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J. Turbomach. 2019;141(9):. doi:10.1115/1.4044212.

The aerodynamic interaction of upstream and downstream blade rows can have a significant impact on the forced response of the compressor. Previously, the authors carried out the forced response analysis of a three-row stator-rotor-stator (S1-R2-S2) configuration from a 3.5-stage compressor. However, since the stator vane counts in both the stators (S1 and S2) were the same, it was not possible to separate the excitations from both the rows as they excited the rotor at the same frequency. Hence, a new configuration was developed and tested in which the stator 1 blade count was changed to 38 and stator 2 blade count was maintained at 44 in order to study the individual influences of the stator on the embedded rotor. By using this method, the excitations from both rows can be determined, and the excitations can be quantified to determine the row having the maximum influence on the overall forcing. To achieve this, two sets of simulations were carried out. The three-row stator-rotor (S1-R2-S2) simulation was carried out at both the 38EO (engine order) and 44EO crossings at the peak efficiency (PE) operating condition. The two-row stator-rotor analysis (S1-R2) was carried out at the 38EO crossing, and the other two-Row (R2-S2) analyses were carried out at the 44EO crossing. The steady aerodynamics was preserved in both the cases. A study was done to determine the contribution of wave reflections from the stator inlet and exit planes to the forcing function. Two conclusions drawn from this study are as follows: (1) the modal force value decreased after the upstream stator was removed, which proved that wave reflections from this stator were significant and (2) the increase in modal force was in-line with experimental observations.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(9):091001-091001-8. doi:10.1115/1.4043577.

The experimental results reported in this contribution address the time-dependent impact of periodically unsteady wakes on the development of profile and end wall boundary layers and consequently on the secondary flow system. Experimental investigations are conducted on an annular 1.5 stage axial turbine rig at Ruhr-Universität Bochum’s Chair of Thermal Turbomachines and Aeroengines. The object under investigation is a modified T106 profile low-pressure turbine (LPT) stator row at a representative exit flow Reynolds number of 200,000. By making use of an annular geometry instead of a linear cascade, the influence of curvilinear end walls, nonuniform, increasing pitch across the span and radial flow migration can be represented. Incoming wakes are generated by a variable-speed driven rotor equipped with cylindrical bars. Special emphasis is put on the wake-induced recurrent formation, suppression, weakening, and displacement of individual vortices and separated flow regimes. For this, based on a comprehensive set of time-resolved measurement data, the interaction of impinging bar wakes and boundary layer flow and thus separation and its periodic manipulation along the passage end walls and on the blade suction surface are studied within the frequency domain.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(9):091002-091002-11. doi:10.1115/1.4043407.

Blade flutter in the last stage is an important design consideration for the manufacturers of steam turbines. Therefore, the accurate prediction method for blade flutter is critical. Since the majority of aerodynamic work contributing to flutter is done near the blade tip, resolving the tip leakage flow can increase the accuracy of flutter predictions. The previous research has shown that the induced vortices in the tip region can have a significant influence on the flow field near the tip. The structure of induced vortices due to the tip leakage vortex cannot be resolved by unsteady Reynolds-averaged Navier–Stokes (URANS) simulations because of the high dissipation in turbulence models. To the best of author’s knowledge, the influence of induced vortices on flutter characteristics has not been investigated. In this paper, the results of detached-eddy simulation (DES) and URANS flutter simulations of a realistic-scale last-stage steam turbine are presented, and the influence of induced vortices on the flutter stability has been investigated. Significant differences for the predicted aerodynamic work coefficient distribution on the blade surface, especially on the rear half of the blade suction side near the tip, are observed. At the least stable interblade phase angle (IBPA), the induced vortices show a destabilizing effect on the blade aeroelastic stability. The motion of induced vortices during blade oscillation is dependent on the blade amplitude, and hence, the aerodynamic damping is also dependent on the blade vibration amplitude. In conclusion, the induced vortices can influence the predicted flutter characteristics of the steam turbine test case.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(9):091003-091003-9. doi:10.1115/1.4043280.

For the development of the latest generation of axial compressors, it is necessary to enlarge the design space by using advanced aerodynamic processes. This enables a further increase in efficiency and performance. The use of a tandem blade configuration in a transonic compressor row provides the possibility to enlarge the design space. It is necessary to address the design aspects a bit more in detail in order to efficiently apply this blading concept to turbomachinery. Therefore, in the current study, the known design aspects of tandem blading in compressors will be summed up under the consideration of the aerodynamic effects and construction characteristics of a transonic compressor tandem. Based on this knowledge, a new transonic compressor tandem cascade (DLR TTC) with an inflow Mach number of 0.9 is designed using modern numerical methods and a multi-objective optimization process. Three objective functions as well as three operating points are used in the optimization. Furthermore, both tandem blades and their arrangement are parameterized. From the resulting database of 1246 members, a final best member is chosen as the state-of-the-art design for further detailed investigation. The aim of the ensuing experimental and numerical investigation is to answer the question, whether the tandem cascade resulting from the modern design process fulfills the described design aspects and delivers the requested performance and efficiency criteria. The numerical simulations within the study are carried out with the DLR flow solver TRACE. The experiments are performed at the transonic cascade wind tunnel of DLR in Cologne. The inflow Mach number during the tests is 0.9, and the AVDR is adjusted to 1.3 (design value). Wake measurements with a three-hole probe are carried out in order to determine the cascade performance. The experimental results show an increase in losses and a reduction of the cascade deflection by about 2 deg compared to the design concept. Nevertheless, the experimental and numerical results allow a good understanding of the aerodynamic effects. In addition, planar PIV was applied in a single S1 plane located at midspan to capture the velocity field in the wake of blade 1 in order to analyze the wake flow in detail and describe its influence on the cascade deflection and loss behavior. Finally, an outlook will be given on what future tandem compressor research should be focused.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(9):091004-091004-12. doi:10.1115/1.4043329.

The flow through a transonic compressor cascade shows a very complex structure due to the occurring shock waves. In addition, the interaction of these shock waves with the blade boundary layer inherently leads to a very unsteady flow behavior. The aim of the current investigation is to quantify this behavior and its influence on the cascade performance as well as to describe the occurring transonic flow phenomena in detail. Therefore, an extensive experimental investigation of the flow in a transonic compressor cascade has been conducted within the transonic cascade wind tunnel of DLR Institute of Propulsion Technology at Cologne. In this process, the flow phenomena were thoroughly examined for an inflow Mach number of 1.21. The experiments investigate both the laminar and the turbulent shock wave boundary layer interaction within the blade passage and the resulting unsteady behavior. The experiments show a fluctuation range of the passage shock wave of about 10% chord for both cases, which is directly linked with a change of the inflow angle and of the operating point of the cascade. Thereafter, Reynolds-averaged Navier–Stokes (RANS) simulations have been performed aiming at the verification of the reproducibility of the experimentally examined flow behavior. Here, it is observed that the dominant flow effects are not reproduced by a steady numerical simulation. Therefore, a further unsteady simulation has been carried out to capture the unsteady flow behavior. The results from this simulation show that the fluctuation of the passage shock wave can be reproduced but not in the correct magnitude. This leads to a remaining weak point within the design process of transonic compressor blades because the working range will be overpredicted. The resulting conclusion of this study is that the use of scale-resolving methods such as LES or the application of DNS is necessary to correctly predict unsteadiness of the transonic cascade flow and its impact on the cascade performance.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(9):091005-091005-10. doi:10.1115/1.4043263.

The effect of film cooling on a transonic squealer tip has been examined in a high speed linear cascade, which operates at engine-realistic Mach and Reynolds numbers. Tests have been performed on two uncooled tip geometries with differing pressure side rim edge radii, and a cooled tip matching one of the uncooled cases. The pressure sensitive paint technique has been used to measure adiabatic film cooling effectiveness on the blade tip at a range of tip gaps and coolant mass flow rates. Complementary tip heat transfer coefficients have been measured using transient infrared thermography, and the effects of the coolant film on the tip heat transfer and engine heat flux were examined. The uncooled data show that the tip heat transfer coefficient distribution is governed by the nature of flow reattachments and impingements. The squealer tip can be broken down into three regions, each exhibiting a distinct response to a change in the tip gap, depending on the local behavior of the overtip leakage flow. Complementary computational fluid dynamics (CFD) shows that the addition of casing motion causes no change in the flow over the pressure side rim. Injected coolant interacts with the overtip leakage flow, which can locally enhance the tip heat transfer coefficient. The film effectiveness is dependent on both the coolant mass flow rate and tip clearance. At increased coolant mass flow, areas of high film effectiveness on the pressure side rim coincide strongly with a net heat flux reduction and in the subsonic tip region with low heat transfer coefficient.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(9):091006-091006-7. doi:10.1115/1.4043786.

The impact of the hub and shroud leakage flows on the compressor efficiency has been investigated for four compressor stages with flow coefficients of 0.017, 0.0265, 0.063, and 0.118 belonging to a family of centrifugal compressor stages, designed for process compressor applications. A very good agreement was observed between the measured and predicted performance when the detailed geometrical features were included in the calculations. The computational fluid dynamics (CFD) calculations indicated that addition of leakage cavities and leakage flows resulted in about 3% drop in stage polytropic efficiency for the highest flow coefficient stage. The detrimental effect of leakages increased to about 8% for the lowest flow coefficient stage investigated here. The increase in the compressor work input due to the disc windage and the leakage recirculation was estimated from the CFD calculations and compared with values obtained using 1D methods, showing a very good agreement between the two. The impact of parasitic losses on compressor efficiency has been investigated and the contribution of various loss sources to the stage efficiency is discussed.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(9):091007-091007-10. doi:10.1115/1.4043644.

Jet engines with boundary layer ingestion (BLI) could offer significant reductions in aircraft fuel burn compared with podded turbofans. However, the engine fans must run continuously with severe inlet distortion, which is known to reduce stability. In this paper, an experimental study has been completed on a low-speed rig fan operating with a BLI-type inlet distortion. Unsteady casing static pressure measurements have been made at multiple locations during stall events. Steady-state, full-annulus area traverses have also been performed at rotor inlet and exit at a near-stall operating point. The reduction in stability caused by BLI is found to be small. It is found that with BLI the fan can operate stably despite the presence of localized regions where the rotor operating point lies beyond the stability boundary measured in clean flow. With the BLI-type distortion applied, the measured rotor incidence varies around the annulus due to nonuniform upstream velocity and swirl. The measured amplitude of unsteady casing pressure fluctuations just prior to stall is found to correlate with the circumferential variation of rotor incidence, suggesting that rotor incidence is a key variable affecting the creation and growth of flow disturbances. In regions of high incidence, disturbances resembling local flow separations are initiated. However, in regions of low or negative incidence, any disturbances decay rapidly. Full rotating stall with BLI occurs when high incidence regions are widespread enough to sustain disturbances which can propagate around the entire annulus.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(9):091008-091008-13. doi:10.1115/1.4043660.

Purge air is injected in cavities at the hub of axial turbines to prevent hot mainstream gas ingestion into interstage gaps. This process induces additional losses for the turbine due to an interaction between the purge and mainstream flow. This paper investigates the flow in a low-speed linear cascade rig with upstream hub cavity at a Reynolds number commonly observed in modern low-pressure turbine stages by the use of numerical simulation. Numerical predictions are validated by comparing against experimental data available. Three different purge mass flow rates are tested using three different rim seal geometries. Numerical simulations are performed using a large-eddy simulation (LES) solver on structured grids. An investigation of the different mechanisms associated with the turbine flow including cavity and purge air is intended through this simplified configuration. The underlying mechanisms of loss are tracked using an entropy formulation. Once described for a baseline case, the influence of purge flow and rim seal geometry on flow mechanisms and loss generation is described with the emphasis to obtain design parameters for losses reduction. The study quantifies loss generation due to the boundary layer on wetted surfaces and secondary vortices developing in the passage. The analysis shows different paths by which the purge flow and rim seal geometry can change loss generation including a modification of the shear layer between purge and mainstream, interaction with secondary vortices, and a modification of the flow behavior close to hub compared with a smooth configuration. The study shows the influence of purge flow rate and swirl on the strengthening of secondary vortices in the passage and the ability of axial overlapping rim seal to delay the development of secondary vortices compared with simple axial gaps.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(9):091009-091009-14. doi:10.1115/1.4043781.

As aero gas turbine designs strive for ever greater efficiencies, the trend is for engine overall pressure ratios to rise. Although this provides greater thermal efficiency, it means that cycle temperatures also increase. One potential solution to managing the increasing temperatures is to employ a cooled cooling air system. In such a system, a purge flow into the main gas path downstream of the compressor will be required to prevent hot gas being ingested into the rotor drive cone cavity. However, the main gas path in compressors is aerodynamically sensitive and it is important to understand, and mitigate, the impact such a flow may have on the compressor outlet guide vanes, pre-diffuser, and the downstream combustion system aerodynamics. Initial computational fluid dynamics (CFD) predictions demonstrated the potential of the purge flow to negatively affect the outlet guide vanes and alter the inlet conditions to the combustion system. The purge flow modified the incidence onto the outlet guide vane, at the hub, such that the secondary flows increased in magnitude. An experimental assessment carried out using an existing fully annular, isothermal test facility confirmed the CFD results and importantly demonstrated that the degradation in the combustor inlet flow resulted in an increased combustion system loss. At the proposed purge flow rate, equal to ∼1% of the mainstream flow, these effects were small with the system loss increasing by ∼4%. However, at higher purge flow rates (up to 3%), these effects became notable and the outlet guide vane and pre-diffuser flow degraded significantly with a resultant increase in the combustion system loss of ∼13%. To mitigate these effects, CFD was used to examine the effect of varying the purge flow swirl fraction in order to better align the flow at the hub of the outlet guide vane. With a swirl fraction of 0.65 (x rotor speed), the secondary flows were reduced below that of the datum case (with no purge flow). Experimental data showed good agreement with the predicted flow topology and performance trends but the measured data showed smaller absolute changes. Differences in system loss were measured with savings of around 10% at the turbine feed ports for a mass flow ratio of 1% and a swirl fraction of 0.65.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(9):091010-091010-13. doi:10.1115/1.4043953.
OPEN ACCESS

Engine durability tests are used by manufacturers to demonstrate engine life and minimum performance when subjected to doses of test dusts, often Arizona Road Dust. Grain size distributions are chosen to replicate what enters the engine; less attention is paid to other properties such as composition and shape. We demonstrate here the differences in the probability of interaction of a particle of a given particle Reynolds number on to a vane if particle shape, vane geometry, and flow Reynolds number are varied and discuss why the traditional definition of Stokes number is inadequate for predicting the likelihood of interaction in these flows. We develop a new generalized Stokes number for nozzle guide vanes and demonstrate its use through application to 2D sections of the General Electric E3 nozzle guide vane. The new Stokes number is used to develop a reduced-order probability curve to predict the interaction efficiency of spherical and nonspherical particles, independent of flow conditions and vane geometry. We show that assuming spherical particles instead of more realistic sphericity of 0.75 can lead to as much as 25% difference in the probability of interaction at Stokes numbers of around unity. Finally, we use a hypothetical size distribution to demonstrate the application of the model to predict the total mass fraction of dust interaction with a nozzle guide vane at design point conditions and highlight the potential difference in the accumulation factor between spherical and nonspherical particles.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(9):091011-091011-13. doi:10.1115/1.4043974.

The Boxprop is a novel, double-bladed, tip-joined propeller for high-speed flight. The concept draws inspiration from the box wing concept and could potentially decrease tip vortex strength compared with conventional propeller blades. Early Boxprop designs experienced significant amounts of blade interference. By performing a wake analysis and quantifying the various losses of the flow, it could be seen that these Boxprop designs produced 45% more swirl than a conventional reference blade. The reason for this was the proximity of the Boxprop blade halves to each other, which prevented the Boxprop from achieving the required aerodynamic loading on the outer parts of the blade. This paper presents an aerodynamic optimization of a 6-bladed Boxprop aiming at maximizing efficiency and thrust at cruise. A geometric parametrization has been adopted which decreases interference by allowing the blade halves to be swept in opposite directions. Compared with an earlier equal-thrust Boxprop design, the optimized design features a 7% percentage point increase in propeller efficiency and a lower amount of swirl and entropy generation. A vortex-like structure has also appeared downstream of the optimized Boxprop, but with two key differences relative to conventional propellers. (1) Its formation differs from a traditional tip vortex and (2) it is 46% weaker than the tip vortex of an optimized 12-bladed conventional propeller.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(9):091012-091012-13. doi:10.1115/1.4043968.

The objective of this study is to examine the effect of the geometrical modification of the land on the overall film cooling effectiveness on the cutback region of a turbine blade model. A room-temperature experiment was conducted, in which nitrogen serves as the cooling stream, and the mainstream flow is air. The adiabatic film cooling effectiveness was mapped employing the pressure-sensitive paint (PSP) technique. Data was acquired at five different blowing ratios (from 0.45 to 1.65) for both the baseline and the modified model. Detailed film cooling effectiveness from PSP measurements in correlation with the flow map in streamwise and spanwise planes from particle image velocimetry (PIV) measurements was performed, characterizing the effect of rounding the edges of the lands. The results show that the rounded edges enable the coolant flow to reach the top surface of the land area more readily, especially at low blowing ratios. Superior coolant coverage on the land surface observed in the PSP measurements are well correlated with the PIV measurements. At the high blowing ratio of 1.65, the round edge of the lands helps regulate the mixing between the coolant and mainstream flows, therefore the film cooling effectiveness in the slot region is also improved.

Commentary by Dr. Valentin Fuster

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