Research Papers

J. Turbomach. 2017;139(7):071001-071001-11. doi:10.1115/1.4035450.

Time-accurate transient blade row (TBR) simulation approaches are required when there is a close flow coupling between the blade rows, and for fundamentally transient flow phenomena such as aeromechanical analysis. Transient blade row simulations can be computationally impractical when all of the blade passages must be modeled to account for the unequal pitch between the blade rows. In order to reduce the computational cost, time-accurate pitch-change methods are utilized so that only a sector of the turbomachine is modeled. The extension of the time-transformation (TT) pitch-change method to multistage machines has recently shown good promise in predicting both aerodynamic performance and resolving dominant blade passing frequencies for a subsonic compressor, while keeping the computational cost affordable. In this work, a modified 1.5 stage Purdue transonic compressor is examined. The goal is to assess the ability of the multistage time-transformation method to accurately predict the aerodynamic performance and transient flow details in the presence of transonic blade row interactions. The results from the multistage time-transformation simulation were compared with a transient full-wheel simulation. The aerodynamic performance and detailed flow features from the time-transformation solution closely matched the full-wheel simulation at fractional of the computation cost.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(7):071002-071002-11. doi:10.1115/1.4035519.

This paper describes the flow mechanisms of rotating stall inception in a multistage axial flow compressor of an actual gas turbine. Large-scale numerical simulations of the unsteady have been conducted. The compressor investigated is a test rig compressor that was used in the development of the Kawasaki L30A industrial gas turbine. While the compressor consists of a total of 14 stages, only the front stages of the compressor were analyzed in the present study. The test data show that the fifth or sixth stages of the machine are most likely the ones leading to stall. To model the precise flow physics leading to stall inception, the flow was modeled using a very dense computational mesh, with several million cells in each passage. A total of 2 × 109 cells were used for the first seven stages (3 × 108 cells in each stage). Since the mesh was still not fine enough for large-eddy simulation (LES), a detached-eddy simulation (DES) was used. Using DES, a flow field is calculated using LES except in the near-wall where the turbulent eddies are modeled by Reynolds-averaged Navier–Stokes. The computational resources required for such large-scale simulations were still quite large, so the computations were conducted on the K computer (RIKEN AICS in Japan). Unsteady flow phenomena at the stall inception were analyzed using data mining techniques such as vortex identification and limiting streamline drawing with line integral convolution (LIC) techniques. In the compressor studied, stall started from a separation on the hub side rather than the commonly observed leading-edge separation near the tip. The flow phenomenon first observed in the stalling process is the hub corner separation, which appears in a passage of the sixth stator when approaching the stall point. This hub corner separation grows with time, and eventually leads to a leading-edge separation on the hub side of the stator. Once the leading-edge separation occurs, it rapidly develops into a rotating stall, causing another leading-edge separation of the neighboring blade. Finally, the rotating stall spreads to the upstream and downstream blade rows due to its large blockage effect.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(7):071003-071003-10. doi:10.1115/1.4035567.

This paper presents the work on part-speed fan flutter due to acoustic reflections from the intake, commonly called “flutter bite.” A simple model for the prediction of the flutter bite is presented. In a previous work by the authors, it was shown that the acoustic effects of the intake are very important and need to be considered during the design of a fan blade. It was also shown that the contribution to blade aerodamping due to blade motion (for the isolated rotor in an infinitely long duct) and intake acoustics is independent and can be analyzed separately. The acoustic reflections from the intake change the damping of the blade by modifying the phase and amplitude of the unsteady pressure at the leading edge of the fan. It will be shown in the paper that, for a given intake, the phase and amplitude of the reflected acoustic waves can be evaluated analytically based on established theories independent of the fan design. The proposed model requires only the design intent of the fan blade and the geometry of the intake, which are available in the early design stages of a new engine, and can predict the operating conditions at which fan flutter is likely to occur.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(7):071004-071004-8. doi:10.1115/1.4035569.

Over recent years, engine designs have moved increasingly toward low specific thrust cycles to deliver significant specific fuel consumption (SFC) improvements. Such fan blades may be more prone to aerodynamic and aeroelastic instabilities than conventional fan blades. The aim of this paper is to analyze the flutter stability of a low-speed/low pressure ratio fan blade. By using a validated computational fluid dynamics (CFD) model (AU3D), three-dimensional unsteady simulations are performed for a modern low-speed fan rig for which extensive measured data are available. The computational domain contains a complete fan assembly with an intake duct and the downstream outlet guide vanes (OGVs), which is a whole low-pressure (LP) domain. Flutter simulations are conducted over a range of speeds to understand flutter characteristics of this blade. Only the first flap (1F) mode is considered in this work. Measured rig data obtained by using the same fan set but with two different lengths of the intake showed a significant difference in the flutter boundary for the two intakes. AU3D computations were performed for both intakes and were used to explain this difference between the two intakes, and showed that intake reflections play an important role in flutter of this blade. This observation indicates that the experiment with the long intake used for the performance test may be misleading for flutter. In the next phase of this work, two possible modifications for increasing the flutter margin of the fan blade were explored: changing the mode shape of the blade and using acoustic liners in the casing. The results show that it is possible to increase the flutter margin of the blade by either decreasing the ratio of the twisting to plunging motion in 1F mode or by introducing deep acoustic liners in the intake. The liners have to be deep enough to attenuate the flutter pressure waves and hence influence the stability. The results indicate the importance of reflection in flutter stability of the fan blade and clearly show that intake duct needs to be included in flutter study of any fan blade.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(7):071005-071005-7. doi:10.1115/1.4035604.

The primary contribution of this research is to clarify the aerodynamic performance of a novel lightweight turbine blade with internal cooling passage and external film cooling, which is invented aiming at drastic weight reduction of a cooled blade. With a considerably thinner airfoil, a significant separation region is formed along the pressure side, and therefore, aerodynamic performance with such a flow field should be investigated. First, the lightweight-cooled airfoil is designed. In the design process, a conventional thick airfoil is first defined as a baseline. With the baseline airfoil, only the mid and rear parts of pressure side profile are redesigned to thin the airfoil without any change in the suction side geometry. The airfoil geometry is optimized so as not to bring significant aerodynamic loss increase. In this numerical optimization, the airfoil shape is gradually changed and evaluated step-by-step. In every step, an adjoint variable method is used to seek better airfoil shape, and then, the generated new shape is evaluated with full Reynolds-averaged Navier–Stokes (RANS) calculation. This iteration is repeated until any further recognizable weight reduction cannot be obtained without sensitive pressure loss increase and/or the airfoil shape reaches some geometrical constraints. The resultant optimized airfoil is approximately 20% lighter than the baseline hollow airfoil without any noticeable change in aerodynamic loss in the numerical solution. Next, the optimized airfoil is tested in a high-speed linear cascade rig to verify its aerodynamic performance. The baseline airfoil is also tested for comparison. The rig is composed of six airfoil passages. The compressed air is supplied to the cascade and discharges to the atmospheric exhaust chamber. The air is also heated up to about 540 K upstream of the cascade. The cascade exit Mach number at the design point is 1.25, while in the experiment other several off-design conditions are also tested to check if there is any Mach number sensitivity. At the design point, the optimized lightweight airfoil shows less total pressure loss compared to the baseline airfoil. Also, at any other off-design Mach number conditions tested, the magnitude of the pressure loss is less with the lightweight airfoil. These results verify that the proposed airfoil does not only bring a considerable weight advantage but also compares favorably with the conventional airfoil in aerodynamic performance.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(7):071006-071006-10. doi:10.1115/1.4035605.

An empirical riblet model for manufactured V-shaped and trapezoidal riblets which is suitable for turbomachinery application is presented. The implementation of the riblet effect employs a correlation-based correction for the damping of the specific dissipation rate ω in the vicinity of the wall which has been previously presented by other researchers. In the current paper, the correlations are extended into the drag-increasing regime and are extended to account for the effect of misalignment of the riblets relative to the flow and for the effect of adverse pressure gradients. In order to account for the latter in modern, massive parallel Reynolds-averaged Navier–Stokes (RANS) codes, a local Clauser parameter has been newly derived. The model is implemented in a three-dimensional (3D) turbomachinery design code and validated with flat plate measurement data and a NACA6510 compressor cascade. The predictions of the experimental values are in very good agreement with the experimental data, showing the capability of the model for designing riblet structured turbomachinery blading.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(7):071007-071007-12. doi:10.1115/1.4035570.

Airfoil trailing-edge cooling is the main focus of this study. The test section was made up of two adjacent trapezoidal channels, simulating the trailing-edge cooling cavity of a gas turbine airfoil and its neighboring cavity. Eleven racetrack-shaped holes were drilled on the partition wall between the two channels to produce 11 cross-over jets that impinged on the rib-roughened wall of the trailing-edge channel. The jets, after impinging on their respective target surface, turned toward the trailing-edge channel exit. Smooth target wall, as a baseline case, as well as four rib angles with the flow of 0 deg, 45 deg, 90 deg, and 135 deg are investigated. Cross-over holes axes were on the trailing-edge channel center plane, i.e., no tilting of the cross-over jets. Steady-state liquid crystal thermography technique was used in this study for a range of jet Reynolds number of 10,000–35,000. The test results are compared with the numerical results obtained from the Reynolds-averaged Navier–Stokes and energy equation. Closure was attained by k–ω with shear stress transport (SST) turbulence model. The entire test rig (supply and trailing-edge channels) was meshed with variable density hexagonal meshes. The numerical work was performed for boundary conditions identical to those of the tests. In addition to the impingement heat transfer coefficients, the numerical results provided the mass flow rates through individual cross-over holes. This study concluded that: (a) the local Nusselt numbers correlate well with the local jet Reynolds numbers, (b) 90 deg rib arrangement, that is, when the cross-over jet axis was parallel to the rib longitudinal axis, produced higher heat transfer coefficients, compared to other rib angles, and (c) numerical heat transfer results were generally in good agreement with the test results. The overall difference between the computational fluid dynamics (CFD) and test results was about 10%.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(7):071008-071008-11. doi:10.1115/1.4035508.

Since the experiment in turbulence research is of very high importance for evaluating turbulence hypothesis, turbulence measurements were carried out in a two-stage two-spool transonic turbine test rig at the Institute for Thermal Turbomachinery and Machine Dynamics in Graz in which the two rotors are counter-rotating with two different rotational speeds. For the current measurement campaign, triple hot-wire probes, which represent a very new measurement technique in this test rig, were used and their results validated with a fast response aerodynamic pressure probe (FRAPP). Based on the data measured with this device, turbulence intensities may be determined using a method called Fourier filtering. If the classical ensemble averaging procedure with only one trigger is applied, the periodic fluctuations of the other rotor will artificially increase the stochastic fluctuations. Therefore, the two trigger signals of the two rotors require a special analysis method, which was established at Graz University of Technology. The results from this method will be compared to the classical triple decomposition, which uses only one trigger signal. With this analysis tool, it is not only possible to evaluate unsteady signals triggered by one of the two rotors, but also the unsteady interactions of the rotors can be determined and investigated.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(7):071009-071009-8. doi:10.1115/1.4035656.

This paper aims for the analysis of experimental and numerical results of windmilling flow topologies far from freewheeling condition. Two major cooling 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. Also, tests were performed on a turbofan engine to confirm some trends observed on the cooling fans. 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 nonrotating elements can be considered as loss generators.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(7):071010-071010-11. doi:10.1115/1.4035510.

This paper presents a novel optimization methodology based on both adjoint sensitivity analysis and trust-based dynamic response surface modeling to improve the performance of a modern turbine of a large civil aero-engine in the presence of high-fidelity geometry configurations. The system has been applied to the nonaxisymmetric hub and tip endwall optimization of a high-pressure turbine stage making use of multirow 3D simulations, parametric modeling, and rapid meshing of real geometry features such as rim seals and modeling of film cooling flows. It has been shown in previous papers that improvements gained using simplified models of the stage are lost when applying the high-fidelity geometry configuration. New results presented in this paper indicate that controlling the purge flow that exits the disk space through the rim seal at the hub of the main annulus is more significant than the reduction of secondary flows in the main passage. For a given rim sealing mass flow rate and whirl velocity, the nonaxisymmetric endwalls are optimized such that the detrimental impact of the sealing flow on the turbine performance is reduced, and hence, the stage efficiency is significantly increased. The traditional optimization approaches based on evolutionary methods or even sequential modifications for defining the endwalls shape are computationally demanding. Since turbomachinery industry continuously strive to reduce the design cycle time, in particular when high-fidelity 3D computational fluid dynamics (CFD) is used, the main body of this paper outlines the novel methods developed to produce a practical design in a very aggressively short design cycle time.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(7):071011-071011-9. doi:10.1115/1.4035606.

In aero-engine applications, centrifugal compressors are often close-coupled with their respective diffusers to increase efficiency at the expense of a reduced operating range. The aim of this paper is to show that state-of-the art steady-state computational fluid dynamics (CFD) simulations can model a hubside cavity between an impeller and a close-coupled diffuser and to enhance the understanding of how the cavity affects performance. The investigated cavity is located at the impeller trailing edge, and bleed air is extracted through it. Due to geometrical limitations, the mixing plane is located in the cavity region. Therefore, the previous analyses used only a cut (“simple”) model of the cavity. With the new, “full” cavity model, the region inside the cavity right after the impeller trailing edge is not neglected anymore. The numerical setup is validated using the experimental data gathered on a state-of-the art centrifugal compressor test-rig. For the total pressure field in front of the diffuser throat, a clear improvement is achieved. The results presented reveal a drop in stage efficiency by 0.5%-points caused by a new loss mechanism at the impeller trailing edge. On the hubside, the fundamentally different interaction of the cavity with the coreflow increases the losses in the downstream components resulting in the mentioned stage efficiency drop. Finally, varying bleed air extraction is investigated with both cavity models. Only the full cavity (FC) model captures the changes measured in the experiment.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(7):071012-071012-10. doi:10.1115/1.4035631.

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 nonaxisymmetric throughflow analysis has been developed to define fan flow with inlet distortion. The turbomachinery is modeled using momentum and energy source distributions that are determined as a function of local flow conditions and specified blade camber surface geometry. Comparison with higher-fidelity computational and experimental results shows the analysis captures the principal flow redistribution and distortion transfer effects associated with BLI. Distortion response is assessed for a range of (i) design flow and stagnation enthalpy rise coefficients, (ii) rotor spanwise work profiles, (iii) rotor–stator spacings, and (iv) nonaxisymmetric stator geometries. Of the approaches examined, nonaxisymmetric stator geometry and increased stage flow and stagnation enthalpy rise coefficients provide the greatest reductions in rotor flow nonuniformity, and may offer the most potential for mitigating performance loss due to BLI inlet distortion.

Commentary by Dr. Valentin Fuster

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