Future Gas Turbine Research Articles

Rotordynamics of a Single-Stage Brush Seal in Isolation: The Effects of Variable Stiffness and Back Plate Geometry

Abstract Brush seals control leakage around rotating components from areas of high to low pressure inside turbomachinery. They are known to contribute to the overall stability of gas turbines, therefore their dynamic behavior is of particular importance to engine designers. Despite this, limited research exists in the literature on the rotordynamic behavior of brush seals. This paper aims to experimentally characterize the leakage and rotordynamic performance of two seals with different bristle diameters tested with both conventional and pressure-relieved back plates with a slight interference. A dynamic test facility was utilized to study the dynamic characteristics of an isolated seal with changes in excitation frequency, rotational speed, and pressure drop. Seal leakage increased with bristle diameter and with the use of the pressure-relieved back plate but reduced with increasing rotational speed for all tests. The direct dynamic coefficients were shown to increase with pressure difference. The back plate geometry influenced the change in stiffness coefficient with rotational speed. The larger bristle diameter resulted in a stiffer seal, however, the damping coefficient reduced with the reduction in packing density. The insight provided by these results will help inform engine manufacturers on the suitability of implementing brush seals in future gas turbine designs.

Geometric and Flow Characterization of Additively Manufactured Turbine Blades With Drilled Film Cooling Holes

Abstract As designers investigate new cooling technologies to advance future gas turbine engines, manufacturing methods that are fast and accurate are needed. Additive manufacturing facilitates the rapid prototyping of parts at a cost lower than conventional casting but is challenged in accurately reproducing small features such as turbulators, pin fins, and film cooling holes. This study explores the potential application of additive manufacturing and advanced hole drill methods as tools to investigate cooling technologies for future turbine blade designs. Data from computed tomography scans are used to nondestructively evaluate each of the cooling features in the blade. The resulting flow performance of these parts is further related to the manufacturing through benchtop flow testing. Results show that while total blade flow is consistent for all additively manufactured cooled blades, flow through smaller regions of the blades shows variations. Shaped film cooling holes manufactured using a high-speed electrical discharge machining method are within tolerance in the metering section but do not expand at the specified angle in the diffuser even though design tolerances are met. In contrast to high-speed EDM, conventional EDM holes are undersized throughout the length of the hole. Due to the additive manufacturing process, the surface roughness was higher on the additively manufactured parts in the current study than has been previously reported for surface roughness of commonly used cast components. The roughness results show high levels on thin walls, particularly at the trailing edge as well as on downskin surfaces. Internal surface roughness is higher than external roughness at most locations on the blade. The results of this study confirm that additive manufacturing along with advanced hole drilling techniques offers faster development of blade cooling designs.

Design of a Thermal Performance Test Equipment for a High-Temperature and High-Pressure Heat Exchanger in an Aero-Engine

For next-generation power systems, particularly aero-gas turbine engines, ultra-light and highly efficient heat exchangers are considered key enabling technologies for realizing advanced cycles. Consequently, the development of efficient and accurate aero-engine heat exchanger test equipment is essential to support future gas turbine heat exchanger advancements. This paper presents the development of a high-pressure and high-temperature (HPHT) heat exchanger test facility designed for aero-engine heat exchangers. The maximum temperature and pressure of the test facility were configured to simulate the conditions of the last-stage compressor of a large civil engine, specifically 1000 K and 5.5 MPa. These conditions were achieved using multiple electric heater systems in conjunction with an air compression system consisting of three turbo compressor units and a reciprocating compressor unit. A commissioning test was conducted using a compact tubular heat exchanger, and the results indicate that the test facility operates stably and that the measured data closely align with the predicted performance of the heat exchanger. A commissioning test of the tubular heat exchanger showed a thermal imbalance of 1.02% between the high-pressure (HP) and low-pressure (LP) lines. This level of imbalance is consistent with the ISO standard uncertainty of ±2.3% for heat dissipation. In addition, CFD simulation results indicated an average deviation of approximately 1.4% in the low-pressure outlet temperature. The close alignment between experimental and CFD results confirms the theoretical reliability of the test bench. The HPHT thermal performance test facility will be expected to serve as a critical test bed for evaluating heat exchangers for current and future gas turbine applications.

Lifing Assessment of Gas Turbine Blade Root Affected by Out-of-Tolerances.

Current and future heavy-duty gas turbines (GTs) are being developed as an alternative or support to renewable energy sources (RESs). Therefore, GTs are subjected to several instances of being switched on and off; thus, the material fatigue limit can be reached in a short time. In such a scenario, possible out-of-tolerances (OoTs) in critical components must be considered. In this paper, OoTs related to critical parameters in the attachment geometry of a rotor blade are considered to estimate their impact on component life through a 2D finite element (FE) analysis. First, a mesh refinement is performed to obtain mesh-independent results; second, OoT geometries are simulated to determine stresses and strains at the blade attachment and disc groove. The mesh refinement process is critical to ensure model accuracy for both nominal and OoT geometries. The results show that OoTs can lead to important reductions in the life of the intended component, both on the blade and disc sides. These results could be useful in updating the maintenance plan for components and could be used for future insights, with further work extending the study to 3D geometry, for example, and evaluating the effect of other geometries.

Nonpremixed Approaches for Fuel Flexible, Low NOx Combustors in High-Efficiency Gas Turbines

Abstract Lean, premixed combustor designs have almost completely replaced non-premixed combustors for industrial gas turbine applications where NOx emissions are regulated. Nonetheless, these designs have also introduced turndown and fuel flexibility constraints and made combustion instabilities and flashback more problematic. Future gas turbine applications will require combustors to accommodate a range of alternative fuels, provide operational flexibility, and compete for services with a host of new technologies, including energy storage and fuel cells. The purpose of this paper is to propose nonpremixed, multi-stage designs for the next generation of high turndown, high fuel flexibility, low NOx combustion designs – referred to here as a Nonpremixed, Rich, Relaxation, Lean (NRRL) combustor. The key concept we explore is non-premixed combustion, followed by additional fuel mixing to locally fuel-rich conditions, a relaxation stage, and then a lean stage. This non-premixed approach can handle essentially any fuel composition, including pure hydrogen, liquid fuels, pure methane, pure ammonia, and any combination in between while breaking the NOx-CO tradeoff and reducing combustion instability risk. This paper provides chemical reactor network modeling calculations to identify key kinetic processes and time scales required for such a concept. This concept has completely inverted sensitivities from lean, premixed systems which prefer short residence times, low pressures, and low temperatures to minimize NO formation. This concept prefers long residence times, high pressures, and high temperatures, indicating a very different set of design trades for part load and off-design performance.

Reaction of a Molten Cr-Si-Base Alloy with Ceramics and a High Entropy Oxide

Due to their higher thermal and chemical stability than other high-temperature materials, chromium-silicon-base (Cr-Si-base) alloys are promising materials for future gas turbines and other high-temperature applications operating under harsh conditions. To enable near-net-shape casting of Cr-Si-base alloys, a compatibility of the alloy melt with the ceramic crucibles and molds is necessary. Additionally, a metal-ceramic contact exists at the interface between thermal barrier coating (TBC) and alloy, where metallic may melts play a role in the case of coating failure and overheating. In this study, molten Cr92Si8 (in at. %) alloy is brought into contact with powders of ceramics commonly used for casting molds or crucibles (e.g. ZrSiO4, Al2O3, 3YSZ), to investigate liquid metal corrosion, interdiffusion, and stabilities. Additionally, the high entropy oxide (Sm0.2Gd0.2Dy0.2Er0.2Yb0.2)2Zr2O7 (HEO), a potential future TBC material, is investigated. Before melting using an electric arc furnace, the powders of the investigated ceramics were mixed with pulverized Cr92Si8 and pressed into alloy-ceramic pairs, to maximize the contact area between molten metal and ceramic. For microstructural investigations and phase analysis, the materials were assessed using scanning electron microscopy (SEM) equipped with energy-dispersive X-ray spectroscopy (EDS), and X-ray diffraction (XRD). The widely used mold material ZrSiO4 and the coating BN were found to decompose, while reaction products of SiO2 and CoAl2O4 with the melt were detected. Al2O3, 3YSZ, and the HEO did not show decomposition or corrosion by the melt. Al2O3, 3YSZ, and the HEO are therefore considered as promising crucible, mold, and TBC materials for Cr-Si-base alloys.

Development of High Temperature Materials for Aero Engine Hot End Components: An Overview

The present paper overviews the development of materials and superalloys for aero engine hot end components to meet the increasing trend of turbine inlet temperature. Requirement of higher and higher thrust is steadily increasing the turbine inlet temperature and the development of nickel-chromium superalloys in early 1940s could not fulfil the material capability for long. Various nickel-base superalloys developed in 1950s and 1960s could increase the life of hot end components by retaining strength and resisting oxidation at extreme temperatures. In the 1960s and 1970s, with almost stagnation in high temperature alloy development, metallurgists changed focus from alloy chemistry to alloy processing which evolved the directional solidification and single crystal casting technologies. At present, almost all fighter class engines and high bypass commercial engines are using nickel and cobalt base superalloys for hot end components and single crystal superalloys particularly for turbine blades. This paper covers the developmental phases of superalloys and casting technologies for engine hot end components. This paper will be an invaluable asset for the researchers as well as for designers of future gas turbine engines.

Advanced Gas Turbine Cooling for the Carbon-Neutral Era

In the coming carbon-neutral era, industrial gas turbines (GT) will continue to play an important role as energy conversion equipment with high thermal efficiency and as stabilizers of the electric power grid. Because of the transition to a clean fuel, such as hydrogen or ammonia, the main modifications will lie with the combustor. It can be expected that small and medium-sized gas turbines will burn fewer inferior fuels, and the scope of cogeneration activities they are used for will be expanded. Industrial gas turbine cycles including CCGT appropriate for the carbon-neutral era are surveyed from the viewpoint of thermodynamics. The use of clean fuels and carbon capture and storage (CCS) will inevitably increase the unit cost of power generation. Therefore, the first objective is to present thermodynamic cycles that fulfil these requirements, as well as their verification tests. One conclusion is that it is necessary to realize the oxy-fuel cycle as a method to utilize carbon-heavy fuels and biomass and not generate NOx from hydrogen combustion at high temperatures. The second objective of the authors is to show the required morphology of the cooling structures in airfoils, which enable industrial gas turbines with a higher efficiency. In order to achieve this, a survey of the historical development of the existing cooling methods is presented first. CastCool® and wafer and diffusion bonding blades are discussed as turbine cooling technologies applicable to future GTs. Based on these, new designs already under development are shown. Most of the impetus comes from the development of aviation airfoils, which can be more readily applied to industrial gas turbines because the operation will become more similar. Double-wall cooling (DWC) blades can be considered for these future industrial gas turbines. It will be possible in the near future to fabricate the DWC structures desired by turbine cooling designers using additive manufacturing (AM). Another conclusion is that additively manufactured DWC is the best cooling technique for these future gas turbines. However, at present, research in this field and the data generated are scattered, and it is not yet possible for heat transfer designers to fabricate cooling structures with the desired accuracy.

Review of Advanced Effusive Cooling for Gas Turbine Blades

Turbine inlet temperature has continuously increased to improve gas turbine performance during the past few decades. Although internal convection cooling and traditional film cooling have contributed significantly to the current achievement, advanced cooling schemes are needed to minimize the coolant consumption and maximize the cooling efficiency for future gas turbines. This paper conducts a comprehensive review of advanced effusive cooling schemes for gas turbine blades. First, the background and the history of turbine blade cooling are introduced. Then, the metrics of effusive cooling efficiency are defined. Next, effusion cooling, impingement/effusion cooling, and transpiration cooling are reviewed. The flow and heat transfer mechanisms of the cooling schemes are emphasized, and the design trends of the cooling schemes are revealed. Finally, the conclusions and future research perspectives are summarized.

Development of an Open-Source Autonomous Computational Fluid Dynamics Meta-Modeling Environment for Small-Scale Combustor Optimization

Abstract This work presents an open-source autonomous computational fluid dynamics (CFD) metamodeling environment (OpenACME) for small-scale combustor design optimization in a deterministic and continuous design space. OpenACME couples several object-oriented programing open-source codes for conjugate-heat transfer, steady-state, multiphase incompressible Reynolds averaged Navier-Stokes CFD-assisted engineering design metamodeling. There are fifteen design variables. Nonparametric rank regression (NPRR), global sensitivity analyses (GSA), and single-objective (SOO) optimization strategies are evaluated. The Euclidean distance (single-objective criterion) between a design point and the utopic point is based on the multi-objective criteria: combustion efficiency (η) maximization and pattern factor (PF), critical liner area factor (Acritical ), and total pressure loss (TPL) minimization. The SOO approach conducts Latin hypercube sampling (LHS) for reacting flow CFD for subsequent local constraint optimization by linear interpolation. The local optimization successfully improves the initial design condition. The SOO approach is useful for guiding the design and development of future gas turbine combustors. NPRR and GSA indicate that there are no leading-order design variables controlling η, pattern factor (PF), Acritical , and TPL. Therefore, interactions between design variables control these output metrics because the output design space is inherently nonsmooth and nonlinear. In summary, OpenACME is developed and demonstrated to be a viable tool for combustor design metamodeling and optimization studies.

Creep detection of Hastelloy X material for gas turbine components with nonlinear ultrasonic frequency modulation

Creep damage is one of the main failure modes in hot-gas‑leading components in gas turbines, which results from high temperatures along with mechanical loads. The aim of this study is to clarify the metallurgical creep behaviour of the Hastelloy X material and detect and evaluate creep damage at an early stage with a nonlinear ultrasonic modulation technique. For this purpose, multiple samples were examined to demonstrate that pores and microcracks in grain boundaries spread from the outside to the inside. Inside the specimen, molybdenum was identified as the main precipitation element. In addition, the chromium diffusion in the outer areas led to the depletion of this element and favoured the formation of pores and microcracks. Failures were proven with nonlinear dual-frequency ultrasound technology. Moreover, two different longitudinal waves were sent into the samples to use harmonic and modulated response frequencies for evaluation. As a result, harmonic frequencies offered a favourable prediction of pore sizes, whereas defined sideband frequencies reacted very sensitively to the damage density and area distribution of the failures. This study offers a method for detecting creep damage with nonlinear ultrasound techniques at an early stage as well as for differentiating between pores, microcracks, dislocations and precipitation. Therefore, the design of future gas turbine components made of Hastelloy X can be adapted with regard to the shown metallurgical behaviour and damage signatures.

Unsteady adiabatic film cooling effectiveness behind shaped holes

Gas turbine as a vital component requires higher inlet temperature to improve its thermal efficiency, and thus a thin coolant film spreads over the blade surface to isolate it from hot gases. However, such coolant intensively interacts with the mainstream, leading to highly unsteady coolant coverage, which damages structural integrity, challenging its reliable and sustainable operations. Consequently, fast-response pressure-sensitive paint technique (fast-PSP) was used to measure the coolant unsteadiness behind the shaped holes, considering plenum and crossflow feeds. A steady-RANS simulation was performed to predict the large-scale flow structures and explain the mean results. The measured standard deviation (SD) and fluctuating components were used to uncover the spatial-temporal features associated with the predicted vortical structures. The mean effectiveness was dramatically influenced by the blowing ratios (M), showing attached flow at a low M and lift-off at a high M. The internal crossflow showed asymmetric spreading, resulting in deteriorated performance behind the holes. The unsteadiness was highly influenced by the energetic vortical structures, which originated from the hole entrance, forming in-hole counter-rotating vortex pair (CRVP) and vortex-tube structures. Meanwhile, the vortex-tube formed a swirl-vortex signature downstream of the hole-trailing-edge, leading to asymmetric CRVP. The swirl-vortex interacted with the CRVP windward-leg and reduced the spreading behind the holes, leading to mainstream ingestions. The combined results suggested considering the internal flow effects, which could help the designers understand the characteristics of unsteady effectiveness; promoting advanced cooling strategies for enhanced protection of future gas turbines.

Use of Bayesian statistics to calculate transient heat fluxes on compressor disks

Future gas turbine engines require improved understanding of the heat transfer between compressor disks and air in compressor cavities under transient operating conditions. Calculation of transient heat fluxes from temperature measurements on compressor disks is a typical ill-posed inverse problem where small uncertainties of measurements can lead to large uncertainties of the calculated fluxes. This paper develops a Bayesian model for the heat flux to reduce the adverse nature of the problem by using a Gaussian prior distribution with Matérn covariance. To efficiently find the maximum a posterior, a neural network was used to solve the heat equation for compressor disks for any choice of parameters, allowing fast evaluation of the solution to the forward model for any heat flux of interest. The power of the Bayesian model is first demonstrated using numerically simulated data. Subsequently, the model is used to calculate heat fluxes from measurements of transient temperature collected from the Compressor Cavity Rig at the University of Bath. During these transient tests, the periphery of the rotating compressor disk was initially heated to a steady-state condition and then cooled rapidly by the ambient air. The fluxes for four transient cycles were calculated, with the operating range of 7.0×105<Reϕ<2.8×106, 0.0<βΔT<0.15, and 0.13<χ, where Reϕ,βΔT, and χ are the rotational Reynolds number, the buoyancy parameter, and the compressibility parameter, respectively. The results show that, for all four cases, the flow and the heat transfer in the closed cavity were initially dominated by buoyancy effects, with heat transferred from the disk to the cavity air in the outer region and reversed in the inner region. The initial heat fluxes at Reϕ=2.1×106 were higher than those at Reϕ=2.8×106 owing to a compressibility effect. During the cooling transient, for cases with Reϕ≤1.4×106, the magnitudes of the heat fluxes gradually decreased and eventually reached virtually zero. This indicated that the flow was first governed by the buoyancy effects and then became stratified. At Reϕ=2.8×106, where the rotational speed was at its maximum, buoyancy-induced flow dominated the entirety of the transient process due to significant frictional heating at the periphery of the rig. The calculated fluxes present evidence for future theoretical and computational modeling of transient disk heat transfer, and the Bayesian model provides guidance for transient temperature data analysis.

On-Design Component-Level Multiple-Objective Optimization of a Small-Scale Cavity-Stabilized Combustor

Abstract This work presents an on-design component-level multiple-objective optimization of a small-scaled uncooled cavity-stabilized combustor. Optimization is performed at the maximum power condition of the engine thermodynamic cycle. The computational fluid dynamics simulations are managed by a supervised machine learning algorithm to divide a continuous and deterministic design space into nondominated Pareto frontier and dominated design points. Steady, compressible three-dimensional simulations are performed using a multiphase realizable k–ε RANS and nonadiabatic flamelet/progress variable combustion model. Conjugate heat transfer through the combustor liner is also considered. There are fifteen geometrical input parameters and four objective functions viz., maximization of combustion efficiency, and minimization of total pressure losses, pattern factor, and critical liner area factor. The baseline combustor design is based on engineering guidelines developed over the past two decades. The small-scale baseline design performs remarkably well. Direct optimization calculations are performed on this baseline design. In terms of Pareto optimality, the baseline design remains in the Pareto frontier throughout the optimization. However, the optimization calculations show improvement from an initial design point population to later iteration design points. The optimization calculations report other nondominated designs in the Pareto frontier. The Euclidean distance from design points to the Utopic point is used to select a “best” and “worst” design point for future fabrication and experimentation. The methodology to perform computational fluid dynamics optimization calculations of a small-scale uncooled combustor is expected to be useful for guiding the design and development of future gas turbine combustors.

Cooling performance analysis and structural parameter optimization of X-type truss array channel based on neural networks and genetic algorithm

This study aims to optimize the structural parameters of an innovative cooling channel filled with a novel X-type truss array structure, so as to improve the cooling performance of the mid-chord region of gas turbine blades. In this study, the design of experiment (DOE) for influence parameters of X-type truss array channels was carried out, the variation ranges of influence parameters are as follows: Reynolds number (ReH, 10,000 to 60,000), truss rod diameter ratio (d/H, 0.1 to 0.3), transverse spacing ratio (Xs/C, 0.1 to 0.3) and streamwise spacing ratio (Zs/C, 0.1 to 0.3). The comprehensive effects of each structural parameter and Reynolds number on the flow and heat transfer performance of X-type truss array channels were analyzed. Then the multi-input and multi-output neural network model with prediction deviations less than 3% for the X-type truss array channels was established. Finally, the optimal structure parameters and arrangement of the X-type truss array channels were obtained by optimization design based on the genetic algorithm method, and have been verified by experimental measurement. The results show that the average Nusselt numbers and the friction coefficients of the X-type truss array cooling channel both rise with increasing d/H, drop with increasing Xs/C, and first rise then drop with increasing Zs/C; the comprehensive thermal coefficients first rise and then drop with increasing d/H, Xs/C, and Zs/C. Based on the maximization of average Nusselt number, the optimal values of d/H and Xs/C are 0.3 and 1, the optimal values of Zs/C increase from 1.179 to 2.201, as ReH increases from 10,000 to 60,000. According to the optimization results of maximization of comprehensive thermal coefficient, the optimal values of d/H, Xs/C and Zs/C increase from 0.186 to 0.248, 1.696 to 2.482 and 1.0 to 1.945 respectively with increasing ReH. Compared with the reference channel, the average Nusselt number of the optimized X-type truss array channel increases by 44.50 to 145.49% and the comprehensive thermal coefficient increases by 4.16–9.56% at different Reynolds numbers. The results may provide new ideas for the design of internal cooling structure in future advanced gas turbine blades.

Low temperature in-situ formation of oxygen barriers by heat treatment of aluminum infiltrated environmental barrier coatings

Environmental barrier coatings (EBCs) with great oxidation-resistant performance in steam are needed for efficient protection of silicon carbide-based composites in future advanced gas turbine engines. Besides a dense structure without rapid penetration paths for oxidants and other corrosives, refractory components such as alumina oxide or oxide-containing compounds also help to repel oxygen through coatings, which can ensure a long service life of EBC when used to infiltrate and block pores. However, due to their melting temperature up to nearly 2000 °C, melting and infiltration will damage EBC and matrix. In this work, metallic aluminum first infiltrates into open pores and microcracks at a low temperature of 750 °C, forming a dense structure, and then reacts in situ to form the refractory phases during heat treatment at 1100 °C, which is far below the melting point of refractory components of nearly 2000 °C. The results demonstrate that the structure remained dense after heat treatment. An intermedia layer composed of alumina and ytterbium aluminum garnet was observed due to the diffusion and reaction of oxygen and aluminum. The dense and refractory diffusion-reaction layer prevents the rapid infiltration and diffusion of oxidants in the engine environment to coating and matrix.

Preheating and premixing effects on NOx emissions in a high-pressure axially staged combustor

NOx emissions remain a primary concern for modern and future gas turbines, particularly as we progress towards a net-zero carbon future. It is critical to identify novel strategies to mitigate environmental pollutants for both modern turbines operated on natural gas and future turbines projected to use carbon-free fuels (such as hydrogen). In the current study, NOx emissions of a reacting methane-air jet in a vitiated crossflow are experimentally investigated in a model axially staged combustor at 5 atm. The combustion products from the main stage combustor flow into the axial stage at temperatures ranging from 1580 to 1650 °C. The axial stage contains a reacting jet in crossflow, which provides an overall temperature rise ranging from 50 to 220 °C to generate exit temperatures similar to gas turbine engines. The focus of this paper is to explore the effects of flame liftoff and ignition timescales on the NOx contribution of this secondary jet. This is done by varying the premixed level and preheat temperature of the axial jet. The effects are explored at different momentum flux ratios (J), jet equivalence ratios (φjet), exit temperatures (Texit), and main stage (headend) temperatures (THE). High-speed CH* chemiluminescence imaging is employed for each test case to determine the flame stabilization point. For a given temperature rise and combustor exit temperature, an increase in ignition delay was observed with a decrease in jet temperature. This provided a NOx benefit, particularly for jets that are injected at or close to an ignitable mixture fraction. Similarly, the coaxial injector caused a significant ignition delay compared to the fully premixed injector; this also led to a NOx benefit. Overall, the results show that a thermally driven delay in the axial stage provides a greater NOx reduction than a mixing delay.

On the unsteady behaviours of the adiabatic endwall film cooling effectiveness

The film cooling technique is introduced in modern gas turbines to protect the blade from the high temperature of the incoming hot gases by forming a thin coolant blank over the blade surface. However, it is known as a jet in crossflow (JICF), where coolant and mainstream interact intensively and generate complex vortices leading to highly unsteady coolant coverage over the blades surface. In this study, a fast-response pressure-sensitive paint technique (fast-PSP) was used to measure the coolant unsteadiness with a high-resolution camera. The measurements were performed in a novel single-passage transonic wind tunnel to uncover the unsteady effectiveness of the endwall surface. Such effectiveness was dramatically influenced by the blowing ratios (M), showing attached flow at a low blowing ratio and lift-off at a high blowing ratio. The effectiveness was asymmetrically distributed due to the pressure gradients, jet compounding angle, and associated complex flows. The unsteady effectiveness was highly influenced by the energetic vortical structures, which interacted with the mainstream flow immediately behind the holes. It was featured by secondary structures (horseshoe, passage, and counter vortices) beside the JICF structures. Meanwhile, the unsteadiness was originated from the middle of the passage behind the holes. It is suggested to pay close attention to the locations of the holes for further optimization. This study could help the designers to understand the characteristics of unsteady effectiveness, promoting advanced cooling strategies for enhanced protection of future gas turbines.

Deposition of Fuel Impurities Within Thermal Barrier Coatings in Gas Turbine Hot Gas Paths

Abstract For the reliable operation of modern gas turbines, thermal barrier coatings (TBCs) need to withstand a wide range of ambient conditions resulting from impurities in inlet air or fuels. When analyzing the deposition of detrimental hot gas constituents, previous efforts largely focus on the investigation of solid and molten deposit interaction with TBCs. Recent literature and observations in gas turbines indicate that not only liquids can penetrate porous TBCs, but the deposition from gas phase inside of pores and cracks is also an aspect of TBC degradation. To investigate this vapor deposition process, a diffusion model has been coupled with a thermodynamic equilibrium solver. The diffusion model calculates vapor transport of trace elements through pores and gaps in the TBC, where the thermodynamic equilibrium solver calculates local thermodynamic equilibria to predict whether deposition takes place. In this work, the model is applied to discuss the deposition properties of calcium. In recent literature, calcium has—in some cases—been reported to deposit inside of TBCs as pure anhydrite (CaSO4). An actual anhydrite finding in the TBC of a stationary gas turbine blade was reproduced applying the introduced model. The vapor deposition is shown to occur within and on top of the TBC, depending on several factors, such as pressure, temperatures, calcium-to-silicon ratio, and calcium-to-sulfur ratio. The successful alignment of conditions in real engines with model results will allow addressing the increasing demand for more fuel- and operational flexibility of current and future gas turbines.

Fully Coupled Turbojet Engine Computational Fluid Dynamics Simulations and Cycle Analyses Along the Equilibrium Running Line

Abstract This work presents fully coupled computational fluid dynamics (CFD) simulations and thermodynamic cycle analyses of a small-scale turbojet engine at several conditions along the equilibrium running line. The CFD simulations use a single mesh for the entire engine, from the intake to the exhaust, allowing information to travel in all directions. The CFD simulations are performed along the equilibrium running line by using the iterative Secant method to compute the fuel flow rate required to match the compressor and turbine power. The freestream pressure and temperature and shaft angular speed are the only inputs needed for the CFD simulations. To evaluate the consistency of the CFD results with thermodynamic cycle results, outputs from the CFD simulations are prescribed as inputs to the cycle model. This approach enables on-design and off-design cycle calculations to be performed without requiring turbomachinery performance maps. In contrast, traditional off-design cycle analyses require either scaling, calculating, or measuring compressor and turbine maps with boundary condition assumptions. In addition, the CFD simulations and the cycle analyses are compared with measurements of the turbojet engine. The CFD simulations, thermodynamic cycle analyses, and measurements agree in terms of total temperature and pressure at the diffuser–combustor interface, air and fuel mass flow rate, equivalence ratio, and thrust. The developed methods to perform CFD simulations from the intake to the exhaust of the turbojet engine are expected to be useful for guiding the design and development of future small-scale gas turbine engines.