Research Papers

J. Turbomach. 2019;141(7):071001-071001-14. doi:10.1115/1.4042565.

The combined action of Coriolis and centrifugal buoyancy forces results in nonuniform heat transfer coefficient on pressure and suction side internal walls, hence leading to nonuniform metal temperatures and increased thermal stresses. The present study addresses the problem of nonuniform heat transfer distribution due to rotation effect and proposes novel designs for serpentine cooling passages, which are arranged along the chord of the blade. The two configurations were four-passage and six-passage serpentine smooth channels. Detailed heat transfer coefficients were measured using transient liquid crystal thermography under stationary and rotating conditions. Heat transfer experiments were carried out for Reynolds numbers ranging from 12,294 to 85,000 under stationary conditions. Rotation experiments were carried out for the Rotation numbers of 0.05 and 0.11. Heat transfer enhancement levels of approximately two times the Dittus–Boelter correlation (for developed flow in smooth tubes) were obtained under stationary conditions. Under rotating conditions, we found that the four-passage configuration had slightly lower heat transfer compared with the stationary case, and the six-passage configuration had higher heat transfer on both the leading and trailing sides compared with the stationary case. The leading and trailing side heat transfer characteristics were near-similar to each other for both the configurations, and the rotating heat transfer was near-similar to the stationary condition heat transfer.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(7):071002-071002-9. doi:10.1115/1.4042569.

Slot film cooling is a popular choice for trailing edge (TE) cooling in high pressure (HP) turbine blades because it can provide more uniform film coverage compared to discrete film cooling holes. The slot geometry consists of a cutback in the blade pressure side connected through rectangular openings to the internal coolant feed passage. The numerical simulation of this kind of film cooling flows is challenging due to the presence of flow interactions such as step flow separation, coolant-mainstream mixing, and heat transfer. The geometry under consideration is a cutback surface at the trailing edge of a constant cross-section aerofoil. The cutback surface is divided into three sections separated by narrow lands. The experiments are conducted in a high-speed cascade in Oxford Osney Thermo-Fluids Laboratory at Reynolds and Mach number distributions representative of engine conditions. The capability of computational fluid dynamics (CFD) methods to capture these flow phenomena is investigated in this paper. The isentropic Mach number and film effectiveness are compared between CFD and pressure sensitive paint (PSP) data. When compared with the steady k − ω shear stress transport (SST) method, scale adaptive simulation (SAS) can agree better with the measurement. Furthermore, the profiles of kinetic energy, production, and shear stress obtained by the steady and SAS methods are compared to identify the main source of inaccuracy in RANS simulations. The SAS method is better to capture the unsteady coolant–hot gas mixing and vortex shedding at the slot lip. The cross flow is found to affect the film significantly as it triggers flow separation near the lands and reduces the effectiveness. The film is nonsymmetric with respect to the half-span plane, and different flow features are present in each slot. The effect of mass flow ratio (MFR) on flow pattern and coolant distribution is also studied. The profiles of velocity, kinetic energy, and production of turbulent energy are compared among the slots in detail. The MFR not only affects the magnitude but also changes the sign of production.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(7):071003-071003-10. doi:10.1115/1.4042648.

Rotation-induced Coriolis and centrifugal buoyancy forces result in significant modification of cooling characteristics of blade pressure and suction side internal walls. The nonuniformity in cooling, coupled with high-speed rotation, results in increased levels of thermal stresses. To address this problem, this study presents two multipassage configurations featuring 45-deg angled turbulators, in four- and six-passage designs. Experiments were carried out under stationary and rotating conditions using transient liquid crystal thermography to measure detailed heat transfer coefficient. It has been shown through experimental data that heat transfer characteristics of the new configurations’ pressure and suction side internal walls were very similar under rotating conditions, at both local and global scales. The heat transfer levels under rotating conditions were also similar to those of the stationary conditions. The contribution of multiple passages connected with 180-deg bends toward overall frictional losses has been evaluated in terms of pumping power and normalized friction factor. The configurations are ranked based on their thermal hydraulic performances over a wide range of Reynolds numbers. The four-passage ribbed configuration had slightly higher heat transfer levels compared with those of the corresponding six-passage ribbed configuration.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(7):071004-071004-18. doi:10.1115/1.4042654.

This paper numerically investigated the impact of the holes and their location on the flow and tip internal heat transfer in a U-bend channel (aspect ratio = 1:2), which is applicable to the cooling passage with dirt purge holes in the mid-chord region of a typical gas turbine blade. Six different tip ejection configurations are calculated at Reynolds numbers from 25,000 to 200,000. The detailed three-dimensional flow and heat transfer over the tip wall are presented, and the overall thermal performances are evaluated. The topological methodology, which is first applied to the flow analysis in an internal cooling passage of the blade, is used to explore the mechanisms of heat transfer enhancement on the tip wall. This study concludes that the production of the counter-rotating vortex pair in the bend region provides a strong shear force and then increases the local heat transfer. The side-mounted single hole and center-mounted double holes can further enhance tip heat transfer, which is attributed to the enhanced shear effect and disturbed low-energy fluid. The overall thermal performance of the optimum hole location is a factor of 1.13 higher than that of the smooth tip. However, if double holes are placed on the upstream of a tip wall, the tip surface cannot be well protected. The results of this study are useful for understanding the mechanism of heat transfer enhancement in a realistic gas turbine blade and for efficient designing of blade tips for engine service.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(7):071005-071005-9. doi:10.1115/1.4042731.

A combined experimental and computational test program, with two low-pressure ratio aero-engine fans, has been used to identify the flow mechanisms at stall inception and the subsequent stall cell growth. The two fans have the same rotor tip clearance, annulus design, and downstream stators, but different levels of tip loading. The measurement data show that both the fans stall via spike-type inception, but that the growth of the stall cell and the final cell size is different in each fan. The computations, reproducing both the qualitative and quantitative behavior of the steady-state and transient measurements, are used to identify the flow mechanisms at the origin of stall inception. In one fan, spillage of tip leakage flow upstream of the leading edge plane is responsible. In the other, sudden growth of casing corner separation blockage leads to stall. These two mechanisms are in accord with the findings from core compressors. However, the transonic aerodynamics and the low hub-to-tip radius ratio of the fans lead to the following two findings: first, the casing corner separation is driven by shock-boundary layer interaction and second, the spanwise loading distribution of the fan determines whether the spike develops into full-span or part-span stall and both types of behavior are represented in the present work. Finally, the axial momentum flux of the tip clearance flow is shown to be a useful indicator of the leakage jet spillage mechanism. A simple model is provided that links the tip loading, stagger, and solidity with the tip clearance axial momentum flux, thereby allowing the aerodynamicist to connect, qualitatively, design parameters with the stall behavior of the fan.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(7):071006-071006-9. doi:10.1115/1.4042645.

The aim of this paper is to introduce design modifications that can be made to improve the flutter stability of a fan blade. A rig fan blade, which suffered flutter in the part-speed range and for which good quality measured data in terms of steady flow and flutter boundary is available, is used for this purpose. The work is carried out numerically using the aeroelasticity code AU3D. Two different approaches are explored: aerodynamic modifications and aero-acoustic modifications. In the first approach, the blade is stabilized by altering the radial distribution of the stagger angle based on the steady flow on the blade. The re-staggering patterns used in this work are therefore particular to the fan blade under investigation. Moreover, the modifications made to the blade are very simple and crude, and more sophisticated methods and/or an optimization approach could be used to achieve the above objectives with a more viable final design. This paper, however, clearly demonstrates how modifying the steady blade aerodynamics can prevent flutter. In the second approach, flutter is removed by drawing bleed air from the casing above the tip of the blade. Only a small amount of bleed (0.2% of the total inlet flow) is extracted such that the effect on the operating point of the fan is small. The purpose of the bleed is merely to attenuate the pressure wave that propagates from the trailing edge to the leading edge of the blade. The results show that extracting bleed over the tip of the fan blade can improve the flutter margin of the fan significantly.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(7):071007-071007-9. doi:10.1115/1.4042647.

Leakage flow between the rotating turbine blade tip and the fixed casing causes high heat loads and thermal stress on the tip and near the tip region. For this study, new squealer tips called partial cavity tips, which combine the advantages of plane and squealer tips, were suggested, and the effects of the cavity shape on the tip heat transfer coefficient and film cooling effectiveness were investigated experimentally in a low-speed linear cascade. The suggested blade tips had a flat surface near the leading edge and a squealer cavity from the mid-chord to trailing edge region to achieve the advantages of both blade tip types. The heat transfer coefficient was measured via the 1-D transient heat transfer technique using an IR camera, and the film cooling effectiveness was obtained via the pressure-sensitive paint (PSP) technique. Results showed that the heat transfer coefficient and film cooling effectiveness on the partial cavity tips strongly depended on the cavity shape. Near the leading edge, the heat transfer coefficients for the partial cavity tip cases were lower than that for the squealer tip case. However, the heat transfer coefficient on the cavity surface was higher for the partial cavity tip cases. The D10 tip showed a similar distribution of film cooling effectiveness to that of the plane (PLN) tip near the leading edge and the double side squealer (DSS) tip near the mid-chord region. However, the overall average film cooling effectiveness of the DSS tip was higher than that of the D10 tip.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(7):071008-071008-10. doi:10.1115/1.4042653.

The present study features a two-pass rectangular channel with an aspect ratio (AR) = 4:1 in the first pass and an AR = 2:1 in the second pass after a 180-deg tip turn. In addition to the smooth-wall case, ribs with a profiled cross section are placed at 60 deg to the flow direction on both the leading and trailing surfaces in both passages (P/e = 10, e/Dh ∼ 0.11, parallel and in-line). Regionally averaged heat transfer measurement method was used to obtain the heat transfer coefficients on all internal surfaces. The Reynolds number (Re) ranges from 10,000 to 70,000 in the first passage, and the rotational speed ranges from 0 to 400 rpm. Under pressurized condition (570 kPa), the highest rotation number achieved was Ro = 0.39 in the first passage and 0.16 in the second passage. The results showed that the turn-induced secondary flows are reduced in an accelerating flow. The effects of rotation on heat transfer are generally weakened in the ribbed case than the smooth case. Significant heat transfer reduction (∼30%) on the tip wall was seen in both the smooth and ribbed cases under rotating condition. Overall pressure penalty was reduced for the ribbed case under rotation. Reynolds number effect was found noticeable in the current study. The heat transfer and pressure drop characteristics are sensitive to the geometrical design of the channel and should be taken into account in the design process.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(7):071009-071009-17. doi:10.1115/1.4042646.

Truncating the exit of a discrete passage centrifugal compressor diffuser is observed to enhance a research compressor's stall line. By interrogating the experimental data along with a set of well-designed Reynolds-Averaged Navier–Stokes computations, this improvement is traced to the reduced impact of secondary flows on the truncated diffuser's boundary layer growth. The secondary flow system is characterized by counter-rotating streamwise vortex pairs that persist throughout the diffuser passage. The vortices originate from two sources: flow nonuniformity at the impeller exit and separation off the leading edge cusps unique to a discrete passage diffuser. The latter detrimentally impacts the diffuser pressure rise capability by accumulating high loss flow along the diffuser wall near the plane of symmetry between the vortices. This contributes to a large passage separation in the baseline diffuser. Using reduced-order modeling, the impact of the vortices on the boundary layer growth is shown to scale inversely with the diffuser aspect ratio, and thus, the separation extent is reduced for the truncated diffuser. Because the diffuser incidence angle influences the strength and location of the vortices, this mechanism can affect the slope of the compressor's pressure rise characteristic and impact its stall line. Stall onset for the baseline diffuser configuration is initiated when the vortex location and the corresponding passage separation transition from pressure to suction side with increased cusp incidence. Conversely, because the extent of the passage separation in the truncated diffuser is diminished, the switch in separation side does not immediately initiate instability.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(7):071010-071010-9. doi:10.1115/1.4043343.

In a boundary layer ingesting (BLI) fan system, the inlet flow field is highly nonuniform. In this environment, an axisymmetric stator design suffers from a nonuniform distribution of hub separations, increased wake thicknesses, and casing losses. These additional loss sources can be reduced using a nonaxisymmetric design that is tuned to the radial and circumferential flow variations at exit from the rotor. In this paper, a nonaxisymmetric design approach is described for the stator of a low-speed BLI fan. First, sectional design changes are applied at each radial and circumferential location. Next, this approach is combined with the application of nonaxisymmetric lean. The designs were tested computationally using full-annulus unsteady computational fluid dynamics (CFD) of the complete fan stage with a representative inlet distortion. The final design has also been manufactured and tested experimentally. The results show that a 2D sectional approach can be applied nonaxisymmetrically to reduce incidence and diffusion factor at each location. This leads to reduced loss, particularly at the casing and midspan, but it does not eliminate the hub separations that are present within highly distorted regions of the annulus. These are relieved by nonaxisymmetric lean where the pressure surface is inclined toward the hub. For the final design, the loss in the stator blades operating with BLI was measured to be 10% lower than that for the original stator design operating with undistorted inflow. Overall, the results demonstrate that the nonaxisymmetric design has the potential to eliminate any additional loss in a BLI fan stator caused by the nonuniform ingested flow field.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(7):071011-071011-9. doi:10.1115/1.4043151.

Magnetic resonance thermometry (MRT) is a maturing diagnostic tool used to measure three-dimensional temperature fields. It has a great potential for investigating fluid flows within complex geometries leveraging medical grade magnetic resonance imaging (MRI) equipment and software along with novel measurement techniques. The efficacy of the method in engineering applications increases when coupled with other well-established MRI-based techniques such as magnetic resonance velocimetry (MRV). In this study, a challenging geometry is presented with the direct application to a complex gas turbine blade cooling scheme. Turbulent external flow with a Reynolds number of 136,000 passes a hollowed NACA-0012 airfoil with internal cooling features. Inserts within the airfoil, fed by a second flow line with an average temperature difference of 30 K from the main flow and a temperature-dependent Reynolds number in excess of 1,800, produces a conjugate heat transfer scenario including impingement cooling on the inside surface of the airfoil. The airfoil cooling scheme also includes zonal recirculation, surface film cooling, and trailing edge ejection features. The entire airfoil surface is constructed of a stereolithography resin—Accura 60—with low thermal conductivity. The three-dimensional internal and external velocity field is measured using an MRV. The fluid temperature field is measured within and outside of the airfoil with an MRT, and the results are compared with a computational fluid dynamics (CFD) solution to assess the current state of the art for combined MRV/MRT techniques for investigating these complex internal and external flows. The accompanying CFD analysis provides a prediction of the velocity and temperature fields, allowing for errors in the MRT technique to be estimated.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(7):071012-071012-9. doi:10.1115/1.4043474.

In this paper, the effect of aerodynamic mistuning on stability of a compressor cascade is studied. The experiments have been carried out at a low-speed test facility of the Technische Universität Berlin. The test section contains a linear cascade with compressor blades that are forced to oscillate in sinusoidal pitching motion. The aerodynamic mistuning is realized by a blade-to-blade stagger angle variation, and three mistuning patterns have been investigated: one-blade mis-staggering, alternating mis-staggering, and random mis-staggering. Mis-staggering can have a stabilizing or destsabilizing effect, but depends strongly on the amount of detuning that alters the flow passage. For positive stagger angle variation for the one-blade and alternating mis-staggering, the trend of the damping curve was maintained, in the sense that the unstable interblade phase angles (IBPAs) remained unstable. For negative stagger angle variation, one IBPA shifted from stable to unstable. For the random pattern, only very moderate changes are observed. The cascade stability was not noticeably affected by the aerodynamic mistuning.

Commentary by Dr. Valentin Fuster

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