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ASME Conference Presenter Attendance Policy and Archival Proceedings

2016;():V02AT00A001. doi:10.1115/GT2016-NS2A.
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This online compilation of papers from the ASME Turbo Expo 2016: Turbomachinery Technical Conference and Exposition (GT2016) represents the archival version of the Conference Proceedings. According to ASME’s conference presenter attendance policy, if a paper is not presented at the Conference by an author of the paper, the paper will not be published in the official archival Proceedings, which are registered with the Library of Congress and are submitted for abstracting and indexing. The paper also will not be published in The ASME Digital Collection and may not be cited as a published paper.

Commentary by Dr. Valentin Fuster

Axial Flow Fan and Compressor Aerodynamics

2016;():V02AT37A001. doi:10.1115/GT2016-56002.

This paper presents a constrained Repetitive Model Predictive Controller (RMPC) implemented as closed-loop flow controller for an experimental compressor stator cascade. The objective of the controller is to decrease the impact of periodic disturbances on the passage flow. The disturbances are generated by movable flaps that are located downstream of the trailing edges of the stator vanes. The flaps emulate the throttling effect of periodically closed combustion tubes in a pulsed detonation engine on the flow over the stator vanes. The RMPC adjusts the actuation amplitude of fluidic sidewall actuators according to the present state of the passage flow. The current flow situation is monitored by pressure sensors that are mounted flush to the surface of one of the stator vanes. This data is fed back in real-time to the RMPC which thereupon modifies the actuation amplitude. By learning from period to period, a control command trajectory is computed that reduces detrimental effects of the periodic disturbance in an optimal manner while respecting the input constraints of the physical system. Five-hole-probe measurements in the wake of the passage are utilized to compare the optimized, transient actuation trajectories to the case of constant amplitude actuation.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A002. doi:10.1115/GT2016-56024.

During the operation of a jet engine, deterioration occurs. This constantly affects the engine performance parameters like exhaust gas temperature (EGT) and thrust specific fuel consumption (TSFC). If the EGT reaches a given limit, the engine has to be overhauled during a shop visit at a MRO company (Maintenance, Repair and Overhaul). Using the example of the high pressure compressor (HPC), the airfoils get analyzed for a few geometric properties and classified as serviceable, repairable and non-repairable. The repairable airfoils go through a repair process without in detail considering the actual geometry. To improve the repair process, tailored maintenance actions are desirable. For this purpose, the aerodynamic properties of the airfoil shall be the key factor for defining the repair actions. Therefore, geometric properties with high influence on the aerodynamic performance have to be known to reduce the amount of measuring time.

This paper will present a Design of Experiments (DoE) for HPC-airfoil geometry variations. Therefore, 550 different stage setups will be generated, simulated and analyzed. The database will be imported to a Kriging Method to generate a meta-model. Afterwards, the impact of the different geometric properties on the aerodynamic performance, like pressure-, work- and loss coefficients, will be analyzed by using the meta-model. The most important parameters will be determined and their impact on the flow will be explained.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A003. doi:10.1115/GT2016-56043.

Flow in turbomachines is generally highly turbulent. The boundary layers, however, often exhibit laminar-to-turbulent transition. Relaminarization from turbulent to laminar flow may also occur. The state of the boundary layer is important since it strongly influences transport processes like skin friction and heat transfer.

It is therefore vitally important for the designer to understand the process of laminar-to-turbulent transition and to determine the position of transition onset and the length of the transitional region. In order to better understand transition and relaminarization it is helpful to study simplified test cases first. Therefore, in this paper the flow along a flat plate is experimentally studied to investigate laminar-to-turbulent transition.

Measurements were performed for the different free-stream velocities of 5 m/s and 10 m/s. Several measurement techniques were used in order to reliably detect the transitional zone: the Preston tube, hot wire anemometry, thermography and Laser Interferometric Vibrometry (LIV). The first two measurement techniques are extensively in use at the institute ITTM and by other research groups. They are therefore used as a reference for validating the LIV measurement results.

An advantage of the LIV technique is that it does not need any seeding of the fluid and that it is non-intrusive. Therefore this measurement technique does not influence the flow, and it can be used in narrow flow passages since there is no blockage, in contrast to probe-based measurement techniques.

Further to the measurements, computational simulations were performed with the Fluent® and CFX® codes from ANSYS®, as well as with the in-house code Linars. The Menter SST k-ω turbulence model with the γ-ReΘ transition model was used in order to test its capability to predict the laminar-to-turbulent transition.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A004. doi:10.1115/GT2016-56045.

The need of increased stall margin is very high for aero gas turbine engines, as they operate under varied operating conditions. A number of different options are being used to increase the stall margin of gas turbine engines. Circumferential casing groove, in the compressor section of a gas turbine engine, is one of such methods. Incorporation of the grooves on the shroud increases the stall margin of the compressor, but this generally gives rise to loss of performance, such as efficiency and pressure ratio. By employing 3D blading techniques for rotor blades as well as stator vanes, performance of a compressor can be increased. 3D blading helps in reducing secondary flow losses and hence increased performance. Sweep and lean are examples of 3D blading, which is very common in any modern gas turbine compressor. A number of literatures are available in public domain, giving detailed understanding of effect of circumferential casing grooves and 3D blade features, but the interaction effect of sweep and casing grooves are not well published in public domain literature. In this work, an effort is made to understand, numerically, the interaction effect of sweep with circumferential grooves, using Computational Fluid Dynamics (CFD). Any numerical tool needs thorough validation before the results of numerical analyses can be used for analyzing the underlying physics. NASA Rotor37 is used to validate current CFD methodology. Mesh sensitivity is carried out to get mesh independence solution. Different turbulence models are used to get the best turbulence model for the problem in hand. 1D averaged performance data as well as hub to shroud variation of various flow parameters are compared to have full confidence on the CFD methodology.

A baseline axial compressor rotor, without sweep and lean is generated, as the first step of this study. This rotor is created by using hub and tip profiles of NASA Rotor37. The profiles are stacked along a radial line through their center of gravities, which has resulted in rotor geometry without any sweep and lean. Modifications are done to the tip profile of the baseline rotor, in terms of stagger angle, to get comparable performance w.r.t. NASA Rotor37. Casing of the NASA Roto37 is used as the redesigned compressor casing. Circumferential casing grooves, with five grooves between leading edge to trailing edge, are created as per industry standards. Meshing and modeling are done according to the best practices developed while validating CFD methodology. It is to be noted that the casing grooves and the main flow domain are meshed with one to one mesh connectivity, in order to avoid any numerical losses due to interface interpolations. This is considered very critical in this work, as the vortices from the tip is expected to have a strong interaction with grooves. This interaction is expected to create high gradients of flow variables in this region. Valuable flow information might be lost, if flow variables are interpolated in this region.

Baseline rotor is analyzed with and without casing grooves from choke to stall at 100% corrected speed. As expected, introduction of casing grooves has resulted in increased stall margin. A number of rotor geometries are created with different amount of sweeps. In the current study, blades are swept in the direction of chord, in order to avoid introduction of any sweep induced lean. The span location, where sweep starts, is also changed to understand the localized and global effect of this blade design features. Results obtained from numerical simulations of these geometries are presented in this paper. The performance and flow features are compared with respect to baseline rotor, with and without circumferential grooves, in an attempt to understand the underlying flow physics.

Topics: Compressors
Commentary by Dr. Valentin Fuster
2016;():V02AT37A005. doi:10.1115/GT2016-56050.

Effects of a large rotor tip gap on the performance of a one and half stage axial compressor are investigated in detail with a numerical simulation based on LES and available PIV data. The current paper studies the main flow physics, including why and how the loss generation is increased with the large rotor tip gap. The present study reveals that when the tip gap becomes large, tip clearance fluid goes over the tip clearance core vortex and enters into the next blade’s tip gap, which is called double-leakage tip clearance flow. As the tip clearance flow enters into the adjacent blade’s tip gap, a vortex rope with a lower pressure core is generated. This vortex rope breaks up the tip clearance core vortex of the adjacent blade, resulting in a large additional mixing. This double-leakage tip clearance flow occurs at all operating conditions, from design flow to near stall condition, with the large tip gap for the current compressor stage. The double-leakage tip clearance flow, its interaction with the tip clearance core vortex of the adjacent blade, and the resulting large mixing loss are the main flow mechanism of the large rotor tip gap in the compressor. When the tip clearance is smaller, flow near the end wall follows more closely with the main passage flow and this double-leakage tip clearance flow does not happen near the design flow condition for the current compressor stage. When the compressor with a large tip gap operates at near stall operation, a strong vortex rope is generated near the leading edge due to the double-leakage flow. Part of this vortex separates from the path of the tip clearance core vortex and travels from the suction side of the blade toward the pressure side of the blade. This vortex is generated periodically at near stall operation with a large tip gap. As the vortex travels from the suction side to the pressure side of the blade, a large fluctuation of local pressure forces blade vibration. Non-synchronous blade vibration occurs due to this vortex as the frequency of this vortex generation is not the same as the rotor. The present investigation confirms that this vortex is a part of separated tip clearance vortex, which is caused by the double-leakage tip clearance flow.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A006. doi:10.1115/GT2016-56087.

The paper reports on numerical investigations into the effects of inlet boundary layer skew on the aerodynamic performance of a low aspect ratio, high turning compressor cascade. The cascade geometry corresponds to the hub section geometry of a low aspect ratio stator of a highly loaded single-stage axial-flow low-speed compressor (fan). The skewed cascade flow simulates the reenergized stator hub flow and brings about a much better performance with inlet skew than without. However, the performance improvements caused by a skewed inlet boundary layer decrease with increasing inlet angle.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A007. doi:10.1115/GT2016-56208.

Inflow distortions in the compression system of a jet engine are becoming increasingly important for research focus. The investigation of the emergence of a distortion, its interaction with the rotor and the resulting impact on the rotor flow is challenging. In this work a separation in the inflow of a transonic compressor was created and the impact on stage aerodynamics investigated. The separation resulted in a total pressure distortion close to the casing within a sector of 120°. Effects were studied both numerically and experimentally in a joint collaboration project. The numerical model consisted of the full rotor-stator compressor stage, the inlet duct and the distortion generator upstream of the stage. This enables both an accurate validation of the numerical results and contributes to a deeper understanding of the flow. The results of both the numerical and experimental studies were in good agreement. The rotor is locally throttled by the inlet separation, resulting in the formation of an additional loss core at the stability limit due to a local aerodynamic overload. Considering classic distortion descriptors like the DC60, it is shown that they are not able to adequately assess the impact of a strong, but small distortion close to the tip of the rotor. The data can be considered as test case for future numerical models as well as for the validation of new analytical models. Furthermore, the results of this study reveal effects in both experimental and numerical studies that would not be realized if only a model of the separation was analyzed.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A008. doi:10.1115/GT2016-56232.

Tip clearance has great influence on the performance of multistage axial compressors including efficiency, pressure rise, mass flow, as well as matching. This paper reports a study into the influence of tip clearance on the performance and matching of a 5-stage axial compressor by a numerical method. Different tip clearances from 0% to 5.0% span which represents the typical range of tip clearance in modern multistage axial compressors were simulated and analyzed. The results show that as tip clearance increases from 0% to 5.0% span, the choked mass flow decreases by about 21.8%, the peak pressure ratio decreases by about 43.1% and the peak efficiency decreases about 14.3 percents. As tip clearance increases, the efficiency of the whole compressor decreases in a parabolic manner not linearly as previous suggested, which is partially attributable to the cantilevered stators considered in this paper and primarily due to the mismatching of different stages. It is of great importance to control the tip clearance. When tip clearance increases, the front stage tends to work near surge condition and the rear stage tends to work near choke condition, which leads to lower efficiency than in the middle stages. A weight was defined to evaluate each stage’s contribution to the whole compressor’s efficiency deficit caused by the increase of tip clearance. Front and rear stages contribute more to the efficiency deficit than the middle stages, which indicates that more attention should be paid on front and rear stages to improve the performance of multistage axial compressors. In order to evaluate the matching of multistage axial compressors with a quantified method, a new parameter named “Peak Efficiency Deviation (PED)” was defined based on the difference between each stage’s operating efficiency and its peak efficiency. The mass flow of multistage axial compressors should be well considered to make the PED parameter to be close to zero as possible. In the most commonly used range of tip clearance from 0.5% to 3.0% span, the PED varies little within 0.4 percent, which is only about 8.4% of the peak efficiency deficit at 1.5% span tip clearance. So, the PED could be small within a wide range of tip clearances if the matching of the compressor is perfect at design tip clearance.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A009. doi:10.1115/GT2016-56403.

An experimental study on the use of synthetic jet actuators for lift control on a generic compressor airfoil is conducted. A wind tunnel model of a NACA 65(2)-415 airfoil, representative of the cross section of an Inlet Guide Vane (IGV) in an industrial gas compressor, is 3D-printed. Nine synthetic jet actuators are integrated within a planar wing section with their slots covering 61% of pressure side of the airfoil span, located 13% chord upstream of the trailing edge. The Helmholtz frequency of the slot is matched closely with the piezoelectric element material frequency. The slot is designed so that the bi-morph actuation creates a jet normal to the airfoil surface. By redirecting or vectoring the shear layer at the trailing edge, the synthetic jet actuator increases lift and decreases drag on the airfoil without a mechanical device or flap. Tests are performed at multiple Reynolds number ranging from Re=150,000 to Re=450,000. The increased lift of the integrated synthetic jet actuator is dependent on the Reynolds number and free stream velocity, the actuation frequency, and angle of attack. For actuation at 1450 Hz the synthetic jet actuator increases lift up to 7%. The synthetic jet increases L/D up to 15%. Velocity contours obtained through PIV show that the synthetic jet turns the trailing edge shear layer similar to a Gurney flap.

Topics: Actuators
Commentary by Dr. Valentin Fuster
2016;():V02AT37A010. doi:10.1115/GT2016-56440.

Reducing the fuel consumption is a main objective in the development of modern aircraft engines. Focusing on aircraft for mid-range flight distances, a significant potential to increase the engines overall efficiency at off-design conditions exists in reducing secondary flow losses of the compressor. For this purpose, Active Flow Control (AFC) by aspiration or injection of fluid at near wall regions is a promising approach.

To experimentally investigate the aerodynamic benefits of AFC by aspiration, a 4½-stage high-speed axial-compressor at the Leibniz Universitaet Hannover was equipped with one AFC stator row. The numerical design of the AFC-stator showed significant hub corner separations in the first and second stator for the reference configuration at the 80% part-load speed-line near stall. Through the application of aspiration at the first stator, the numerical simulations predict the complete suppression of the corner separation not only in the first, but also in the second stator. This leads to a relative increase in overall isentropic efficiency of 1.47% and in overall total pressure ratio of 4.16% compared to the reference configuration.

To put aspiration into practice, the high-speed axial-compressor was then equipped with a secondary air system and the AFC stator row in the first stage. All experiments with AFC were performed for a relative aspiration mass flow of less than 0.5% of the main flow. Besides the part-load speed-lines of 55% and 80%, the flow field downstream of each blade row was measured at the AFC design point. Experimental results are in good agreement with the numerical predictions. The use of AFC leads to an increase in operating range at the 55% part-load speed-line of at least 19%, whereas at the 80% part-load speed-line no extension of operating range occurs.

Both speed-lines, however, do show a gain in total pressure ratio and isentropic efficiency for the AFC configuration compared to the reference configuration. Compared to the AFC design point, the isentropic efficiency ηis rises by 1.45%, whereas the total pressure ratio Πtot increases by 1.47%. The analysis of local flow field data shows that the hub corner separation in the first stator is reduced by aspiration, whereas in the second stator the hub corner separation slightly increases. The application of AFC in the first stage further changes the stage loading in all downstream stages. While the first and third stage become unloaded by application of AFC, the loading in terms of the De-Haller number increases in the second and especially in the fourth stage. Furthermore, in the reference as well as in the AFC configuration, the fourth stator performs significantly better than predicted by numerical results.

Topics: Compressors
Commentary by Dr. Valentin Fuster
2016;():V02AT37A011. doi:10.1115/GT2016-56455.

For most technical applications, simulations of the Reynolds-Averaged-Navier-Stokes equations has become a standard analysis tool, since it brings a good compromise between computational accuracy and costs. However, turbulence models have to be implemented to close the system of differential equations. To study the effects of turbulent boundary conditions on the prediction of the secondary flow field in a linear compressor cascade with tip clearance, the state of the art RANS solver TRACE in conjunction with Wilcox’ k-ω-turbulence model is used. Besides a stagnation point anomaly prevention, no turbomachinery specific modifications of the turbulence model are applied. Transition is not considered. The current investigations focus on the influence of the imposed turbulent inlet quantities (k0, ω0) on the development of the wall-bounded flow in the cascade. The turbulent kinetic energy k is basically described as a function of the turbulence intensity level measured in an equivalent experimental setup. For the reconstruction of turbulent fluctuations beyond measuring accessibility in the vicinity of the wall, an analytical approach is proposed and validated with DNS data of turbulent flat plate and fully developed channel flows. To identify the influence of different dissipation rates ω on the characteristics of the secondary flow, the free stream turbulent length scale LtFS is varied in four steps ranging from 0.2 up to 5 millimeters. Additionally, the effects of different span-wise length scale distributions across the inlet flow boundary layer are considered.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A012. doi:10.1115/GT2016-56469.

Interpolation is a common procedure in scientific practice. In many cases, linear interpolation is tacitly applied in order to present measurements or numerical results. The application of sophisticated interpolation methods is rather rare, even though they can help to extract more information from a measurement, especially if the resolution is relatively low. This paper presents an interpolation method and a validation procedure, which is demonstrated on experimental data from a low pressure compressor of a twin spool turbofan engine.

As a research test vehicle the Larzac 04 turbofan engine is used, which is equipped with extensive instrumentation and control systems. This publication focuses on the high speed pressure transducers mounted directly above the tip of the first LPC rotor blades. Besides the installation of those sensors in a jet engine, the processing, the storage, and the analysis of the data is challenging, as the sampling rate is extraordinary high (1 MHz). Those sampling rates result in a high circumferential resolution (1 MHz ≙ 2.69 samples per mm at 90% rel. spool speed) compared to the relatively low geometrical resolution of the nine axial sensors (0.23 samples per mm). Therefore, sophisticated interpolation methods are applied on the data points in axial direction, aiming to get the best possible result from the acquired data. A major part of this paper is the discussion and validation of several different interpolation methods. Especially with the Akima sub-spline interpolation, very promising results were obtained.

Finally, some of the results from the fast response pressure transducers are presented. The casing wall pressure distribution and some flow phenomena at the design operating point are compared to the corresponding near stall conditions.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A013. doi:10.1115/GT2016-56519.

Highly irreversible flows can arise from the rotor tip region which cause compressor performance degradation. The understanding of the flow irreversibility and the spatial distribution in the rotor passage may provide hints to retrofit rotor geometries as well as design efficiency-friendly casing treatments. This paper presents a numerical investigation on the spatial distributions of entropy generation through a control volume analysis on a transonic compressor. The typical loss sources can be classified as the viscous shear on solid boundaries, the shockwaves at the leading edge and within the passage, the tip leakage vortex and the turbulence mixing in the rotor wake in the tip region, however, it is hard to distinguish their individual contributions because they always interact with each other. To avoid this difficulty, this paper targets at the spatial distribution of entropy generation since it can tell the local losses quantitatively as the results of local flow structures and their interactions. Tip clearance is used as a controlled parameter for the investigation. By varying the tip clearance over the blade chord, the tip leakage flow is energized and thus examined first, followed by the other changes of inflow structures reacting to the tip clearance variation. The changes in entropy generation distributions are then carefully compared. Based on the results, the dominative contributor on entropy generation and its impact on total loss are identified. Finally, new casing designs with stepped tip gaps that may ameliorate the entropy generation of the tip region are discussed and numerically validated.

Topics: Compressors , Entropy
Commentary by Dr. Valentin Fuster
2016;():V02AT37A014. doi:10.1115/GT2016-56554.

Axial compressors can obtain substantial improvement on stall margin by using axial-slot casing treatments. However, this type of casing treatment usually yields large peak efficiency penalty due to the interaction between the slots and rotor tip region where the tip leakage flow plays an important role. Therefore, as a main factor that influences the peak efficiency, the tip leakage loss was examined in this paper with a variety of slot geometries. Unsteady numerical simulations were performed on both low speed and transonic compressors with axial skewed slot casing treatments with different geometric parameters. In addition, an equation which can be applied to evaluate the tip leakage loss under casing treatment cases was derived from Denton’s leakage mixing model. The leakage loss can be expressed in terms of the cube of the tip leakage flow rate. Combined with the simulation results, the effects of the number, depth and width of the slots on both the leakage loss and peak efficiency deficit were investigated. For the transonic compressor, the impacts of shock wave and its interaction with the tip leakage flow /vortex were assessed as well. Lastly, two axial-slot casing treatments with an isosceles-trapezoid shaped opening were designed to reduce the loss in the rotor tip region. It was shown that the newly designed axial-slot casing treatments were capable of improving the peak efficiency of both compressors.

Topics: Compressors
Commentary by Dr. Valentin Fuster
2016;():V02AT37A015. doi:10.1115/GT2016-56577.

This paper aims for the analysis of experimental and numerical results of windmilling flow topologies far from freewheeling condition. Two fans were investigated: a baseline design and an innovative one meant to reach good performance in both compressor and turbine modes. Experiments are conducted with global and local characterizations to determine energy recovery potential and local loss mechanisms. The numerical study is carried out with mixing plane steady simulations, the results of which are in fair agreement with experimental data. The difference of local topology between freewheeling and highly loaded windmill demonstrates that classical deviation rules such as Carter’s are not well-suited to highly loaded windmilling flows. Finally, under certain conditions, the minor influence of the stator on the rotor topology indicates that non rotating elements can be considered as loss generators.

Topics: Flow (Dynamics) , Fans
Commentary by Dr. Valentin Fuster
2016;():V02AT37A016. doi:10.1115/GT2016-56589.

Humpback whales possess bumpy tubercles on the leading edge of their flippers. Due to these leading edge tubercles, the whales are able to perform complex underwater maneuvers agilely. Inspired by the flippers, this paper applies sinusoidal-like tubercles to the leading edge of the blade in an annular compressor cascade, and presents a numerical investigation to explore the effects of tubercles with the aim of controlling the corner separation and reducing losses.

A preliminary study by steady 3D RANS simulations is performed. The aerodynamic performance and the behavior of the corner separation are investigated in the baseline compressor cascade. Subsequently, cascades with leading edge tubercles are numerically simulated. A crucial geometry parameter of the tubercles, wavelength, is varied to obtain different configurations. The influence of the parameter is concluded from the comparison of the performance attained by these configurations. Also, several configurations, which are typical in loss characteristics, are selected for further DES simulations so as to obtain more flow details, especially at the separation region. Flow visualizations show that leading edge tubercles could induce the formation of counter-rotating streamwise vortices. The interaction between the streamwise vortices and corner separation is emphatically investigated. By analysis of all the results obtained, this paper tries to figure out the mechanism of leading edge tubercles in loss reduction and separation delay in an annular compressor cascade.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A017. doi:10.1115/GT2016-56681.

It is known that the tip-clearance gap plays a pivotal role in determining the performance of an aero-engine compressor, both in terms of stability and pressure rise. However, the exact effect of tip-clearance size on stall margin over the range of clearances experienced during engine operation is not known, and most designers substitute rules of thumb for real knowledge when developing new compressors.

Eccentricity in the tip-clearance is also known to affect compressor performance, though again little work has been done to quantify the penalties. It is generally assumed that the stall margin of an eccentric machine will be approximately equal to that of a concentric machine with a clearance equal to the maximum eccentric clearance. Results given in this paper show a stabilising effect of the small tip-clearance sector of an eccentric compressor on the large tip-clearance sector, so the penalty on stall margin is not as large as commonly assumed. It is shown that the stall margin penalty in a single-stage eccentric machine is only 50–60% of that which a concentric compressor with clearance equal to the maximum clearance would exhibit. In addition, the effect of eccentricity is shown to diminish as the average clearance is increased.

In this paper, experiments and computational modelling are used to examine the effects of eccentric tip-clearance on flowfield redistribution and compressor performance. In particular, the three-dimensional nature of the flowfield generated by an eccentric gap is shown for the first time. The purpose of this work is not to provide ‘hard-and-fast’ design rules for eccentric compressors (this cannot be done on the basis of single-stage measurements), but to provide a starting point for a better physical understanding of the problem.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A018. doi:10.1115/GT2016-56682.

This paper presents an optimized rotor as part of a 3-blade row optimization (IGV-rotor-stator) of a high-pressure compressor. It is based on modifying blade angles and advanced control of curvature of the airfoil camber line. The effects of these advanced blade techniques on the performance of the transonic 1.5-stage compressor were calculated using a 3D Navier-Stokes solver combined with a vortex/vorticity dynamics diagnosis method. The optimized rotor produces a 3-blade row efficiency improvement over the baseline of 1.45% while also improving stall margin. The throttling range of the compressor is expanded largely because the shock in the rotor tip area is further downstream than that in the baseline case at the operating point. Additionally, optimizing the 3-blade row block while only adjusting the rotor geometry ensures good matching of flow angles allowing the compressor to have more range. The flow diagnostics of the rotor blade based on vortex/vorticity dynamics indicate that the boundary-layer separation behind the shock are verified by on-wall signatures of vorticity and skin-friction vector lines. In addition, azimuthal vorticity and boundary vorticity flux (BVF) are shown to be two vital flow parameters of compressor aerodynamic performance that directly relate to the improved performance of the optimized transonic compressor blade.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A019. doi:10.1115/GT2016-56752.

A linear to nonlinear transition during the spike stall process of an axial flow compressor rotor is presented. Recently, some researchers thought that spike stall inception is directly induced by the tip leakage flow. However, the authors utilized unsteady full annular simulations and found that a second-order disturbance appeared two revolutions before the breakdown of the tip leakage flow in an axial rotor, associate with spike stall inception while the tip leakage flow is still stable. This second-order disturbance grew rapidly in the next two revolutions and the process was unlike the low order disturbance development in modal stall inception. The response of the compression system was still linear in this process. The rapidly developing second-order disturbance made the tip leakage flow unstable, leading to the start of spike stall inception. The response of the compression system became nonlinear in this process.

Topics: Rotors , Axial flow
Commentary by Dr. Valentin Fuster
2016;():V02AT37A020. doi:10.1115/GT2016-56756.

Recirculating casing treatment (RCT) was studied in a subsonic axial flow compressor experimentally and numerically. The RCT was parameterized with the injector throat height and circumferential coverage percentage (ccp) to investigate its influence on compressor stability and on the overall performance in the experimentation. The injector throat height varied from 2 to 6 times the height of the rotor tip clearance, and the ccp ranged from 8.3% to 25% of the casing perimeter. Various RCT configurations were achieved with a modular design procedure. The rotor casing was instrumented with fast-response pressure transducers to detect the stall inception, rotational speed of stall cells, and pressure flow fields. Whole-passage unsteady simulations were also implemented for the RCT and solid casing to understand the flow details. Results indicate that both the compressor stability and overall performance can be improved through RCT with appropriate geometrical parameters. The effect of injector throat height on the stability depends on the choice of ccp, i.e., interaction effect exists. In general, the RCT with a moderate injector throat height and a large circumferential coverage is the optimal choice. Phase-locked pattern of the casing wall pressure reveals a weakened tip leakage vortex under the effect of RCT compared with the solid casing. The numerical results show that the RCT has a substantial effect on tip blockage even when the blade passages break away from the domain of RCT. The reduction of tip blockage induced by the tip leakage vortex is the main reason for the extension of stable operation range. The unsteadiness of double-leakage flow is detected both in the experiment and in numerical simulations. The pressure fluctuations caused by double-leakage flow are depressed with RCT. This observation indicates reduced losses related with the double-leakage flow. Although the stall inception is not changed by implementing RCT, the stall pattern is altered. The stall with two cells is detected in RCT compared with the solid casing with only one stall cell.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A021. doi:10.1115/GT2016-56757.

For compressor blade tip stall, one effective way of extending stable operating range is with the application of circumferential grooved casing treatment and its validity was proved by a lot of experimental and numerical investigations. The emphases of most circumferential grooved investigations are focused on the influence of groove depth and groove number on compressor stability, and there is few investigations dealt with the center offset degree of circumferential grooves casing treatment. Hence, an axial compressor rotor with casing treatment (CT) was investigated with experimental and numerical methods to explore the effect of center offset degree on compressor stability and performance.

In the work reported here, The center offset degree is defined as the ratio of the central difference between rotor tip axial chord and CT to the axial chord length of rotor tip. When the center of CT is located within the upstream direction of the center of rotor tip axial chord, the value of center offset degree is positive. The experimental and numerical results show that stall margin improvement gained with CT is reduced as the value of center offset degree varies from 0 to 0.33 or −0.33, and the CT with −0.33 center offset degree achieves the lowest value of stall margin improvement at 53% and 73% design rotational speed. The detailed analysis of the flow-field in compressor tip indicates that there is not positive effect made by grooves on leading edge of rotor blade tip when the value of center offset degree is −0.33. As the mass flow of compressor reduces further, tip clearance leakage flow results in the outlet blockage due to the absence of the positive action of grooves near blade tip tail when the value of center offset degree is 0.33. Blockage does not appear in rotor tip passage owing to utilizing the function of all grooves with CT of 0 center offset degree.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A022. doi:10.1115/GT2016-56795.

Plasma actuation is a novel method for axial compressor flow control with advantages of short response time and broad frequency range. Numerical simulation of tip leakage vortex control in a low speed axial compressor with pulsed plasma actuation is performed. Millisecond pulsed dielectric barrier discharge plasma actuation with different frequencies are generated on the inner wall of compressor casing at the rotor leading edge. Scale adaptive hybrid Reynolds-averaged Navier-Stokes/large eddy simulation method based on shear stress transport turbulence model is adopted. The plasma actuation is simplified as a body force in the simulation. Results show that the frequency has a strong influence on the control effect of pulsed plasma actuation. Pulsed plasma actuation with frequency of 0.25 blade passing frequency (BPF), 0.5 BPF and 1.0 BPF extend the compressor’s stability range effectively. The mechanism is tip leakage vortex oscillation in the stream wise direction through coupling between unsteady plasma actuation and tip leakage flow. However, pulsed plasma actuation with frequency of 0.125 BPF, 2 BPF and 3 BPF fails to improve the stability range. The mechanism of pulsed plasma actuation at 2 BPF and 3 BPF is similar to that with steady plasma actuation, which is only stream wise boundary layer acceleration. The oscillation of tip leakage vortex in the stream wise direction can’t occur. For the pulsed plasma actuation at 0.125 BPF, its frequency is too low to get enough control effect.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A023. doi:10.1115/GT2016-56830.

Numerical investigations on the control effects of synthetic jets are conducted upon a highly loaded compressor stator cascade. The influence of forcing parameters including actuation frequency, jet amplitude and slot location are analyzed in detail with the single-slit synthetic jet. Besides, a new slot arrangement is put forward for the purpose of effectively controlling flow separation.

Simulation results validate the remarkable effectiveness of the single-slit synthetic jet on controlling flow separation. Owing to the coupling effect between the jet and the main flow, the actuation appears to be most efficient under the characteristic frequency of the main flow passing through the airfoil. Additionally, with the increase of jet momentum coefficient, the control effect is enhanced at first and then decreased, depending on the two aspects: the improvements of aerodynamic performance by momentum injection and the additional flow losses caused by the jet. Compared to other actuator configurations, the segment synthetic jet with three sections can more effectively deflect the end-wall cross flow and thus impede the development of corner vortex, which helps to restrain the accumulation of low momentum fluid towards the corner, emphasizing the importance of slot arrangement. Accordingly, under the optimum condition, the total pressure loss coefficient gains a 15.8% reductions and the static pressure rise coefficient is increased by 5.01%.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A024. doi:10.1115/GT2016-56843.

This paper presents a numerical investigation of secondary flow control in a high speed compressor cascade for different incoming flow incidences by means of endwall vortex generator jets (VGJs). The inlet Reynolds number is 560,000 in corresponding to an inlet Mach number of 0.67. Based on the detail analysis of the flow field and cascade performance, two effect mechanisms of the vortex induced by the VGJ are proposed. The first is to enhance the mixing between the endwall boundary layer and the mainstream. The second is to block the cross flow as an air obstacle. Therefore, the low energy fluids accumulation in the corner region could be decreased significantly, weakening the separation on the suction side and reducing the losses effectively. This benefit becomes more obvious with the increase of the incidence from i = −2° to 4°. Additionally, a more uniform flow angle as well as static pressure profile along the blade height is obtained at the cascade outlet. The maximum loss reduction is up to 12.9% while i = 4° with a jet mass flow ratio of 0.2%. However, the unfavorable impact of the VGJs is also detected in the up-washed region, where the loss is increased by the mixing processes between the mainstream fluids and the low energy fluids. For the case i = −4°, a strengthened induced vortex is generated due to the increased angle between the jet and incoming flow, resulting in loss increase in the up-washed region. Besides, a more rapid corner boundary layer development appears in the rear part of the passage, contributing to severe separation and loss enhancement, which suggests that the VGJ should be switched off for this incidence. Therefore, the advice to the application of the VGJ according the incidence is further obtained.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A025. doi:10.1115/GT2016-56855.

Based on the previous research about the combined flow control method which was carried out by applying the endwall steady VGJ to the bowed compressor cascades to reduce the secondary flow loss, for the consideration that the pulsed jets may save the mass flow required for control, therefore the unsteady VGJs over the different actuation frequencies and blowing ratios were investigated in detail. Under the conditions of same jet geometry parameters, the improvements to the fluid fields in the bowed compressor cascades caused by the pulsed jets are less than that induced by the steady cases. With the pulsed VGJ, for the positively bowed blade, the enhancement of the time-averaged aerodynamic performance can be achieved when the blowing ratio is greater than 0.6, but all of the unsteady conditions in this research can improve the flow field in the negatively bowed blade. The time-averaged total losses decrease by 1.6% and 7.0% at most for the positively and negatively bowed blades, respectively. The mechanisms by which the endwall pulsed vortex generator jets delay flow separation and reduce loss were explored. The results show that, being different from the single vortex produced in steady VGJ, the pulsed case generates a pair of streamwise vortices with the opposite sense of rotation. One vortex suppresses the development of the secondary flow, but the other one increases the size of the passage vortex. Furthermore, for the endwall pulsed VGJ, the changes of the blowing ratio plays a more important role in improving the flow fields in the bowed cascades than that of the actuation frequency.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A026. doi:10.1115/GT2016-56888.

Aspirated compressor is a promising design concept to promote the working capacity of compression system and thus may result in an intensive interactions with adjacent rows, however, few extensive unsteady flow analysis have been conducted on this type of highly-loaded compressor. This work presents a three-dimensional numerical simulation of a low speed 1.5 stage low-reaction aspirated compressor (LRAC). The boundary layer suction is only implemented in the outlet vane to control the corner stall and boundary layer separation while no active flow control method is applied in the rotor. The total aspiration flow rate is around 3%. Both aspiration slot and plenum were integrated into the computational domain. Two operating points were selected with the aim to investigate the unsteady effects on the performance of the LRAC and to provide a preliminary unsteady description for this type of aspirated compressor. It is found that compared with the differences in 1D total aerodynamic parameters, evident departures are found in the radial distributions of stage outlet flow parameters between the steady and time-averaged results. For the unsteady case, the radial distributions of pitchwise-averaged parameters become more uniform due to the redistributed aspiration flow rate and the convection behavior of the rotor wake. For the aspiration scheme in the blade, although one-side-aspiration manner is applied, the aspiration flow rate presents a C-type distribution in the radial direction, and this tendency becomes more prominent for the forward slot and time accurate results. Besides, the fluctuation of aspiration flow rate is mainly focused on the upper span due to the intensive rotor outlet secondary flow here. Moreover, the potential effect of aspirated stator on rotor is also examined preliminarily. It is found that for the LRAC investigated in this paper, the rotor suction side is apt to be influenced by the downstream aspirated stator. Finally, some suggestions on the design of the aspirated compressor are provided.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A027. doi:10.1115/GT2016-56891.

This investigation discusses the impact of a non-steady outflow condition on the compressor stator flow in an annular cascade which is periodically chocked through a rotating disc in the wake, to simulate the expected conditions for a pulsed detonation engine (PDE). A 2D controlled diffusion airfoil of the highly loaded linear stator cascade by [1] has been transferred to the annular compressor test rig to compare results under non-steady conditions via multi-colored oil flow visualization on the suction side and pressure measurements in the wake of the blades. Three different Strouhal numbers of the choking device are investigated and analyzed by phase averaged pressure measurements downstream of the stator to visualize the unsteady flow characteristics. Triggered by the changed incidence angle due to the choking, separation on the suction side and in the hub region form a periodic event depending on the position of the blockage device. Active flow control (AFC) is implemented by means of side wall actuation at the hub to improve flow conditions. Pressure measurements show that the turning of the blades can be raised and a static pressure rise is gained by the AFC while periodic choking is active.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A028. doi:10.1115/GT2016-56934.

The pitch-wise uniformity of inlet flow was proved by experiment in low-speed plane cascade, and simulations were performed for investigation and improvement in this study after that. The deflector, which was connected to the blade strictly, was proved to be the cause of the problem according to the simulation, and unexpected inflection of inlet flow were proved at the same time. It was also proved that a gap, which was placed between the blade and the deflector, could weaken the influence of deflector. However, the leakage that caused by the gap would also lead to some negative effects that attentions must be paid to. Finally, with cascade moved half-pitch along x-axis and one extra blade added, more acceptable result between 1 and 4 relative pitch-wise was showed in one of the improved cases (Case B). Better inlet condition with more periodic pressure distribution and acceptable incident angle were gained.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A029. doi:10.1115/GT2016-56966.

Tip blowing and axial slot casing treatments have shown their ability to enhance the stability of a transonic axial compressor with different effects on efficiency. For an effective application of these casing treatments, a good knowledge of the influence of the casing treatment on the rotor flow field is important. There is still a need for more detailed investigations, in order to understand the interaction between the treatment and the near casing 3D flow field. For transonic compressor rotors this interaction is more complex, as super- and subsonic flow regions alternate while interacting with the casing treatment.

In the present study, an axial slot and a tip blowing casing treatment, which have been developed and optimized for the same tip critical transonic axial compressor rotor (reference rotor) by Streit et al. [1] and Guinet et al. [2], are subject of the investigation. Both casing treatment types showed their capabilities to enhance the compressor stability without losing by means of CFD simulations. Since the higher compressor stability allows a higher blade loading, Streit et al. reduced the blade number of the rotor. Thus, the efficiency was increased due to the reduction of friction losses. However, applying the tip blowing casing treatment to the reduced rotor shows a negative effect on the efficiency.

Both casing treatment types recirculate flow from a downstream to an upstream location of the rotor and reinject it to enhance the near casing flow field. Although the working principle of the two casing treatment types are similar, the transfer of the casing treatments from the reference to the reduced rotor show different trends in efficiency. Therefore, the effect of recirculation cannot explain the difference in efficiency. Hence, applying axial slots must include additional flow features, compared to recirculation channels.

Compensating effects as in circumferential groove casing treatments and other flow interactions between the near casing flow field and the slot flow are considered. These additional mechanisms of the axial slot casing treatment will be identified and isolated by comparing the two different casing treatment types. The numerical simulations are carried out on a 1.5 stage transonic axial compressor using URANS simulations.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A030. doi:10.1115/GT2016-57050.

The flows in the tip regions of two rotors with blades of similar geometry but different tip clearance are studied experimentally to determine the effect of gap on the flow structure at different operating conditions. The experiments have been performed in the JHU optically index-matched facility, where the refractive index of the fluid is matched with that of the acrylic rotor blades and casing, facilitating unobstructed Stereo Particle Image Velocimetry (SPIV) measurements. 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 tip gap sizes are 0.49% and 2.3% of the blade chordlength, and measurements are performed for two flow rates, the lower of which is just above stall conditions. The presence and trajectories of the tip leakage vortex (TLV) and secondary structures are visualized by recording high speed movies of cavitation at lower pressures. The results consist of performance curves, distributions of velocity, circumferential vorticity and turbulent kinetic energy, as well as the strength and trajectory of vortices. Increasing the tip gap reduces the static-to-static pressure coefficient for all flow conditions. For the higher flow rate, a wider tip gap has several effects: (i) It delays the rollup of the TLV and its detachment from the suction side (SS) corner of the blade, presumably due to the larger distance from the endwall casing and the ‘image vortex’. (ii) It alters the blade loading and reduces the circulation shed from the blade. (iii) It delays the onset of TLV bursting in the aft part of the rotor passage. (iv) For both gaps, the endwall boundary layer separates at the point where the leakage flow meets the opposite-direction main passage flow. For the wide gap, the separated layer with opposite sign vorticity remains above the TLV; while for the narrow gap, the TLV entrains this layer around itself. And (v) consistent with the major differences in flow structure, the spatial distributions and magnitudes of all the turbulence intensity are also very different. Trends and flow structure are quite different at pre-stall conditions. Most notably, TLV rollup is still delayed for the wide gap, but vortex bursting and associated arrival of multiple secondary structures to the pressure side (PS) of the next blade occur earlier. Consequently, the turbulence level on both sides of the blade tip is substantially higher, and remnants of the previous TLV are ingested into the next tip gap.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A031. doi:10.1115/GT2016-57054.

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.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A032. doi:10.1115/GT2016-57081.

Rotating stall is a primary limit to compressor performance, and the reasonable estimation of stall-onset point is very useful in compressor design. Extensive investigations have been conducted in the past few decades to develop analytical models and numerical methods for stall-onset prediction, and much progress has been made in the understanding of flow mechanism for rotating stall. In spite of the robust prediction ability of stall-onset condition, the unsteady 3-D computations are still time-consuming for industrial applications. Analytical models are able to provide a fast estimation of compressor stall onset. However, empirical correlations are usually needed in the analytical models, which leads to a decrement in the accuracy and application scope of the models. Especially for high speed compressors, tip clearance effects hasn’t been evaluated reasonably in the previous analytical models, which actually plays a very important role in determining the stall-onset point. Therefore, new analytical models accounting for tip clearance effects will be promising in estimating the stall-onset more precisely. It’s the requirement for a new analytical model that motivates the present work.

In the present work, the unsteady flow simulation of a transonic compressor rotor at near stall condition was performed to clarify the relations between tip clearance flow oscillations and compressor stall-onset in transonic axial compressor rotors. The interaction between tip clearance and incoming flow is simplified to a 2-D analogy of free-stream and counter-flow wall jet interaction. Momentum balance analysis is applied to identify the position of tip clearance/incoming flow interface, together with a prediction method of tip leakage vortex core trajectory. The effects of the in-passage shock on tip clearance/incoming flow interaction is taken into account by applying an upstream deflection of the interface, and this deflection is also observed in the computational flow field at near stall conditions. As a combination of the above-mentioned aspects, a model is proposed to define the critical point for tip clearance flow spillage from blade leading edge, which corresponds to the stall-onset point on compressor performance curves. Validations against numerical results prove that the model is capable of including tip clearance effects on stall-onset point. Parametric study of the model shows that blade tip offloads with increasing tip clearance, reducing the inverse momentum flux of tip clearance flow. As a result, the stalling flow coefficient appears to be less sensitive to tip clearance variation, which accords with the published experimental results.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A033. doi:10.1115/GT2016-57082.

In the current work, pressure distribution on the surface of blade and endwall is investigated experimentally in a compressor cascade for capturing some flow features by analyzing pressure profile aerodynamics. Pressure is measured using pressure sensitive paint (PSP) technique. Before the tests, some system errors including temperature-dependence of paint, in-situ calibration, and Self-illumination among neighboring structure are studied in details. In addition, the spatial layout scheme of LEDs and CCDs for each test is designed by self-developing optical theoretical method that based on the LTS (light Tracing Simulation). In the post-processing, an in-situ calibration is conducted to convert the intensity of luminescence to pressure, and each test uses independent calibration equation. Finally, discrete pressure is also measured by pressure taps to compare with PSP. All of the tests are performed at the AOA (Attack of Angle) of −2.5°∼10°, and Mach number of 0.4∼0.8. The results show that the pressure distribution at low Mach number presents a good PSP performance in terms of signal-to-noise-ratio, and apparent pressure step gradient is captured at high Mach number. Meanwhile, the pressure distribution at high incidence reveals the unsteady of separated flow. By combining the end-wall plot with profile plot, the influence of corner-flow on main flow in the passage is obtained.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A034. doi:10.1115/GT2016-57276.

In this paper, experimental investigations of a linear compressor cascade have been performed with Reynolds number of 2.4×105 to analyze the flow mechanism of hub-corner separation with end-wall jet and suction control. The vortices are measured with a quasi-three dimensional test system in different axial planes consisting of vorticity distribution and secondary flow structure. The experimental results give detailed insight into the performance of the principle vortices with different flow control methods. Corner separation losses could remarkably decrease with the jets and suction position near the asymptotic line of separation lines on suction surface. The flow control position plays a great role in affecting the corner separation losses while it is a more sensitive factor in the case of jets rather than the suction. It is evidenced that the combined flow control would get a higher decrease in the total pressure loss coefficient while an additional benefit in the reduction of losses has been gained in the case of a combined actuator layout.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A035. doi:10.1115/GT2016-57311.

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.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A036. doi:10.1115/GT2016-57333.

This paper presents a novel one-dimensional design method based on the radial equilibrium theory and constant span-wise diffusion factor to redesign of NASA rotor 67 just aerodynamically with a higher pressure ratio at the same design point. A one-dimensional design code is developed to obtain the meridional plane and blade to blade geometry of rotor to reach the three-dimensional view of rotor blades. To verify the redesigned rotor, its flow numerical simulation is carried out to compute its performance curve. The experimental performance curve of NASA rotor 67 is used for validation of the numerical results. Structured mesh with finer grids near walls is used to capture flow field and boundary layer effects. RANS equations are solved by finite volume method for rotating zones and stationary zones. The numerical results of the new rotor show about 9% increase in its pressure ratio at both design and off design mass flow rate. The new rotor has a higher outlet velocity through its upper span improving bypass ratio of a turbofan engine. To prove the new fan ability of producing more bypass ratio, a thermodynamic analysis is conducted. The results of this analysis show 13% increase in bypass ratio and 5.7% decline in specific fuel consumption in comparison to NASA rotor 67.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A037. doi:10.1115/GT2016-57372.

In this work, unsteady numerical results and high-frequency measurements are investigated from nominal to loaded operating points with the objective to contribute to the understanding of pre-stall rotating disturbances. A 3.5 stages high speed axial multistage compressor is investigated on a 2 MW test rig in the laboratory of fluid mechanics (LMFA) at Ecole Centrale de Lyon, France. The compressor has been built by Snecma, and is representative of modern high-pressure rear blocks of a modern aircraft engine. The unsteady numerical results predict a rotating disturbance in the tip flow field of the rotor 2 at the loaded operating point. It causes a frequency which is not a multiple of the periodicity of the compressor, and is rotating at about 72% of the shaft speed. The mesh independency of this disturbance is ensured. The analysis of the circumferential and axial propagation of the disturbance reveals a rotating instability like phenomenon. Most characteristic is the very important periodic oscillation of the tip leakage vortex trajectory, leading to a modulation of the leakage flow in the neighboring tip gap. The influence of the neighboring blade rows is investigated by filtering their unsteady contribution by means of mixing planes up and/or downstream of the rotor 2. In either case, the rotating disturbance is found to be still present. There are no traces of this rotating disturbance in the high-frequency measurements investigated at a near surge operating point. A spike like stall inception and almost instantaneous surge inception is identified. The mis-prediction of the tip region flow field in the rotors 2 and 3 is believed to cause the mis-prediction of the pre-stall disturbance.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A038. doi:10.1115/GT2016-57396.

Near the endwalls of multi-stage compressor blade rows, there is a spanwise region of low momentum, high entropy fluid which develops due to the presence of annulus walls, leakage flows and corner separations. Off-design this region, known as the endwall flow region, often grows rapidly and in practice sets the compressor’s operating range. By contrast, over the operating range of the compressor, the freestream region of the flow is not usually close to its diffusion limit and has little effect on overall range. In light of these two distinct flow regions within a bladerow, this paper considers how velocity triangles in the endwall region should be designed to give a more balanced spanwise failure across the blade span.

In the first part of the paper, the sensitivity of the operating flow range of a single blade row to variations in realistic multistage inlet conditions and endwall geometry is investigated. It is shown that the operating range of the blade row is largely controlled by the size and structure of the endwall ‘repeating stage’ inlet boundary layer and not the detailed local geometry within the blade row.

In the second part of the paper the traditional design process is ‘flipped’. Instead of redesigning a blade’s endwall geometry to cope with a particular inlet profile into the blade row, the endwall region is redesigned in the multi-stage environment to ‘tailor’ the inlet profile into downstream blade rows. This is shown to allow an extra degree of freedom not usually open to the designer. This extra degree of freedom is exploited to balance freestream and endwall operating range, resulting in a compressor having an increased operating range of ∼20%. If this increased operating range is traded with reduced blade count, it is shown that a design efficiency improvement of Δη∼0.5% can be unlocked.

Topics: Compressors , Design
Commentary by Dr. Valentin Fuster
2016;():V02AT37A039. doi:10.1115/GT2016-57467.

The front fan of a turbofan aircraft engine often operates under distorted inlet flow conditions. This distortion is caused by either flight operating conditions, such as a crosswind or boundary layer ingestion, or due to its nacelle installation. These flow conditions negatively impact the aerodynamic performance of the compression system. Moreover, the asymmetry of the flow causes non-uniform circumferential pressure distortions which can trigger a strong aeromechanical response in the fan blades.

Numerical simulation can contribute to the design process if it can accurately predict the aerodynamic performance penalties and the loads experienced by the fan blades, thereby identifying potential problems early in the design phase. This requires accurate accounting of the pressure loads on the fan from the upstream inlet distortion and the potential effect of the downstream stator row. The loads are inherently transient in nature, requiring solutions on the full wheel geometry. However, full wheel modeling is expensive and not practical early in the design cycle. In this work, an efficient modeling strategy is proposed for an axial compressor fan with a downstream stator row (NASA Stage 67, rotor/stator) undergoing inlet distortion.

A multi-frequency frozen gust analysis using the Fourier-Transformation (FT) pitch-change method is utilized to solve this flow problem on a reduced geometry (two rotor-passages only). A once-per-revolution inlet distortion modeled as a cosine variation in total pressure is imposed upstream of the rotor. The influence of the stator row on the fan is accounted for within a transient simulation by imposing a 360 degree profile at the exit of the rotor. The profile from the stator row is obtained previously from a steady-state simulation using a multiple mixing-plane approach. In this approach the stator potential flow and the pressure variation in the stator row due to inflow distortion are accounted for.

The paper compares the reduced geometry model with full wheel transient predictions, thereby demonstrating the efficiency of the proposed method both in terms of accuracy and solution speedup. Important aerodynamic performance parameters as well as flow field solution monitors are compared to assess the viability of this modeling strategy.

Topics: Compressors , Modeling
Commentary by Dr. Valentin Fuster
2016;():V02AT37A040. doi:10.1115/GT2016-57550.

This paper describes the use of the Free-Form-Deformation [1] parameterisation method to create a novel blade shape for a highly loaded, transonic axial compressor. The novel geometry makes use of pre-compression (via an S-shaping of the blade around mid-span) to weaken the shock and improve the aerodynamic performance.

It has been known for some time that reducing the pre-shock Mach number of transonic compressors (via pre-compression) can improve their efficiency [2]. However, early attempts at this in the 60s [3] showed undesirable results (such as bi-stable operation), leading the design community to shy away from using pre-compression [4]. This issue is re-addressed here. It is shown that using modern simulation, optimisation and a 3D design, large amounts of pre-compression may be employed without the negative effects that plagued early attempts.

This paper shows how Free-Form-Deformation offers superior flexibility over traditionally used parameterisation methods. The novel design (produced via an efficient optimisation method) is presented and the resulting flow analysed in detail. The efficiency benefit is over 2%, surpassing other results in the literature for the same geometry. The pre-compression effect of the S-shape is analysed and explained, and the entropy increase across the shock (along the mid-blade line) is shown to be reduced by almost 80%. Adjoint surface sensitivity analysis of the datum and optimised designs is presented, showing that the S-shape is located in the region predicted to be most significant for changes in efficiency. Finally the off-design performance of the blade is analysed across the rotor characteristics at various speeds.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A041. doi:10.1115/GT2016-57617.

The unsteady and three-dimensional nature of the flow past axial compressor blading poses substantial challenges to the design of the main flow passage. High aspect ratio blading is amenable to the approach of splitting the design task between the cascade and the meridional planes. However, the three-dimensional flows increasingly affect the stage aerodynamic performance with decreasing blade aspect ratios. At very high load conditions, corner vortices can grow to two-thirds of the blade span, under the influence of the pitchwise pressure gradient, causing significant blockage and loss. A survey of treatments for three-dimensional flows highlights a variety of approaches, including longitudinal and tangential slots for suction and blowing, fences, turning vanes, fillets, and grooves. The merits and issues exposed by past implementations of these end-wall treatments are summarized. Considered together, these arrangements display a variable and open approach, which points towards an opportunity for considering a more common framework, led by a greater understanding of the flow physics. Preliminary work on the parametrization of end-wall grooves has highlighted some promising topological features of end walls generated by using the Beta distribution function as the guide curve. This seemingly unexplored application of the Beta function to axial compressor end wall design promises a better fit with the pitchwise periodic axial compressor geometry than other guide curve functions considered herein and used in the past.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A042. doi:10.1115/GT2016-57703.

This paper presents results of an experimental investigation of the impacts of total pressure distortions on the performance of a transonic turbofan. Fan rotor response to screen-produced total pressure distortions was measured by two high-frequency response probes (Kulite sensors). These probes measured the transient flow behind the fan rotor. Ensemble averaging based on a once-per-revolution encoder enabled separation of the pressure signal into its deterministic and stochastic components. The deterministic portion of the signal provided information about the effects of the distortion on the wake structure of the individual fan blades. The stochastic component of the signal gave indications of the effects of distortion on turbulence production (and losses) in the fan.

The qualitative trends provided by these high-frequency measurements show that the blade loading variations caused the wake depths and thicknesses to increase in the low pressure inlet regions. In addition, the turbulence production increased, peaking in the blade wakes. These behaviors indicate that the response of the fan was governed by a trailing edge separation that moves forward and backward on the upper blade surface in a dynamic manner.

This paper also includes discussion of the challenges associated with obtaining quantitative data with this type of probe. The results were significantly affected by probe frequency response and dynamic flow angle sensitivity. Resolving the blade wake structure and turbulence information requires probe tolerance to high-gradient flows and large flow angle fluctuations, with very high frequency response. The paper incorporates a discussion of improved probe design approaches to advance future research.

Topics: Pressure , Wakes , Blades
Commentary by Dr. Valentin Fuster
2016;():V02AT37A043. doi:10.1115/GT2016-57718.

This paper presents the often-neglected correlation terms for uncertainty propagation and shows their importance for rigorous uncertainty analysis. It shows that these terms can be leveraged, often through simple experimental design changes to substantially reduce uncertainty. The efficiency equation is particularly suited to this approach, with order of magnitude uncertainty reductions possible. The paper will present this effect from both a theoretical study and from data measured behind a transonic fan during a distortion response program.

The efficiency equation is particularly suited to improving uncertainty by leveraging the correlation terms because the temperature uncertainty contributes most of the overall uncertainty. Systematic error sources usually dominate thermocouple measurement uncertainty because the slow time response leads to small random errors. The partial derivatives with respect to inlet and outlet total temperature have opposite signs. Correlations between these systematic errors are relatively easy to force through using a single data acquisition system and in-situ calibration standard. This paper presents performance comparisons of these effects for thermocouples. For a 1.35 pressure ratio, 85% efficiency fan, the uncertainty can be reduced by correlation from 85% ± 6.7% to 85% ± 0.48% for typical K-type thermocouples.

Topics: Pressure , Fans , Uncertainty
Commentary by Dr. Valentin Fuster
2016;():V02AT37A044. doi:10.1115/GT2016-57726.

A highly loaded, optimized compressor tandem cascade for low speed application (Ma1 = 0.175) [1, 2] was investigated numerically and experimentally using the standard particle image velocimetry measurement technique. The investigated regions are divided in the inlet, the overlap and the outlet region and are investigated for three different inlet angles.

For all inlet angles investigated the results show a homogeneous velocity distribution. The analysis of the measured inlet angle shows a strong suction effect of the cascade starting far in front of the leading edge of the front blade. The acceleration in the overlap region and the change of the velocity distribution is measured for all inlet angles investigated and clarifies the conclusions resulting of former blade surface static pressure measurements [1]. Although the used PIV technique is only capable to resolve two velocity components, the results of the spanwise measured outlet flow fields allow a good analysis of the secondary flow and complete the standard measurement results [2]. The comparison of the experimental flow field results with the numerical results reveals the known numerical limits of secondary flow prediction regarding the extension and intensity [2].

Commentary by Dr. Valentin Fuster
2016;():V02AT37A045. doi:10.1115/GT2016-57727.

Changes in loss generation associated with altering the rotor tip blade loading of an embedded rotor-stator compressor stage are assessed with unsteady three-dimensional computations, complemented by control volume analyses. Tip-fore-loaded and tip-aft-loaded rotor blades are designed and assessed to provide variation in rotor tip blade loading distributions for determining if aft-loading rotor tip would yield a stage performance benefit in terms of a reduction in loss generation. Aft-loading rotor blade tip delays the formation of tip leakage flow resulting in a relatively less mixed-out tip leakage flow at the rotor outlet and a reduction in overall tip leakage mass flow, hence a lower loss generation; however, the attendant changes in tip flow angle distribution are such that there is an overall increase in the flow angle mismatch between tip flow and main flow leading to higher loss generation. The latter outweighs the former so that rotor passage loss from aft-loading rotor tip is marginally higher unless a constraint is imposed on tip flow angle distribution so that associated induced loss is negligible; a potential strategy for achieving this is proposed. Tip leakage flow, which is not mixed-out at the rotor outlet, enters the downstream stator, where it can be recovered. The tip leakage flow recovery process yields a higher benefit for a relatively less mixed-out tip leakage flow from aft-loading a rotor blade tip. These characterizing parameters together determine the attendant loss associated with rotor tip leakage flow in a compressor stage environment. A revised design hypothesis is thus as follows: rotor should be tip-aft-loaded and hub-fore-loaded while stator should be hub-aft-loaded and tip-fore-loaded with tip/hub leakage flow angle distribution such that it results in no additional loss. For the compressor stage being assessed here, an estimated 0.15 points enhancement in stage efficiency is possible from aft-loading rotor tip only. In the course of assessing the benefit from unsteady tip leakage flow recovery in the downstream stator, it was determined that tip clearance flow is inherently unsteady with a time-scale distinctly different from the blade passing time. The disparity between the two timescales: (i) defines the periodicity of the unsteady rotor-stator flow, which is an integral multiple of blade passing time; and (ii) causes tip leakage vortex to enter the downstream stator at specific pitchwise locations for different blade passing cycles, a tip leakage flow phasing effect. Despite the inherent unsteadiness from tip leakage flow, the recovery process is demonstrated to be beneficial on a time-averaged basis.

Topics: Compressors , Rotors , Blades
Commentary by Dr. Valentin Fuster
2016;():V02AT37A046. doi:10.1115/GT2016-57797.

Incidents of partial or total thrust loss at cruise due to engine icing (mainly ice crystals and super-cooled water droplets at altitudes greater than 10,000 feet) have been recorded over the past several years. As air traffic continues to increase in subtropical areas where high moisture laden air is present at subfreezing conditions, engine icing probability increases. There is a need to better understand compressor dynamics under icing conditions which will help designers to develop more accurate and fast ice detection systems and anti-icing mechanisms. Stage re-matching occurs due to heat exchange between air and ice which dictates the stall inception stage in the compressor. It has been shown that compressor stages re-match under icing conditions — front stages are choked while rear stages throttle due to ice melting and evaporation. Such an analysis uses various empirical models to represent ice-breakup and water-splash processes as ice/water particles interact with rotors/stators. The following paper presents a compressor stall sensitivity analysis around different splash models. The effect of droplet splash at both rotor and stator blades, blade solidity effect and trailing edge shed effect are modeled. A representative 10 stage high speed compressor section operating near design point (100% Nc) is used for the study. Results show that T3 drop and overall compressor operability is a function of evaporating stages and droplet-blade interaction models influence them. A comprehensive compressor stability envelope has been evaluated for different models. It is observed that droplet-blade interaction behavior influences overall compressor stability and stall-margin predictions can vary by as much as 25% with different models. Therefore, there is a need for better calibration and continual improvement of empirical models to capture compressor inter-stage dynamics and stage re-matching accurately under ice/water ingestion.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A047. doi:10.1115/GT2016-57808.

This paper describes a new conceptual framework for three-dimensional turbomachinery flow analysis and its use to assess fan stage attributes for mitigating adverse effects of inlet distortion due to boundary layer ingestion (BLI). A non-axisymmetric throughflow method has been developed to describe the fan flow field with inlet distortion. In this the turbomachinery is modeled using momentum and energy source distributions that are determined as a function of local flow conditions and a specified blade camber surface geometry. Comparison with higher-fidelity computational and experimental results shows that the method captures the principal flow redistribution and distortion transfer effects associated with BLI. Distortion response is assessed for a range of (i) rotor spanwise work profiles, (ii) rotor-stator spacings, and (iii) non-axisymmetric stator geometries. For the parameters examined, changes in axisymmetric design result in trades between rotor and stator distortions, or between different radial sections of a given blade row with marginal overall gain. Of the approaches examined, non-axisymmetric stator exit flow angle distributions were found to provide the greatest reduction in rotor flow distortion and thus may offer the most potential for mitigating decreases in performance due to BLI inlet distortion.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A048. doi:10.1115/GT2016-57896.

It has been shown in several cases that casing treatments can improve the surge margin of a compressor. The question that is often left unanswered when designing and optimizing a casing treatment for a given compressor stage is whether the casing treatment can still improve the performance and surge margin when used on an aerodynamically improved rotor.

The current work includes the influence of the rotor geometry and casing geometry in the analysis. Several optimizations with equal objectives, to improve surge margin and efficiency, were performed. First the different components were optimized separately. The possible impact of optimizing the rotor, the casing geometry and the casing treatment on surge margin and efficiency are compared in terms of global performance and aerodynamic effects. It is then analyzed how combinations of optimized geometries perform. Of particular interest is how the circumferential grooves perform when used in combination with the optimized blade geometry.

The optimizations were performed using state of the art CFD and optimization procedures. Constraints on boundary conditions, mass flow rates and stresses were applied in order not to change the behavior of the compressor at design point significantly or to deteriorate the structural mechanics.

Topics: Compressors , Surges
Commentary by Dr. Valentin Fuster
2016;():V02AT37A049. doi:10.1115/GT2016-57917.

Aircraft engines ingest airborne particulate matter, such as sand, dirt, and volcanic ash, into their core. The ingested particulate is transported by the secondary flow circuits via compressor bleeds to the high pressure turbine and may deposit resulting in turbine fouling and loss of cooling effectiveness. Prior publications focused on particulate deposition and sand erosion patterns in a single stage of a compressor or turbine. The current work addresses the migration of ingested particulate through the high pressure compressor and bleed systems. This paper describes a 3D CFD methodology for tracking particles along a multi-stage axial compressor and presents particulate ingestion analysis for a high pressure compressor section. The commercial CFD multi-phase solver ANSYS CFX R has been used for flow and particulate simulations. Particle diameters of 20, 40, and 60 microns are analyzed. Particle trajectories and radial particulate profiles are compared for these particle diameters. The analysis demonstrates how the compressor centrifuges the particles radially towards the compressor case as they travel through the compressor; the larger diameter particles being more significantly affected. Non-spherical particles experience more drag as compared to spherical particles and hence a qualitative comparison between spherical and non-spherical particles is shown.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A050. doi:10.1115/GT2016-57972.

In the design of modern jet engines the need for accurate loss prediction techniques is ever present. The most common tool currently in use is Reynolds Averaged Navier-Stokes model which provides good estimation at design conditions but can struggle with off design conditions. With accuracy being such an important requirement, an alternative method such as Large Eddy Simulation presents an opportunity to improve and assess the off design performance.

Although still limited by computational resources, the use of Large Eddy Simulations in conjunction with more detailed loss analysis methods forms a powerful tool for assessing and improving current Reynolds Averaged Navier-Stokes techniques. The simulations performed here are an incidence sweep at off-design conditions with free stream turbulence. The results of the two methodologies are compared with the use of loss breakdown analysis and the best practice of applying the loss breakdown technique to compressors is outlined.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A051. doi:10.1115/GT2016-57985.

This paper solves the filtered Navier-Stokes equations to simulate stall inception of NASA compressor transonic Stage 35 with delayed detached eddy simulation (DDES). A low diffusion E-CUSP Riemann solver with a 3rd order MUSCL scheme for the inviscid fluxes and a 2nd order central differencing for the viscous terms are employed. A full annulus of the rotor-stator stage is simulated with an interpolation sliding boundary condition (BC) to resolve the rotor-stator interaction. The tip clearance is fully gridded to accurately resolve tip vortices and their effect on stall inception. The DDES results show that the stall inception of Stage 35 is initialized by a weak harmonic disturbance with the length scales of the full annulus and grows rapidly with two emerging spike like disturbance. The two spike disturbances propagate in counter rotational direction with about 42% of rotor speed. The spike stall cells cover about 6 blades. They lead to two stall cells grown circumferentially and inwardly.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A052. doi:10.1115/GT2016-58009.

Applications such as boundary-layer-ingesting fans, and compressors in turboprop engines require continuous operation with distorted inflow. A low-speed axial fan with incompressible flow is studied in this paper. Previous work in the literature has shown that the same flow mechanisms contributing to the response of a fan to distortion are at play in incompressible and transonic flows. The objective is to determine how fan performance scales as the type and severity of inlet distortion varies at the design flow coefficient.

A distributed source term approach to modeling the rotor and stator blade rows is used in numerical simulations in this paper. The approach has been shown to capture overall stage performance and flow field behavior with distortions having length scales much longer than the blade pitch. The approach requires only knowledge of the blade geometry, but the model does not include viscous losses. As a result, efficiency is not assessed but instead a metric based on changes in diffusion factor is defined which is conjectured to be related to efficiency changes.

Distortions in stagnation pressure, swirl, and stagnation temperature are considered. By studying the distortions individually, it is found that the diffusion metric scales linearly with the intensity of the distortions (i.e. the ratio of minimum to maximum values) but that changes in distortion location relative to the fan axis produce nonlinear changes in the diffusion metric. Combinations of distortions are also studied and it is found that the diffusion metric associated with the combined distortion can be predicted using a summation procedure for the metrics associated with the individual constituent distortions. The mechanism found to govern the effectiveness of this summation procedure is the incidence distortions at rotor and stator inlet.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A053. doi:10.1115/GT2016-58117.

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.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A054. doi:10.1115/GT2016-58141.

Flow separations often take place in the junction of blades and endwalls and limit seriously the aerodynamic loading increase of turbomachinery, which are caused mainly by mixing of the boundary layers on blades and endwall surfaces and the transverse secondary flow generated by the pressure difference between the pressure and suction side.

Firstly, focusing on a linear diffusion cascade with 42 degrees turning angle, it can be found that the transverse secondary flow can be reduced by inviscid hub and the flow separation is eliminated further through the numerical comparison between the viscous and inviscid hub cases. So the transverse secondary flow is the dominate factor for the flow separation in this cascade. We should try to control the transverse secondary flow to reduce the flow separation.

Secondly, based above analysis, the flow separation can be controlled effectively if we can cut off the secondary flow. So nine kinds of streamwise groove schemes are designed and analyzed. It can be seen that the streamwise grooves at the end wall inhibit obviously the transverse secondary flow but the flow structure change is different at different span. There is an optimum combination of width and height of groove, and the height is more important than width.

Thirdly, the detailed flow analysis of best scheme with smaller width, moderate height are carried out. It can decrease the separation zone scope at the corner zone, reduce the energy loss coefficient and also reduce the flow loss.

Commentary by Dr. Valentin Fuster
2016;():V02AT37A055. doi:10.1115/GT2016-58145.

Half-annulus unsteady numerical simulations have been conducted with a 60-deg total pressure circumferential distortion in a transonic axial-flow fan. The effects of inlet distortion on the performance, stability and flow field of the test case are investigated and analyzed. Results show that the incidence angles are reduced when the blades are entering into the distorted region. Conversely, distortion increases the incidence angles onto the blades when they are leaving the distorted section. Results further reveal that the time-averaged flow field at the tip of the blade is similar with and without distortion. However, the distortion applied is found to have detrimental effects on both the stability and performance. The impacts of both annular and discrete tip injection on the endwall flow field are further studied in the current work. It is shown that endwall injection reduces the incidence angles onto the blades. Consequently, the passage shock and the leakage flow are pushed rearward, which postpones stall initiation.

Commentary by Dr. Valentin Fuster

Ducts and Component Interactions

2016;():V02AT40A001. doi:10.1115/GT2016-56199.

Flow in annular ducts is sensitive to the presence of downstream blockages which can cause flow non-uniformities propagating far upstream of the blocking body. These effects can be exacerbated in swirling flows where a cascade of uniform guide vanes is present upstream of the blockage. This work uses two- and three-dimensional boundary singularity methods to model and optimise a guide vane cascade geometry to minimise the upstream velocity distortion. Starting from a uniform cascade, the geometry is modified to provide a uniform upstream velocity distribution and minimised blade-to-blade loading in two dimensions. The new geometry is then extrapolated to a three-dimensional annulus. A three-dimensional tool is used to further modify the geometry in three dimensions to minimise the velocity distortion in the whole annulus upstream of the cascade.

Topics: Flow (Dynamics)
Commentary by Dr. Valentin Fuster
2016;():V02AT40A002. doi:10.1115/GT2016-56327.

In this paper, the modelling of leakages through a compressor stator penny cavity, and their effect on the aerodynamics within the compressor are studied. The penny, sometimes also referred to as ‘button’, is the cylindrical platform feature of a variable stator normally found between a vane’s airfoil and spindle. The pennies nominally lie recessed into the compressor endwalls at hub and casing, with a surrounding clearance to ensure the vane’s stagger angle can be adjusted. RANS-simulations, with these clearances included, have shown a significant impact from the penny cavity leakages on compressor efficiency and surge line. Neglecting this secondary flow path through the penny cavities results in an under prediction of the losses close to the endwalls.

The prediction of the penny cavity effect on the stator row is based on a Reynolds-Averaged-Navier-Stokes (RANS) study, using a hybrid structured-unstructured mesh to provide adequate resolution of the local flow phenomena.

The complex geometry and pressure field result in flows that are unevenly distributed within the penny cavity. The outflow or leakage is focused in a concentrated area leading to a high local velocity that strongly impacts the stator losses and turning. Since such geometries lie beyond the normal validated cases, the modelling uncertainties are discussed and the plausibility of the results is checked. In order to provide an experimental database and validate the turbulent mixing of leakage and main flow, which is seen as the main contributor to loss production, a validation test case — ‘Jet-In-Crossflow’ was chosen. As well as the standard RANS code, this validation case was run as a time-accurate high-order Lattice Boltzmann (LBM) simulation (PowerFLOW), using Very-Large-Eddy-Simulation (VLES) turbulence modelling. The LBM simulation showed significant unsteady flow features and was considerably closer to the test data than the RANS calculations.

A future test campaign, currently being prepared at the annular cascade test facility of the Institute of Jet Propulsion and Turbomachinery (IST) at RWTH Aachen university, will be briefly presented. This focuses on investigating the penny flows in a typical engine design.

Commentary by Dr. Valentin Fuster
2016;():V02AT40A003. doi:10.1115/GT2016-56476.

The flow leaving the high pressure turbine should be guided to the low pressure turbine by an annular diffuser, which is called as the intermediate turbine duct. Flow separation, which would result in secondary flow and cause great flow loss, is easily induced by the negative pressure gradient inside the duct. And such non-uniform flow field would also affect the inlet conditions of the low pressure turbine, resulting in efficiency reduction of low pressure turbine. Highly efficient intermediate turbine duct cannot be designed without considering the effects of the rotating row of the high pressure turbine.

A typical turbine model is simulated by commercial computational fluid dynamics method. This model is used to validate the accuracy and reliability of the selected numerical method by comparing the numerical results with the experimental results. An intermediate turbine duct with eight struts has been designed initially downstream of an existing high pressure turbine. On the basis of the original design, the main purpose of this paper is to reduce the net aerodynamic load on the strut surface and thus minimize the overall duct loss. Full three-dimensional inverse method is applied to the redesign of the struts. It is revealed that the duct with new struts after inverse design has an improved performance as compared with the original one.

Commentary by Dr. Valentin Fuster
2016;():V02AT40A004. doi:10.1115/GT2016-56639.

Three different optimization methods using genetic algorithms have been developed, aiming to achieve better aggressive intermediate turbine duct (ITD) performance. To overcome defects of simple genetic algorithms, a niche genetic algorithm is used, for its better adaptability to multi-peak function. These three methods are two-dimensional optimization for pursuing the highest static pressure coefficient; two-dimensional optimization via controlling static pressure coefficient; and a further three-dimensional optimization. The second method introduces a restrainer to make sure the maximum value of static pressure coefficient gradient less than a limitation. A simulation case, an ITD with eight struts, was implemented to demonstrate the capabilities of the presented optimization methods. Compared to the baseline ITD, the results show as follow. The first method obtains a best static pressure coefficient but a severe separation. The second optimization method with static pressure coefficient gradient control can definitely suppress separation in the ITD, the second method also obtains better static pressure coefficient and the lowest total pressure loss coefficient. The third method, a further three-dimensional optimization can obtain better ITD overall performance because of a more realistic simulation. Nevertheless, the second two-dimensional optimization method can get good enough results while it is apparently much more time-saving compared to three-dimensional one, which make it more suitable for engineering applications.

Commentary by Dr. Valentin Fuster
2016;():V02AT40A005. doi:10.1115/GT2016-56904.

Currently in an aircraft gas turbine engine, the turbomachinery and combustor components are designed in relative isolation and the effect of the upstream and downstream components on each other’s flow are not fully captured in the design process. The objective of this work is to carry out a multi-code integrated unsteady simulation of Compressor-Combustor components with each zone simulated using its own specialised CFD flow solver. The multi-code URANS technique is simple, based on files and involves the generation of new 2D boundary conditions for the required flow field at each time step. A driver based on a Python script automates the entire process. This paper shows the method first validated in a simple vortex shedding 2D case and then extended to a cold flow URANS simulation matching an isothermal compressor/combustor rig experiment. An external coupler code is invoked that produces unsteady, spatially varying, inlet conditions for the downstream components. The simulation results are encouraging as the mass, momentum and energy losses across the interface are less than 1%. The multi-code unsteady simulation produces wake profiles closer to the experiment than the coupled steady RANS simulation. The present study shows a reasonable agreement with the experimental PIV and hot-wire data thus demonstrating the potential of the multi-code integrated simulation technique.

Commentary by Dr. Valentin Fuster
2016;():V02AT40A006. doi:10.1115/GT2016-58120.

A detailed numerical simulation is presented to investigate the new de-swirling methods and their effect on the mixing mechanisms of a turbofan mixer with 12 lobes. The numerical simulation employed a commercial solver, ANSYS CFX, using k-ω SST model. The core-to-bypass temperature ratio and pressure ratio were set to 2.59, and 0.97 respectively, giving the Mach number of 0.66 and bypass ratio of 2.65 at mixing nozzle outlet. The inlet swirl typically accelerates the jet-flow mixing by enhancing the vortices intensity and interaction, but leakage swirling flow can cause a three-dimensional separation bubble and the recirculation zone resulting in the dramatic increasing the total pressure loss and thrust loss. Removal of the leakage swirling flow between the lobes’ trough and centre-body was the key to limit the negative influence of inlet swirl.

Two IGV design were investigated, DS1 and DS2. DS1 was installed at the upstream of the lobed mixer, could remove the negative effect of inlet swirl properly, but also inhibited the active role of the inlet swirl. The total pressure and thrust loss reduced by 0.31% and 3.8%, respectively, but the mixing efficiency also decreased by 1.72%. DS2, an integrated strut with the lobed mixer design, not only ensured the structure strength of the lobed mixer, but also reduced the length and weight of the exhaust system. This method suppressed the flow separation bubble on centre-body to some extent, and eliminated the recirculation zone downstream of the cenrebody, resulting in the total pressure loss decrease of 0.31% and thrust gain of 3.63%. On the other hand, the method DS2 also made full use of the inlet swirl to enhance the jet-flow mixing, resulting in the mixing efficiency increased 1.54% compared with that of the DS1 case. Under the off-design conditions with the incidence angle of ±10°, the aerodynamic performance of the DS2 cases didn’t changed too much such as the DS1 cases.

Topics: Swirling flow
Commentary by Dr. Valentin Fuster

Noise and Innovative Noise Reduction (With Aircraft Engine)

2016;():V02AT41A001. doi:10.1115/GT2016-56020.

A strong focus in the development of modern aircraft engines is the reduction of the engine tonal core noise. For the development of efficient noise reduction techniques, a detailed understanding of the sound transmission throughout all turbomachinery components of the engine is mandatory. In this paper an excitation system is developed to generate turbomachinery-specific sound fields by controlling their circumferential and radial mode order. The excitation system consists of two rows of eight loudspeakers distributed circumferentially around the outer duct wall. This paper gives a detailed description of the analytically- and numerically-supported design methodology of an optimized excitation system, as well as an optimized microphone array mounted flush with the outer duct wall. A sensitivity analysis of the loudspeaker array and of the microphone array with respect to distance and frequency is then carried out numerically. To analyze the microphone signals and to deconstruct the propagating sound field into its modal components, a Radial Mode Analysis (RMA) is carried out. To ensure high-quality RMA results, the axial distribution of the microphones is optimized with respect to the condition number of the array’s transfer matrix. The procedure explained in this paper shall help guide the development of acoustic excitation and microphone array systems for experiments to better understand sound propagation in turbomachinery and flow ducts.

Commentary by Dr. Valentin Fuster
2016;():V02AT41A002. doi:10.1115/GT2016-56453.

This paper is a continuation of a series of study on the mechanism of the broadband noise reduction for turbomachinery blade using trailing edge serrations. The noise reduction potential of turbine blade with trailing edge serrations is experimentally assessed as well as the various parameters on the noise reduction effect. Special focus is put on whether the trailing edge serrations affect turbine cascade tailing edge noise in the same way as they do on the isolated airfoil. Five different trailing edge serrations were designed for a turbine linear cascade to investigate the effects of serration geometry parameter on the noise reduction. A linear microphone array was used to quantify the difference of sound source levels of turbine cascade with and without trailing edge modifications. The experiment was carried out at various velocities and the Reynolds number (based on cascade inlet velocity and chord) ranges from 1.3×105 to 3.3×105. The experiment results show that trailing edge serrations can reduce turbine trailing edge noise in a wide frequency range that we are interested (from 1600Hz to 10000Hz) and a maximum noise reduction of about 5dB is obtained in the mid frequency range (2000Hz to 4000Hz). The results show that the serration length has an important effect on the noise reduction effect and the longer serration in the experiment lead to more noise reduction. However, serration wavelength has only a little effect on the noise reduction although the wider trailing edge serrations tested in the experiment can achieve slightly more noise reduction. This is quite different from that for airfoils. At all the velocities tested, the cascade trailing edge noise is effectively reduced and the maximum noise reduction occurs at St=2fh/U≈1.

Commentary by Dr. Valentin Fuster
2016;():V02AT41A003. doi:10.1115/GT2016-56673.

The major techniques for measuring jet noise have significant drawbacks, especially when including engine installation effects such as jet-flap interaction noise. Numerical methods including low order correlations and Reynolds-Averaged Navier-Stokes (RANS) are known to be deficient for complex configurations and even simple jet flows. Using high fidelity numerical methods such as Large Eddy Simulation (LES) allow conditions to be carefully controlled and quantified. LES methods are more practical and affordable than experimental campaigns. The potential to use LES methods to predict noise, identify noise risks and thus modify designs before an engine or aircraft is built is a possibility in the near future. This is particularly true for applications at lower Reynolds numbers such as jet noise of business jets and jet-flap interaction noise for under-wing engine installations. Hence, we introduce our current approaches to predicting jet noise reliably and contrast the cost of RANS-Numerical-LES (RANS-NLES) with traditional methods. Our own predictions and existing literature are used to provide a current guide, encompassing numerical aspects, meshing and acoustics processing. Other approaches are also briefly considered. We also tackle the crucial issues of how codes can be validated and verified for acoustics and how LES based methods can be introduced into industry. We consider that hybrid RANS-(N)LES is now of use to industry and contrast costs, indicating the clear advantages of eddy resolving methods.

Commentary by Dr. Valentin Fuster
2016;():V02AT41A004. doi:10.1115/GT2016-56773.

Aerodynamic performance and noise level of AC (Air-Conditioner) indoor unit fan system has always been receiving much concern. But most previous studies have been focused on aerodynamic performance and noise level, and less attention is paid to the noise radiation problem. It is essential to identify and quantify the noise radiation, so as to control and reduce the noise level effectively.

The objective of the present investigation is to identify the influential factors of the Blade Passing Frequency (BPF) noise radiation and investigate noise directivity distribution of two double-suction centrifugal flow fans with different fan blades used in AC indoor unit. A hybrid numerical approach is developed for predicting the unsteady fan flow and noise radiation. Unsteady fan flow is solved by Detached Eddy Simulation (DES) method. Then the areoacoustic calculation is performed based on the predicted unsteady flow field, and the obtained results are used as dipole source inputs in acoustic Finite Element Model (FEM) for predicting the noise radiation of the centrifugal fans.

Experimental tests of aerodynamic and aeroacoustic performance are conducted respectively for the fan. A comparison is made between the experimental and measured results. It is demonstrated that the predicted aeroacoustic and noise radiation both agree with the measured ones and the fan blade geometry is much influential to acoustic prediction. Finally, directivity of BPF noise radiation to noise level is evaluated for two different fans.

Commentary by Dr. Valentin Fuster
2016;():V02AT41A005. doi:10.1115/GT2016-56969.

A detailed understanding of the sound propagation and transmission within the engine and adjacent ducts is mandatory for the development of efficient noise reduction techniques for the tonal sound field produced by the turbomachinery components of aircraft engines. For this purpose, experimental acoustic investigations are needed. In the first part of this paper, an acoustic excitation system for the generation of acoustic spinning modes with circumferential mode order one and varying radial mode order, as well as a microphone array optimized for a radial decomposition of the sound field have been systematically designed.

To verify the excitation method and the design of the excitation system, corresponding experimental measurements are carried out in an acoustic wind tunnel. Amongst others, the sound power of the specific excited acoustic modes of order (1,0) and (1,1) are compared with the respective powers achieved with a non-specific sound field excitation. To test the range of flexible use of the sound generator, measurements are carried out over a wide frequency range. It is shown that the intended modes can be controlled at the design frequency of the sound generator as well as off-design frequencies. However, the dominance of the excited modes strongly depends on the number of cut-on modes and the excitation frequency as non-linear resonance effects may interfere. Furthermore, the benefit of an increased number of loudspeaker rows for stable mode excitation is discussed. The experimental results are supported by numerical simulations.

Commentary by Dr. Valentin Fuster
2016;():V02AT41A006. doi:10.1115/GT2016-57137.

The aerodynamic performance and noise of the blade are two important aspects which people pay much attention nowadays in the design of turbine machinery such as centrifugal fan and axial flow fan. In this paper, the three-dimensional model of the long-eared owl wing is established based on the section theory and fitting formula firstly. And then, unsteady aerodynamic and acoustic characteristics of the bionic blade are numerically investigated using Large-Eddy Simulation (LES) and the Ffowcs Williams-Hawkings (FW-H) equation based on Lighthill’s acoustic theory. The results indicate that the deeply concaved lower surface near the wing root plays a significant role in improving the lift-to-drag ratio. The lift coefficient and drag coefficient of the bionic blade is analyzed by comparing two-dimensional and three-dimensional results. The cross section profiles near the wing root possess the larger lift coefficients and the lesser drag coefficients, even than the three-dimensional long-eared owl wing. The size of the separation bubble grows at increasing angle of attack. The 40% cross-section profile of the long-eared owl wing could increase the distance between the corresponding vortex centers with wall surface thus reducing the range of the vortex shedding near the wall effectively. The iso-Q surfaces show that the location of the vortex shedding and the movement of separation bubble. When the angle of attack α is 5°, the aerodynamic noise generated by the bionic blade is lower than in other angle of attack condition. The minimum value of the sound pressure level (SPL) is even 17.9dB on the y-direction. In the range of 5°–15°, the strength and size of the vortex motion increase with the increase of the angle of attack. The far-field noise suggests the directivities of dipole noise. The range of the separation bubbles act as the most influence of the noise generation. The sound pressure level (SPL) of bionic blade at α = 5° is less than other conditions and the minimum value is even 17.9dB. The thin airfoil near the wingtip could decrease the pressure fluctuation from the blade surface that can reduce the unsteady aerodynamic noise. It means the unique structure of the long-eared owl wing can suppresses the unsteady pressure fluctuation on the surface which could decrease the noise generated by the wing surface.

Commentary by Dr. Valentin Fuster
2016;():V02AT41A007. doi:10.1115/GT2016-57209.

The noise originating from the core of an aero-engine is usually difficult to quantify and the knowledge about its generation and propagation is less advanced than that for other engine components. In order to overcome the difficulties associated with dynamic measurements in the crowded core region, dedicated experiments have been set up in order to investigate the processes associated with the generation of noise in the combustor, its propagation through the turbine and the interaction of these two components, which may produce additional — so-called indirect combustion — noise. In the current work, a transonic turbine stage installed at the Laboratorio di Fluidodinamica delle Macchine of the Politechnico di Milano was exposed to acoustic, entropic, and vortical disturbances. The incoming and outgoing sound fields were analyzed in detail by two large arrays of microphones. The mean flow field and the disturbances were carefully mapped by several aerodynamic and thermal probes. The results include transmission and reflection characteristics of the turbine stage, latter one was found to be much lower than usually assumed. The modal decomposition of the acoustic field in the upstream and downstream section show beside the expected rotor-stator interaction modes additional modes. At the frequency of entropy or respectively vorticity excitation, a significant increase of the overall sound power level was observed.

Commentary by Dr. Valentin Fuster
2016;():V02AT41A008. doi:10.1115/GT2016-57352.

The results of the first booster stage tone noise numerical investigation for a model of low pressure compressor are presented. The investigation was performed using the frequency domain numerical method of multistage turbomachines tone noise simulation, developed in CIAM (Central Institute of Aviation Motors) and implemented in the 3DAS (3 Dimensional Acoustics Solver) in-house solver. The model under consideration included high bypass ratio fan, stator, booster and bypass duct. Calculation was performed at the approach operating conditions. Far field directivities for two tones in the forward hemisphere were obtained. One tone corresponded to the blade passing frequency of the first stage rotor, the other - to the sum of this frequency with the blade passing frequency of the fan. The results of the computation were compared with the experimental data obtained in the CIAM C-3A acoustic test facility. In general satisfactory correspondence between calculation and experiment was obtained.

Commentary by Dr. Valentin Fuster
2016;():V02AT41A009. doi:10.1115/GT2016-57405.

In large modern turbochargers, compressors often constitute the main source of noise, with a frequency spectrum typically dominated by tonal noise at the blade passing frequency (BPF) and its harmonics. In transonic operation, inflow BPF noise is mainly generated by rotor locked shock fronts. These and the resulting acoustic fields can be predicted numerically with reasonable accuracy. Outflow noise, while also dominated by BPF tones, is linked to more complex source mechanisms. Its modal structure and the relationships between sources and modal sound pressure levels (SPL) are less well understood. Perhaps this is linked to the intrinsically non-axisymmetric geometries, which results in the need for full stage simulations if high accuracy is of paramount importance.

In order to shed some light on outflow noise generation, a transient simulation of a 360° model of a radial compressor stage, including a vaned diffuser and a volute, was carried out using state-of-the-art CFD. Additionally, experimental data was gathered at a multitude of data points downstream of the volute exit for post processing and modal analysis. The sources and the propagation were calculated directly. Optimized values for tempo-spatial acoustic wave resolution and buffer layer design were chosen, based on extensive studies on simplified models. Two grid refinement levels were used to check grid convergence and time step size independence of the results was ensured.

Numerical and experimental data match within 1% for total pressure ratio, volume flow and exit total temperature for the studied operating point. Both show the same modal content at the 1st BPF and indicate the presence of the same single dominating mode. The numerical results underpredict overall sound power levels (PWL) at the 1st BPF by 6.6dB. This difference is expected to decrease with further grid refinement and improved accounting for numerical damping. At the 2nd BPF, the experimental data show a significant broadening of the modal content with homogeneous modal PWL distributions. The multitude of modes leads to the generation of complex interference patterns, which shows that single-point acoustic measurements are often inadequate for component noise qualification and should be substituted by modal techniques. The dominating dipole sound sources are found in narrow areas around the vane leading edges and the rotor blade trailing edges. Because of the non-axisymmetric geometry, vane dipole source strengths become a function of circumferential position. The unsteady shedding of vortices from the vane suction surfaces is identified as a further possible source mechanism. However, the contributions of structural vibrations and mode scattering due to small manufacturing imperfections remain unclear.

Commentary by Dr. Valentin Fuster
2016;():V02AT41A010. doi:10.1115/GT2016-57612.

As a result of dissipative and dispersive properties of numerical methods the accuracy of direct noise prediction degrades with increasing distance from the source. Hybrid approaches are hence applied to predict the acoustic farfield, which rely on an integration of disturbance quantities extracted from the nearfield flow solutions. In order to elaborate the influence of the integration limits on sound prediction, different approaches of the Ffowcs Williams and Hawkings method are systematically applied to an advanced counter rotating propfan configuration within the present study. The solutions of various permeable and impermeable integration surfaces are analysed with respect to nearfield sound radiation and compared against direct sound predictions from the compressible Reynolds-averaged Navier-Stokes solutions which likewise serve as input for the extrapolation routine. Due to the flexibility of the routine, source terms and zones can be selectively excluded from the surface integration, allowing a systematic identification of the origin of dissimilar sound prediction. Subsequent farfield analyses are used to conclude on the propagation and persistency of differences identified in the nearfield predictions.

Commentary by Dr. Valentin Fuster
2016;():V02AT41A011. doi:10.1115/GT2016-57754.

In the present paper, a comparison of the numeric and experimental results obtained from the acoustic mode analysis for a turning mid turbine frame (TMTF) is presented. The investigated turning mid turbine frame is part of the two-stage two-spool test turbine located at the Institute for Thermal Turbomachinery and Machine Dynamics of Graz University of Technology. In this specific test turbine a transonic turbine stage (HP) is followed by a low pressure turbine stage (LP) consisting of a TMTF with 16 highly 3D-shaped turning struts and a counter-rotating low pressure rotor.

The experimental dataset is obtained by a measurement section downstream of the low pressure rotor which is instrumented with acoustic sensors. This microphone array is wall flush mounted in the outer casing which is traversable over 360 degrees in circumferential direction. The numerical setup consists of the whole test turbine including the experimental measurement section downstream of the low pressure rotor. Since the periodicity of the test setup equals 90 degrees the CFD calculations were performed accordingly using the unsteady inhouse Navier-Stokes code LINARS.

For both, numerical and experimental datasets the same post-processing tools are used in order to perform the acoustic mode analysis of the unsteady data. At first a comparison in terms of noise generation and propagation of the results is done by the frequency spectra, the emitted sound pressure and sound power level of both rotors independently. Since the emitted sound pressure level rises to a maximum at the first blade passing frequency of the HP rotor as well as the LP rotor the further analysis focuses on these two specific frequencies only. Therefore the acoustic field at those frequencies is characterized by azimuthal and radial modes. For a correct comparison between the numerical and experimental results numeric data taken from the same geometric locations as the microphones’ positions (thus the measurement locations) is processed.

Hence, this paper provides a deep insight into the capability of using unsteady CFD calculation in combination with the acoustic mode analysis in order to obtain the noise generation and propagation in turbines.

Topics: Acoustics , Turning , Turbines
Commentary by Dr. Valentin Fuster
2016;():V02AT41A012. doi:10.1115/GT2016-58175.

An end-to-end LES/FW-H noise prediction model has been demonstrated and validated with acoustic and flowfield data from a dual stream nozzle with pylon experiment conducted at NASA GRC using their Jet Engine Simulator (JES) geometry. Results show a large region of high turbulent kinetic energy (TKE) in the wake of the pylon. Acoustic Source Localization (ASL) studies using our numerical phased array methodology show this wake region to be the principle location of low frequency noise sources while higher frequency sources occur nearer to the nozzle lips. Numerical simulations have also been conducted on Jet-Surface Interaction (JSI) effects of a supersonic jet exhausting parallel to a finite surface. Time-averaged LES data and far-field noise predictions have been obtained for multiple surface locations as well as for an isolated jet nozzle. For upstream observers located below the surface, results show an increase in low-frequency noise over what was predicted for the isolated nozzle due to JSI effects and decrease in high-frequency noise due to shielding. This was significantly more pronounced for an over-expanded jet than for an under-expanded jet, an effect that was primarily attributed to the shorter core length of the over-expanded jet.

Commentary by Dr. Valentin Fuster

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