and been analyzed for the present study. The effects of different hub gap sizes of the cantilever stator in combination with the tandem bladed axial compressor stage are investigated in order to explore passive flow control mechanism near the hub. The goal of the work is to get further insights into the aerodynamic aspects of flow using a detailed flow field analysis. The numerical study was performed using ANSYS TurboGrid® for mesh generation and the commercial package ANSYS CFX® 18.0 was used as solver for steady-state simulation. Stationary hub boundary conditions have been employed for the stator in all 3 cases [baseline, 1% and 2% (of span) part clearance]. For no clearance case, the regions of momentum deficit were observed in the vicinity of the hub endwall and suction surface of the stator. The region keeps growing along both streamwise and spanwise direction as a low momentum bubble is formed near trailing edge. This low momentum bubble seems to be transported along the span and moved more towards the suction surface. The solution strategy explored to mitigate the effect of hub corner separation by adapting hub clearance. The role played by secondary flow in feeding the low momentum flow along the span is seen to be moderated by the high momentum leakage flow from the pressure side. The hub leakage flow from the blade pressure side reenergized the low momentum fluid on the suction side refraining it to travel along the span and mitigate its effect by suppressing the separation tendency near end wall region. The formation of large size bubble gets reduced in overall size both in the circumferential and span-wise direction. This phenomenon compels the low momentum flow to pass along the low span region. Numerically obtained results provide an insightful mechanism of the interaction of secondary flow structures and the influence of hub clearance flow. Hub corner stall, which is the consequence of low momentum fluid sweeping across the blade passage near the end wall got wiped out in the presence of hub clearance. This phenomenon diminishes the extent and overall effect of the hub corner stall. The interaction of hub leakage vortex and passage vortex leads to mitigation of overall secondary flow adverse effects. As a result, performance improvement at design flow conditions have been elucidated by implementation of cantilever stator. The peak pressure operation is dominated by mid-span flow complexities and as a result cantilevered stator doesn’t show much improvements. Nevertheless, the improvements in design point operating conditions do justify the study for gaining physical insights.

Control of Three-Dimensional Separations in Axial Compressors by Tailored Boundary Layer Suction

2007 ◽  
Vol 130 (1) ◽  
Author(s):  
Semiu A. Gbadebo ◽  
Nicholas A. Cumpsty ◽  
Tom P. Hynes
Keyword(s):  
Boundary Layer ◽  
Static Pressure ◽  
Three Dimensional ◽  
Pressure Rise ◽  
Surface Flow ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Blade Passage ◽  
Boundary Layer Suction ◽  
End Wall

One of the important ways of improving turbomachinery compressor performance is to control three-dimensional (3D) separations, which form over the suction surface and end wall corner of the blade passage. Based on the insights gained into the formation of these separations, this paper illustrates how an appropriately applied boundary layer suction of up to 0.7% of inlet mass flow can control and eliminate typical compressor stator hub corner 3D separation over a range of operating incidence. The paper describes, using computational fluid dynamics, the application of suction on the blade suction surface and end wall boundary layers and exemplifies the influence of end wall dividing streamline in initiating 3D separation in the blade passage. The removal of the separated region from the blade suction surface is confirmed by an experimental investigation in a compressor cascade involving surface flow visualization, surface static pressure, and exit loss measurements. The ensuing passage flow field is characterized by increased blade loading (static pressure difference between pressure and suction surface), enhanced average static pressure rise, significant loss removal, and a uniform exit flow. This result also enables the contribution of the 3D separation to the overall loss and passage blockage to be assessed.

Incidence Effect on the Aero-Thermal Performance of a Film Cooled Nozzle Vane Cascade

2019 ◽  
Vol 141 (5) ◽  
Author(s):  
H. Abdeh ◽  
G. Barigozzi ◽  
A. Perdichizzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Thermal Performance ◽  
Cooling System ◽  
Negative Impact ◽  
Surface Flow ◽  
Intensity Level ◽  
Inlet Flow ◽  
Flow Angle ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane ◽  
End Wall

In the present paper, the influence of inlet flow incidence on the aerodynamic and thermal performance of a film cooled linear nozzle vane cascade is fully assessed. Tests have been carried out on a solid and a cooled cascade. In the cooled cascade, coolant is ejected at the end wall through a slot located upstream of the leading edge plane. Moreover, a vane showerhead cooling system is also realized through four rows of cylindrical holes. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at a constant inlet Mach number of 0.12 and nominal cooling condition, varying the inlet flow angle in the range ±20 deg. The aero-thermal characterization of vane platform was obtained through five-hole probe and end wall adiabatic film cooling effectiveness measurements. Vane load distributions and surface flow visualizations supported the discussion of the results. A relevant negative impact of positive inlet flow incidence on the cooled cascade aerodynamic and thermal performance was detected. A negligible influence was instead observed at negative incidence, even at the lowest tested value of −20 deg.

Direct numerical simulations of Rayleigh–Bénard convection in water with non-Oberbeck–Boussinesq effects

2019 ◽  
Vol 881 ◽  
pp. 1073-1096 ◽  
Author(s):  
Andreas D. Demou ◽  
Dimokratis G. E. Grigoriadis
Keyword(s):  
Numerical Simulations ◽  
Two Dimensions ◽  
Direct Numerical Simulations ◽  
Wall Region ◽  
Number Range ◽  
Bénard Convection ◽  
Benard Convection ◽  
Rayleigh Benard Convection ◽  
Rayleigh Benard ◽  
Near Wall

Rayleigh–Bénard convection in water is studied by means of direct numerical simulations, taking into account the variation of properties. The simulations considered a three-dimensional (3-D) cavity with a square cross-section and its two-dimensional (2-D) equivalent, covering a Rayleigh number range of $10^{6}\leqslant Ra\leqslant 10^{9}$ and using temperature differences up to 60 K. The main objectives of this study are (i) to investigate and report differences obtained by 2-D and 3-D simulations and (ii) to provide a first appreciation of the non-Oberbeck–Boussinesq (NOB) effects on the near-wall time-averaged and root-mean-squared (r.m.s.) temperature fields. The Nusselt number and the thermal boundary layer thickness exhibit the most pronounced differences when calculated in two dimensions and three dimensions, even though the $Ra$ scaling exponents are similar. These differences are closely related to the modification of the large-scale circulation pattern and become less pronounced when the NOB values are normalised with the respective Oberbeck–Boussinesq (OB) values. It is also demonstrated that NOB effects modify the near-wall temperature statistics, promoting the breaking of the top–bottom symmetry which characterises the OB approximation. The most prominent NOB effect in the near-wall region is the modification of the maximum r.m.s. values of temperature, which are found to increase at the top and decrease at the bottom of the cavity.

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