Gas Turbine Engine Performance Research Articles

Performance and Emission Characteristics of a Small Gas Turbine Engine Using Hexanol as a Biomass-Derived Fuel.

The global transition to renewable energy has amplified the need for sustainable aviation fuels. This study investigates hexanol, a biomass-derived alcohol, as an alternative fuel for small-scale gas turbines. Experimental trials were conducted on a JETPOL GTM-160 turbine, assessing blends of 25% (He25) and 50% (He50) hexanol with kerosene (JET A) under rotational velocities ranging from 40,000 to 110,000 RPM. The parameters measured included thrust-specific fuel consumption (TSFC), turbine inlet and outlet velocities, and the emission indices of NOx and CO. The results demonstrated that the He25 and He50 blends achieved comparable thermal efficiency to pure JET A at high rotational velocities, despite requiring higher fuel flows due to hexanol's lower heating value. CO emissions decreased significantly at higher velocities, reflecting improved combustion efficiency with hexanol blends, while NOx emissions exhibited a slight increase, attributed to the oxygen content of the fuel. This study contributes a novel analysis of hexanol-kerosene blends in gas turbines, offering insights into their operational and emission characteristics. These findings underscore hexanol's potential as an environmentally friendly alternative fuel, aligning with global efforts to reduce fossil fuel dependency and carbon emissions.

Numerical simulation of micro gas turbine engine based on time-marching one-dimensional method

The advancement of gas turbine technology serves as a crucial indicator of high-tech and technological prowess, with widespread applications in aviation, marine vessels, power generation, and clean energy. The precision of the performance simulation model stands as a pivotal component in the development of gas turbines. The present study deals with the one-dimensional modeling and validation of gas turbines by utilizing a time-marching scheme, in order to establish the correlation between the engine components and full-engine performance. The flow through the turbomachinery was solved by the one-dimensional model. The combustor was modelled as a compressible actuator disk. The accuracy of the turbomachinery solver was validated using the KJ66 compressor and turbine, and the predicted results demonstrated consistency with the experimental data or the three-dimensional simulations. The characteristic interface method was developed based on Riemann invariants, in order to provide the flow properties on both sides of the combustor. The results indicated that compared to the component-by-component approach, the developed characteristic interface method significantly reduced the computational time. Finally, the developed one-dimensional method was employed to simulate the performance of the KJ66 micro gas turbine engine. The predicted performance for the KJ66 gas turbine engine was consistent with the trend in the experiment results. The novelty of the present study was its modeling approach, and its ability to eliminate the dependence on the general component map, and obtain meridional distributions of flow parameters, which cannot be achieved in traditional zero-dimensional thermal cycle calculations. These results can provide an accurate and powerful tool for the analysis and optimization of component and full-engine performance in the preliminary design stage.

Robust acceleration schedule design for gas turbine engine using multilayer perceptron network with adaptive sample class weighting

The acceleration schedule is crucial for generating acceleration control references for the gas turbine engine (GTE) and ensuring optimal acceleration performance. However, under harsh operating conditions, GTEs may encounter difficult-to-diagnose pressure sensor faults, which lead to inaccurate control references and decreased acceleration performance. This paper proposes a data-driven robust acceleration schedule (RAS) design method to tackle this issue, including fault data augmentation and adaptive sample class weighting (ASCW). Fault data augmentation generates multiple pressure fault sample classes from a normal acceleration schedule dataset. Due to the redundant pressure sensor configuration, the RAS can reconstruct these classes and generate accurate control references when a single pressure sensor fails. The ASCW employs a multilayer perceptron network to reconstruct the normal and fault sample classes accurately. Proportional integral regulation adjusts their weights during training to ensure balanced reconstruction precision. Simulation cases were conducted to verify the effectiveness of the RAS under actual GTE conditions, including engine-model mismatches, performance deterioration, flight envelope, and measurement uncertainty. The results demonstrate that the RAS ensures superior acceleration performance of GTEs across the full flight envelope in both normal and single pressure sensor fault scenarios. Additionally, the ASCW achieves reconstruction precision of 0.068 %, 0.080 %, 0.080 %, 0.082 %, and 0.077 % for normal and fault sample classes, respectively, prioritizing the precision of the normal sample class and balancing the precision of fault sample classes.

Experimental analysis and modeling of gas turbine engine performance: Design point and off-design insights through system of equations solutions

In this article, a systematic approach is presented for simulating turbojet engine performance at both design and off-design points. The construction of gas turbine engines involves a core, with additional components added based on the proposed methodology. Accurate formulation of component characteristics and their matching forms the basis for the system of equations needed for solving off-design engine performance. The developed algorithm's initial step involves comparing results with experimental data obtained from 230 N engines. Experimental testing utilizes a 780 N engine under diverse operating conditions, allowing determination of thrust values, fuel consumption, and outlet gas temperatures. An engine characteristic curve is derived under static conditions at ground level and validated against ground test results in Tehran. Unknown parameters are calibrated using a dedicated model, enabling the extraction of engine performance parameters across different corrected speeds. These results are then compared with GasTurb outputs. Comparative analysis between the proposed model's simulation and experimental data demonstrates high accuracy, affirming the reliability of the presented model.

Моделирование скругления острых кромок наклонных охлаждающих отверстий лопаток турбин методом электроэрозионного фрезерования на многоосевых станках с ЧПУ

The article views efficiency improving issues, quality and accuracy of sharp edge smoothing for numerous angular fans in the turbine blades of gas turbine engines made of heat-resistant nickel alloys and used in modern growing passenger aviation transport based on the development of ED milling new technology for modern multi-axis CNC machines with a rotating hollow electrode and a high-pressure fluidsupply of a special flush through the electrode body. An automated process for sharp edge machining has been developed for the production of more than 200 differently placed angular fans with a diameter of 0,45…0,6 mm in one blade instead of labor-intensive manual benchwork, where various needles, abrasive and metal rotary files are used. It is found that modeling and the development of special software resulted in a uniform radius of smoothing with high accuracy due to automatic ED milling operation along the entire trailing edge of each angular hole in the turbine blades at a multi-axis machining electro discharge center. The use of a rounded edge of the outlet ports provides more efficient cooling of the turbine blades and higher performance of gas turbine engines. At the same time, it is found that with a significant efficiency improving in blade sharp edge machining, a higher quality of the rounded edge surfaces in the holes is ensured. The results of the conducted research are recommended for the adoption of technologies in the aircraft engine industry.

Approximation of the Profile of Gas Turbine Engine Blades

Introduction. Increasing the durability of gas turbine engine (GTE) blades is achieved through the use of special protective coatings on their surface. For the development of such coatings, the basic source information is the geometric profile of the blade section. To transfer a given blade cross-section profile to the appropriate CAD/CAM system or engineering analysis package, parametric modeling methods are used to automate this operation. However, the known approaches to creating a parametric model of a blade profile are not without a number of disadvantages, and a generally accepted method for creating it does not currently exist. The research was aimed at creating a technique for approximating the profile of gas turbine engine blades, convenient for use in the subsequent analysis of the operating conditions of special coatings on the surface of the blades.Materials and Methods. When constructing parametric models of the profile of gas turbine engine blades, a method based on the orthogonal Legendre polynomials was used. This made it possible to provide high accuracy of approximation and construction of a continuous mapping for the parameters of the blade profile approximation. A Python application was created for automated processing of source profiles. It provided the calculation of the coefficients of approximating polynomials for the contour lines of the blade, visualization of the calculation results, and creation of a dxf file based on the points of approximating functions to transfer it to the CAD system. Next, geometric models of blades were used to solve the problem of a stationary aerodynamic flow around a blade. The results of solving this problem were used to study the effect of the blade profile on its cooling in an aerodynamic flow.Results. As an example, three options of blade profiles belonging to different types of GTE were considered. It was shown that for all three studied profiles, the proposed technique provided obtaining parametric models that maintained high accuracy in constructing approximating lines, which was confirmed by the values of the determination coefficients close to unity. To illustrate the possibility of using the obtained models, examples of solving the gas dynamic problem with a potential flow around a blade in a stationary aerodynamic flow were given. The distributions of pressure and temperature on the surface of the blade were calculated using the finite element method.Discussion and Conclusion. The calculation results show that the proposed technique of approximating the profile of the GTE blade, based on the use of orthogonal polynomials, is a convenient tool to automate the creation of a geometric model of the blade and compare different types and profiles of blades, solving the corresponding gas dynamic problems. At the same time, for a given blade profile and GTE operating conditions, it is possible to obtain the distribution of temperatures and forces acting on the surface of the blade, which is required for predicting the durability of special coatings.

Determining Steady-State Operation Criteria Using Transient Performance Modelling and Steady-State Diagnostics

Data from the steady-state operation of gas turbine engines are used in gas path diagnostic procedures. A method to identify steady-state operation is thus required. This paper initially explains and demonstrates the factors that cause a deviation in engine health when transient data are used for diagnosis and shows that there is a threshold in the slope of time traces, below which the variation in engine health parameters is acceptable. A methodology for deriving a criterion for steady-state operation based on actual flight data is then presented. The slope of the exhaust gas temperature variation with time and the size of its time-series window, from which this slope is determined, are the required parameters that must be specified when applying this criterion. It is found that the values of these parameters must be selected so that a sufficient number of steady-state points are available without compromising the accuracy of the diagnostic procedure.

Fault feature signal extraction method for rolling bearings in gas turbine engines based on threshold parameter decision screening

Vibrations during the operation of gas turbine engines (GTEs) include shocks from the combustion process as well as vibrations from high-speed rotating components, which may mask or distort the characteristic signals of rolling bearing failure. Compared to those of ordinary rolling bearings, the fault characteristics of rolling bearings in GTEs are more difficult to extract. In response to the difficulty of extracting characteristic signals of rolling bearing faults in gas turbine engines due to the effect of environmental noise, a method for extracting characteristic features of GTE rolling bearings is proposed. First, the particle swarm optimization (PSO) algorithm is used to optimize the parameters in the variational mode decomposition (VMD) algorithm, resulting in K0 modal components. VMD can not only suppress the modal aliasing phenomenon, but also decompose signals at different scales, thereby providing multi-resolution information. Second, a new parameter harmonization equation is introduced, which balances the relationship between the kurtosis and correlation coefficients, and is fused into a single parameter P. Subsequently, high signal-to-noise ratio (SNR) signals are selected ted based on threshold parameter criteria, and these high-SNR signals are used to generate new vibration signals. Finally, the envelope spectrum is used to extract fault characteristics of bearings. According to the results, the fault feature signal extraction (FFSE) method can extract fault characteristics in rolling bearings under both simple and complex transmission paths.

Evaluation Of Foreign Object Damage On The Fan Blades with Microscopic Techniques

Foreign object damage (FOD) is a common problem in gas turbine engines, particularly in the fan and compressor blades. FOD generally occurs when hard body particles are ingested into an aero-engine fan and compressor blade. High-speed impact on fan titanium alloy blades leads to foreign object damage, which weakens the fatigue performance of the blades. It is important to note that FOD can also cause damage that is not visible to the naked eye, such as internal cracks or fatigue damage. Therefore, regular inspections and maintenance are crucial to detect and prevent FOD damage in gas turbine engines. Also using a microscope for the inspection of blades is important. Proper examination of blades damaged by foreign or domestic object can shed light on blade status and prevent further damage in the future. Here are some of the benefits of using a microscope for blade inspection:
 Microscopes can provide high magnification and resolution, allowing for detailed inspection of the blade surface and any defects or damage.
 Microscopes can help detect and identify small cracks, chips, or other forms of damage that may not be visible to the naked eye.
 Microscopes can provide accurate measurements of the blade's dimensions, which can be important for quality control and maintenance purposes.
 Microscopes can help identify the root cause of any damage or wear on the blade, which can inform future design and manufacturing decisions.
 Microscopes can provide visual documentation of the blade's condition, which can be useful for tracking changes over time and for communicating with other stakeholders.
 Overall, using a microscope for blade inspection can help ensure the safety, reliability, and performance of gas turbine engines and other equipment that relies on blades.
 In this study, a comparison was made between visual inspection, borescope control, stereo microscope and SEM-EDX microscope methods used in the detection and removal of foreign material damage. The advantages and disadvantages of these 4 methods are comparatively examined and presented to the attention of the reader.

Assessment of a virtual boundary concept for efficient modeling of turbine film cooling

Film cooling is one of key technologies advancing the performance of gas turbine engines, which is extensively applied in the hot turbine component for thermal protection by injecting cooling air from numerous film holes. To achieve an optimal film cooling scheme, many iterations of the position, the diameter, and the orientation of the holes through computation simulations are required to well fit the aerodynamic profiles of the turbine blades. However, a conventional computational simulation that resolves each individual hole and its coolant supply channel implies substantial costs of computation resources. In this study, a novel film cooling modeling method that uses a virtual boundary to represent film holes is assessed by comparing to the conventional computational simulation in two representative test cases, involving flat-plate and turbine endwall inlet film cooling. Relevant experiments are carried out to provide experimental data for verifying the approach at various coolant injection rates. Generally, the virtual boundary model produces rather equivalent results to those predicted by the conventional simulation method. Though coolant concentration downstream of the holes is over-estimated by the virtual boundary model due to neglected in-hole mixing, the overall flow fields and film cooling performance are well captured, showing the feasibility of the novel strategy in accurately modeling film cooling in a much more efficient way by saving computational efforts. This novel concept thereby allows designers to modify the film hole geometries independently of the blade geometry and thus, is helpful to accelerate the design process of a film-cooled turbine or to assess multiple film cooling configurations more quickly, as compared to a conventional computation simulation.

Internal Flow Characteristics of Novel Turbine Performance Enhancement Method through Specific Fluid Injection

Gas turbine engines are an essential component for many industries, such as aerospace, marine propulsion, energy generation, etc., with modern engines capable of achieving high powers, efficiencies and reliability. However, these high performances are achieved on a narrow interval of working regimes; when operating at partial loads, a drastic decrease in performance is expected. In order to mitigate this drawback, a novel injection method has been proposed to improve axial turbine and gas turbine engine performance at these regimes. The method consists of the injection of a fluid in specific sections of turbine vanes to accelerate the flow, modify the velocity triangles and increase the generated power at partial loads. In this paper, the authors discuss the internal flow characteristics of the injection channels using numerical studies to determine the flow fields for different working regimes. The results show that the power generated by the rotor can be improved by 10% to 21% for different operating regimes without considering the internal geometries. The introduction of internal flow configurations led to smaller improvements in power generation, obtaining injection system pressure losses of between 10% and 20%. The paper concludes that the flow through channels is not uniform, with smaller dimensions of the internal geometries leading to higher pressures and an increased influence of the injection system.

Transient Tip Clearance Prediction Model Considering Transient Radial Temperature Distribution of Discs in a Gas Turbine Engine

Abstract The tip clearance of turbomachinery components, such as compressor and turbine, significantly impacts the gas turbine engine performance, including the thrust and fuel consumption. Therefore, an appropriate prediction of the turbomachinery tip clearance is indispensable in improving the accuracy of engine performance evaluations. The traditional 1D transient tip clearance prediction model does not consider the radial temperature distribution of the disk (also called disc). However, as the engine performance keeps improving, the radial temperature gradient of the disk also increases. In order to maintain the prediction accuracy within an acceptable range, a generic transient tip clearance prediction model is developed in this paper. Based on the special treatment of the axial and radial convective heat transfer boundary regions of the disk, a 2D transient heat transfer model is established, and the radial temperature of the disk is predicted. At the same time, the interaction effect of the thermal expansion between the two different layers of the shroud is studied to improve the prediction accuracy further. In addition, the centrifugal deformation of the blade and disk with variable-thickness characteristics is considered in the new transient model. Compared with the traditional 1D transient prediction model, the model developed in this paper further improves the predicting accuracy of the transient tip clearance, which is verified by the high-fidelity fluid–solid–thermal coupling method. The prediction model provides a useful tool for evaluating the transient tip clearance. It may also benefit the future design strategy of active clearance control.

Методика оцінки залежності власних частот коливань лопаток газових турбін від анізотропії монокристалу

The subject of study of this article is one of the key tasks that arises during the development and operation of aviation gas turbine engines, namely ensuring the strength of their parts. The turbine blade is one of the most heavily loaded parts of an aircraft engine. Therefore, at the stage of designing these blades, it is necessary to be able to identify and avoid possible resonance modes that appear during the flight cycle of an aircraft gas turbine engine. Such resonant oscillations of gas turbine blades can occur due to time-varying gas-dynamic forces from the action of the gas flow and are periodic in nature, as they are determined by the frequency of rotation of the rotor. In order to regulate the frequency characteristics of the blades to prevent dangerous resonant forms of oscillations that arise under the action of various harmonics of the exciting force during variable modes of operation of the aircraft engine, it is necessary to carry out a complex of various technological or structural changes. At the current stage, turbine blades of aircraft engines are manufactured by the method of single-crystal casting. As is known, single crystals have anisotropic properties, namely they manifest themselves as materials with the characteristics of cubic symmetry. A very urgent task is to investigate the effect of the anisotropy of the elastic characteristics of monocrystalline blades on their natural frequencies and forms of oscillations. In this research, the authors developed a method for determining the elastic characteristics of a single crystal, namely Young's modulus of elasticity, Poisson's ratio and pure shear modulus, which is based on available experimental data for typical heat-resistant alloys. With the help of finite element analysis on the example of a typical model of a cooled blade of a gas turbine, its modal analysis was carried out and its resonance diagram was constructed. The trend of changing the natural frequencies and forms of blade oscillations when the elastic constants of the single crystal change due to the rotation of the crystallographic system of directions was also investigated. The three-dimensional model of the blade was built using the capabilities of the NX graphics complex, all calculations were performed in the Maple computing complex, and the capabilities of the ANSYS software complex were used for finite element analysis.

Synergistic Effect of Physical Faults and Variable Inlet Guide Vane Drift on Gas Turbine Engine

This study presents a comprehensive analysis of the impact of variable inlet guide vanes and physical faults on the performance of a three-shaft gas turbine engine operating at full load. By utilizing the input data provided by the engine manufacturer, the performance models for both the design point and off-design scenarios have been developed. To ensure the accuracy of our models, validation was conducted using the manufacturer’s data. Once the models were successfully validated, various degradation conditions, such as variable inlet guide vane drift, fouling, and erosion, were simulated. Three scenarios that cause gas turbine degradation have been considered and simulated: First, how would the variable inlet guide vane drift affect the gas turbine performance? Second, how would the combined effect of fouling and variable inlet guide vane drift cause the degradation of the engine performance? Third, how would the combined effect of erosion and variable inlet guide vane drift cause the degradation of the engine performance? The results revealed that up-VIGV drift, which is combined fouling and erosion, shows a small deviation because of offsetting the isentropic efficiency drop caused by fouling and erosion. It is clearly observed that fouling affects more upstream components, whereas erosion affects more downstream components. Furthermore, the deviation of performance and output parameters due to the combined faults has been discussed.

Thermodynamic analysis of the thermal and exergetic performance of a mixed gas-steam aero derivative gas turbine engine for power generation

A thermodynamic analysis is performed for an aero derivative gas turbine engine which utilizes steam injection to increase its efficiency. The target was to explore the performance of a high efficiency gas turbine unit for electric power generation without downstream Rankine cycle. A Rankine cycle for exhaust heat recovery is unattractive because of its large response time and cost of investment. The main purpose of this research was to develop a better understanding of how the optimal cycle efficiency is reached, when the steam for injection is generated by use of the turbine exhaust heat. The STIG cycle becomes attractive for grid stabilization because of its low CAPEX and small footprint and response time.A thermodynamic model has been developed to simulate the simple cycle gas turbine, steam generation and effects of steam injection. Reference input parameters for the model are taken for the GE LM6000 turbine as publicly available. The performance of the engine without steam injection as predicted by the model is compared with literature for validation and compares well.The performance of the STIG cycle as a function of operation parameter steam mass flow and design parameters pressure ratio and turbine inlet temperature is investigated and the optimal parameter settings determined. It is found that this type of cycle shows a very specific parameter setting for optimal efficiency. By using steam injection for the chosen turbine and its parameters an efficiency gain of around 11% points and an output power augmentation of 45% can be achieved.

An analysis approach for micro gas turbine engine's performance by experiment and numerical simulation

The corrected special maps of components have revealed an outstanding role on valuating the off-design performance of micro gas turbine engines. The main purpose of this paper is to investigate the steady-state performance at design and off-design point of a micro turbojet engine capable of generating 400 N thrust. At present work, the experiments were conducted and the thrust and specific fuel consumption are 39.48 daN and 1.62kg/(daN h) at design point, respectively. Importantly, a present methodology, that is adjusting the values of Beta and N/sqrt(T) in specific map of the components when performing the off-design calculation of this engine, is proposed in this research, where results are compared with the data from experiment and other literature. Exemplarily, according to a comparison with the experimental results, there is less than 1% of the deviation in thrust at design point and 3% of the difference in the thrust at off-design point with the corrected specific map when performing the numerical simulation. This method enables evaluating the operating characteristics of both on design and off-design point, providing more accurate trend analysis of performance parameters in terms of engine compared to the researches which seldom used the scaling map.

Improving Film Cooling Efficiency with Lobe-Shaped Cooling Holes: An Investigation with Large-Eddy Simulation

Shaped cooling holes have received considerable attention in recent years due to their potential to improve heat transfer while minimizing pressure drop. In this study, the effects of lobe-shaped cooling holes on film cooling performance and turbulent flow characteristics are investigated using large-eddy simulation (LES). The LES approach was validated by comparing the calculated film cooling effectiveness with experimental data. The results show that well-designed lobe-shaped cooling holes can improve the cooling efficiency by up to 37% and reduce the total pressure drop. Additionally, the cooling jets of lobe-shaped cooling holes have a smaller angle to the wall surface, indicating better wall attachment. However, poorly designed lobe-shaped cooling holes can reduce the cooling efficiency. Overall, this study suggests that lobe-shaped cooling holes are a promising technology for improving gas turbine engine performance.

Optimization Model of the Shell Capsules Geometry for a System for Diagnosing Damage to Gas Turbine Blades in Non-stationary

Scientific research is devoted to the problem of diagnosing damage in the blades of gas turbine engines in non-stationary conditions. The paper proposes a model for a more accurate calculation of the parameters of the crack detection system using a system of capsules, inside of which, under the influence of pressure, there is a substance exhibiting ionizing properties. In the study, the relations for determining the uneven pressure due to the action of centrifugal tensile forces during the rotation of the turbine blade are obtained. The relations for calculating the minimum required diameter of the capsule to ensure the rupture of the capsule when cracks are opened for the effective operation of the crack detection system are obtained. Calculations of the pressure inside the capsule and the geometry of the capsule at different speeds of rotation of the turbine blade are carried out. An estimate of the error in calculating the pressure inside the capsule is given in the case of not taking into account the action of centrifugal forces. The obtained dependences will significantly optimize the system for detecting damage to turbine blades, and increase the efficiency and safety of the operation of gas turbine engines.

Computational Optimization of a Loosely-Coupled Strategy for Scale-Resolving CHT CFD Simulation of Gas Turbine Combustors

The accurate prediction of heat fluxes and, thus, metal wall temperatures of gas turbine (GT) combustor liners is a complicated and numerically expensive task. Computational Fluid Dynamics (CFD) support for the design of cooling systems is essential to ensure safe and proper operation of the entire gas turbine engine. Indeed, it is well known how complicated, and, at the same time, expensive it is to carry out experimental campaigns inside combustors operating under working conditions, and, therefore, pressurized and having high temperatures. The correct prediction of thermal fluxes in a CFD simulation depends on the proper modeling of all the involved phenomena and their interactions with each other. For this reason, Conjugate Heat Transfer (CHT) simulations are mandatory in gas turbine cooling system applications. Multiphysics and multiscale simulations, based on loosely-coupled approaches, have emerged as extremely effective numerical tools, providing enormous computational time savings, as compared with standard CHT simulations. The fundamental advantage of such approaches is based on the fact that each heat transfer mechanism is solved with the most suitable numerical setup, which leads to the use of spatial and temporal resolutions following the characteristic time scales of each phenomenon to be solved. For industrial applications, where the availability of numerical resources is limited and, at the same time, the timelines with which to obtain results are rather tight, having robust and easy-to-use loosely-coupled solutions available for the design of combustion chamber cooling systems would be extremely valuable. In this context, the objective of this work was to perform an initial optimization step for the multiphysics and multiscale tool, U-THERM3D, developed at the University of Florence to revise the coupling strategy workflow with a view to making the numerical tool faster and easier to use. The revised methodology was applied to the RSM gas turbine combustor model test case developed with cooperation between the Universities of Darmstadt, Heidelberg, Karlsruhe, and the DLR. In particular, all experimental tests were conducted by the Institute of Reactive Flows and Diagnostics (Reaktive Strömungen und Messtechnik) of the Department of Mechanical Engineering at TU Darmstadt, from which the gas turbine combustor model takes its name. The newly obtained results were compared and analyzed, both qualitatively and in terms of computational time savings, with those previously achieved with the current version of the U-THERM3D tool already studied by the authors and available in the literature. Moreover, an analysis of computing times was carried out relative to the super-computing center used for the different adopted methodologies.

An optical technique based on silicate glass sintering for temperature mapping

AbstractThermal paints have been used for decades by the gas turbine engine community to map surface temperature with low resolution. A novel thermal paint based on the sintering of a lead‐silicate glass powder was developed that can map maximum temperature with high resolution (±5°C) over a 60°C range beginning at the glass transition temperature (). The paint exhibited excellent adhesion to nickel‐based superalloy components due to similar coefficients of thermal expansion between the superalloy and glassy ceramic coating. An optical transition was qualitatively and quantitatively observed using scanning electron microscopy, ultraviolet‐visible (UV‐VIS) reflectance spectroscopy, and visual inspection. UV‐VIS reflectance spectroscopy was used to confirm the optical transition observed by the naked eye and quantitatively assess the transition of the thermal paint with high resolution. This technique for obtaining high resolution experimental temperature maps can aid the performance, efficiency, and reliability of gas turbine engines.