bluuebunny/crossref_metadata_2025 huggingface dataset:
Experimental and Numerical Investigation of High-Temperature Multi-Axial Fatigue
Authors: Ramesh Babu Harish , Böcker Marco , Raddatz Mario , Henkel Sebastian , Biermann Horst , Gampe Uwe
Published: September 2021

Gas turbines and aircraft engines are dominated by cyclic operating modes with fatigue-related loads. This may result in the acceleration of damage development on the components. Critical components of turbine blades and discs are exposed to cyclic thermal and mechanical multi-axial fatigue. In the current work, planar-biaxial Low-Cycle-Fatigue (LCF) tests are conducted using cruciform specimens at different test temperatures. The influence on the deformation and lifetime behaviour of the nickel-base disk alloy Inconel 718 is investigated at selected cyclic proportional loading cases, namely shear and equi-biaxial. The calculation of the stress and strain distribution of the cruciform specimens from the experimental data is difficult to obtain due to complex geometry and temperature gradients. Therefore, there is a need for Finite Element (FE) Simulations. A viscoplastic material model is considered to simulate the material behaviour subjected to uniaxial and the selected planar-biaxial loading conditions. At first, uniaxial simulation results are compared with the uniaxial experiment results for both batches of IN718. Then, the same material parameters are used for simulating the biaxial loading cases. The prediction of FE simulation results is in good agreement with the experimental LCF test for both shear and equi-biaxial loadings. The equivalent stress amplitude results of the biaxial simulation are compared with the uniaxial results. Furthermore, the lifetime is calculated based on the stabilized cycle from the simulation and by using Crossland and Sines multi-axial stress-based approaches. The Crossland model predicts fatigue life significantly better than the Sines model. Finally, the simulated lifetime results are compared with the experimental lifetime.
Prediction of Dynamic Behavior of a Single Shaft Gas Turbine Using NARX Models
Authors: Asgari Hamid , Ory Emmanuel
Published: September 2021

Gas turbines are internal combustion engines widely used in industry as main source of power for aircrafts, turbo-generators, turbo-pumps and turbo-compressors. Modelling these engines can help to improve their design and manufacturing processes, as well as to facilitate their operability and maintenance. These eventually lead to manufacturing of gas turbines with lower costs and higher efficiency at the same time. The models may also be employed to unfold nonlinear dynamics of these systems.

The aim of this study is to predict the dynamic behavior of a single shaft gas turbine by using open-loop and closed-loop NARX models, which are subsets of artificial neural networks. To set up these models, datasets of significant variables of the gas turbine are used for training, test and validation processes. For this purpose, a comprehensive code is developed in MATLAB programming environment. In addition to the open-loop model, a closed-loop model is set up for multi-step prediction. The results of this study demonstrate the capability of the NARX models in reliable prediction of gas turbines’ dynamic behaviors over different operational ranges.
Analysis of Loading and Vibration Histories on Natural Frequencies and Modal Damping of Blades With Friction at Root Contact Interfaces
Authors: Chen Junjie , Zang Chaoping , Zhou Biao , Petrov E. P.
Published: September 2021

Gas turbine engines change the rotation speed during operation in a wide range from zero to the speed corresponding to the cruise flight and the history of the rotation speed variation is individual for each flight. During the engine operation the bladed disks pass different resonances frequencies and may experience significant vibration amplitudes. The vibrations can affect the interaction conditions at friction contact interfaces, including contact stresses distribution and their contact status. As a result, the contact conditions can differ from those that are established at the initial bladed disk assembling and these conditions are dependent on the history of the vibration that a bladed disk experiences during every engine operation cycle. The variation of the contact conditions at friction contact interfaces affects modal properties and modal damping characteristics of a bladed disk, and it is important to assess the possible variation of these properties during the engine exploitation.

In this paper, a transient analysis is performed to simulate blade vibration under different loading histories occurred in a flight mission. The analysis is performed for different histories of the blade vibration and a method is proposed for the analysis of modal damping in the blade root joints under the influence of different loading histories. The influence of different loading histories, friction contact parameters, the vibration excitation levels and resonating modes on modal characteristics and modal damping factors is studied, The pre-stress effects due to blade assembling are also explored.
On Small Scale LNG Concepts
Authors: Kurz Rainer , Ji Min , Beck Griffin , Allison Timothy C.
Published: September 2021

The different economics of small scale LNG plants put more emphasis on capital expenses over process efficiency, and thus favors simpler refrigeration cycles. We typically find reverse Brayton cycles, or SMR (Single mixed refrigerant) cycles. These cycles have specific requirements to the compression equipment, and typically have smaller drivers, either electric drives or gas turbines. The relationship between output, driver size, and process preferences is explained. The type of compressors, and expanders needed are discussed, together with thoughts and the driver preferences. This includes the different control methods that can be used, both for the cycle adaptation, as well as the related control of the compressors, expanders, valves and drivers. Equipment performance maps are created to highlight the required different operating conditions. This result allows for subsequent optimization discussions.
Additively Manufactured Porous Geometries for Hybrid Turbine Cooling
Authors: Fier Nathan D. , Bogard David G.
Published: September 2021

Discrete film cooling holes are limited by subtractive manufacturing techniques and experience depreciating performance when operating above critical velocity ratios. This study introduces an alternative method of bringing coolant to the surface of the blade via finite strips of porous material interlaced throughout the blade, made possible by advances in additive manufacturing (AM). Both experimental and computational studies were performed on the porous hybrid configuration to characterize downstream and off-wall performance, where experimental adiabatic effectiveness values were achieved using a plastic, fused deposition printed lattice structure. The method of bringing coolant onto the surface of the blade through an additively manufactured porous region experienced downstream adiabatic effectiveness values similar to slots while providing better structural stability. Additionally, the hybrid configuration outperformed shaped film cooling holes by injecting an ultra-thin layer of coolant that was evenly distributed span-wise across the blade. When operating at VRhybrid = 0.052 and L/d = 2 the hybrid configuration produced spatially averaged values 30% greater than the shaped holes while using equivalent coolant mass flow rate. Also, for an L/d = 10, the spatially averaged adiabatic effectiveness, for the hybrid configuration, is a factor of three greater than for shaped film cooling holes, while requiring a five times greater coolant mass flow rate. Finally, the RANS computational model accurately predicted downstream effectiveness values, at low velocity ratios, within experimental uncertainty but showed inaccuracies when predicting off wall effectiveness values and at higher velocity ratios.
Effect of Differential Tip Clearance on the Performance of a Tandem Rotor
Authors: Kumar Amit , Chhugani Hitesh , More Shubhali , Pradeep A. M.
Published: September 2021

Tandem blade is an interesting concept that promises a higher total pressure rise per stage. Owing to two separate tip leakage vortices and their interaction, losses are likely to increase particularly near the tip region. Although, rotors are designed with optimum tip clearance, the clearance changes during engine operation as well as during its service life. In the case of tandem rotors, the forward and the aft rotors can have different tip clearances. This will also impact the performance of the stage. Six different tip clearances have been investigated. ANSYS CFX is used for steady RANS computational analysis. The results suggest that the performance of the tandem rotor is highly sensitive to the forward rotor tip clearance. Higher tip clearance adversely affects the total pressure rise and operation stability of the tandem rotor. At design mass flow rate, the performance degradation for tandem configuration with the higher tip clearance (Case2, Case 3, Case 5, and Case 6), is attributed to the vortex breakdown of TLV1, which leads to the sudden expansion of the blockage region near the rotor tip. Vortex breakdown primarily depends upon the swirling strength of TLV1 and TLV2 as well as on the adverse pressure gradient. Near the stall point, the role of the adverse pressure gradient becomes more dominant in the vortex breakdown.
Investigations on the Aerothermal Performance Uncertainty Quantification of the Turbine Blade Squealer Tip
Authors: Huang Ming , Li Zhigang , Li Jun , Song Liming
Published: September 2021

The first-stage rotor squealer tip is a key area in the gas turbine for both aerodynamic performance and heat transfer characteristics, which should be carefully designed. However, harsh operating conditions near the rotor squealer tip can cause the geometry of the squealer tip to degrade, and manufacturing inaccuracies can also cause the squealer tip geometry to deviate from the ideal design. In this work, an uncertainty quantification (UQ) method is proposed using the non-intrusive polynomial chaos expansion method and Smolyak sparse grids. Then coupled with three dimensional (3D) Reynolds-Averaged Navier-Stokes (RANS) solutions, an uncertainty quantification procedure is carried out for aerodynamic and heat transfer performance of GE-E3 rotor blade squealer tip. A parameter sensitivity analysis using the Sobol Indice method is carried out to identify the key parameters for aerothermal performance of the squealer tip. Wherein, the inlet total temperature and the blowing ratio are considered as flow condition uncertainty parameters and tip clearance is considered as geometrical uncertainty parameters.

The uncertainty analysis results show that under the influence of the uncertain geometry and operating conditions, the heat flux of squealer tip basically conforms to the normal distribution and the statistical mean value of it increased by 13.56% relative to the design value and the probability of it deviating from the design value by 10% is as high as 65.68% The statistical average of the squealer tip film cooling effectiveness is reduced by 29.52% compared to the design value, and the probability of it deviating from the design value by 10% is as high as 91.83%. The result of sensitivity analysis reveals that the uncertainty of the aerodynamic characteristics of the squealer tip is almost entirely caused by the tip clearance which accounts for 88.02% of the variance of the leakage flow rate. The inlet total temperature has almost no effect on the uncertainty of the aerodynamic performance. However, it is the dominant variable for the uncertainty of the heat transfer performance considering that its variance indexes for tip heat flux QTip and film cooling effectiveness η are 84.87% and 24.87% respectively. Compared with the main effects, the influence of the interaction effects among the variables on the squealer tip aerothermal performance is almost negligible.
Numerical Investigations on the Aerothermal Performance and Film Cooling Effectiveness of Turbine Vane Endwall at Inlet Swirl Conditions
Authors: Li Zhiyu , Zhang Kaiyuan , Li Zhigang , Li Jun
Published: September 2021

Lean burn premix technology of gas turbines results in that the turbine combustor outlet has strong swirl flow characteristics which directly influences the inlet flow condition of the first turbine vane downstream from combustor and raise the thermal load of endwall. The aerothermal performance and film cooling effectiveness of first turbine vane endwall at different inlet swirl conditions is numerically investigated in this paper. The flow pattern, Nusselt number distribution and film cooling effectiveness of turbine vane endwall at the uniform and three kinds of swirl numbers (0.6, 0.8, 1.0) inflow conditions with clockwise and anticlockwise swirl orientations are analyzed using three-dimensional Reynolds-Averaged Navier-Stokes (RANS) and SST k-ω turbulence model solutions. The obtained results show that the flow field is apparently influenced by inlet swirl conditions. The separation in passage is clearly suppressed at clockwise swirl inflow conditions but anticlockwise make the flow pattern more complex. Inlet swirl can increase the overall Nu on the endwall. Especially in the area upstream of leading edge and area between the first and second row of film holes whose Nu are increased by 4 and 1.5 times compared with uniform inflow, respectively. Not only swirl number but also orientation can affect the film cooling effectiveness distribution of the vane endwall. The better film cooling effectiveness distribution and higher film cooling effectiveness downstream of each film hole row can be achieved at the clockwise swirl inflow conditions by weakening the accumulation of coolant near the suction side of the turbine vane endwall. Compared with uniform inflow, average film cooling effectiveness of endwall between the second row and third row is increased from 0.21 to 0.27 at clockwise SN = 1.0 for the maximum increase. The detailed flow field and aerothermal performance of the turbine vane endwall at different swirl inflow conditions is also discussed and illustrated.
Modelling of Turbulent Premixed CH4/H2/Air Flames Including the Influence of Stretch and Heat Losses
Authors: Kutkan Halit , Amato Alberto , Campa Giovanni , Ghirardo Giulio , Tay Wo Chong Luis , Æsøy Eirik
Published: September 2021

This paper presents a RANS turbulent combustion model for CH4/H2/air mixtures which includes the effect of heat losses and flame stretch. This approach extends a previous model concept designed for methane/air mixtures and improves the prediction of flame stabilization when hydrogen is added to the fuel. Heat loss and stretch effects are modelled by tabulating the consumption speed of laminar counter flow flames in a fresh-to-burnt configuration with detailed chemistry at various heat loss and flame stretch values. These computed values are then introduced in the turbulent combustion model by means of a turbulent flame speed expression which is derived as a function of flame stretch, heat loss and H2 addition. The model proposed in this paper is compared to existing models on experimental data of spherical expanding turbulent flame speeds. The performance of the model is further validated by comparing CFD predictions to experimental data of an atmospheric turbulent premixed bluff-body stabilized flame fed with CH4/H2/air mixtures ranging from pure methane to pure hydrogen.
Center Body Burner for Sequential Combustion: Superior Performance at Lower Emissions
Authors: Ciani A. , Wood J. P. , Maurer M. , Bunkute B. , Pennell D. , Riazantsev S. , Früchtel G.
Published: September 2021

Modern gas turbines call for an ultra-high firing temperature and fuel flexibility while keeping emissions at very low levels. Sequential combustion has demonstrated its advantages toward such ambitious targets.

A sequential combustion system, as deployed in the GT26 and GT36 engines, consists of two burners in series, the first one optimized to provide the optimum boundary condition for the second one, the sequential burner. This is the key component for the achievement of the required combustor performance dictated by F and H class engines, including versatile and robust operation with hydrogen-based fuels.

This paper describes the key development considerations used to establish a new sequential burner surpassing state-of-the-art hardware in terms of emission reduction, fuel flexibility and load flexibility. A novel multi-point injector geometry was deployed based on combustion and fluid dynamic considerations to maximize fuel / air mixing quality at minimum pressure loss. Water channel experiments complemented by CFD describe the evolution of the fuel / air mixture fraction through the mixing section and combustion chamber to enable operation with major NOx reduction. Furthermore, Laser Doppler Anemometry and Laser Induced Fluorescence were used to best characterize the interaction between hot-air and fuel and the fuel / air mixing in the most critical regions of the system. To complete the overview of the key development steps, mechanical integrity and manufacturing considerations based on additive manufacturing are also presented. The outcome of 1D, CFD and fluid dynamic experimental findings were then validated through full-scale, full-pressure combustion tests. These demonstrate the novel Center Body Burner is enabling operation at lower emissions, both at part load and full load conditions. Furthermore, the validation of the burner was also extended to hydrogen-based fuels with a variety of hydrogen / natural gas blends.
Evolutionary Algorithm for Enhanced Gas Path Analysis in Turbofan Engines
Authors: Rootliep T. O. , Visser W. P. J. , Nollet M.
Published: September 2021

Adaptive modelling (AM) based Gas Path Analysis (GPA) is a powerful diagnostic and prognostic technique for turbofan engine maintenance. This involves the assessment of turbofan component condition using thermodynamic models that can iteratively adapt to measurements values in the gas path by changing component condition parameters. The problem with this approach is that newer turbofan engines such as the General Electric GEnx-1B have fewer gas path sensors installed causing the AM equation systems to become underdetermined. To overcome this problem, a novel approach has been developed that combines the AM model with an Evolutionary Algorithm (EA) optimization scheme and applies it to multiple operating points. Additionally, these newer turbofan engines provide performance data continuously during flight. Information on variable geometry and bleed valve position, active clearance control state and power off-take is included and can be accounted for to further enhance AM model accuracy. A procedure is proposed where the selection of operating points is based on steady-state stability requirements, cycle model operating point uncertainty and parameter outlier filtering. The Gas turbine Simulation Program (GSP) is used as the non-linear GPA modelling environment. A Multiple Operating Point Analysis (MOPA) is chosen to overcome the problem of underdetermination by utilizing multiple data sets at different operating points. The EA finds the best fit of health parameter deviations by minimizing the multi-point objective function using the GSP AM model. A sub-form of the EA class named Differential Evolution (DE) has been chosen as the optimizer. Like all EAs, DE is a parallel direct search method in which a population of parameter vectors evolves following genetic operations towards an optimum output candidate.

The resulting hybrid GPA tool has been verified by solving for different simulated deterioration cases of a GSP model. The tool can identify the direction and magnitude of condition deviation of 10 health parameters using 6 gas path sensors. It has subsequently been validated using historical in-flight data of the GEnx-1B engine. It has demonstrated successful tracking of engine component condition for all 10 health parameters and identification of events such as turbine blade failure and water washes. The authors conclude that the tool has proven significant potential to enhance turbofan engine condition monitoring accuracy for minimizing maintenance costs and increasing safety and reliability.
Aero-Thermal Characterization of Accelerating and Diffusing Passages Downstream of Rotating Detonation Combustors
Authors: Braun James , Paniagua Guillermo , Ferguson Donald
Published: September 2021

Cycle benefits of rotating detonation engines show up to five percentage points of efficiency gain for low-pressure ratio engines. An optimal integration between the combustor and the turbine needs to be guaranteed to realize this potential gain. The rotating detonation combustor (RDC) exhausts transonic flow with shocks rotating at frequencies ranging from a few to tens of kilohertz depending on the number of present waves. Hence, the turbine design requires precise knowledge of the fluctuations and losses downstream of the combustor. This paper focuses on the quantification of fluctuations and losses for accelerating and diffusing passages. The analysis of the combustor is performed via reactive unsteady Reynolds Averaged Navier-Stokes (URANS) simulations. The unsteady RANS equations are solved via CFD++ from Metacomp with a one-step reaction mechanism for an H2-air mixture. The resolving of the boundary layer is achieved with a structured mesh of around 36 million cells. Inlet pressure of 10 bar and two different back pressures are applied to the combustor to model the interconnection with downstream turbines. Finally, we present and assess a methodology to reduce the computational time to model these passages ten times.
Relative Effects of Velocity- and Mixture-Coupling in a Thermoacoustically Unstable, Partially-Premixed Flame
Authors: Karmarker Ashwini , O’Connor Jacqueline , Boxx Isaac
Published: September 2021

Combustion instability, which is the result of a coupling between combustor acoustic modes and unsteady flame heat release rate, is a severely limiting factor in the operability and performance of modern gas turbine engines. This coupling can occur through different coupling pathways, such as flow field fluctuations or equivalence ratio fluctuations. In realistic combustor systems, there are complex hydrodynamic and thermo-chemical processes involved, which can lead to multiple coupling pathways. In order to understand and predict the mechanisms that govern the onset of combustion instability in real gas turbine engines, we consider the influences that each of these coupling pathways can have on the stability and dynamics of a partially-premixed, swirl-stabilized flame. In this study, we use a model gas turbine combustor with two concentric swirling nozzles of air, separated by a ring of fuel injectors, operating at an elevated pressure of 5 bar. The flow split between the two streams is systematically varied to observe the impact on the flow and flame dynamics. High-speed stereoscopic particle image velocimetry, OH planar laser-induced fluorescence, and acetone planar laser-induced fluorescence are used to obtain information about the velocity field, flame, and fuel-flow behavior, respectively. Depending on the flow conditions, a thermoacoustic oscillation mode or a hydrodynamic mode, identified as the precessing vortex core, is present. The focus of this study is to characterize the mixture coupling processes in this partially-premixed flame as well as the impact that the velocity oscillations have on mixture coupling. Our results show that, for this combustor system, changing the flow split between the two concentric nozzles can alter the dominant harmonic oscillation modes in the system, which can significantly impact the dispersion of fuel into air, thereby modulating the local equivalence ratio of the flame. This insight can be used to design instability control mechanisms in real gas turbine engines.
Describing the Mechanism of Instability Suppression Using a Central Pilot Flame With Coupled Experiments and Simulations
Authors: Li Jihang , Kwon Hyunguk , Seksinsky Drue , Doleiden Daniel , O’Connor Jacqueline , Xuan Yuan , Akiki Michel , Blust James
Published: September 2021

Pilot flames are commonly used to extend combustor operability limits and suppress combustion oscillations in low-emissions gas turbines. Combustion oscillations, a coupling between heat release rate oscillations and combustor acoustics, can arise at the operability limits of low-emissions combustors where the flame is more susceptible to perturbations. While the use of pilot flames is common in land-based gas turbine combustors, the mechanism by which they suppress instability is still unclear. In this study, we consider the impact of a central jet pilot on the stability of a swirl-stabilized flame in a variable-length, single-nozzle combustor. Previously, the pilot flame was found to suppress the instability for a range of equivalence ratios and combustor lengths. We hypothesize that combustion oscillation suppression by the pilot occurs because the pilot provides hot gases to the vortex breakdown region of the flow that recirculate and improve the static, and hence dynamic, stability of the main flame. This hypothesis is based on a series of experimental results that show that pilot efficacy is a strong function of pilot equivalence ratio but not pilot flow rate, which would indicate that the temperature of the pilot gases as well as the combustion intensity of the pilot flame play more of a role in oscillation stabilization than the length of the pilot flame relative to the main flame. Further, the pilot flame efficacy increases with pilot flame equivalence ratio until it matches the main flame equivalence ratio; at pilot equivalence ratios greater than the main equivalence ratio, the pilot flame efficacy does not change significantly with pilot equivalence ratio. To understand these results, we use large-eddy simulation to provide a detailed analysis of the flow in the region of the pilot flame and the transport of radical species in the region between the main flame and pilot flame. The simulation, using a flamelet/progress variable-based chemistry tabulation approach and standard eddy viscosity/diffusivity turbulence closure models, provides detailed information that is inaccessible through experimental measurements.
Numerical Investigation of Axial Thrust Control in a Multistage Canned-Motor Pump With Pump-Out-Vanes
Authors: Wang Yintao , Zhang Min , Patil Abhay , Morrison Gerald
Published: September 2021

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.
Influence of Supersonic Nozzle Design Parameters on the Unsteady Stator-Rotor Interaction in Radial-Inflow Turbines for Organic Rankine Cycles
Authors: Cappiello Alessandro , Tuccillo Raffaele
Published: September 2021

Organic Rankine Cycle (ORC) technology represents an interesting option for improving the efficiency of existing power plants and industrial processes as well as exploiting renewable and renewable-equivalent energy sources.

The use of Radial-Inflow Turbine (RIT) for ORC plant sizes below 100 kW is promising, although the application remains challenging. In fact, the single stage arrangement imposed by economic constraints and hence the large expansion ratio, together with the large molecular weight, which characterizes organic fluids, usually result in highly supersonic flows, so making the use of transonic stators often mandatory.

Particularly, the influence of RIT stator design parameters on losses and the level of unsteadiness seen by the subsequent rotor is still scarcely addressed in published literature.

Previous work by the authors investigated the effect of some stator design parameters on stator loss and downstream circumferential uniformity.

The present work investigates the effect of the convergent-divergent stators design parameters and the resulting downstream flow field non-uniformity on the unsteady stator-rotor interaction and loss generation in ORC Radial-Inflow Turbines.

To this end, two stator and rotor configurations which differ by the stator design parameters (i.e., discharge metal angle and number of vanes) have been tested by means of 3D unsteady CFD calculations accounting for real-gas properties.

The results show that larger stator-rotor interaction is present for the case featuring higher vane count and lower outlet metal, which also features the largest fluctuations of power output and pressure force on blade, together with a substantially lower average total-to-static efficiency.
High Temperature Solid Particle Erosion in a Melt-Infiltrated SiC/SiC Ceramic Matrix Composite
Authors: Presby Michael J.
Published: September 2021

Ceramic matrix composites (CMCs) are an enabling propulsion material system that offer weight benefits over current Ni-based superalloys, and have higher temperature capabilities that can reduce cooling requirements. Incorporating CMCs into the hot section of gas-turbine engines therefore leads to an increase in engine efficiency. While significant advancements have been made, challenges still remain for current and next-generation gas-turbines; particularly when operating in dust-laden or erosive environments. Solid particles entrained in the gas flow can impact engine hardware resulting in localized damage and material removal due to repeated, cumulative impacts. In this study, the erosion behavior of a melt-infiltrated (MI) silicon carbide fiber-reinforced silicon carbide (SiC/SiC) CMC is investigated at high temperature (1,200 °C) in a simulated combustion environment using 150 μm alumina particles as erodent. Particle impact velocities ranged from 100 to 200 m/s and the angle of impingement varied from 30° to 90°. Erosion testing was also performed on α-SiC to elucidate similarities and differences in the erosion response of the composite compared to that of a monolithic ceramic. Scanning electron microscopy (SEM) was used to study the post-erosion damage morphology and the governing mechanisms of material removal.
Improving Vibration Response of Radial Turbine in Variable Geometry Turbochargers With CFD Analysis
Authors: Gupta Bipin , Yoshida Toyotaka , Ogawa Shinji , Danmoto Yosuke , Yoshimoto Takashi
Published: September 2021

Recent advancements in internal combustion engine for efficient fuel combustion, such as application of miller cycle, where the closing of engine intake valve is purposely delayed to provide more cooling of air-fuel mixture during compression stroke for better engine efficiency, has led to a requirement for turbochargers to function at a wider operating range and higher compression ratio. One of the methods which have been largely accepted is the use of variable geometry turbochargers. As compared to diesel engine, operating conditions for gasoline engine require the turbine to operate at higher exhaust temperature, which increases the risk of damaging the rotor.

This paper discusses a detailed flow analysis of the effect of tip leakage and nozzle vane wake flow on surface pressure distribution of the turbine rotor, especially at the severe condition when vane trailing edge and rotor leading edge are in proximity. It was observed in steady and unsteady CFD simulations that the origination and propagation of tip leakage flow can be varied depending on the blade loading at the rotor leading edge, and the major interaction of nozzle wake can be switched from pressure surface to suction surface as rotor blade crossed a nozzle vane, which can drastically affect the alternating aerodynamic stresses. The sensitivity to this phenomenon has been evaluated by calculating the safety factor.

The authors modified the rotor design to weaken the effect of tip leakage flow in order to suppress variations in rotor surface pressure as it crosses the nozzle vane. It significantly reduced the alternating stress and increased the safety factor at vibration mode 2 from 0.3 to 9.3 and mode 3 from 0.6 to 3.2 respectively.
Global Optimisation of a Transonic Fan Blade Through AI-Enabled Active Subspaces
Authors: Lopez Diego I. , Ghisu Tiziano , Shahpar Shahrokh
Published: September 2021

The increased need to design higher performing aerodynamic shapes has led to design optimisation cycles requiring high-fidelity CFD models and high-dimensional parametrisation schemes. The computational cost of employing global search algorithms on such scenarios has typically been prohibitive for most academic and industrial environments. In this paper, a novel strategy is presented that leverages the capabilities of Artificial Neural Networks for regressing complex unstructured data, while coupling them with dimensionality reduction algorithms. This approach enables employing global-based optimisation methods on high-dimensional applications through a reduced computational cost. This methodology is demonstrated on the efficiency optimisation of a modern jet engine fan blade with constrained pressure ratio. The outcome is compared against a state-of-the-art adjoint-based approach. Results indicate the strategy proposed achieves comparable improvements to its adjoint counterpart with a reduced computational cost, and can scale better to multi-objective optimisation applications.
Characterization of a Newly Designed Test Bench for Investigations of Flame-Wall-Interaction
Authors: Dalshad Rahand , Sander Tobias , Pfitzner Michael
Published: September 2021

For the thermal design of combustion chambers and turbine blades in jet engines, a detailed knowledge of the combustion and of the heat loads to the walls is necessary. In general, high operating temperatures and reduced combustor size are striven for in order to increase engine efficiency and reduce weight. Consequently, the components are exposed to temperatures above the melting point of the materials and there is a growing risk of in-complete combustion within the combustion chambers. To study these effects, we set up a new test bench for fundamental investigation of chemical near-wall reactions at atmospheric pressure. First results of gaseous, non-premixed near-wall CH4/air and H2/air flames are presented. Optical methods such as two-line laser-induced fluorescence themometry and OH* chemiluminescence were applied. Further, the heat release to the wall was determined by means of inverse heat conduction calculation using the data of implemented thermocouples.
Influence of the Tool Wear on the Quality and Service Life of Gears for the Geared Turbofan Technology Machined by Five-Axis Milling
Authors: Lakner Thomas , Zachert Christoph , Greschert René , Schraknepper Daniel , Bergs Thomas
Published: September 2021

The geared turbofan technology is one essential way to reduce the fuel consumption, the environmental footprint and the noise pollution of civil aircrafts. An added gearbox between the fan and the low-pressure compressor that reduces the fan speed, which allows higher bypass ratios, achieves the mentioned benefits of geared turbofans. To withstand the high mechanical loads, large double helical gears are used. Gear hobbing and gear grinding require large tool maneuvering spaces. This leads to a larger required space between the single gears of the double helical gear. As a result, the gears are larger and heavier, which leads to a reduced economy of the aircraft. The tool maneuvering space of five-axis milling with solid carbide end mills is much smaller. This enables the design of smaller, lighter and more efficient aircraft engines. However, manufacturing these gears in tolerances better than IT5 is very challenging on five-axis milling machine tools. This paper presents investigations about finish machining of hardened gears on five-axis machine tools. In the investigations performed, varying tool substrates and tool coatings have been investigated together with tool travel paths in order to reduce the tool wear, which is key to achieve the demanded tolerances. Finally, the five-axis milled gears were compared to conventionally manufactured gears on test benches to enable statements regarding the expectable service lives of the manufactured gears.
A Comparison Between ORC and Supercritical CO2 Bottoming Cycles For Energy Recovery From Industrial Gas Turbines Exhaust Gas
Authors: Ancona M. A. , Bianchi M. , Branchini L. , De Pascale A. , Melino F. , Peretto A. , Torricelli N.
Published: September 2021

Gas turbines are often employed in the industrial field, especially for remote generation, typically required by oil and gas production and transport facilities. The huge amount of discharged heat could be profitably recovered in bottoming cycles, producing electric power to help satisfying the onerous on-site energy demand. The present work aims at systematically evaluating thermodynamic performance of ORC and supercritical CO2 energy systems as bottomer cycles of different small/medium size industrial gas turbine models, with different power rating. The Thermoflex software, providing the GT PRO gas turbine library, has been used to model the machines performance. ORC and CO2 systems specifics have been chosen in line with industrial products, experience and technological limits.

In the case of pure electric production, the results highlight that the ORC configuration shows the highest plant net electric efficiency. The average increment in the overall net electric efficiency is promising for both the configurations (7 and 11 percentage points, respectively if considering supercritical CO2 or ORC as bottoming solution). Concerning the cogenerative performance, the CO2 system exhibits at the same time higher electric efficiency and thermal efficiency, if compared to ORC system, being equal the installed topper gas turbine model. The ORC scarce performance is due to the high condensing pressure, imposed by the temperature required by the thermal user. CO2 configuration presents instead very good cogenerative performance with thermal efficiency comprehended between 35 % and 46 % and the PES value range between 10 % and 22 %. Finally, analyzing the relationship between capital cost and components size, it is estimated that the ORC configuration could introduce an economical saving with respect to the CO2 configuration.
Extension of Harmonic Balance Approach for Large-Eddy Simulation of Unsteady Flows in Cascade
Authors: Iwamoto Yuma , Teramoto Susumu , Okamoto Koji
Published: September 2021

A full scale-resolving simulation of cascades flutter is time consuming because of computational inefficiency owing to its low nondimensional frequencies. To improve the efficiency and reliability of the numerical analyses for such flows, we propose an efficient scale-resolving simulation method dedicated to time-periodic flows by extending the harmonic balance approach to a large-eddy simulation. This method combines convergence calculations of the steady-state problem based on the harmonic balance method for periodic components, and the nonlinear time-marching method for small scale turbulent fluctuations. Using the proposed method, deterministic periodic components and stochastic turbulent fluctuations are calculated simultaneously, and the effect of turbulent fluctuations on deterministic periodic components is directly calculated without using turbulence models. In this paper, we present the algorithm of the simulation technique and the progress of validation calculations for channel flow excited in the streamwise direction.
Advanced Exergy Analysis of the Supercritical Carbon Dioxide Power Cycle for Waste Heat Recovery of Gas Turbine
Authors: Li Bo , Wang Shun-sen , Song Liming
Published: September 2021

In this paper, the supercritical carbon dioxide power cycle used to recover the waste heat of gas turbine is investigated by means of conventional exergy analysis and advanced exergy analysis. Firstly, the thermodynamic parameters of carbon dioxide cycle in design stage are determined by single-objective optimization with net power output as objective function. Then, conventional exergy analysis is carried out on the partial heating cycle under real, unavoidable and ideal conditions. After that, advanced exergy analysis, in which the exergy destruction is divided into endogenous / exogenous part and avoidable / unavoidable part is adopted to reveal the improvement potential of the system and illustrate the interaction among the components. According to the calculation results, a total amount of 3.55MW (47.33%) exergy destruction could be reduced by the improvement of component efficiency. Endogenous exergy destruction is higher than exogenous exergy destruction in all components. Based on the results of conventional exergy analysis, the high-temperature heater should be paid attention in order to reduce exergy destruction. However, according to the results of advanced exergy analysis, the technical improvement of turbine should be emphasized due to its high endogenous-avoidable exergy destruction. Meanwhile, for the components with high unavoidable exergy destruction, external systems should be employed to exploit the underutilized energy and enhance the system performance.
Analysis of Swirl Number Effects on Effusion Flow Behaviour Using Time Resolved PIV
Authors: Lenzi T. , Picchi A. , Andreini A. , Facchini B.
Published: September 2021

The analysis of the interaction between the swirling and cooling flows, promoted by the liner film cooling system, is a fundamental task for the design of turbine combustion chambers since it influences different aspects such as emissions and cooling capability. In particular high turbulence values, flow instabilities, and tangential velocity components induced by the swirling flow deeply affect the behavior of effusion cooling jets, demanding for dedicated time-resolved near-wall experimental analysis. The experimental set up of this work consists of a non-reactive single-sector linear combustor test rig scaled up with respect to engine dimensions; the test section was equipped with an effusion plate with standard inclined cylindrical holes to simulate the liner cooling system. The rig was instrumented with a 2D Time-Resolved Particle Image Velocimetry system, focused on different field of views. The degree of swirl for a swirling flow is usually characterized by the swirl number, Sn, defined as the ratio of the tangential momentum flux to axial momentum flux. To assess the impact of such parameter on the near-wall effusion behavior, a set of three different axial swirlers with swirl number equal to Sn = 0.6 - 0.8 - 1.0 were designed and tested in the experimental apparatus. An analysis of the main flow field by varying the Sn was first performed in terms of average velocity, RMS, and Tu values, providing kinetic energy spectra and turbulence length scale information. In a second step, the analysis was focused on the near-wall regions: the strong effects of Sn on the coolant jets was quantified in terms of vorticity analysis and jet oscillation.
Design and Testing of a Multi-Stage IP Turbine for Future Geared Turbofans
Authors: Torre D. , García-Valdecasas G. , Puente A. , Hernández D. , Luque S.
Published: September 2021

The multi-stage intermediate pressure turbine (IPT) is a key enabler of the thermodynamic cycle in geared turbofan engine architectures, where fan and turbine rotational speeds become decoupled by employing a power gearbox between them. This allows for the separate aerodynamic optimization of both components, an increase in engine bypass ratios, higher propulsive efficiency, and lower specific fuel consumption. Due to significant aerodynamic differences with conventional low pressure turbines (LPTs), multi-stage IPT designs present new aerodynamic, mechanical and acoustic trade-offs. This work describes the aerodynamic design and experimental validation of a fully featured three-stage IP turbine, including a final row of outlet guide vanes.

Experiments have been conducted in a highly engine-representative transonic rotating wind tunnel at the CTA (Centro de Tecnologías Aeronáuticas, Spain), in which Mach and Reynolds numbers were matched to engine conditions. The design intent is shown to be fully validated. Efficiency levels are discussed in the context of a previous state-of-the-art LPT, tested in the same facility. It is argued that the efficiency gains of IPTs are due to higher pitch-to-chord ratios, which lead to a reduction in overall profile losses, and higher velocity ratios and lower turning angles, which reduce airfoil secondary flows and three-dimensional losses.
A Novel Airfoil Design Method and its Application to the Optimization Design of a 3-Stage Transonic Compressor
Authors: Wang Xiaochen , Ren Xiaodong , Li Xuesong , Wu Hong , Gu Chunwei
Published: September 2021

In this paper, a novel airfoil design method is proposed by introducing multiple physical characteristics to the process of parameterization. The new design method is to operate the camber and thickness distributions in a way that the chordwise mapping information is adjusted without using local geometrical parameters. A proper design space is expected to be constructed, in which the laws of optimization can be simply transmitted from a single airfoil to the whole blade and even stages. The capability of the method is firstly validated in the Pareto optimization of a 2D compressor airfoil, and then specialized for the optimization of a rotor in a 3-stage transonic compressor. The 3-stage transonic compressor studied in this paper is representative of the inlet stages of an industrial gas turbine, and the shock structure of the third rotor (R3) presents the peak strength near the mid span. The optimized design reduces the peak Mach number on the suction surface, and increases the efficiency of the third stage and the whole compressor by nearly 1% and 0.3% respectively. The mechanism of the loss optimization is analyzed in detail.