Analysis of Transonic Bladerows With Non-Uniform Geometry Using Spectral Method

2020 ◽  
Author(s):  
Feng Wang ◽  
Luca di Mare
Keyword(s):  
Flow Field ◽  
Spectral Method ◽  
Computational Cost ◽  
Leading Edge ◽  
Wave Number ◽  
Efficient Tool ◽  
Flow Solver ◽  
Engine Order ◽  
Harmonic Flow ◽  
Good Agreement

Abstract Turbomachinery blade rows can have non-uniform geometries due to design intent, manufacture errors or wear. When predictions are sought for the effect of such non-uniformities, it is generally the case that whole assembly calculations are needed. A spectral method is used in this paper to approximate the flow fields of the whole assembly but with significantly less computation cost. The method projects the flow perturbations due to the geometry non-uniformity in an assembly in Fourier space, and only one passage is required to compute the flow perturbations corresponding to a certain wave-number of geometry variation. The performance of this method on transonic blade rows is demonstrated on a modern fan assembly. Low engine order and high engine order geometry non-uniformity (e.g. “saw-tooth” pattern) are examined. The non-linear coupling between the flow perturbations and the passage-averaged flow field is also demonstrated. Pressure variations on the blade surface and the potential flow field upstream of the leading edge from the proposed spectral method and the direct whole assembly solutions are compared. Good agreement is observed on both quasi-3D and full 3D cases. A lumped approach to compute deterministic fluxes is also proposed to further reduce the computational cost of the spectral method. The spectral method is formulated in such a way that it can be easily implemented into an existing harmonic flow solver by adding an extra source term, and can be potentially used as an efficient tool for aeromechanical and aeroacoustics design of turbomachinery blade rows.

ANALYSIS OF TRANSONIC BLADEROWS WITH NON-UNIFORM GEOMETRY USING SPECTRAL METHOD

2021 ◽  
pp. 1-27
Author(s):  
Feng Wang ◽  
Luca di Mare
Keyword(s):  
Flow Field ◽  
Spectral Method ◽  
Computational Cost ◽  
Leading Edge ◽  
Wave Number ◽  
Efficient Tool ◽  
Flow Solver ◽  
Engine Order ◽  
Harmonic Flow ◽  
Good Agreement

Abstract Turbomachinery blade rows can have non-uniform geometries due to design intent, manufacture errors or wear. When predictions are sought for the effect of such non-uniformities, it is generally the case that whole assembly calculations are needed. A spectral method is used in this paper to approximate the flow fields of the whole assembly but with significantly less computation cost. The method projects the flow perturbations due to the geometry non-uniformity in an assembly in Fourier space. Only one passage is required to compute the flow perturbations corresponding to a certain wave-number of geometry variation. The performance of this method on transonic blade rows is demonstrated on a modern fan assembly. Low and high engine order geometry non-uniformity (e.g. “saw-tooth” pattern) are examined. The non-linear coupling between the flow perturbations and the passage-averaged flow field is also demonstrated. Pressure variations on the blade surface and the potential flow field upstream of the leading edge from the proposed method and the direct whole assembly solutions are compared. Good agreement is observed on both quasi-3D and full 3D cases. A lumped approach to compute deterministic fluxes is also proposed to further reduce the computational cost of the spectral method. The spectral method is formulated in such a way that it can be easily implemented into an existing harmonic flow solver by adding an extra source term, and can be used as an efficient tool for aeromechanical and aeroacoustics design of turbomachinery blade rows.

Reducing Instrumentation Errors Caused by Circumferential Flow-Field Variations in Multistage Axial Compressors

2020 ◽  
Vol 142 (9) ◽  
Author(s):  
M. Chilla ◽  
G. Pullan ◽  
S. Gallimore
Keyword(s):  
Flow Field ◽  
Large Scale ◽  
Sampling Error ◽  
Leading Edge ◽  
Stagnation Pressure ◽  
Wave Number ◽  
Stagnation Temperature ◽  
Sampling Errors ◽  
Mean Flow ◽  
Axial Compressors

Abstract The effects of blade row interactions on stator-mounted instrumentation in axial compressors are investigated using unsteady numerical calculations. The test compressor is an eight-stage machine representative of an aero-engine core compressor. For the unsteady calculations, a 180-deg sector (half-annulus) model of the compressor is used. It is shown that the time-mean flow field in the stator leading edge planes is circumferentially nonuniform. The circumferential variations in stagnation pressure and stagnation temperature, respectively, reach 4.2% and 1.1% of the local mean. Using spatial wave number analysis, the incoming wakes from the upstream stator rows are identified as the dominant source of the circumferential variations in the front and middle of the compressor, while toward the rear of the compressor, the upstream influence of the eight struts in the exit duct becomes dominant. Based on three circumferential probes, the sampling errors for stagnation pressure and stagnation temperature are calculated as a function of the probe locations. Optimization of the probe locations shows that the sampling error can be reduced by up to 77% by circumferentially redistributing the individual probes. The reductions in the sampling errors translate to reductions in the uncertainties of the overall compressor efficiency and inlet flow capacity by up to 50%. Recognizing that data from large-scale unsteady calculations are rarely available in the instrumentation phase for a new test rig or engine, a method for approximating the circumferential variations with single harmonics is presented. The construction of the harmonics is based solely on the knowledge of the number of stators in each row and a small number of equispaced probes. It is shown how excursions in the sampling error are reduced by increasing the number of circumferential probes.

Time-Space Spectral Method for Rotor–Rotor/Stator–Stator Interactions

2019 ◽  
Vol 141 (11) ◽  
Author(s):  
Xiuquan Huang ◽  
Ding Xi Wang
Keyword(s):  
Flow Field ◽  
Spectral Method ◽  
Harmonic Balance ◽  
Harmonic Balance Method ◽  
Frequency Domain Method ◽  
Test Case ◽  
Flow Solver ◽  
Time Space ◽  
Space Harmonics

Abstract The paper presents a time-space spectral method for an efficient analysis of rotor–rotor/stator–stator interactions in the framework of the time spectral form harmonic balance method. The method treats time and space harmonics in a coherent way and allows for easy choice of time and passage through the introduction of pseudo-shaft frequency and composite frequency. The proposed method can accommodate passage to passage variation of time-averaged flow field and amplitude of unsteady flow field of a given frequency as needed for rotor–rotor/stator–stator interactions. Minimum change is required to extend an existing harmonic balance flow solver to incorporate the proposed method. The proposed method is the most concise and the simplest one of its kind so far. The first three blade rows of a two-stage fan are used as a test case to demonstrate the validity and effectiveness of the proposed method. Numerical results demonstrate the necessity of including rotor–rotor/stator–stator interaction in an analysis using a frequency-domain method and the capability of the proposed method for such a purpose. It is also concluded from the case study that extra spatial and temporal harmonics are needed to adequately analyze a rotor–rotor/stator–stator interaction.

Reducing Instrumentation Errors Caused by Circumferential Flow Field Variations in Multi-Stage Axial Compressors

2019 ◽  
Author(s):  
M. Chilla ◽  
G. Pullan ◽  
S. Gallimore
Keyword(s):  
Flow Field ◽  
Large Scale ◽  
Sampling Error ◽  
Leading Edge ◽  
Stagnation Pressure ◽  
Wave Number ◽  
Stagnation Temperature ◽  
Sampling Errors ◽  
Mean Flow ◽  
Axial Compressors

Abstract The effects of blade row interactions on stator-mounted instrumentation in axial compressors are investigated using unsteady numerical calculations. The test compressor is an 8-stage machine representative of an aero-engine core compressor. For the unsteady calculations, a 180deg sector (half-annulus) model of the compressor is used. It is shown that the time-mean flow field in the stator leading edge planes is circumferentially non-uniform. The circumferential variations in stagnation pressure and stagnation temperature respectively reach 4.2% and 1.1% of the local mean. Using spatial wave number analysis, the incoming wakes from the upstream stator rows are identified as the dominant source of the circumferential variations in the front and middle of the compressor, while towards the rear of the compressor, the upstream influence of the eight struts in the exit duct becomes dominant. Based on three circumferential probes, the sampling errors for stagnation pressure and stagnation temperature are calculated as a function of the probe locations. Optimization of the probe locations shows that the sampling error can be reduced by up to 77% by circumferentially redistributing the individual probes. The reductions in the sampling errors translate to reductions in the uncertainties of the overall compressor efficiency and inlet flow capacity by up to 50%. Recognizing that data from large-scale unsteady calculations is rarely available in the instrumentation phase for a new test rig or engine, a method for approximating the circumferential variations with single harmonics is presented. The construction of the harmonics is based solely on the knowledge of the number of stators in each row and a small number of equi-spaced probes. It is shown how excursions in the sampling error are reduced by increasing the number of circumferential probes.

Numerical simulation of swirling flow effect on the first stage vane film cooling distribution

2016 ◽  
Vol 07 (03) ◽  
pp. 1650031
Author(s):  
Hong Yin
Keyword(s):  
Flow Field ◽  
Film Cooling ◽  
Swirling Flow ◽  
Cooling System ◽  
Leading Edge ◽  
Suction Side ◽  
Local Area ◽  
Swirl Flow ◽  
Flow Effect ◽  
Recirculation Flow

In advanced gas turbine technology, lean premixed combustion is an effective strategy to reduce peak temperature and thus, NO[Formula: see text] emissions. The swirler is adopted to establish recirculation flow zone, enhancing mixing and stabilizing the flame. Therefore, the swirling flow is dominant in the combustor flow field and has impact on the vane. This paper mainly investigates the swirling flow effect on the turbine first stage vane cooling system by conducting a group of numerical simulations. Firstly, the numerical methods of turbulence modeling using RANS and LES are compared. The computational model of one single swirl flow field is considered. Both the RANS and LES results give reasonable recirculation zone shape. When comparing the velocity distribution, the RANS results generally match the experimental data but fail to at some local area. The LES modeling gives better results and more detailed unsteady flow field. In the second step, the RANS modeling is incorporated to investigate the vane film cooling performance under the swirling inflow boundary condition. According to the numerical results, the leading edge film cooling is largely altered by the swirling flow, especially for the swirl core-leading edge aligned case. Compared to the pressure side, the suction side film cooling is more sensitive to the swirling flow. Locally, the film cooling jet is lifted and turned by the strong swirling flow.

On the Flow Fields of Inertial Impactors

1974 ◽  
Vol 96 (4) ◽  
pp. 394-400 ◽  
Author(s):  
V. A. Marple ◽  
B. Y. H. Liu ◽  
K. T. Whitby
Keyword(s):  
Flow Field ◽  
Stokes Equations ◽  
Relaxation Method ◽  
Water Model ◽  
Navier Stokes ◽  
Visualization Technique ◽  
Navier Stokes Equations ◽  
Theoretical Results ◽  
Plate Distance ◽  
Good Agreement

The flow field in an inertial impactor was studied experimentally with a water model by means of a flow visualization technique. The influence of such parameters as Reynolds number and jet-to-plate distance on the flow field was determined. The Navier-Stokes equations describing the laminar flow field in the impactor were solved numerically by means of a finite difference relaxation method. The theoretical results were found to be in good agreement with the empirical observations made with the water model.

Numerical Investigation of Intermittent Corner Separation in a Linear Compressor Cascade

2016 ◽  
Author(s):  
Wei Ma ◽  
Feng Gao ◽  
Xavier Ottavy ◽  
Lipeng Lu ◽  
A. J. Wang
Keyword(s):  
Flow Field ◽  
Large Scale ◽  
Leading Edge ◽  
Suction Side ◽  
Detached Eddy Simulation ◽  
Compressor Cascade ◽  
Linear Compressor ◽  
Corner Separation ◽  
Eddy Simulation ◽  
Linear Compressor Cascade

Recently bimodal phenomenon in corner separation has been found by Ma et al. (Experiments in Fluids, 2013, doi:10.1007/s00348-013-1546-y). Through detailed and accurate experimental results of the velocity flow field in a linear compressor cascade, they discovered two aperiodic modes exist in the corner separation of the compressor cascade. This phenomenon reflects the flow in corner separation is high intermittent, and large-scale coherent structures corresponding to two modes exist in the flow field of corner separation. However the generation mechanism of the bimodal phenomenon in corner separation is still unclear and thus needs to be studied further. In order to obtain instantaneous flow field with different unsteadiness and thus to analyse the mechanisms of bimodal phenomenon in corner separation, in this paper detached-eddy simulation (DES) is used to simulate the flow field in the linear compressor cascade where bimodal phenomenon has been found in previous experiment. DES in this paper successfully captures the bimodal phenomenon in the linear compressor cascade found in experiment, including the locations of bimodal points and the development of bimodal points along a line that normal to the blade suction side. We infer that the bimodal phenomenon in the corner separation is induced by the strong interaction between the following two facts. The first is the unsteady upstream flow nearby the leading edge whose angle and magnitude fluctuate simultaneously and significantly. The second is the high unsteady separation in the corner region.

scholarly journals Aerothermal Investigations of Mixing Flow Phenomena in Case of Radially Inclined Ejection Holes at the Leading Edge

1999 ◽  
Author(s):  
Dieter E. Bohn ◽  
Karsten A. Kusterer
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Gas Turbines ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Blade Surface ◽  
Cooling Fluid ◽  
Flow Phenomena ◽  
Vortex Systems

A leading edge cooling configuration is investigated numerically by application of a 3-D conjugate fluid flow and heat transfer solver, CHT-Flow. The code has been developed at the Institute of Steam and Gas Turbines, Aachen University of Technology. It works on the basis of an implicit finite volume method combined with a multi-block technique. The cooling configuration is an axial turbine blade cascade with leading edge ejection through two rows of cooling holes. The rows are located in the vicinity of the stagnation line, one row is on the suction side, the other row is on the pressure side. The cooling holes have a radial ejection angle of 45°. This configuration has been investigated experimentally by other authors and the results have been documented as a test case for numerical calculations of ejection flow phenomena. The numerical domain includes the internal cooling fluid supply, the radially inclined holes and the complete external flow field of the turbine vane in a high resolution grid. Periodic boundary conditions have been used in the radial direction. Thus, end wall effects have been excluded. The numerical investigations focus on the aerothermal mixing process in the cooling jets and the impact on the temperature distribution on the blade surface. The radial ejection angles lead to a fully three dimensional and asymmetric jet flow field. Within a secondary flow analysis it can be shown that complex vortex systems are formed in the ejection holes and in the cooling fluid jets. The secondary flow fields include asymmetric kidney vortex systems with one dominating vortex on the back side of the jets. The numerical and experimental data show a good agreement concerning the vortex development. The phenomena on the suction side and the pressure side are principally the same. It can be found that the jets are barely touching the blade surface as the dominating vortex transports hot gas under the jets. Thus, the cooling efficiency is reduced.

Instability of a gravity current within a soap film

2014 ◽  
Vol 753 ◽  
Author(s):  
Raymond E. Goldstein ◽  
Herbert E. Huppert ◽  
H. Keith Moffatt ◽  
Adriana I. Pesci
Keyword(s):  
Fluid Flow ◽  
Gravity Current ◽  
Leading Edge ◽  
Soap Film ◽  
Plateau Border ◽  
Dynamic Variations ◽  
Characteristic Wavelength ◽  
Overall Stability ◽  
Stability Regime ◽  
Good Agreement

AbstractOne of the simplest geometries in which to study fluid flow between two soap films connected by a Plateau border is provided by a catenoid with a secondary film at its narrowest point. Dynamic variations in the spacing between the two rings supporting the catenoid lead to fluid flow between the primary and secondary films. When the rings are moved apart, while keeping their spacing within the overall stability regime of the films, after a rapid thickening of the secondary film the excess fluid in it starts to drain into the sloped primary film through the Plateau border at which they meet. This influx of fluid is accommodated by a local thickening of the primary film. Experiments described here show that after this drainage begins the leading edge of the gravity current becomes linearly unstable to a finite-wavelength fingering instability. A theoretical model based on lubrication theory is used to explain the mechanism of this instability. The predicted characteristic wavelength of the instability is shown to be in good agreement with experimental results. Since the gravity current advances into a film of finite, albeit microscopic, thickness this situation is one in which the regularization often invoked to address singularities at the nose of a thin film is physically justified.

Comparison of Experimental and Numerical Simulations of Flow Within an Arterio-Venous Graft

2000 ◽  
Author(s):  
Paul F. Fischer ◽  
Seung Lee ◽  
Francis Loth ◽  
Hisham S. Bassiouny ◽  
Nurullah Arslan
Keyword(s):  
Flow Field ◽  
Steady Flow ◽  
Laser Doppler Anemometry ◽  
Transition To Turbulence ◽  
Flow Conditions ◽  
Venous Graft ◽  
Venous Anastomosis ◽  
Av Graft ◽  
Good Agreement

Abstract This was a study to compare computational and experimental results of flow field inside the venous anastomosis of an arteriovenous (AV) graft. Laser Doppler anemometry (LDA) measurements were conducted inside an upscaled end-to-side graft model under steady flow conditions at Reynolds number 1820 which is representative of the in vivo flow conditions inside a human AV graft. The distribution of the velocity and turbulence intensity was measured at several locations in the plane of the bifurcation. This flow field was simulated using computation fluid dynamics (CFD) and shown to be in good agreement. Under steady flow conditions, the flow field demonstrated an unsteady character (transition to turbulence).

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