We report on a series of highly resolved large-eddy simulations of the LS89 high-pressure turbine (HPT) vane, varying the exit Mach number between Ma=0.7 and 1.1. In order to accurately resolve the blade boundary layers and enforce pitchwise periodicity, we for the first time use an overset mesh method, which consists of an O-type grid around the blade overlapping with a background H-type grid. The simulations were conducted either with a synthetic inlet turbulence condition or including upstream bars. A quantitative comparison shows that the computationally more efficient synthetic method is able to reproduce the turbulence characterictics of the upstream bars. We further perform a detailed analysis of the flow fields, showing that the varying exit Mach number significantly changes the turbine efficiency by affecting the suction-side transition, blade boundary layer profiles, and wake mixing. In particular, the Ma=1.1 case includes a strong shock that interacts with the trailing edge, causing an increased complexity of the flow field. We use our recently developed entropy loss analysis (Zhao and Sandberg, GT2019-90126) to decompose the overall loss into different source terms and identify the regions that dominate the loss generation. Comparing the different Ma cases, we conclude that the main mechanism for the extra loss generation in the Ma=1.1 case is the shock-related strong pressure gradient interacting with the turbulent boundary layer and the wake, resulting in significant turbulence production and extensive viscous dissipation.

Gas turbines have been widely used. With the continuous improvement of the performance of gas turbines, the turbine inlet temperature has greatly exceeded the heat resistance limit of the turbine blade material, so advanced cooling technology is required. The film cooling effectiveness distribution over the blade under the effect of wake was obtained by Pressure Sensitive Paint (PSP) technique. The test blade has 5 rows of chevron film holes on the pressure side, 3 rows of cylindrical film holes on the leading edge and 3 rows of chevron film holes on the suction side. The mainstream Reynolds number is 130,000 based on the blade chord length, and the mainstream turbulence intensity is 2.7%. The upstream wake was simulated by the spoken-wheel type wake generator. The film cooling effectiveness was measured at three wake Strouhal numbers (0, 0.12 and 0.36) and three mass flux ratios (MFR1, MFR2 and MFR3). The results show that the increase of mass flux ratio leads a decrease of the film cooling effectiveness on the suction surface. In the wake condition, the effect of mass flux ratio is weakened. Wake leads a marked decrease of the film cooling effectiveness over most blade surface except for the surface near leading edge on the pressure surface. In the high mass flux ratio condition, the effect of wake on the film cooling effectiveness is weakened on the suction surface and strengthened on the pressure surface.

The flow field in a compressor is circumferentially non-uniform due to the wakes from upstream stators, the potential field from both upstream and downstream stators, and blade row interactions. This non-uniform flow impacts stage performance as well as blade forced vibrations. Historically, experimental characterization of the circumferential flow variation is achieved by circumferentially traversing either a probe or the stator rows. This involves the design of complex traverse mechanisms and can be costly. To address this challenge, a novel method is proposed to reconstruct compressor nonuniform circumferential flow field using spatially under-sampled data points from a few probes at fixed circumferential locations. The paper is organized into two parts. In the present part of the paper, details of the multi-wavelet approximation for the reconstruction of circumferential flow and use of the Particle Swarm Optimization algorithm for selection of probe positions are presented. Validation of the method is performed using the total pressure field in a multi-stage compressor representative of small core compressors in aero engines. The circumferential total pressure field is reconstructed from 8 spatially distributed data points using a triple-wavelet approximation method. Results show good agreement between the reconstructed and the true total pressure fields. Also, a sensitivity analysis of the method is conducted to investigate the influence of probe spacing on the errors in the reconstructed signal.

Understanding basic aerodynamic and thermodynamic processes in engine components is critical to achieving higher efficiencies and lower fuel consumption in aircraft engines. To aid in this process, a linear compressor cascade was investigated in the high-speed cascade wind tunnel of the Institute of Jet Propulsion to quantify the influence of heat transfer on the temperature distribution in the wake and, finally, the profile loss. For this purpose, a patented five-hole probe with an integrated thermocouple was developed and applied for steady measurements. Additionally, a hot-wire measurement set-up was implemented to receive temperature fluctuations via the constant current mode as well as velocity fluctuations via the constant temperature mode. A novel method for a two-way temperature and velocity correction for the two types of hot-wire measurement is presented. Good agreement between the measurement data of the five-hole probe and averaged data from hot-wire anemometry was found. The temperature distribution indicates the occurrence of energy separation which in some cases is overlain with the effects of heat transfer. In addition, the analysis of unsteady fluctuations of temperature and velocity give more detailed information about the vortex shedding in the wake, including the size of the vortices. Finally, this is the first discussion of energy separation at a compressor cascade combined with overlain effects of heat transfer on the blade surface.

This paper presents the simultaneous application of fast-response pressure transducers and unsteady pressure-sensitive paint (unsteady PSP) for the precise determination of pressure amplitudes and phases up to 3000 Hz. These experiments have been carried out on a low-pressure turbine blade cascade under engine-relevant conditions (Re, Ma, and Tu) in the high-speed cascade wind tunnel. Periodic blade/vane interactions were simulated at the inlet to the cascade using a wake generator operating at a constant perturbation frequency of 500 Hz. The main goal of this paper is the detailed comparison of amplitude and phase distributions between both flow sensing techniques at least up to the second harmonic of the wake generator’s fundamental perturbation frequency (i.e., 1000 Hz). Therefore, a careful assessment of the key drivers for relative deviations between measurement results as well as a detailed discussion of the data processing is presented for both measurement techniques. This discussion outlines the mandatory steps which were essential to achieve the quality as presented down to pressure amplitudes of several Pascal even under challenging experimental conditions. Apart from the remarkable consistency of the results, this paper reveals the potential of (unsteady) PSP as a future key flow sensing technique in turbomachinery research, especially for cascade testing. The results demonstrate that PSP was able to successfully sense pressure dynamics with very low fluctuation amplitudes down to 8 Pa.

A particular turbine cascade design is presented with the goal of providing a basis for high quality investigations of endwall flow under high-speed conditions with unsteady inflow. The key feature of the design is an integrated two-part flat plate serving as a cascade endwall at part-span, which enables a variation of the inlet endwall boundary layer conditions. The new design is applied to the T106A low pressure turbine cascade for endwall flow investigations in the High-Speed Cascade Wind Tunnel of the Institute of Jet Propulsion at the Bundeswehr University Munich. Measurements are conducted under realistic flow conditions (M 2th = 0.59, Re 2th = 2 · 10 5 ) in three cases of varying endwall boundary layer conditions with and without periodically incoming wakes. The endwall boundary layer is characterized by 1D-CTA measurements upstream of the blade passage. Secondary flow is evaluated by five-hole-probe measurements in the turbine exit flow. A strong similarity is found between the time-averaged effects of unsteady inflow conditions and the effects of changing inlet endwall boundary layer conditions regarding the attenuation of secondary flow. Furthermore, the experimental investigations show that all design goals for the improved T106A cascade are met.

Arrays of staggered pin fins are a typical geometry found in the trailing edge region of modern airfoils. If coolant is supplied by bleeding from the mid-section of the airfoil instead of provided through the root, the channel length is insufficiently long to reach a fully developed flow which is commonly found from the fifth row downstream. This present study focuses on the developing section (four rows) of a staggered array with a height-to-diameter ratio of 2 and a spanwise and streamwise spacing of 2.5, respectively. Measurements are conducted at Reynolds numbers of 10,000 and 30,000 based on the maximum velocity and pin diameter. Stereoscopic particle image velocimetry (PIV) is used to describe the flow field and turbulence characteristics in the wake of the first and third row pin. It is found that the dominating vortical structures depend highly on the Reynolds number. A transient thermochromic liquid crystal (TLC) technique is used to obtain local heat transfer coefficients on the endwall which are then discussed in the context with the vortical structures. The structure of the horseshoe vortex and the transient wake shedding behaves differently in the first and third row. The interaction of both vortex systems affects directly the endwall heat transfer. The results are supplemented by a thorough discussion of TLC and PIV uncertainty.

This article compares experimental and numerical data for a low-speed high-lift low pressure turbine (LPT) cascade under unsteady flow conditions. Three Reynolds numbers representative of LPTs have been tested, namely, 5 × 10 4 , 10 5 , and 2 × 10 5 ; at two reduced frequencies, f r = 0.5 and 1, also representative of LPTs. The experimental data were obtained at the low-speed linear cascade wind tunnel at the Polytechnic University of Madrid using hot wire, Laser Doppler Velocimetry (LDV), and pressure tappings. The numerical solver employs a sixth-order compact scheme based on the flux reconstruction method for spatial discretization and a fourth-order Runge–Kutta method to march in time. The longest case ran 550 h on 40 GPUs to reach a statistically periodic state. Pressure coefficients around the profile, boundary layer profiles and exit cross section distributions of velocity, pressure loss defect, shear Reynolds stress, and angle are compared against high-quality experimental data. Cascade loss and exit angle have also been compared against the experimental data. Very good agreement between experimental and numerical data is seen. The results demonstrate the suitability of the present methodology to predict the aerodynamic properties of unsteady flows around LPT linear cascades accurately.

In this paper, we study the effect of rotor-stator axial gap on midspan compressor loss using high-fidelity scale-resolving simulations. For this purpose, we mimic the multi-stage environment using a new numerical method that recycles wake unsteadiness from a single blade passage back into the inlet of the computational domain. As a result, a type of repeating-passage simulation is obtained such as observed by an embedded blade-row. We find that freestream turbulence levels rise significantly as the size of the rotor-stator axial gap is reduced. This is because of the effect of axial gap on turbulence production, which becomes amplified at smaller axial gaps and drives increases in dissipation and loss. This effect is found to raise loss by between 5.5% and 9.5% over the range of conditions tested here. This effect significantly outweighs the beneficial effects of wake recovery on loss.

This paper presents the large Eddy simulation (LES) of a propeller representative of the first rotor of a counter rotative open rotor (CROR) configuration based on a multiple frequency phase-lagged approach in conjunction with a proper orthogonal decomposition (POD) data storage. This method enables to perform unsteady simulations on multistage turbomachinery configurations including multiple frequency flows with a reduction of the computational domain composed of one single blade passage for each row. This approach is advantageous when no circumferential periodicity occurs in the blade rows of the configuration and a full 360 deg simulation would be required. The data storage method is based on a POD decomposition replacing the traditional Fourier series decomposition (FSD). The inherent limitation of phase-shifted periodicity assumption remains with POD data storage but this compression method alleviates some issues associated with the Fourier transform, especially spectrum issues. The paper is first dedicated to compare the flow field obtained with the LES with phase-lagged condition against full-matching URANS, LES simulations, and experimental data available around the blade and in the wake of the rotor. The study shows a close agreement of the phase-lagged LES simulation with other simulations performed and a thicker wake compared with the experiments with lower turbulent activity. The analysis of the losses generated in the configuration, based on an entropy formulation and a splitting between boundary layer and secondary flow structures, shows the strong contribution of the blade boundary layer in the losses generated.

Loss analysis is a valuable technique for improving the thermodynamic performance of turbomachines. Analyzing loss in terms of the “mechanical work potential” (Miller, R.J., ASME Turbo Expo 2013 , GT2013-95488) provides an instantaneous and local account of the thermal and aerodynamic mechanisms contributing to the loss of thermodynamic performance. This study develops the practical application of mechanical work potential loss analysis, providing the mathematical formulations necessary to perform loss analysis using practical Reynolds-averaged Navier–Stokes (RANS) or large eddy simulations (LES). The analysis approach is demonstrated using RANS and LES of a linear compressor cascade, both with and without incoming wakes. Spatial segmentation is used to attribute loss contributions to specific regions of the flow, and phase-averaging is performed in order to associate the variation of different loss contributions with the periodic passage of wakes through the cascade. For this un-cooled linear cascade, viscous dissipation is the dominant source of loss. The analysis shows that the contribution of the viscous reheat effect depends on the operating pressure of the compressor stage relative to the ambient “dead state” pressure—implying that the optimal blade profile for a low-pressure compressor stage may be different from the optimal profile for a high-pressure compressor stage in the same engine, even if the operating conditions for both stages are dynamically similar.

In this paper, the effects of an array of herringbone riblets with different riblet geometry (height and spacing) and start locations on the pressure losses in a cascade of diffuser blades are investigated over a range of low Reynolds numbers (0.50 × 10 5 –1.00 × 10 5 ). The herringbone riblets with a given geometry are found to produce a profound modification to the wake structure above certain critical Reynolds numbers. It is also found that within the range of parameters tested an increase in riblet height and riblet spacing results in an onset of significant control effect at a lower Reynolds number, which is accompanied by a slight reduction in zone-averaged loss coefficient and flow turning angle. An upstream shift of the start position of the riblet array along the blades enables the riblets to become effective at a lower Reynolds number at the expense of a reduced loss reduction and flow turning angle. A semi-empirical relationship between the ratio of riblet height to local baseline boundary layer displacement thickness and the critical Reynolds number is established using the present experimental data. A preliminary methodology for designing the herringbone riblets to ensure an effective control of 2D flow separations around the mid-span of diffuser blades over a specified range of Reynolds numbers is also proposed.

The mechanisms of blade row interaction affecting rotor film cooling are identified to make recommendations for the design of film cooling in the real, unsteady turbine environment. Present design practice makes the simplifying assumption of steady boundary conditions despite intrinsic unsteadiness due to blade row interaction; we argue that if film cooling responds nonlinearly to unsteadiness, the time-averaged performance will then be in error. Nonlinear behavior is confirmed using experimental measurements of flat-plate cylindrical film cooling holes, mainstream unsteadiness causing a reduction in film effectiveness of up to 31% at constant time-averaged boundary condition. Unsteady computations are used to identify the blade row interaction mechanisms in a high-pressure turbine rotor: a “negative jet” associated with the upstream vane wake, and frozen and propagating vane potential field interactions. A quasi-steady model is used to predict unsteady excursions in momentum flux ratio of rotor cooling holes, with fluctuations of at least ±30% observed for all hole locations. Computations with modified upstream vanes are used to vary the relative strength of wake and potential field interactions. In general, both mechanisms contribute to rotor film cooling unsteadiness. It is recommended that the designer should choose a cooling configuration that behaves linearly over the expected unsteady excursions in momentum flux ratio as predicted by a quasi-steady hole model.

The effects of blade row interactions on stator-mounted instrumentation in axial compressors are investigated using unsteady numerical calculations. The test compressor is an eight-stage machine representative of an aero-engine core compressor. For the unsteady calculations, a 180-deg sector (half-annulus) model of the compressor is used. It is shown that the time-mean flow field in the stator leading edge planes is circumferentially nonuniform. The circumferential variations in stagnation pressure and stagnation temperature, respectively, reach 4.2% and 1.1% of the local mean. Using spatial wave number analysis, the incoming wakes from the upstream stator rows are identified as the dominant source of the circumferential variations in the front and middle of the compressor, while toward the rear of the compressor, the upstream influence of the eight struts in the exit duct becomes dominant. Based on three circumferential probes, the sampling errors for stagnation pressure and stagnation temperature are calculated as a function of the probe locations. Optimization of the probe locations shows that the sampling error can be reduced by up to 77% by circumferentially redistributing the individual probes. The reductions in the sampling errors translate to reductions in the uncertainties of the overall compressor efficiency and inlet flow capacity by up to 50%. Recognizing that data from large-scale unsteady calculations are rarely available in the instrumentation phase for a new test rig or engine, a method for approximating the circumferential variations with single harmonics is presented. The construction of the harmonics is based solely on the knowledge of the number of stators in each row and a small number of equispaced probes. It is shown how excursions in the sampling error are reduced by increasing the number of circumferential probes.

Compressor corner stall is a phenomenon difficult to predict with numerical tools but essential to the design of axial compressors. Predictive methods are beneficial early in the design process to understand design and off-design limitations. Prior numerical work using Navier–Stokes computational methods has assessed the prediction capability for corner stall. Reynolds-averaged Navier–Stokes (RANS) simulations using several turbulence models have shown to over-predict the region of corner hub stall where large eddy simulations (LES) and detached eddy simulations (DES) approaches improved the airfoil surface and wake pressure loss prediction. A linear compressor cascade designed and tested at Ecole Centrale de Lyon provides a good benchmark for the evaluation of the accuracy of numerical methods for corner stall. This paper presents results obtained with Lattice-Boltzmann method (LBM) coupled with very large-eddy simulations (VLES) approach of PowerFLOW and compares them with the experimental and numerical work from Ecole Centrale de Lyon. The ability to achieve equivalent accuracy at a lower computational cost compared to LES scale resolving methods can enable multi-stage design and off-design compressor predictions. A methodical approach is taken by first accurately simulating the upstream flow conditions. Geometric trips are modeled upstream on the endwalls to match both the mean and fluctuating inflow boundary layer conditions. These conditions were then applied to the simulation of the linear compressor cascade. The benchline experimental study includes trips on both the pressure and suction of the airfoil. These trips are also included for the current simulation. The significance of capturing both inflow conditions and including trips on the airfoil is assessed. Detailed comparisons are then made to airfoil loading and downstream losses between experiment and previous RANS and LES simulations. LBM-VLES corner stall results of pitchwise averaged total pressure match LES agreement relative to experimental data at 50 times lower computational cost.

The development and verification of new turbulence models for Reynolds-averaged Navier–Stokes (RANS) equation-based numerical methods require reliable experimental data with a deep understanding of the underlying turbulence mechanisms. High accurate turbulence measurements are normally limited to simplified test cases under optimal experimental conditions. This work presents comprehensive three-dimensional data of turbulent flow quantities, comparing advanced constant temperature anemometry (CTA) and stereoscopic particle image velocimetry (PIV) methods under realistic test conditions. The experiments are conducted downstream of a linear, low-pressure turbine cascade at engine relevant high-speed operating conditions. The special combination of high subsonic Mach and low Reynolds number results in a low density test environment, challenging for all applied measurement techniques. Detailed discussions about influences affecting the measured result for each specific measuring technique are given. The presented time mean fields as well as total turbulence data demonstrate with an average deviation of Δ T u < 0.4 % and Δ C / C ref < 0.9 % an extraordinary good agreement between the results from the triple sensor hot-wire probe and the 2D3C-PIV setup. Most differences between PIV and CTA can be explained by the finite probe size and individual geometry.

Synthetic jets are produced by devices that enable a suction phase followed by an ejection phase. The resulting mean mass budget is hence null and no addition of mass in the system is required. These particular jets have especially been considered for some years for flow control applications. They also display features that can become of interest to enhance heat exchanges, for example, for wall cooling issues. Synthetic jets can be generated through different mechanisms, such as acoustics by making use of a Helmholtz resonator or through the motion of a piston as in an experience mounted at Institut Pprime in France. The objective of this specific experiment is to understand how synthetic jets can enhance heat transfer in a multi-perforated configuration. As a complement to this experimental setup, large-eddy simulations are produced and analyzed in the present document to investigate the flow behavior as well as the impact of the synthetic jets on wall heat transfer. The experimental system considered here consists in a perforated heated plate, each perforation being above a cavity where a piston is used to control the synthetic jets. Placed in a wind tunnel test section, the device can be studied with a grazing flow and multiple operating points are available. The one considered here implies a grazing flow velocity of 12.8 m s −1 , corresponding to a Mach number around 0.04, and a piston displacement of 22 mm peak-to-peak at a frequency of 12.8 Hz. These two latter parameters lead to a jet Reynolds number of about 830. A good agreement is found between numerical results and experimental data. The simulations are then used to provide a detailed understanding of the flow. Two main behaviors are found, depending on the considered mid-period. During the ejection phase, the flow transitions to turbulence and the formation of characteristic structures are observed; the plate is efficiently cooled. During the suction phase, the main flow is stabilized; the heat enhancement is particularly efficient in the hole wakes but not between them, leading to a heterogeneous temperature field.

The boundary layer developing over the suction side of a low-pressure turbine cascade operating under unsteady inflow conditions has been experimentally investigated. Time-resolved particle image velocimetry (PIV) measurements have been performed in two orthogonal planes, the blade-to-blade and a wall-parallel plane embedded within the boundary layer, for two different wake-reduced frequencies. Proper orthogonal decomposition (POD) has been used to analyze the data and to provide an interpretation of the most significant flow structures for each phase of the wake passing cycle. Detailed information on the most energetic turbulent structures at a particular phase is obtained with a newly developed procedure that overcomes the limit of classical phase average. The synchronization of the measurements in the two planes allows the computation of the characteristic dimension of boundary layer streaky structures that are responsible for transition. The largest and most energetic structures are observed when the wake centerline passes over the rear part of the suction side, and they appear practically the same for both reduced frequencies. The passing wake forces transition leading to the breakdown of the boundary layer streaks. Otherwise, the largest differences between the low and high reduced frequency are observed in the calmed region. The postprocessing of these two planes allowed computing the spacing of the streaky structures and making it nondimensional by the boundary layer displacement thickness observed for each phase. The nondimensional value of the streaks spacing is about constant, irrespective of the reduced frequency.

Modern turbomachinery faces increased performance demands in terms of efficiency, compactness, and pressure-rise. Advancements in computational technology have allowed numerical methods to become the backbone of design development efforts. However, the unique complexities of centrifugal compressor flow-fields pose difficult computational problems. As such, advanced experimental methods must be used to obtain high-quality data sets to further inform, improve, and validate computational methods in complex flow regimes. A recent experimental work on a high-speed centrifugal compressor has provided detailed, unsteady, three-component velocity data using laser Doppler velocimetry. A passage vortex is present, and its nascent tied to the increased incidence at mid-span associated with impeller wake flow. This vortex begins in the hub-pressure side corner and grows to fill the passage and become temporally stable. The vortex development is unsteady in nature, and the unsteady effects persist 40% downstream of the throat. Distinct jet and wake flow patterns from the impeller also do not agglomerate until 40% downstream of the throat. Additionally, the critical impact of the unsteady flow development on the time-averaged flow-field is explained.

A novel non-intrusive method has been developed to monitor rotor blade vibration using unsteady casing pressure. The present blade vibration monitoring technique utilizes casing unsteady pressure sensors that can detect the pressure waves associated with blade vibration. Spinning mode theory was used to identify the specific frequencies and nodal diameters (NDs) of the spinning pressure waves associated with the blade vibration. A dual temporal-spatial analysis method has been developed to extract the specific frequency components using Fourier transforms, and the specific ND component was extracted using a circumferential mode-fitting algorithm. An experimental study was done in the Purdue 3-stage axial research compressor to verify the new rotor blade vibration monitoring method against the blade tip timing (BTT) method. During the experiment, the compressor was swept through the resonant crossing speed corresponding to the first torsion (1T) vibratory mode of the embedded rotor, while the unsteady casing pressure data and BTT data were simultaneously acquired. Utilizing as few as two sensors, the pressure wave due to blade forced vibration was extracted. A constant scaling factor between the resultant pressure wave strength and blade deflection amplitude was calculated for two different loading conditions. The close match between blade vibration-generated pressure wave strength and blade deflection amplitude through the resonant range provides the validation for the new rotor blade vibration monitoring method. This is the first time in the open literature that blade vibration-related pressure waves have been extracted from casing pressure sensor arrays and used to quantify blade vibration.