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Research Papers

J. Turbomach. 2018;140(5):051001-051001-10. doi:10.1115/1.4038876.

This paper introduces a new approach for the preliminary design and aerothermal analysis of centrifugal impellers using a relative diffusion effectiveness parameter. The relative diffusion effectiveness is defined as the ratio of the achieved diffusion to the maximum available diffusion in an impeller. It represents the quality of the relative diffusion process in an impeller. This parameter is used to evaluate impeller performance by correlating the relative diffusion effectiveness with the impeller isentropic efficiency using the experimental data acquired on a single-stage centrifugal compressor (SSCC). By including slip, which is appropriate considering it is an inviscid effect that should be included in the determination of maximum available diffusion in the impeller, a linear correlation between impeller efficiency and relative diffusion effectiveness resulted for all operating conditions. Additionally, a new method for impeller preliminary design was introduced using the relative diffusion effectiveness parameter, in which the optimal design is selected to maximize relative diffusion effectiveness. While traditional preliminary design methods are based on empirical loss models or empirical knowledge for selection of diffusion factor (DF) in the impeller, the new method does not require any such models, and it also provides an analytical approach for the selection of DF that gives optimal impeller performance. Validation of the method was performed using three classic impeller designs available in the open literature, and very good agreement was achieved. Furthermore, a sensitivity study shows that the method is robust in that the resulting flow angles at the impeller inlet and exit are insensitive to a wide range of blockage factors and various slip models.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(5):051002-051002-13. doi:10.1115/1.4038871.

Shaped film cooling holes are used as a cooling technology in gas turbines to reduce metal temperatures and improve durability, and they generally consist of a small metering section connected to a diffuser that expands in one or more directions. The area ratio (AR) of these holes is defined as the area at the exit of the diffuser, divided by the area at the metering section. A larger AR increases the diffusion of the coolant momentum, leading to lower average momentum of the coolant jet at the exit of the hole and generally better cooling performance. Cooling holes with larger ARs are also more tolerant of high blowing ratio conditions, and the increased coolant diffusion typically better prevents jet lift-off from occurring. Higher ARs have traditionally been accomplished by increasing the expansion angle of the diffuser while keeping the overall length of the hole constant. The present study maintains the diffuser expansion angles and instead increases the length of the diffuser, which results in longer holes. Various ARs have been examined for two shaped holes: one with forward and lateral expansion angles of 7 deg (7-7-7 hole) and one with forward and lateral expansion angles of 12 deg (12-12-12 hole). Each hole shape was tested at numerous blowing ratios to capture trends across various flow rates. Adiabatic effectiveness measurements indicate that for the baseline 7-7-7 hole, a larger AR provides higher effectiveness, especially at higher blowing ratios. Measurements also indicate that for the 12-12-12 hole, a larger AR performs better at high blowing ratios but the hole experiences ingestion at low blowing ratios. Steady Reynolds-averaged Navier–Stokes simulations did not accurately predict the levels of adiabatic effectiveness, but did predict the trend of improving effectiveness with increasing AR for both hole shapes. Flowfield measurements with particle image velocimetry (PIV) were also performed at one downstream plane for a low and high AR case, and the results indicate an expected decrease in jet velocity due to a larger diffuser.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(5):051003-051003-11. doi:10.1115/1.4038872.

The estimation of boundary layer losses requires the accurate specification of the freestream velocity, which is not straightforward in centrifugal compressor blade passages. This challenge stems from the jet-wake flow structure, where the freestream velocity between the blades cannot be clearly specified. In addition, the relative velocity decreases due to adverse pressure gradient. Therefore, the common assumption of a single freestream velocity over the blade surface might not be valid in centrifugal compressors. Generally in turbomachinery, the losses in the blade cascade boundary layers are estimated, e.g., with different loss coefficients, but they often rely on the assumption of a uniform flow field between the blades. To give guidelines for the estimation of the mentioned losses in highly distorted centrifugal compressor flow fields, this paper discusses the difficulties in the calculation of the boundary layer thickness in the compressor blade passages, compares different freestream velocity definitions, and demonstrates their effect on estimated boundary layer losses. Additionally, a hybrid method is proposed to overcome the challenges of defining a boundary layer in centrifugal compressors.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(5):051004-051004-11. doi:10.1115/1.4038868.

The aeroelastic prediction of blade forcing is still a very important topic in turbomachinery design. Usually, the wake from an upstream airfoil and the potential field from a downstream airfoil are considered as the main disturbances. In recent years, it became evident that in addition to those two mechanisms, Tyler–Sofrin modes, also called scattered or spinning modes, may have a significant impact on blade forcing. It was recently shown in literature that in multirow configurations, not only the next but also the next but one blade row is very important as it may create a large circumferential forcing variation, which is fixed in the rotating frame of reference. In the present paper, a study of these effects is performed on the basis of a quasi three-dimensional (3D) multirow and multipassage compressor configuration. For the analysis, a harmonic balancing code, which was developed by DLR Cologne, is used for various setups and the results are compared to full-annulus unsteady calculations. It is shown that the effect of the circumferentially different blade excitation is mainly contributed by the Tyler–Sofrin modes and not to blade-to-blade variation in the steady flow field. The influence of various clocking positions, coupling schemes and number of harmonics onto the forcing is investigated. It is also shown that along a speed-line in the compressor map, the blade-to-blade forcing variation may change significantly. In addition, multirow flutter calculations are performed, showing the influence of the upstream and downstream blade row onto aerodynamic damping. The effect of these forcing variations onto random mistuning effects is investigated in the second part of the paper.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(5):051005-051005-9. doi:10.1115/1.4038869.

The prediction of aerodynamic blade forcing is a very important topic in turbomachinery design. Usually, the wake from the upstream blade row and the potential field from the downstream blade row are considered as the main causes for excitation, which in conjunction with relative rotation of neighboring blade rows, give rise to dynamic forcing of the blades. In addition to those two mechanisms, the so-called Tyler–Sofrin (or scattered or spinning) modes, which refer to the acoustic interaction with blade rows further up- or downstream, may have a significant impact on blade forcing. In particular, they lead to considerable blade-to-blade variations of the aerodynamic loading. In Part I of the paper, a study of these effects is performed on the basis of a quasi-three-dimensional multirow and multipassage compressor configuration. Part II of the paper proposes a method to analyze the interaction of the aerodynamic forcing asymmetries with the already well-studied effects of random mistuning stemming from blade-to-blade variations of structural properties. Based on a finite element (FE) model of a sector, the equations governing the dynamic behavior of the entire bladed disk can be efficiently derived using substructuring techniques. The disk substructure is assumed as cyclically symmetric, while the blades exhibit structural mistuning and linear aeroelastic coupling. In order to avoid the costly multistage analysis, the variation of the aerodynamic loading is treated as an epistemic uncertainty, leading to a stochastic description of the annular force pattern. The effects of structural mistuning and stochastic aerodynamic forcing are first studied separately and then in a combined manner for a blisk of a research compressor without and with aeroelastic coupling.

Topics: Excitation , Blades , Disks
Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(5):051006-051006-10. doi:10.1115/1.4038878.

Variable geometry turbine (VGT) has been widely applied in internal combustion engines to improve engine transient response and torque at light load. One of the most popular VGTs is the variable nozzle turbine (VNT) in which the nozzle vanes can be rotated along the pivoting axis and thus the flow passage through the nozzle can be adjusted to match with different engine operating conditions. One disadvantage of the VNT is the turbine efficiency degradation due to the leakage flow in the nozzle endwall clearance, especially at small nozzle open condition. With the purpose to reduce the nozzle leakage flow and to improve turbine stage efficiency, a novel split sliding variable nozzle turbine (SSVNT) has been proposed. In the SSVNT design, the nozzle is divided into two parts: one part is fixed and the other part can move along the partition surface. When sliding the moving vane to large radius position, the nozzle flow passage opens up and the turbine has high flow capacity. When sliding the moving vane to small radius position, the nozzle flow passage closes down and the turbine has low flow capacity. As the fixed vane does not need endwall clearance, the leakage flow through the nozzle can be reduced. Based on calibrated numerical simulation, there is up to 12% turbine stage efficiency improvement with the SSVNT design at small nozzle open condition while maintaining the same performance at large nozzle open condition. The mechanism of efficiency improvement in the SSVNT design has been discussed.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(5):051007-051007-9. doi:10.1115/1.4038879.

From the efforts of many researchers and engineers related to the S-CO2 Brayton cycle technology development, the S-CO2 Brayton cycle is now considered as one of the key power technologies for the future. Since the S-CO2 Brayton cycle has advantages in economics due to high efficiency and compactness of the system, various industries have been trying to develop baseline technology on the design and analysis of the S-CO2 Brayton cycle components. According to the previous researches on the S-CO2 Brayton cycle component technology, the treatment of a thermodynamic property near the critical point of CO2 is one of the main concerns since conventional design and analysis methodologies cannot be used for the near critical point region. Among many thermodynamic properties, the stagnation to static condition conversion process is important since the flow in a compressor is at high flow velocity. In this paper, the impact of various stagnation to static conversion methods on the S-CO2 compressor design near the critical point will be evaluated. From the evaluation, the limitation of a certain stagnation to static conversion method will be discussed to provide a guideline for the future S-CO2 compressor designers.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(5):051008-051008-11. doi:10.1115/1.4039051.

The original Spalart–Allmaras (SA) model is known to predict premature stall when applied to fan or compressor, which is in line with the observation of other researchers who use the SA model. Therefore, to improve the prediction of the stall boundary, the original SA model was modified by scaling the source term based on the local pressure gradient and the velocity helicity of the flow. Furthermore, a generalized wall function valid for nonzero wall pressure gradient was implemented to improve the accuracy of boundary conditions at the solid wall. This work aims to produce a turbulence model which can be used to model flows near the stall boundary for the transonic fan rotors on relatively coarse grids of around 600k points per passage. Initially, two fan rotors with different design and operating speeds were used to optimize the new parameters in the modified turbulence model. The optimization was based on improving the correlation between measured and numerical radial profiles of the pressure ratio. Thereafter, steady computations were performed for two other fans (by using the same parameters), and the predictions were compared with the experimental data for all the four fan rotors. Numerical results showed a significant improvement over those obtained with the original SA model, when compared against the measured data. Finally, for completeness, it was decided to test the performance of the modified model by comparing the result with measured data for a simple canonical case.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(5):051009-051009-8. doi:10.1115/1.4039052.

This paper investigates the flow near the stall boundary for a low-speed/low-pressure ratio fan. Three-dimensional, Reynolds-averaged Navier–Stokes computations are performed for a modern low speed fan rig for which extensive measured data are available. Simulations are conducted at 80% corrected speed, for which the measured constant speed characteristic contains a part with positive slope. It is shown in this paper that by using an unsteady whole assembly approach, it is possible to predict the flow for all the points on the measured constant speed characteristic (including those on the positive slope part), which is not achievable by using a single passage strategy as it would result in premature “numerical stall.” The results of the computations reveal that for the operating points on the positive slope part of the characteristic, the flow structure becomes asymmetric and hence requires a whole assembly numerical model. The type of asymmetry which appears at lower flow coefficients is similar to the multicell, part span rotating stall, which can occur on the front stages of core compressors at stable operating conditions. The numerical results showed a good correlation with the measured data in terms of stall characteristics.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(5):051010-051010-13. doi:10.1115/1.4038907.

In gas turbines with can combustors, the trailing edge (TE) of the combustor transition duct wall is found upstream of every second vane. This paper presents an experimental and numerical investigation of the effect of the combustor wall TE on the aerothermal performance of the nozzle guide vane. In the measurements carried out in a high-speed experimental facility, the wake of this wall is shown to increase the aerodynamic loss of the vane. On the other hand, the wall alters secondary flow structures and has a protective effect on the heat transfer in the leading edge-endwall junction, a critical region for component life. The different clocking positions of the vane relative to the combustor wall are tested experimentally and are shown to alter the aerothermal field. The experimental methods and processing techniques adopted in this work are used to highlight the differences between the different cases studied.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(5):051011-051011-10. doi:10.1115/1.4038908.

A transonic centrifugal compressor was aerodynamically optimized by means of a numerical optimization process. The objectives were to increase the isentropic efficiency and to reduce the acoustic signature by decreasing the amplitude of pre-shock pressure waves at the inlet of the compressor. The optimization was performed at three operating points on the 100% speed line in order to maintain choke mass flow and surge margin. At the design point, the specific work input was kept equal. The baseline impeller was designed by using ruled surfaces due to requirements for flank milling. To investigate the benefits of arbitrary blade surfaces, the restrictions of ruled surfaces were abolished and fully three-dimensional (3D) blade profiles allowed. In total, therefore, 45 parameters were varied during the optimization. The combined geometric and aerodynamic analysis reveals that a forward swept leading edge (LE) and a concave suction side at the tip of the LE are effective design features for reducing the shock strength. Beyond that, the blade shape of the optimized compressor creates a favorable impeller outlet flow, which is the main reason why the performance of the vaneless diffuser improves. In total, a gain of 1.4% points in isentropic total-to-static efficiency, evaluated by computational fluid dynamics (CFD) at the exit plane of the vaneless diffuser, is achieved.

Commentary by Dr. Valentin Fuster

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