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NARROW
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1-20 of 29
Toshinori Watanabe
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Proceedings Papers
Proc. ASME. GT2018, Volume 2A: Turbomachinery, V02AT45A019, June 11–15, 2018
Paper No: GT2018-76327
Abstract
This paper focuses on the buoyancy-induced unsteady flow phenomenon inside high-pressure compressor disk cavities. In order to understand the flow structure in a realistic configuration, a 10-stage core compressor of the NASA/GE Energy Efficient Engine is adopted as a computational target. The numerical flow simulation is conducted on a full annulus model, where the temperature distribution on the wall is modeled based on the core test results. The time-averaged flow fields are obtained by detached eddy simulation (DES) and two-dimensional axisymmetric Reynolds-averaged Navier-Stokes (RANS) simulation, and the difference is discussed in detail. The DES result showed large-scale, vortical structures with significant radial velocity fluctuations especially in the rear part of the compressor. These fluctuations create radial arm-like structure in the temperature distribution in the cavity, and greatly enhance the mixing between the bore coolant and hot air near the cavity wall. In addition, it is observed that the hot air discharged from the cavities creates a large cell at bore region, which extends across several rear stages. Although the present study successfully illustrates the entire structure of unsteady flow in heated compressor disk cavities including full stages, a more detailed validation will be needed to further confirm the applicability of DES for the targeted flow.
Proceedings Papers
Proc. ASME. GT2018, Volume 2A: Turbomachinery, V02AT39A018, June 11–15, 2018
Paper No: GT2018-75916
Abstract
Casing boundary layer effectively places a limit on the pressure rise capability achievable by the compressor. The separation of the casing boundary layer not only produce flow loss but also closely related to the compressor rotating stall. The motivation of this paper is to present a viewpoint that the casing boundary layer should be paid attention to in parallel with other flow factors on rotating stall trigger. This paper illustrates the casing boundary layer behavior by displaying its separation phenomena with the presence of tip leakage vortex at different flow conditions. Skin friction lines and the corresponding absolute streamlines are used to demonstrate the three-dimensional flow patterns on and near the casing. The results depict a Saddle, a Node and several tufts of skin friction lines dividing the passage into four zones. The tip leakage vortex is enfolded within one of the zones by the separated flows. All the flows in each blade passage are confined within the passage as long as the compressor is stable. The casing boundary layer of a transonic compressor is also examined in the same way, which results in qualitatively similar zonal flows that enfolds the tip leakage vortex. This research develops a new way to study the casing boundary layer in rotating compressors. The results may provide a first-principle based explanation to stalling mechanisms for compressors that are casing sensitive.
Journal Articles
Journal:
Journal of Turbomachinery
Article Type: Research-Article
J. Turbomach. May 2018, 140(6): 061003.
Paper No: TURBO-17-1196
Published Online: April 18, 2018
Abstract
A fast-response pressure-sensitive paint (PSP) technique was applied to the measurement of unsteady surface pressure of an oscillating cascade blade in a transonic flow. A linear cascade was used, and its central blade was oscillated in a translational manner. The unsteady pressure distributions of the oscillating blade and two stationary neighbors were measured using the fast-response PSP technique, and the unsteady aerodynamic force on the blade was obtained by integrating the data obtained on the pressures. The measurements made with the PSP technique were compared with those obtained by conventional methods for the purpose of validation. From the results, the PSP technique was revealed to be capable of measuring the unsteady surface pressure, which is used for flutter analysis in transonic conditions.
Proceedings Papers
Proc. ASME. GT2017, Volume 7B: Structures and Dynamics, V07BT36A013, June 26–30, 2017
Paper No: GT2017-64211
Abstract
A fast-response pressure-sensitive paint (PSP) technique was applied to the measurement of unsteady surface pressure of an oscillating cascade blade in a transonic flow. A linear cascade was used, and its central blade was oscillated in a translational manner. The unsteady pressure distributions of the oscillating blade and two stationary neighbors were measured using the fast-response PSP technique, and the unsteady aerodynamic force on the blade was obtained by integrating the data obtained on the pressures. The measurements made with the PSP technique were compared with those obtained by conventional methods for the purpose of validation. From the results, the PSP technique was revealed to be capable of measuring the unsteady surface pressure, which is used for flutter analysis in transonic conditions.
Proceedings Papers
Proc. ASME. GT2017, Volume 2B: Turbomachinery, V02BT41A031, June 26–30, 2017
Paper No: GT2017-64195
Abstract
This paper describes basic ideas for the design of a new computational fluid dynamics (CFD) code for large-eddy simulation (LES), as well as presenting fundamental validation and demonstration cases. The developed code combines highorder structured solver with overset mesh technique, which enables to control mesh density easily with keeping high mesh quality. The validations were conducted over two different flow regimes. The mesh size criteria for the present code were identified from the systematic mesh study of wall-bounded channel flow. In addition, the wake profile after the circular cylinder as a validation of free shear flow showed good agreement with the reference data. Then the validated code was applied to the flow field of highly loaded stator vane. The developed solver reproduced the position of separation bubble accurately however, the velocity profiles on the vane needed improvement by introducing free stream turbulence. From the sequence of validation and application studies, it was concluded that the present code can be used for the researches of fundamental turbulent flow field in turbomachinery.
Proceedings Papers
Proc. ASME. GT2016, Volume 2A: Turbomachinery, V02AT37A053, June 13–17, 2016
Paper No: GT2016-58117
Abstract
Computational analysis has been conducted on the NASA Rotor 37 transonic compressor with various tip clearance gap heights. Using steady rotor-only analysis, the change in overall performance, basic flow characteristics, and near-casing phenomena have been carefully observed. The results have clarified that the peak efficiency of the compressor decreases almost linearly with the increase in gap height. Meanwhile, the stall margin was prone to deterioration in cases of significantly small or significantly large clearance gaps. The peak stall margin was attained when the gap was set to 75% of the original height. Focusing on the flow structures, the tip leakage flow and tip leakage vortex seemed to be dominant loss sources in the case of a large tip clearance gap. On the other hand, trailing edge separation at the blade tip was the major loss source in case of a small tip clearance gap. The difference in the near-casing flow structure also determined the onset process of numerical instability. In case of a large tip clearance gap, the advance of the interface between the main flow and tip leakage flow seemed to cause an accumulation of blockage in the region near the casing, possibly triggering the tip-initiated stall. In the case of a small tip clearance gap, interaction among the wall separation, blade tip trailing edge separation, and shockwave /boundary layer interaction was significant. These phenomena appeared to play a major role in the onset of numerical instability in the blade tip region.
Proceedings Papers
Proc. ASME. GT2016, Volume 7B: Structures and Dynamics, V07BT34A017, June 13–17, 2016
Paper No: GT2016-57295
Abstract
This paper aims at quantifying the stabilization effect of mistuning in transonic fan flutter. The results are used to support the evaluation of flutter boundary and to clarify the reason for the mismatch observed in the numerical predictions reported in our previous study. Mistuning is modeled by the deviation of blade-mode frequency, and the stability analysis of vibrating blades is formulated as an eigenproblem of the equation of motion including self-excited aerodynamic force obtained by fluid-structure interaction simulations. Statistics about the modal properties are obtained by Monte Carlo simulation. The change in the averaged damping rate and flutter boundary is evaluated in a wide range of mistuning levels and operating conditions. Nominal levels of mistuning due to manufacturing tolerance have little effect to the flutter boundary because the decline in aerodynamic damping is very steep. Therefore, the accuracy associated with the computational fluid dynamics is likely to have caused the mismatch in the flutter boundary. Histograms of modal properties show that the inter-blade phase angle and blade amplitudes in flutter mode can be highly scattered, even if the level of mistuning is nominal. For largely mistuned cases, new crests which do not exist in nominal cases appear in the eigenvalue histogram. They were found to be highly-localized, single-blade dominant mode.
Proceedings Papers
Proc. ASME. GT2015, Volume 2B: Turbomachinery, V02BT39A040, June 15–19, 2015
Paper No: GT2015-43849
Abstract
Leading edge separation of thin airfoil cascade in subsonic flow at large angle of incidence was simulated by implicit large eddy simulation (ILES) and Reynolds averaged Navier-Stokes (RANS) simulations with various turbulence models. In the ILES simulations with fine grids, the time-averaged surface pressure qualitatively agreed with the experimental data. The RANS and ILES simulations on the coarse mesh failed to capture a peak of pressure near the leading edge. From spectrum analysis, it was observed that the flow-field was turbulent in the separation bubble. In the failed RANS simulations, the separation bubble was much longer and the turbulence energy near the leading edge was much lower than those in the ILES results. The development of lambda-shaped vortex structures and their sudden weakening near the reattachment point was observed in the unsteady simulations. Two possible modifications to existing turbulence models in RANS simulations were proposed based on the comparison of turbulence energy between the ILES and RANS results. It is shown that these modifications improve the bubble length and C p distributions of RANS simulations, though further validation and modeling are needed for the application to realistic cases.
Proceedings Papers
Proc. ASME. GT2015, Volume 4B: Combustion, Fuels and Emissions, V04BT04A022, June 15–19, 2015
Paper No: GT2015-43364
Abstract
Eulerian-Lagrangian hybrid method is implemented for the prediction of liquid atomization phenomena produced by 2 liquid water jets impinging by an angle of 40 deg. in quiet ambient air. To calculate the flow fields with liquid/gas interface, Eulerian analyses are conducted inside a fixed computational grid system. After the atomization occurs, every droplet is converted to a spherical particle. The motion of particles are tracked in Lagrangian form. For the validation of the developed Eulerian-Lagrangian hybrid method, flow visualization by using a high-speed video camera is carried out. To obtain quantitative values of spray characteristics, the liquid mass flux distribution in space is measured by utilizing a patternator. Numerical and experimental results of atomization process and mass flux distribution of spray show a similarity, and thus the developed method is evaluated that it has potential to predict spray characteristics produced by liquid sheet atomization. The developed numerical method can calculate unsteady spray distributions not only at the plane close to the injector but also far downstream. The spray mass flux distribution in the transient state, which is hard to measure by experiment, is demonstrated.
Proceedings Papers
Proc. ASME. GT1999, Volume 4: Manufacturing Materials and Metallurgy; Ceramics; Structures and Dynamics; Controls, Diagnostics and Instrumentation; Education; IGTI Scholar Award; General, V004T03A008, June 7–10, 1999
Paper No: 99-GT-050
Abstract
The unsteady aerodynamic characteristics of oscillating thin turbine blades were studied both experimentally and numerically to obtain the comprehensive knowledge on the aerodynamic damping of the blades operating in transonic flows. The experiment was carried out in a linear cascade tunnel by use of the influence coefficient method. The two flow conditions were adopted, namely, a near-design condition and an off-design condition with a higher back pressure. In the results for the near-design case, a strong vibration instability was observed in the positive side of the interblade phase angle. In the off-design case, however, the instability did not appear for almost all the interblade phase angles. A drastic change was found in the phase angle of unsteady aerodynamic force between the two cases, which change was a governing factor for the oscillation instability. Numerical simulation based on 2-D Euler equation revealed that the phase change came from the change in phase of the unsteady surface pressure across the shock impingement point on the blade suction surface in the off-design case. The numerical results also showed that the aerodynamic damping increased with increasing reduced frequency, and that the oscillation instability disappeared.
Proceedings Papers
Proc. ASME. GT2014, Volume 2A: Turbomachinery, V02AT37A044, June 16–20, 2014
Paper No: GT2014-26691
Abstract
The effect of a single circumferential casing groove on the stability enhancement of two different transonic compressors has been examined with CFD analysis. The differences in flow field and stall inception mechanism between two rotors are presented with principal focus on passage blockage and tip leakage flow behavior. Detailed observation showed that the blockage flow which leads the compressor to stall was different between each other. A parametric study conducted with respect to the axial location of the groove has clarified that the effect which groove has on the tip leakage flow behavior changes according to the blade tip loading and the design tip clearance gap at the location where the groove is applied. When the casing treatment was applied to the compressors with different instability mechanism, whether the casing treatment could enhance the stability of compressor or not was not only dependant on the extent of the influence which it had on the flow field but also on whether it could affect the original stall-initiating phenomena at the adequate location.
Proceedings Papers
Proc. ASME. GT2014, Volume 7B: Structures and Dynamics, V07BT35A019, June 16–20, 2014
Paper No: GT2014-26702
Abstract
Fully coupled steady fluid-solid interaction (FSI) and flutter simulations were conducted on a NASA Rotor 67 transonic experimental fan to demonstrate the capability of application for capturing various aeroelastic phenomena in turbomachinery. The effect of blade deformation on the aerodynamic performance was investigated by steady FSI. Aeroelastic modes were determined using the modal identification technique for the vibration of the cascade. The proposed identification method successfully estimated aeroelastic modes without significant uncertainty. Aeroelastic eigenvalues were localized around the structural modes in vacuum forming the “mode family”, and there was negligible change in their frequency. The calculated aerodynamic coupling between the structural modes was small. Based on the reconstructed local unsteady aerodynamic force, the major damping sources in the 1F mode family were determined to be the shock motion and supersonic region near the leading edge. From these results, it was confirmed that the developed FSI method was applicable to the analysis of unsteady characteristics of blades in multimode oscillation.
Proceedings Papers
Proc. ASME. GT2013, Volume 2: Aircraft Engine; Coal, Biomass and Alternative Fuels; Cycle Innovations, V002T01A022, June 3–7, 2013
Paper No: GT2013-95180
Abstract
Jet noise reduction is essential for next-generation environmentally-friendly supersonic transport. In the present study, experimental and numerical investigations were performed to clarify the effect of microjet injection on supersonic jet noise and flow field. The experiments were focused on supersonic jet with Mach number up to 1.39, issuing from a rectangular nozzle with high aspect ratio. The experiments varied several parameters including main nozzle pressure ratio, total pressure of microjet, number of microjets and microjet injection angle. Far-field sound pressure measurement was performed, and the characteristics of noise reduction, including its directivity, were investigated. On the other hand, the flow field was visualized with a Schlieren technique in order to understand the mechanism of noise reduction. The unsteady behavior of the shock structure and the shear layer were investigated based on the visualization results. To investigate the effect of microjets on the 3-dimensional flow field, steady RANS analysis of the flow field was performed under various conditions of the main jet and the microjets.
Proceedings Papers
Proc. ASME. GT2013, Volume 1A: Combustion, Fuels and Emissions, V01AT04A046, June 3–7, 2013
Paper No: GT2013-94677
Abstract
With increasing focus on environmental effects and the need for fuel diversity in gas turbines, good liquid atomization is increasingly important. It is known that impinging atomization is able to produce fine drops by impingement of fast liquid jets. However, the atomization characteristics deteriorate at lower injection velocities. In this study, for improving atomization characteristics under a wide range of injection velocity, an effective technique is verified utilizing a small amount of gas (microjet) injection. The microjet is supplied from a pressurized reservoir independent of the liquid supply system, and it is injected from the center of the liquid nozzles toward the impingement point. To clarify the flow field and the mechanism of the effectiveness, experimental visualizations and drop size measurements are carried out. It is found that atomization is remarkably promoted when the dynamic pressure of microjet overcomes that of the liquid at the impingement point. By the microjet injection with only 1% of liquid mass flow rate, Sauter mean diameter (SMD) becomes one-tenth of the original SMD. In addition, optimized atomization efficiency is successfully achieved when the dynamic pressure of the microjet is two times that of the liquid at the impingement point.
Proceedings Papers
Proc. ASME. GT2013, Volume 6A: Turbomachinery, V06AT35A019, June 3–7, 2013
Paper No: GT2013-94988
Abstract
The effect of circumferential single grooved casing treatment on the stability enhancement of NASA Rotor 37 has been examined with CFD analysis. Stall inception mechanism of Rotor 37 is presented first with principal focus on the tip leakage flow behavior, passage blockage, and the vortical flow structures. Detailed observation showed that the combined interaction of the stagnated flow of tip leakage vortex breakdown and the jet-like leakage flow from the mid-chord region leads to the blade tip-initiated stall inception. The result of numerical parametric study is then demonstrated to show the effect of varying the axial location and the depth of a circumferential single groove. The evaluation based on stall margin improvement showed a higher potential of deeper grooves in stability enhancement, and the optimal position for the groove to be located was indicated to exist near the leading edge of the blade.
Journal Articles
Journal:
Journal of Turbomachinery
Article Type: Research-Article
J. Turbomach. March 2014, 136(3): 031017.
Paper No: TURBO-13-1132
Published Online: October 25, 2013
Abstract
The effect of circumferential single grooved casing treatment on the stability enhancement of NASA Rotor 37 has been examined with computational fluid dynamics analysis. Stall inception mechanism of Rotor 37 is presented first with principal focus on the tip leakage flow behavior, passage blockage, and the vortical flow structures. Detailed observation showed that the combined interaction of the stagnated flow of tip leakage vortex breakdown and the jetlike leakage flow from the midchord region leads to the blade tip-initiated stall inception. The result of numerical parametric study is then demonstrated to show the effect of varying the axial location and the depth of a circumferential single groove. The evaluation based on stall margin improvement showed a higher potential of deeper grooves in stability enhancement, and the optimal position for the groove to be located was indicated to exist near the leading edge of the blade.
Proceedings Papers
Proc. ASME. GT2012, Volume 1: Aircraft Engine; Ceramics; Coal, Biomass and Alternative Fuels; Controls, Diagnostics and Instrumentation, 173-183, June 11–15, 2012
Paper No: GT2012-68821
Abstract
Jet noise reduction is essential for environmentally-friendly civil transport. Since jet noise becomes very intense in the case of supersonic aircraft, noise reduction is crucial topic for the realization of next-generation supersonic transport. In the present study, experimental investigations were performed to clarify the effect of microjet injection on supersonic jet noise and flow field. The experiments were focused on supersonic jet with Mach number up to 1.47, which was generated from a rectangular nozzle with high aspect ratio. Far-field acoustic measurements were conducted for widely ranged microjet conditions to understand the influence of the condition on characteristics of supersonic jet noise and flow field. For understanding the unsteady behavior of the flow field and the relation with noise reduction, flow field visualization was performed with schlieren technique using a high-speed camera.
Proceedings Papers
Proc. ASME. GT2012, Volume 2: Combustion, Fuels and Emissions, Parts A and B, 1527-1536, June 11–15, 2012
Paper No: GT2012-70087
Abstract
A consistent theoretical model is proposed and validated for calculating droplet diameters and size distributions. The model is derived based on the energy conservation law including the surface free energy and the Laplace pressure. Under several hypotheses, the law derives an equation indicating that atomization results from kinetic energy loss. Thus, once the amount of loss is determined, the droplet diameter is able to be calculated without the use of experimental parameters. When the effects of ambient gas are negligible, injection velocity profiles of liquid jets are the essential cause of the reduction of kinetic energy. The minimum Sauter mean diameter produced by liquid sheet atomization is inversely proportional to the injection Weber number when the injection velocity profiles are laminar or turbulent. A non-dimensional distribution function is also derived from the mean diameter model and Nukiyama-Tanasawa’s function. The new estimation methods are favorably validated by comparing with corresponding mean diameters and the size distributions, which are experimentally measured under atmospheric pressure.
Proceedings Papers
Motoaki Utamura, Hiroshi Hasuike, Kiichiro Ogawa, Takashi Yamamoto, Toshihiko Fukushima, Toshinori Watanabe, Takehiro Himeno
Proc. ASME. GT2012, Volume 3: Cycle Innovations; Education; Electric Power; Fans and Blowers; Industrial and Cogeneration, 155-164, June 11–15, 2012
Paper No: GT2012-68697
Abstract
Power generation with a supercritical CO 2 closed regenerative Brayton cycle has been successfully demonstrated using a bench scale test facility. A set of a centrifugal compressor and a radial inflow turbine of finger top size is driven by a synchronous motor/generator controlled using a high-speed inverter. A 110 W power generating operation is achieved under the operational condition of rotational speed of 1.15kHz, CO 2 flow rate of 1.1 kg/s, and respective thermodynamic states (7.5 MPa, 304.6 K) at compressor and (10.6 MPa, 533 K) at turbine inlet. Compressor work reduction owing to real gas effect is experimentally examined. Compressor to turbine work ratio in supercritical liquid like state is measured to be 28% relative to the case of ideal gas. Major loss of power output is identified as rotor windage. It is found the isentropic efficiency depends little on compressibility coefficient. Off design performance of gas turbine working in supercritical state is well predicted by a Meanline program. The CFD analysis on compressor internal flow indicates that the presence of backward flow around the tip region might create a locally depressurized region leading eventually to the onset of flow instability.
Proceedings Papers
Proc. ASME. GT2011, Volume 1: Aircraft Engine; Ceramics; Coal, Biomass and Alternative Fuels; Wind Turbine Technology, 233-242, June 6–10, 2011
Paper No: GT2011-46150
Abstract
Jet noise reduction is essential for realization of environmentally-friendly and highly-efficient supersonic jet engines for future civil transport. In the present study, experimental and numerical investigations were conducted to clarify the effect of microjet injection on supersonic jet noise. The experiments were focused on supersonic jet with Mach number up to 1.49 that was generated from a rectangular nozzle with high aspect ratio. Far field acoustic measurements were executed and the spectra and sound pressure data of jet noise were obtained. In order to understand the mechanism of noise reduction, flow field visualization was performed with shadowgraph technique. CFD analysis was conducted as well to observe the flow field and to estimate thrust loss due to the microjet injection.