bluuebunny/crossref_metadata_2025_split huggingface dataset:
High-Pressure Laminar Flame Speeds and Markstein Lengths of Syngas Flames Diluted in Carbon Dioxide and Helium
Authors: Turner Mattias A. , Petersen Eric L. 
Published: October 2022

Abstract
New laminar flame speed and burned-gas Markstein length data for H2-CO-O2-CO2-He mixtures have been measured from spherically expanding flames. Experiments were conducted at 10 atm and room temperature for H2:CO ratios ranging from 2:1 to 1:4 and for overall CO2 mole fractions from 0 to 30%. CO2 dilution had little effect on Markstein length, but CO2 dilutions of 10%, 20%, and 30% caused average reductions in flame speed of 47%, 73%, and 89% respectively, regardless of H2:CO ratio. The study was designed to isolate the dilution effect of CO2 on flame speed, and a detailed analysis using the FCO2 method was used to show that the chemical-kinetic participation of CO2 was responsible for up to 20% of the reduction in flame speed. Hence, the majority (80% or more) of the reduction in flame speed due to CO2 is from the thermal effect. Accurate flame speed predictions were produced by five different chemical kinetics mechanisms for most conditions, with the slight exception of high-CO, high-CO2 mixtures. A thorough sensitivity analysis highlighted the larger effect of CO2 dilution on the important kinetics reactions than the effect of changing H2:CO. Sensitivity analysis also showed that the chain branching reaction H2O+O⇌OH+OH could be modified (albeit beyond its uncertainty) to achieve more accurate flame speed predictions, but also indicated that further improvement of flame speed modeling would require changes to many lesser reactions.
Integrated Aerodynamic and Structural Blade Shape Optimisation of Axial Turbines Operating With Supercritical Carbon Dioxide Blended With Dopants
Authors: Abdeldayem Abdelrahman S. , White Martin T. , Paggini Andrea , Ruggiero Marco , Sayma Abdulnaser I. 
Published: October 2022

Abstract
Within this study, the blade shape of a large-scale axial turbine operating with sCO2 blended with dopants is optimised using an integrated aerodynamic-structural 3D numerical model, whereby the optimisation aims at maximising the aerodynamic efficiency whilst meeting a set of stress constraints to ensure safe operation. Specifically, three candidate mixtures are considered, namely CO2 blended with titaniumtetrachloride (TiCl4), hexafluorobenzene (C6F6) or sulfur dioxide (SO2), where the selected blends and boundary conditions are defined by the EU project, SCARABEUS. A single passage axial turbine numerical model is setup and applied to the first stage of a large-scale multi-stage axial turbine design. The aerodynamic performance is simulated using a 3D steady-state viscous computational fluid dynamic (CFD) model while the blade stress distribution is obtained from a static structural finite element analysis (FEA). A genetic algorithm is used to optimise parameters defining the blade angle and thickness distributions along the chord line while a surrogate model is used to provide fast and reliable model predictions during optimisation using genetic aggregation response surface. The uncertainty of the surrogate model represented by the difference between the surrogate model results and the CFD/FEA model results is evaluated using a set of verification points and found to be less than 0.3% for aerodynamic efficiency and 1% for both the mass flow rate and the maximum equivalent stresses. The comparison between the final optimised blade cross-sections have shown some common trends in optimising the blade design by decreasing stator and rotor trailing edge thickness, increasing stator thickness near the trailing edge, decreasing rotor thickness near the trailing edge and decreasing the rotor outlet angle. Further investigations of the loss breakdown of the optimised and reference blade designs are presented to highlight the role of the optimisation process in reducing aerodynamic losses. It has been noted that the performance improvement achieved through shape optimisation is mainly due to decreasing the endwall losses of both stator and rotor blades.
Dynamic Modelling and Simulation of a 100 kW Micro Gas Turbine Running With Blended Methane/Hydrogen Fuel
Authors: Banihabib Reyhaneh , Assadi Mohsen 
Published: October 2022

Abstract
The current shift from centralized to decentralized power generation with renewables as prime movers necessitates the integration of reliable small-scale power supply units to compensate for the intermittency of renewables. Micro gas turbines’ (MGTs) characteristics such as high reliability and low maintenance, along with flexible operation and quick load-following capabilities have made them a dependable source for the modern power generation industry and for households. MGTs are small-scale gas turbine units with a power range lower than 500 kW that can operate with low-calorific fuels such as biofuels and syngas as well as conventional fossil fuels and zero-carbon fuels.
The utilization of MGTs in innovative cycle layouts or varying types of feeding fuels is increasing, which requires the evaluation of system dynamics to ensure the safe operation of the engine and its components. Moreover, the role of MGTs as a backup for the intermittent renewable inputs means that they operate under more transient conditions rather than constant power production mode. Therefore, a reliable dynamic model of an MGT is required to investigate the dynamic response of the engine under various transient modes to ensure safe operation. Moreover, utilizing a dynamic model is vital in the designing process of MGT-based cycles in order to evaluate the behaviour of coupled components in off-design conditions and to optimize the controller parameters. To that end, developing a dynamic model of the MGT cycle that is accurate enough to predict the dynamic response of the engine and its components and fast enough to be utilized in design iterations is necessary.
In this paper, a high-fidelity model for real-time simulation of an MGT, based on a lumped and nonlinear representation of gas turbine components is presented. The model for a recuperated T100 MGT was constructed in Simscape, the object-oriented environment of MATLAB for modelling physical systems. MGT components were modelled as lumped volumes with dynamic equations of mass, momentum, and energy balance along with component-characteristic maps describing the evolution of the flow passing through them. Results from simulations were validated by experimental data collected from a real engine operating under different load conditions. Experimental tests and numerical simulations were conducted for pure methane as well as for blended methane/hydrogen as feeding fuels.
Transient Analysis of a Micro Gas Turbine With Fuel Composition Change
Authors: Raggio Martina , Bellotti Daria , Ferrari Mario Luigi 
Published: October 2022

Abstract
The aim of this study is to analyze the effect of fuel composition change on the performance and dynamic behavior of a T100 micro gas turbine. Fuel flexibility is a key asset of micro gas turbines and can support the energy transition when alternative fuels obtained from renewable energy sources are used. While the current research is mainly focusing on improving combustor performance, it is also required to investigate the effects of different fuel blends on the overall system. During transient operation, the use of fuels with a low LHV requires an increased fuel mass flow that can potentially reduce the compressor surge margin. Conversely, sudden composition changes of high LHV fuel can cause temperature peaks, detrimental for the turbine and recuperator life.
In this paper, a validated transient model for the T100 machine has been used to simulate the injection of different fuels in the natural gas line (feeding the burner). Injection of hydrogen and ammonia, two promising carbon-free fuels, have been tested. A transient analysis was performed with this tool to monitor the main performance parameters with the aim to (i) verify compressor safe operations during different working conditions and (ii) to identify appropriate fuel composition change schedules to ensure Turbine Outlet Temperature values within an acceptable range, while keeping the original fuel control system.
Assessment of a Conjugate Heat Transfer Method on an Effusion Cooled Combustor Operated With a Swirl Stabilized Partially Premixed Flame
Authors: Amerini A. , Paccati S. , Mazzei L. , Andreini A. 
Published: October 2022

Abstract
Computational fluid dynamics play a crucial role in the design of cooling systems in gas turbine combustors due to the difficulties and costs related to experimental measurements performed in pressurized reactive environments. Despite the massive advances in computational resources in the last years, reactive unsteady and multi-scale simulations of combustor real operating conditions are still computationally expensive.
Modern combustors often employ cooling schemes based on effusion technique, which provides uniform protection of the liner from hot gases, combining the heat removal by means of heat sink effect with liner coverage and protection by film cooling. However, a large number of effusion holes results in a relevant increase of computational resources required to perform a CFD simulation capable of correctly predicting the thermal load on the metal walls within the combustor. Moreover, a multi-physics and multi-scale approach is mandatory to properly consider the different characteristic scales of the several heat transfer modes within combustion chambers to achieve a reliable prediction of aero-thermal fields within the combustor and wall heat fluxes and temperatures. From this point of view, loosely-coupled approaches permit a strong reduction of the calculation time, since each physics is solved through a dedicated solver optimized according to the considered heat transfer mechanism. The object of this work is to highlight the capabilities of a loosely-coupled unsteady multi-physics tool (U-THERM3D) developed at the University of Florence within ANSYS Fluent.
The coupling strategy will be employed for the numerical analysis of the TECFLAM effusion cooled swirl burner, an academic test rig well representative of the working conditions of a partially premixed combustion chamber equipped with an effusion cooling system, developed by the collaboration of the Universities of Darmstadt, Heidelberg, Karlsruhe, and the DLR. The highly detailed numerical results obtained from the unsteady multi-physics and multi-scale simulation will be compared with experimental data to validate the numerical procedure.
Validation of Steady-State Aerodynamics in a Transonic Linear Cascade at Near Stall Conditions
Authors: Tavera Guerrero Carlos , Glodic Nenad , Groth Pieter 
Published: October 2022

Abstract
A new set of steady-state aerodynamics data from the recently build KTH transonic linear cascade (TLC) is presented. The operational point is representative of an open source virtual compressor (VINK) operating near stall at 70% speed line. The inlet Mach number (M = 0.81) is above its airfoil critical value with high incidence angle, producing a leading edge separation bubble. The test object in the TLC has a 2% tip gap, and the reference measurement plane is at 85% span. Results show repeatability, and periodicity. Upstream flow field was mapped by a non-intrusive, Laser-2-Focus (L2F) method. The blade loading was recovered by a set of pressure taps. Flow field properties downstream the test section were recovered by a five-hole probe. A flow structures comparison was performed between numerical and experimental data. Numerical results were obtained from the commercial software Ansys CFX, where different turbulence models were tested: SST, SST with reattachment method (RM) and SST with turbulent transition (γ – θ). The closest match to the experimental data was recovered by SST-γ – θ RM. Separation onset and flow properties downstream are strongly affected by the selected turbulence transition model. The results are aimed to work as step towards validation for CFD in separated regions. These data is the first of a series of test campaigns for experimental validation for aerodynamic damping in separated flows.
Recrystallization of René N4 and N5
Authors: van Esch Hans , Pietersen Stijn , Bridges Alex , Scheibel John , Zuber Robert , Greaves Wayne 
Published: October 2022

Abstract
Recrystallization of equiaxed, directionally solidified and single crystal cast nickel base superalloys can occur during some of the manufacturing and repair processes. While equiaxed, and to a lesser extent, directionally solidified cast nickel-based superalloys have grain boundary strengthening this is nonexistent for single crystal superalloys. Therefore, it is especially important to avoid or limit recrystallization during manufacture and repair of these nickel base single crystal gas turbine components.
This study performed by TEServices, EPRI, MD&A and Sulzer concerns the effect of recrystallization when compressive stresses are introduced during manufacturing and repair processes (grit cleaning and shot peening) when followed by (full solution, partial solution and stress relieve) heat treatments. The objective of this paper is to obtain a better understanding the recrystallization behavior of two single crystal superalloys, René N4 and N5, during the repair processes (grit cleaning, shot peening and heat treatments) and determine how to minimize the depth of recrystallization. The impact of recrystallization on mechanical strength (stress rupture and fatigue) on René N4 material was limitedly tested.
This study concludes that for both single crystal alloys, René N4 and N5, the less compressive stresses that are introduced during manufacturing and repair, and the lower temperature of exposure during follow up heat treatments, results in a lower recrystallization layer depth.
Experimental and Numerical Study on Effusion Cooling Configuration for the Swirl CMC Combustor Platform
Authors: Du Kun , Chen Qihao , Liang Tingrui , Jia Yihao , Liu Cunliang 
Published: October 2022

Abstract
With advantageous thermal, strength, and weight properties at high temperatures, Ceramic Matrix Composite (CMC) is regarded as the potential structural material for modern aeroengines with the extreme increase of operating temperature. However, the anisotropic thermal conductivities caused by the weaving type call for a novel cooling structure design since it brings a completely different thermal conduction performance inside the hot components compared to the superalloy. The infrared thermographic experiment was carried out based on the SiC/SiC composite platform prepared by a 2-D plain weave braid structure with the CVI process with three staggered effusion hole configurations in this paper to explore the cooling characteristic of the CMC. The surface overall cooling effectiveness (ϕ) of the platform was measured under a temperature ratio of 1.5 (Tg/Tc = 1.5) and seven mass flow ratios. And a separate numerical simulation with different swirl effects was conducted as a supplement for the combustor operation condition. The results indicate that the thermal conductivity along the thickness direction is of great importance for the platform cooling. The overall cooling effectiveness of the CMC platform was smaller in comparison with the superalloy platform because of its smaller through-thickness thermal conductivity. Moreover, the concentrated coolant outflow is beneficial for the CMC cooling performance given that the area-averaged overall cooling effectiveness disparity of the two materials was narrowed with the smaller hole spacing. The swirling flow strongly inhibits the outflow of coolant near the swirling core but expands the lateral coverage of the film simultaneously. Such behavior alters the film cooling potential as well as the cooling pattern. The difference between CMC and superalloy was increased with the low swirl situation and decreased with the high swirl situation.
Clearance Sensitivity Mitigation in Small Core Compressors
Authors: Evans Simon , Nolan Sean , Botros Barbara , Port Bianca , Chin Peter , Kulkarni Sameer , Stephens Julia 
Published: October 2022

Abstract
As future aircraft become lighter and more aerodynamically efficient, thrust requirements will decrease, reducing the core size of the engine. Furthermore, in the pursuit of improved fuel burn, engine overall pressure ratio and bypass ratio will increase, further driving down engine core size. These drivers together mean that the core size, or corrected mass flow rate at the compressor exit, for future single-aisle aircraft applications will shrink below 3.0 lbm/s. Traditionally, this small core compressor size is in the domain of axi-centrifugal designs, machines that are typically less efficient and limited to pressure ratios of ∼25 due to stress and thermomechanical fatigue in the centrifugal impeller. In this light, NASA, Pratt & Whitney and the Raytheon Technologies Research Center embarked upon a program to develop technologies to enable an all-axial high-pressure compressor with a core size below 3.0 lbm/s and an overall pressure ratio greater than 50.
One of the challenges with an all-axial high-pressure compressor at this core size is the small span at the rear of the compressor. As core size is scaled down, the rotor tip clearances do not scale with span, leading to significant efficiency penalties. This paper documents a numerical and experimental effort to mitigate this penalty through design optimization and technology insertion, enabling an estimated 5 to 10% fuel burn reduction relative to 2020 best-in-class. The experimental program consisted of two builds of a high-speed rig: a baseline build and a second build to insert technology to manage large rotor tip gap/span ratios. The results demonstrated a reduction in the sensitivity of the compressor to large clearance ratios, which would be likely at the end-of-life of a small core application. The test campaign demonstrated that, with the insertion of technology, small core compressors can maintain a similar efficiency to current best-in-class large core size compressors.
In addition to advancing the state-of-the-art of technology, the program also advanced the modeling standards for multistage compressors with large clearance-to-span ratios. The validation of this approach is described in the paper.
Approximating Gas Turbine Combustor Exit Temperature Distribution Factors Using Spatially Under-Sampled Measurements
Authors: Lou Fangyuan 
Published: October 2022

Abstract
In gas turbine engines, the temperature field at the exit plane of the combustor is highly unsteady and complex, with large radial and circumferential variations. The circumferential temperature variations (hot-streaks) are caused by the discrete nature of fuel and dilution air jets, and combustor lining coolant flow results in a strong radial temperature gradient. Two widely adopted parameters include the radial temperature distribution factor (RTDF) and the overall temperature distribution factor (OTDF) to quantify combustor exit temperature nonuniformity. The state-of-the-art approach of characterizing the combustor exit temperature nonuniformity is to carry out an extensive traverse along the circumferential direction at a spatial resolution of two to three degrees. With hundreds of measurements in place, the combustor OTDF and RTDF can be obtained. Though the approach is practical, this involves the design of complex traverse mechanisms and can be costly. To address this challenge, this paper presents a novel method for predicting combustor exit TDFs using much sparser measurements. The approach’s effectiveness was examined using three engine representative combustor exit temperature measurements covering a single-burner, double-burner, and entire annulus. In all cases, the hot-streaks-related features are well resolved in the reconstructed temperature field, and the multi-wavelet approximation method yields almost identical RTDF to experiment with majority variations of less than 1.0% using sparse measurements. In addition, the method’s robustness was examined, and considerations for the implementation of the method were provided.
Gas Turbine Transition Duct Design Considerations
Authors: Satheesh Aparna , Manoharan S. , Soundiramourty Sendil , Babu S. 
Published: October 2022

Abstract
The gas turbine twin shaft variant requires an aerodynamic coupling between its high-pressure gas generator module and low-pressure power module. These two modules are spaced apart for aerodynamic and mechanical reasons. Typically, this is achieved by having a transition duct (TD) in between these two modules which smoothly allows the gas flow from HP turbine to LP turbine. To fulfill this purpose, it includes a radial inner shell and a radial outer shell defining a flow passage and radial struts which connects inner shell and outer shell. Cross-section of these struts are airfoil shaped to facilitate smooth flow of flue gas through TD. In addition, TD allows flow passage for cooling medium, support seals to prevent leakage at various interface locations and holds thermo-well for temperature measurement. Considering thermal and mechanical loads on the TD for standard operating conditions, Low cycle fatigue, creep and oxidation calculations are performed to ensure durability of TD design and necessary design changes are made to fulfill product requirement. Suitable material is selected, and manufacturing method finalized based on the design analysis. While standard offerings of gas turbine fulfill majority of requirements, there are some off-design or customized requirements which require detailed study. To fulfill these types of customized requirements, standard offerings of gas turbine are further analyzed for changed thermal boundary condition. These types of customizations facilitate running gas turbine with different options with a clear understanding of how it impacts durability. Through these studies, necessary maintenance factors are established to arrive at revised maintenance intervals based on severity of change. Low cycle fatigue, creep and oxidation calculations are performed with changed boundary condition also and maintenance factor is determined. This paper is intended to describe the steps followed in determining durability, establish maintenance factor and provide service recommendations.
Influence From Nozzle Guide Vane Wakes and Inlet End-Wall Boundary Layers on Turbine Rear Structure Aerodynamics
Authors: Nylander Pär , Deshpande Srikanth , Larsson Jonas 
Published: October 2022

Abstract
CFD results can be improved by imposing accurate inlet boundary conditions. A previous paper presented extensive measurements from an engine realistic Turbine Rear Structure (TRS), complemented with CFD results using normal radial profiles (1D) at the inlet [10]. This paper extends stat study with two separate studies, investigating the effect of the upstream Nozzle Guide Vane (NGV) wakes and the inlet end-wall boundary layers. In the first study, simulations are done using full 2D inlet boundary condition, imposing the NGV wakes on the inlet. The second study investigates different inlet end-wall boundary layers. Comparisons with measurements show for some aspects that the upstream Nozzle Guide Vane (NGV) wakes and the inlet end-wall boundary layers are important.
Predicting upstream forcing from the TRS on the LPT rotor requires a 2D inlet boundary condition. Also, a strong interaction between the incoming NGV wakes and the secondary flow loss-regions at the outlet is found. However, flow quantities like blade loading, outlet swirl, and OGV wakes are well predicted using regular 1D radial inlet profiles.
How the inlet end-wall boundary layers are modelled have significant impact on secondary flows and outlet swirl. If the full boundary layers are prescribed, the secondary flows are over-predicted. This gives under turning and premature separations. With no inlet boundary layers, secondary flows are well captured and gives better separations, improved outlet swirl and wake predictions. The recommendation is therefore to remove the inlet boundary layers when using the Transition SST k-ω γ-Reθ model in TRS CFD simulations.