The Reynolds number effects on performance of a Controlled Diffusion Airfoil (CDA) geometry in a linear compressor cascade setup are presented. A set of highly-resolved Large Eddy Simulations at Reynolds numbers from 300,000 to 1,000,000 at a Mach number of 0.7 are performed, with one of the simulations including incident wakes at engine relevant reduced frequency of 1.56. The focus of this work is to investigate loss variation of a modern blade geometry by applying both the novel entropy analysis introduced by (Zhao & Sandberg, GT2019-90126) and the mechanical work potential analysis by Miller (GT2013-95488) in the form presented by Leggett et al. (GT2019-91253). The loss analysis, applied both to the whole blade, and to carefully selected subvolumes, including suction and pressure surface boundary layers, highlights regions of increased loss production and the effects of Reynolds number on the location and strength of different flow processes, such as transition and separation. The results are further broken down to show the contribution to loss of individual terms of the analyses and the changes that occur due to changes in Reynolds number. The effects of incident wakes are also captured using this analysis technique and detailed in the paper.

We present the first wall-resolved high-fidelity simulations of high-pressure turbine (HPT) stages at engine-relevant conditions. A series of cases have been performed to investigate the effects of varying Reynolds numbers and inlet turbulence on the aerothermal behavior of the stage. While all of the cases have similar mean pressure distribution, the cases with higher Reynolds number show larger amplitude wall shear stress and enhanced heat fluxes around the vane and rotor blades. Moreover, higher-amplitude turbulence fluctuations at the inlet enhance heat transfer on the pressure-side and induce early transition on the suction-side of the vane, although the rotor blade boundary layers are not significantly affected. In addition to the time-averaged results, phase-lock averaged statistics are also collected to characterize the evolution of the stator wakes in the rotor passages. It is shown that the stretching and deformation of the stator wakes is dominated by the mean flow shear, and their interactions with the rotor blades can significantly intensify the heat transfer on the suction side. For the first time, the recently proposed entropy analysis has been applied to phase-lock averaged flow fields, which enables a quantitative characterization of the different mechanisms responsible for the unsteady losses of the stages. The results indicate that the losses related to the evolution of the stator wakes is mainly caused by the turbulence production, i.e. the direct interaction between the wake fluctuations and the mean flow shear through the rotor passages.

The paper presents the results of aeromechanical design of a large-scale model stage for a high-efficient low-noise fan designed for an advanced civil geared turbofan engines with ultra-low rotational speeds of rotor blades (313.4 m/s), high flow specific capacity (up to 202 kg/m 2 /s) and high bypass ratio (13.5). Total pressure ratio in the bypass duct of the fan model stage is 1.38. To ensure the experimental studies, characteristics are calculated from choking to a surge line within a wide range of rotational speeds. For the studies of the experimental fan model (EFM), a design project is developed and used in manufacturing a fan stage with 0.7-m rotor diameter for tests at the C-3A acoustic test facility. The manufacturing technology for blades made of polymer composite materials (PCM) is of particular importance. Rotor blades of the geared fan model are made of PCM. The analysis of experimental data and their comparison with the computation results within the range of corrected rotational speeds from 0.325 to 1.0 are presented. At first, only gas-dynamic and strength characteristics of the stage are studied. The analysis shows a good agreement of calculated integral parameters with the experimental data. Acoustic performances of the EFM will be studied later on.

A numerical experiment of intentionally reduced complexity is used to demonstrate a method to classify flight missions in terms of the operational severity experienced by the engines. In this proof of concept, the general term of severity is limited to the erosion of the core flow compressor blade and vane leading edges. A Monte Carlo simulation of varying operational conditions generates a required database of 10000 flight missions. Each flight is sampled at a rate of 1 Hz. Eleven measurable or synthesizable physical parameters are deemed to be relevant for the problem. They are reduced to seven universal non-dimensional groups which are averaged for each flight. The application of principal component analysis allows a further reduction to three principal components. They are used to run a support-vector machine model in order to classify the flights. A linear kernel function is chosen for the support-vector machine due to its low computation time compared to other functions. The robustness of the classification approach against measurement precision error is evaluated. In addition, a minimum number of flights required for training and a sensible number of severity classes are documented. Furthermore, the importance to train the algorithms on a sufficiently wide range of operations is presented.

The clearance that exists between the casing and turbine blade tips is one of the key drivers of efficiency in gas turbine engines. For this reason, engine manufacturers utilize precise manufacturing techniques and may employ clearance control systems to minimize tip clearances to reduce associated losses. Despite these efforts, turbines typically exhibit some nominal casing ovality or rotor-casing eccentricity, and changes to blade tip clearance during operation commonly occur due to thermal and mechanical stresses. The present study investigates non-axisymmetric tip clearance effects by creating a rotor-casing eccentricity in a one-stage axial test turbine operating in a continuous-duration mode at engine relevant conditions with engine representative hardware. A magnetic levitation bearing system was leveraged to move the turbine shaft to vary the rotor-casing eccentricity without test section disassembly. The results of this study indicate that rotor-casing eccentricity does not affect overall turbine efficiency over the range that was tested, but does locally influence efficiency and the rotor exit flow field. Comparisons of flow angle and secondary flow kinetic energy agreed with previous studies and existing analytical methods, respectively. Collectively, these results indicate that tip clearance can be studied locally on an eccentric rotor.

This paper presents experimental and numerical CFD studies of the aerodynamics of a turbine rear structure (TRS). The TRS test geometry is an engine-realistic state-of-the-art design with a polygonal outer case, recessed engine mount bumps, and three different vane types: regular vanes, bump vanes in bump sectors, and thick vanes. Using three different sector types simultaneously was found to be crucial for the inlet boundary conditions. Experiments were performed in a modern rotating test facility with an LPT stage upstream of the TRS. A Reynolds number of 350,000 was used, representative of a TRS in a narrow-body geared turbofan engine. The TRS performance was analyzed both at on- and off-design conditions and a thorough side-by-side comparison of CFD and experiments was performed. Static-pressure-distributions, turning and outlet flow-angles, wakes and losses, and surface-flow visualizations and outlet total pressure contours are presented. The thick vane showed good aerodynamic performance, similar to the regular vane. For the bump vane, the mount bumps were found to generate additional local separations and secondary flows, resulting in extra losses. In the regions with strong secondary flows CFD over-predicts the wakes, whereas the wakes around midspan, where secondary flows have a smaller influence, are predicted well.

Turbocharging has become a fundamental technology to realize engine downsizing, which is an attractive strategy for low carbon vehicles in the near term. The stable operation of turbocharger compressors at low and high flow rates is crucial to provide peak torque demand and rated power for turbocharged automotive engines. The scroll or volute is a key component in centrifugal compressors as its design not only impacts the compressor efficiency but also affects the operating range. This component causes a distorted pressure field upstream which can contribute to stall on the impeller, inducing surge. As the flow inside the volute is fully three dimensional and turbulent, a better understanding of flow mechanisms is key to enable a volute design methodology. In this study, a centrifugal compressor stage has been modelled numerically and validated by experimental results, to identify the geometric parameters of the volute which contribute to the main flow losses. By solving Reynolds average Navier-Stokes (RANS) equations using a commercial code, the threedimensional flow field of the compressor was modelled. Based on detailed analysis of this flow field, and the impact of various geometric parameters, an optimized volute was developed. The results showed that the total-to-total isentropic efficiency and surge margin could be improved by 1.5% and 4.5%, respectively at design speed.

Engine operating cost contributes a major share to an aircraft’s direct operating cost. Thus, the knowledge of the current and future state of their engines is a major concern to any airline operator. To be able to schedule shop visits, state-of-the-art diagnostic and prognostic tools including CFD methods are employed. These RANS-based turbulence and transition models are used to predict the overall efficiency and operational behavior of the engine components. Aerofoil surfaces undergo dynamic change during operation and surface roughness increases in complex non-homogeneous ways due to corrosion, erosion, and fouling processes; depending on the engine component and the environmental condition encountered. The influence of real fouling based roughness on the boundary layer transition is investigated experimentally and numerically within this study. For this purpose, the rotor midspan from the second HPC rotor of the CFM56 is used as the basis for experimental and numerical investigations. Realistic fouling based roughness is applied and investigated both in a cascade tunnel and a low speed compressor rig. The results shown here indicate that laminar boundary layers and their transition to turbulence must be included in the RANS model combination used. Furthermore, it is necessary to consider roughness effects in the respective turbulence and transition model. While the consideration of roughness for the turbulence models has already found wide acceptance, the results in this work motivate the additional extension of the transition model to include roughness effects.

Future aircraft are expected to have a high level of fuselage-engine integration, exposing the propulsor to non-uniform inflow conditions, in which the performance and stability of the engine fan can be severly affected. This paper proposes a study of a transonic fan subject to inlet distortion by employing steady RANS simulations and mean line calculations. The steady CFD results, although insufficient to correctly measure the interaction of the distortion across the cascade passages, are used as a benchmark to compare the prediction of a leaner low-order tool, adopting a superposition of several 1D calculations in the discretised compressor approach. A continuous total pressure deficit is imposed at the inlet and the distorted operating points are compared to the clean inflow case. The performance drop at the peak efficiency point is similar for the two models, altough the mean-line 1D method fails in closely matching the distribution in the upper span of the blade. The qualitatively similar response of the low-order approach to the distorted inflow should allow for its use in preliminary blade design exploration once improved for high-speed flow prediction.

GE Aeroderivative Gas Turbines are derived from GE’s Aircraft engine family and converted to Land and Marine applications. As these Aeroderivative Gas Turbines are relatively smaller in size for similar power capacity in comparison to Heavy Duty Gas Turbines, there is a great need of developing efficient and compact turbine enclosure ventilation system for proper cooling and ventilation. Ineffective ventilation flow distribution inside the gas turbine enclosure causes engine circumferential nonuniform temperature distribution and it allows the formation of explosive fuel gas pockets inside the enclosure in the unlikely event of fuel leaks. Also, the engine nonuniform circumferential temperature gradient has an adverse impact on the operational efficiency and/or the mechanical integrity of the turbine. Proper cooling and ventilation system will also protect the sensitive equipment, like fuel valve actuators or any other instruments inside the turbine enclosure due to excessive radiation heat from engine hot surfaces, mainly near combustor and power turbine region. All the expected but significant engine leakages into the enclosure are estimated and considered for selection of right size, type, placing of ventilation fan at different operating conditions like full load, part load, elevation and ambient conditions. For first step, a 3D Computational Fluid Dynamics (CFD) analyses were done for turbine enclosure with mass/volumetric flow rate, temperature, and pressure boundary conditions to understand the flow/temperature distribution inside the enclosure. Radiation boundary conditions are applied on the engine casing external surfaces, enclosure walls and roof. The convective heat transfer from the hot surfaces are computed by CFD model based on the velocity and temperature predictions. In next step, from CFD analysis, identified the poor ventilation/stagnation zones using low velocity and recirculation areas close to the gas fuel components. After identifying the poor ventilation regions, a gas leak was introduced to see whether gas cloud volume is within the ISO 21789 – 2009 limits. From gas leak analysis results, enclosure outlet IR detector settings were decided and implemented to monitor the gas leak amount and feedback to control system in the form of ALARM/SHUTDOWN so that Gas Turbine operates safely.

Resonance in aerospace is a phenomenon that engineers have been trying to predict and avoid for a long time. Acoustic resonance is only a part in this field. When it was previously studied, it was mostly in connection with long slender gaps at the fuselage of aircrafts. Lately it has become a focus in the development of highly efficient aero engines. Bleed systems in the compressor part of engines are needed but not easy to place aerodynamically. Additionally, these bleed systems have complex geometries. These geometries coupled with the operational range of modern aircraft from low to high subsonic Mach numbers can create unwanted acoustic resonances. This paper is part of project study of these resonances. Here the bleed geometry is simplified to an open box cavity that is studied experimentally in order to measure its acoustic behavior in low to high subsonic flow. The experimental data is compared to theoretical prediction models to create a baseline for future studies. The results show a good agreement between Rossiter prediction and experiments for a shallow cavity of L/D = 4. Deeper cavities with a length to depth ratio of one and 0.5 represent more organ pipe resonance phenomena. This is especially governed by the geometry of the cavity itself and the height of the test section. All cavities experience a shift in modes depending on the operating point. This mode shift pattern is similar for deeper cavities. However, the operating range can be divided into four sections in which a mode shift occurs for all cavities.

The gas turbine engine design involves multi-disciplinary, multi-fidelity iterative design-analysis processes. These highly intertwined processes are nowadays incorporated in automated design frameworks to facilitate high-fidelity, fully coupled, large-scale simulations. The most tedious and time-consuming step in such simulations is the construction of a common geometry database that ensures geometry consistency at every step of the design iteration, is accessible to multi-disciplinary solvers and allows system-level analysis. This paper presents a novel design-intent-driven geometry modelling environment that is based on a top-down feature-based geometry model generation method. In the proposed object-oriented environment, each feature entity possesses a separate identity, denotes an abstract geometry, and carries a set of characteristics. These geometry features are organised in a turbomachinery feature taxonomy. The engine geometry is represented by a tree-like logical structure of geometry features, wherein abstract features outline the engine architecture, while the detailed geometry is defined by lower-level features. This top-down flexible arrangement of feature-tree enables the design intent to be preserved throughout the design process, allows the design to be modified freely and supports the design intent variations to be propagated throughout the geometry automatically. The application of the proposed feature-based geometry modelling environment is demonstrated by generating a whole-engine computational geometry. This geometry modelling environment provides an efficient means of rapidly populating complex turbomachinery assemblies. The generated engine geometry is fully scalable, easily modifiable and is re-usable for generating the geometry models of new engines or their derivatives. This capability also enables fast multi-fidelity simulation and optimisation of various gas turbine systems.

The need to reduce fuel-burn and CO 2 emissions, is pushing turbofan engines towards geared architectures with very high bypass-ratios and small ultra-high-pressure ratio core engines. However, this increases the radial offset between compressor spools and leads to a more challenging design for the compressor transition ducts. To minimise weight, these ducts must achieve the radial turning in as short a length, but this leads to strong curvature induced pressure gradients, increased aerodynamic loading and likelihood of flow separation. For the duct connecting the low-pressure fan to the engine core this is further complicated by the poor-quality flow generated at the fan hub which is characterised by low total pressure and large rotating secondary flow structures. In a previous paper the authors numerically designed modifications to an existing test facility such that the rotor would produce these large structures. The current paper presents an experimental evaluation of the new rotor design and examines the effect of the increased loss cores on the performance of a set of engine sector stators (ESS) or outlet guide vanes (OGV) and an engine representative compressor transition duct. Aerodynamic data were collected via miniature five-hole probes, for the time-averaged pressure and velocity field, and phase-locked hot-wire anemometry to capture the rotating secondary flows. Analysis of the experimental data showed that these structures promoted mixing through the ESS increasing the momentum exchange between the core and boundary layer flows. Measurements within the duct showed a continued reduction in the hub-wall boundary layer suggesting that the duct has been moved further from separation. Consequently, a more aggressive duct with 12.5% length reduction was designed and tested with the data confirming that the more aggressive duct remained fully attached. Total pressure loss data suggested a slight increase in loss over the vane row but that was offset by a reduced loss in the duct due to improved flow quality and reduced length. Overall, the 12.5% length reduction represents a significant cumulative effect in terms of reduced fuel burn and CO 2 over the operational life of an engine.

The turbine center frame (TCF) is an inherent component of modern turbofan aircraft engines, used for facilitating the large radius change between the high-pressure (HPT) and low-pressure (LPT) turbines. Secondary flow features that develop in the TCF result in total pressure loss of the mainstream flow and a subsequent performance reduction for the whole of the engine. Purge flows from the HPT interact with these flow features affecting their development and strength. Understanding the details of this interaction is therefore of paramount importance for the design of more efficient engines of the future. This paper presents a detailed investigation of the interaction of purge flows from the hub and shroud cavities upstream and downstream of the HPT rotor with the secondary flow features in a TCF. The investigation was conducted using aerodynamic and seed gas concentration measurements in an engine-representative HPT-TCF setup and under engine-realistic operating conditions. The upstream purge flows interact with the flow-field of the rotor, and especially with the upper and lower passage vortices where they are mainly entrained, forming “zones-of-influence” that occupy the upper and lower 35% of the span at the TCF inlet. Dilution of these purge flows occurs through vortex-to-vortex interactions and in-plane flow migrations driven by the vortices. At the outlet of the TCF, the upstream purge flows form effectiveness bands that encapsulate the various counter-rotating vortices near the hub and shroud. This indicates that these counter-rotating vortices were formed at the inlet of the TCF, in a flow that already includes the upstream purge flows. The downstream purge flows exit the hub and shroud cavities forming effectiveness boundary layers at the inlet of the TCF of thickness equal to around 15% of the span. The circumferential distribution of these purge flows is however asymmetric, owing to the also asymmetric static pressure distribution at the inlet of the TCF, as a result of the effect of the propagated flow-field of the stator vanes. At the outlet of the TCF, the distribution of the downstream hub purge appears as distinct effectiveness lobes with the same periodicity as the HPT vanes. The formation of the lobes is as a result of intense interaction between the counter-rotating vortex pairs and the downstream hub purge flow. The viscous shear mixing due to this interaction is also the cause for the low total pressure in the regions influenced by the lobes. The distribution of the downstream shroud purge appears as alternating regions of high and low effectiveness as a result of radially inwards and outwards flow migrations caused by the shearing actions of the counter rotating vortices near the shroud. These migrations are the cause of regions with the lowest total pressure at the outlet of the TCF.

Axial thrust is one of the critical factors that affect the pump’s continuously operating reliability. Among all the available methods for axial thrust controlling, Pump Out Vanes (POVs) are an easy and effective way. Different from a single-stage pump with a scroll, an in-line multistage pump will have a leakage flow channel from the return channel. With this leakage channel, the working environment of the POVs will be significantly different from a single-stage pump. In this paper, the first stage of a multistage pump with both POVs and casing ribs (vortex breakers) is studied by CFD simulation to evaluate their effect on the axial thrust, pump stage performance, and the internal leakage flow. Because of the similar POV working environment in the multistage pump, the conclusion from one stage can be generalized for the rest stages. In this study, 5 models with different POV outer radius and height are simulated in Ansys Fluent with k-ε turbulence model and transient rotor-stator sliding mesh method. The results show that POVs with suitable geometry can provide good axial thrust control over a wide pump operating range while the stage efficiency can be strongly affected due to the increased turbulence and interstage leakage flow, which is contradicting some previous researcher’s conclusion based on the study of a single-stage centrifugal pump.

Optimizing the design of aviation propulsion systems using computational fluid dynamics is essential to increase their efficiency and reduce pollutant as well as noise emissions. Nowadays, this design process is increasingly aided by computational fluid dynamic methods for which and with the adequate modeling approach it is possible to perform meaningful unsteady computations of the various components of a gas-turbine engine. However, these simulations are often carried out independently of each other and only share averaged quantities at the component interfaces minimizing the impact and interactions between components. The present work investigates the interactions between fan, compressor and annular combustion chamber at takeoff conditions by simulating a 360 azimuthal degrees large-eddy simulation of over 2100 million cells of the DGEN-380 demonstrator. In that case, the domain includes: 14 fan blades; 42 outlet-guide vanes (OGV); the impeller with 11 main blades and 11 splitter blades; a radial and an axial diffuser with 22 and 55 vanes, respectively; and the annular combustion chamber with a contouring casing and 13 swirlers on the back of the chamber. At take-off conditions it is found that the compressor operates in transonic conditions in the rotating frame of reference of the impeller and a shock is formed at the leading edge of the main blades which propagates upstream towards the fan and it is perceived at half the impeller blade-passing frequency (BPF). Preliminary results also show that pressure fluctuations at the impeller BPF generated by the interaction of the impeller blades with the diffuser vanes are propagated through the axial diffuser and enter the combustion chamber through the dilution holes and the swirler. The objective of this paper is to provide a deeper analysis of the interactions between components through the use of the novel operator-based analysis called dynamic mode tracking method (DMT). Indeed, this method facilitates the analysis of three-dimensional results despite the billion-size mesh and the complexity of the simulation, since it extracts modes at specific frequencies on-the-fly within the code. The frequencies corresponding to the fan, impeller and half the impeller BPF are analyzed in the domain and compared against traditional and more computationally demanding methods like the well-known Dynamic Mode Decomposition or the Direct Fourier transform.

Computational fluid dynamics (CFD) is increasingly used during the design phase of turbomachinery. Reducing the cost of such computations is one of the major challenges in the industrial field. The multi-physics phenomena and the multidisciplinary interactions needed for the design of the engine components are difficult to be faced with the classical design methods. In addition, the interest of manufacturers and operators of turbomachines in the increase in performance when degradation processes occur is growing. In the aeronautical sector degradation is among the most critical issues, as it can lead to the in-flight shutdown of the engine. In a bid to tackle these big problems, new design methods based on approximation techniques have been developed. These techniques are called surrogate models and are currently the most used design methods for the aerodynamic design of aircraft engines. In this work, the assessment of a surrogate surface built for the purpose of optimizing an HPT vane in degrading conditions is performed. Using machine learning and statistical techniques, a sensitivity analysis is conducted in order to reduce the problem dimensions. The results of the sensitivity analysis are used for a study of the surrogate, with the purpose of obtaining design guidelines when deterioration effects are considered in the design phase. The main outcome of this study is a map, that outlines the best design zone defined by the combination of the most influential parameters.

The design and analysis of the secondary air system (SAS) of gas turbine engine is a complex and time-consuming process because of its complicated geometry topology. The conventional SAS design-analysis model generation process is quite tedious, time consuming. It is still heavily dependent on human expertise and thus incurs high time-cost. This paper presents an automated, whole-engine SAS flow network model generation methodology. During the SAS preprocessing step, the method accesses a pre-built whole-engine geometry model created using a novel, in-house, feature-based geometry modelling environment. It then transforms the engine geometry features into the features suitable for SAS flow network analysis. The proposed method not only extracts the geometric information from the computational geometry but also retrieves additional non-geometric attributes such as, rotational frames, boundary types, materials and boundary conditions etc. Apart from ensuring geometric consistency, this methodology also establishes a bi-directional information exchange protocol between engine geometry model and SAS flow network model, which enables making engine geometry modifications based on SAS analysis results. The application of this feature mapping methodology is demonstrated by generating the secondary air system (SAS) flow network model of a modern three-shaft gas turbine engine. This capability is particularly useful for the integration of geometry modeler with the simulation framework. The present SAS model is generated within a few minutes, without any human intervention, which significantly reduces the SAS design-analysis time-cost. The proposed method allows performing a large number of whole-engine SAS simulations, design optimisations and fast re-design activities.

A novel turbocompressor test facility has been designed for helium-neon gas mixtures and its specific features are presented. To account for the heat transfer originating from the motor coolant, a surrogate model has been derived. By combining these results with additional numerical ones, a similitude study is conducted quantifying the individual effects and contributions on efficiency of the Reynolds number, tip Mach number and specific heat ratio. Decoupling the effect of the different parameters shows that their respective contribution on efficiency variation is highly correlated to the Reynolds number actual value. The negative contribution of the tip Mach number and the positive effect of Reynolds number can be used to explain the efficiency variation with increasing tip Mach number. Specific heat ratio variation leads to minor changes in polytropic efficiency except at low tip Mach numbers.

One-dimensional (1D) modelling is critical for turbomachinery unsteady performance prediction and system response assessment of internal combustion engines. This paper uses a novel 1D modelling (TURBODYNA) and proposes two additional features for the application to a twin-entry turbocharger turbine. Compared to single-entry turbines, twin-entry turbines enhance turbocharger transient response and reduce engine exhaust valve overlap periods. However, out-of-phase high frequency pulsating pressure waves lead to an unsteady mixing process from the two flows and pose great challenges to traditional 1D modelling. The present work resolves the mixing problem by directly solving mass, momentum and energy conservation equations during the mixing process instead of applying constant pressure assumption at the limb-rotor joint. Comparisons of TURBODYNA and an experimentally validated CFD suggest that TURBODYNA can not only provide a very good agreement on turbine performance, but also accurately capture unsteady features due to flow field inertial and pressure wave propagation. Levels of accuracy achieved by TURBODYNA have proved superior to traditional 1D modelling on turbine performance and the generality of the current 1D modelling has been explored by extending the application to another turbine featuring distinct characteristics.