Design Of Gas Turbine Engines Research Articles

JSC “ODK” experience in implementing digital twin technologies in the design of gas turbine engines

The prerequisites for the formation, basic principles and approaches of the GTE design technology used in JSC “UEC” based on the concept of “digital twin” are considered. The components of the DT technology are fixed, three types of computer models included in the perimeter of the digital twin: models “as designed”, “as manufactured”, “as tested”, are considered. The necessity for constructing automated computational chains is noted and a brief description of the process is given.

Analysis of a turbofan engine using a parametric loop

PurposeIn this study, the performance analysis of the split flow turbofan engine with afterburners has been examined using the parametric cycle analysis method. The purpose of this study is to examine how engine performance is impacted by design parameters and flight ambient values and to develop a software in this context.Design/methodology/approachSoftware has been developed using the open-source PYTHON programming language to perform performance analysis. Mach number, compressor/fan pressure ratio, bypass ratio and density were used as parameters. The effects of these variables on engine performance parameters were investigated.FindingsParametric cycle analysis has been calculated for different flight conditions in the range of 0–2 M and 0–15,000 m altitude for turbofan engines. With this study, basic data were obtained to optimize according to targeted flight conditions.Practical implicationsAs a result of the performance analysis, the association between the flight conditions and design parameters of engine were determined. A software has been developed that can be used in the design of supersonic gas turbine engines for fast and easy simulation of the design parameters.Originality/valueThe variables used in the literature have been analyzed, and the results of the studies have been incorporated into the developed software, which can be used in innovative engine design. Software is capable to be developed further with the integration of new algorithms and models.

Grain boundary microstructure-property relationships in the cast & wrought Ni-based superalloy René 41 with boron and carbon additions

Cast & wrought Ni-based superalloys are required for manufacturing high-performance components in gas turbine engines. Their excellent high-temperature properties and corrosion resistance are brought about by their complex microstructure of a γ-matrix, intra- and intergranular γ' precipitates, carbides, and borides. The most highly-alloyed superalloy grades provide the high-temperature strength required for the design of next generation gas turbine engines but their processability remains challenging due to grain boundary (GB) cracking. A common mitigation approach is the addition of B and C to improve GB cohesion and strength. However, their detailed roles in GB segregation and precipitation and the link to GB properties during processing remains largely unexplored.We provide a systematic study on decorated GBs in the cast & wrought Ni-based superalloy René 41 with B and C additions. Tensile testing reveals that B additions reduce its ductility due to GB cracking but slow down grain growth during heat treatments more effectively, than C additions. GB-M2B, M6C, M23C6, and GB-γ' precipitates decorating GBs are sequentially precipitated. Their structures and compositions are shown to be unaffected by B and C additions. However, significant differences in the interfacial segregation of solutes are revealed. B is captured by GB-M2B and therefore unavailable to improve GB cohesion. C depletion and enrichment of solutes at interphase boundaries are identified as origin for the reduced ductility and thus, GB cracking. The findings are summarized as microstructural model and advance the detailed understanding of the processing – microstructure – properties relationships of cast & wrought Ni-based superalloys.

Advanced Simulation Technologies and Testing Methodologies Applied in the Design and Development of Aero Gas Turbine Engines

An overview of Gas Turbine Engine Development, KAVERI, the power plant for the Light Combat Aircraft (Tejas) and a description of the validated numerical tools, aerothermodynamic and mechanical test facilities of the Gas Turbine Research Establishment (GTRE) are presented. The KAVERI engine compressor, combustion system, turbine, analytical tools, blade heat transfer, exhaust system, Integrated Nozzle Actuation System (INAS), Non-intrusive Stress Measurement System (NSMS) and thermal paint tests are described in detail.

A versatile volume-based modeling technique of distributed local quadratic convergence for aeroengines

For advanced aero-engine design and research, modeling and simulation in a digital environment is indispensable, especially for engines of complicated configurations, such as variable cycle engines (VCE) and adaptive cycle engines (ACE). Also, in the research of future smart engines, reliable real-time digital twins are paramount. However, the 2 dominant methods that used in solving the simulation models, Newton-Raphson (N-R) method and volume-based method, are not fully qualified for the study requirements, because neither of them reaches the satisfactory balance of convergence rate and calculating efficiency. In this study, by deeply analyzing the mathematical principle of these 2 methods, a novel modeling and solving method for aero-engine simulation, which integrates the advantages of both N-R and volume-based methods, is established. It has distributed architecture and local quadratic convergence rate. And a novel modeling method for variable area bypass injectors (VABI) is put forward. These facilitate simulation of various configurations of aero-engines. The modeling cases, including a high bypass-ratio (BPR) turbofan and an ACE, illustrate that the novel technique decreases the iterations by about two-thirds comparing with volume-based method, while the success rate of convergence remains over 99%. This proves its superiority in both convergence and calculating efficiency over the conventional ones. This technique can be used in advanced gas turbine engine design and control strategy optimization, and study of digital twins.

Preliminary Design of a Mini Gas Turbine via 1D Methodology

To address the increasing interest towards more environmentally friendly naval transportation and the introduction of IMO2020 restrictions on pollutant emissions onboard ships, the present work details the preliminary design of a mini gas turbine engine, i.e., a gas turbine engine with an output power up to 5 MW, for onboard energy generation. In comparison to conventional propulsion systems, gas turbine units benefit from known compactness, which can be further enhanced by employing single-stage uncooled radial machines, according to similar works in the field. As such, the present paper aims to set up a complete procedure that allows a reliable and fast (i.e., requiring a limited computational effort) preliminary design of one-stage centrifugal compressors and radial turbines operating at a high pressure ratio via the use of classical one-dimensional theory. The aerodynamic design outputs in terms of forces and torques are then used to perform a preliminary mechanical design of the shaft by means of a one-dimensional finite element model with commercial software to estimate the corresponding shaft line stress. Despite some necessary geometrical and modeling simplification of the design problem, which results in the unavailability of detailed information on individual components, the employed procedure nevertheless allows a comprehensive overview of the possibilities in terms of maximum machine performance achievable at an early design stage with the associated limited computational requirements. The design procedure and the geometry achieved for the application are presented along with aerodynamic and structural results.

Film Cooling and Cold Streaks Tracking on a Fully Cooled Nozzle Guide Vane Under Representative Combustor Outflow Conditions

Abstract The pronounced nonuniform temperature distribution in the core engine flow path is a recurring problem of gas turbine engine design process. Specifically, turbine entry conditions are usually characterized by severe temperature distortions, often referred to as hot and cold streaks, ascribed to combustor burners and combustor liners cooling systems. Temperature distortions remain an issue even at the exit section of the nozzle guide vane (NGV), where additional cold streaks coming from the vane film-cooling system are injected into the flow. A precise knowledge of thermal field and its evolution is thus essential to mitigate their impact on turbine performance and lifetime. Various studies focus on the description of streaks migration through a direct investigation of the thermal field, providing an effective evaluation of the global phenomenon. As a deeper understanding is often required, experimental techniques based on the detection of tracer gases can be successfully adopted. In this study, a realistic combustor outlet swirl profile was imposed on a fully cooled NGV cascade to analyze both film-cooling behavior and cold streaks migration and redirection. A concentration probe based on the fluorescence behavior of an oxygen sensor, fully characterized in a previous work, was here employed to track the position of the film cooling flows at the NGV cascade exit plane, while the adiabatic film-cooling effectiveness was evaluated on the NGV surfaces employing the pressure sensitive paint (PSP) technique. Overall, the swirling structure strongly affects both the film-cooling behavior and cold streaks migration through and downstream the vane. The importance of examining the unsteady aspect is also highlighted to better estimate actual components operating temperatures. A global understanding of the occurring phenomena is therefore provided, as well as significant pieces of information that can be extremely useful for the design phases of both the NGV and the following rotor cascade.

ВЛИЯНИЕ ПОГРЕШНОСТЕЙ ИЗГОТОВЛЕНИЯ ДЕТАЛЕЙ НА ДИСБАЛАНСЫ РОТОРА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ ПРИ ЕГО СБОРКЕ

Increased vibrations of a gas turbine engine reduce the reliability and overall service life of the engine. The main reason for the appearance of increased vibrations is the dynamic imbalance (dynamic imbalance) of one or more engine rotors. Dynamic un-balance in its turn is a cumulative manifestation of static and moment unbalance of the rotor. One of the reasons for the appearance of unacceptable values of rotor imbalances is the imperfection of the geometric pa-rameters of the parts, that is, the manufacturing errors of these parts. Despite the fact that at present there are strict requirements for the accuracy of manufacturing parts of a gas turbine engine, it is impossible to manufacture a part without errors. When joining several rotors, the summing errors in the manufacture of parts may lead to a shift in the center of mass of the rotors, as well as to a misalignment of their axes, which in turn leads to an increase in the imbalances of the assembled rotor. As a result of the vibration of the assembled engine significantly exceed the permissible. The problem of the influence of errors in the manufacture of parts on the unbalances of the gas turbine engine rotor during its assembly is analyzed. A review analysis of the methods for solving this problem, presented in modern scientific papers, is car-ried out. A technique for 3D modeling of the GTE rotor assembly is given, taking into account the beat vectors to predict the static and moment imbalances of the assembled rotor. Based on the simulation results, the displacements of the center of mass of the analyzed rotor and the distortions of its axis are calculated for different positions of the beat vectors. The conclusion is made about the expediency of applying the considered methodology in the design of modern gas turbine engine.

Virtual Gas Turbines Part I: A Top-Down Geometry Modeling Environment for Turbomachinery Application

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

Identification of a mathematical model for a single-shaft power-generating GTE

The design and development processes of gas turbine engines rely on the usage of mathematical models representing the physics of engine functioning processes. One way of increasing the validity of a mathematical model is its identification based on engine test results. The identification of mathematical models of modern power-generating gas turbine engines (GTEs) presents a demanding and time-consuming task due to the necessity to identify the main controlled engine parameters determined in the course of experimental studies depending on a large number of the parameters that are not controlled during the experiment. In this regard the actual direction of reducing the labour intensity of the process of mathematical model identification is using identification program complexes. The object of the study was to solve the problem of structural-parametrical identification of the power-generating GTE functioning model detailing the turbine flow path calculations to the level of blade rows in order to obtain the GTE mathematical model that describes the characteristics of a real engine with given accuracy. To achieve the objective, the following problems were solved: variable parameters, controlled parameters and characteristics, ranges of their variations were selected from the total number of the mathematical model input data, the objective functions were defined; the task of the parametric identification according to the results of bench tests through GTE operating modes was performed; analytical approximating dependences for correcting coefficients (variable parameters) were obtained; structural-parametric identification of the mathematical model was performed. The novelty of the obtained results is the identification of the mathematical model of the nonlinear component GTE of the second level performed without model linearization (without its level lowering) by using the Optimum software packages. The methodological approach for the parametric identification of the mathematical model is proposed. This approach allows reducing the number of variable parameters under the modes lower that the maximum. It shows that the identified model allows obtaining the prediction results of the GTE parameters and characteristics through operating modes with a deviation of no more than 1.4% from the experimental data and, therefore, it will allow reduction of terms and an increase in the quality of power unit development.

Nonlinear generalized predictive controller based on ensemble of NARX models for industrial gas turbine engine

New design and operation of modern gas turbine engines (GTEs) are becoming more and more complex where several limitations and control modes should be fulfilled at the same time to accomplish a safe and ideal performance for the engine. For this purpose, a constrained multi-input multi-output (MIMO) non-linear model predictive controller (NMPC) based on neural network model is designed to fulfill the control requirements of a Siemens SGT-A65 three-spool aero-derivative gas turbine engine (ADGTE) used for power generation. However, the implementation of NMPC in real time has two challenges: Firstly, the design of an accurate non-linear model, which can run many times faster than real time. Secondly, the usage of a rapid and reliable optimization algorithm to solve the optimization problem in real time. To solve these issues, the constrained MIMO NMPC is created based on the generalized predictive control (GPC) algorithm as a result of its clarity, ease of use, and capacity to deal with problems in one algorithm. In addition, seven ensembles of eight multi-input single-output (MISO) non-linear autoregressive network with exogenous inputs (NARX) models are used as a base model for the GPC controller to predict the future process outputs. Estimation of free and forced responses of the GPC based on the neural network (NN) model of the plant each sampling time without performing instantaneous linearization is proposed in this study, which reduces the NMPC optimization problem to a linear optimization problem at each sampling step. In addition, the Hildreth's quadratic programming algorithm is used to solve the quadratic optimization problem within the NMPC controller, which offers ease of use and reliability in real time applications. To demonstrate the performance of the NNGPC controller developed in this study, we have compared the performance of the neural network generalized predictive control (NNGPC) controller to the existing controller of the SGT-A65 engine. The simulation results show that the NNGPC has demonstrated output responses with less oscillatory behavior and smoother control actions to the sudden variation in the electric load disturbance than those observed in the existing min-max controller. However, the min-max controller has faster response than that of the NNGPC controller.

Reduction of the finite element model of a gas turbine engine rotor using CMS technique

This paper is devoted to the problems of modeling the vibration behavior of rotors of gas turbine engines. The necessity of modeling their dynamic behavior is caused by their strong influence on the strength characteristics of the engine and its resource. The complexity of solving this problem lies in the fact that the design of modern gas turbine engines has many features that do not allow the use of only one-dimensional models, and the application of three-dimensional models requires a lot of computing resources. Thus, it is suggested to use methods of reduction of finite element models for solving the problems in three-dimensional statement. This paper uses the mode component synthesis (CMS) method to reduce the dimensionality of the rotor model. It also compares the speed of solving modal analysis tasks and determining critical frequencies using complete and reduced models. The quality of reduction is determined. Conclusions about the efficiency of the reduction technique aimed at solving rotor dynamics problems are made.

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.

Development of novel mathematical models of plate heat exchanger for the task of optimization main parameters of the aviation gas turbine engine with heat recovery

Nowadays, the task of increasing the economic and technical efficiency of the gas turbine engine and decrement of the specific weight and specific fuel consumption is one of the main targets in the development of the aviation gas turbine engine. A promising approach for decrement of specific fuel consumption and obtaining high thermal efficiency of the gas turbine engine (30% and above) is based on the concept of the recuperative cycle of the gas turbine engine. The article presents the computer-aided calculation for analysing the heat exchanger surface and developing novel mathematical models of heat exchanger in terms of weight goodness, flow passage goodness, and optimum weight and flow passage, furthermore, the study presents the novel mathematical models of the pressure drops in the air and gas channels of the heat exchanger. 18 different heat transfer surfaces for heat exchanger cores have been evaluated and the results are presented. The same hydraulic diameter assumed for all heat transfer surfaces and only their thermo-hydraulic performances evaluated during the calculation design. To assess the reliability of the obtained models, the results of the design calculations of the developed models compared with the data of other authors and with the data of the existing gas turbine engine with heat exchanger. The obtained models focus on mathematical design calculation for optimizing the main thermodynamic parameters of a gas turbine engine coupled with a heat exchanger at the stage of conceptual design of aviation gas turbine engine with heat recovery of exhaust gas.

A data-driven kinematic model of a ducted premixed flame

Reduced-order models of flame dynamics can be used to predict and mitigate the emergence of thermoacoustic oscillations in the design of gas turbine and rocket engines. This process is hindered by the fact that these models, although often qualitatively correct, are not usually quantitatively accurate. As automated experiments and numerical simulations produce ever-increasing quantities of data, the question arises as to how this data can be assimilated into physics-informed reduced-order models in order to render these models quantitatively accurate. In this study, we develop and test a physics-based reduced-order model of a ducted premixed flame in which the model parameters are learned from high-speed videos of the flame. The experimental data is assimilated into a level-set solver using an ensemble Kalman filter. This leads to an optimally calibrated reduced-order model with quantified uncertainties, which accurately reproduces elaborate nonlinear features such as cusp formation and pinch-off. The reduced-order model continues to match the experiments after assimilation has been switched off. Further, the parameters of the model, which are extracted automatically, are shown to match the first-order behavior expected on physical grounds. This study shows how reduced-order models can be updated rapidly whenever new experimental or numerical data becomes available, without the data itself having to be stored.

Development of Novel Computer Program for Cycle Analysis of Turbojet Engine

A method to simulate the gas turbine cycle and performance is developed. This paper intent to describe a digital computer code to make it useful for other researchers. This program is written in Python language for analyzing the steady-state, parametric cycle performance of turbojet engines. This can be used to analyze one- and two-spool turbojet engines without any modification to the basic program. The influence of initial parameters, component characteristics and flight condition on performance characteristics of gas turbine during operation are shown. The program results are compared and validated with those from an existing GSP (Gas Turbine Simulation Program) software. The major advantage of this new method is that it frees the programmer from having to minimize the number of equations which require iterative solution. As a result, some of the approximations normally used in engine simulations can be eliminated. The outcomes of this analysis form a strong base for further analysis to predict the performance of the gas turbine engine with reasonable accuracy for design and fabrication of gas turbine engines using this performance code.

Development of Novel Computer Program for Cycle Analysis of Turbojet Engine

A method to simulate the gas turbine cycle and performance is developed. This paper intent to describe a digital computer code to make it useful for other researchers. This program is written in Python language for analyzing the steady-state, parametric cycle performance of turbojet engines. This can be used to analyze one- and two-spool turbojet engines without any modification to the basic program. The influence of initial parameters, component characteristics and flight condition on performance characteristics of gas turbine during operation are shown. The program results are compared and validated with those from an existing GSP (Gas Turbine Simulation Program) software. The major advantage of this new method is that it frees the programmer from having to minimize the number of equations which require iterative solution. As a result, some of the approximations normally used in engine simulations can be eliminated. The outcomes of this analysis form a strong base for further analysis to predict the performance of the gas turbine engine with reasonable accuracy for design and fabrication of gas turbine engines using this performance code.

Static Stiffness Parametric Studies of Newly Conceptualized Inter Shaft Squeeze Film Damper (ISSFD) Rings

Most of the modern gas turbine engine designs requiring super critical operations will often have squeeze film dampers (SFDs) on the high pressure spool bearing as an inevitable feature. However, Squeeze Film Dampers are not very common for inter-shaft bearing applications and are still in the research and development stage. The current effort concentrates on static stiffness parametric studies based on newly conceptualized intershaft squeeze film damper (ISSFD) ring design suitable for space constrained intershaft bearing application of a typical two spool configured system. The work involves static stiffness evaluation of three and four grooves ISSFD rings by varying the groove angle/groove length/subtended angle of groove & hence the overlapping angle of grooves by conducting static tests on a dedicated static stiffness evaluation test rig designed and fabricated for the purpose. The pattern of circumferential variation of stiffness of each ring is studied and the results revealed a typical trend of decreasing stiffness in ISSFD ring with the increase in groove angle and hence the overlapping angle of grooves. The static stiffness parametric studies has resulted in developing an ideal profile of ISSFD ring having six grooves that exhibited axisymmetric pattern of circumferential variation of stiffness. When used in a space constrained intershaft bearing plane of a typical two spool gas turbine engine, the ideal ISSFD ring profile would exhibit the best damping performance. 
 Keywords: Inter Shaft Squeeze Film Damper (ISSFD), Gas Turbine Engine, Static stiffness, two spools, grooves.

Simulation of Volcanic Ash Ingestion Into a Large Aero Engine: Particle–Fan Interactions

Volcanic ash (VA) clouds in flight corridors present a significant threat to aircraft operations as VA particles can cause damage to gas turbine engine components that lead to a reduction of engine performance and compromise flight safety. In the last decade, research has mainly focused on processes such as erosion of compressor blades and static components caused by impinging ash particles as well as clogging and/or corrosion effects of soft or molten ash particles on hot section turbine airfoils and components. However, there is a lack of information on how the fan separates ingested VA particles from the core stream flow into the bypass flow and therefore influences the mass concentration inside the engine core section, which is most vulnerable and critical for safety. In this numerical simulation study, we investigated the VA particle–fan interactions and resulting reductions in particle mass concentrations entering the engine core section as a function of particle size, fan rotation rate, and for two different flight altitudes. For this, we used a high-bypass gas-turbine engine design, with representative intake, fan, spinner, and splitter geometries for numerical computational fluid dynamics (CFD) simulations including a Lagrangian particle-tracking algorithm. Our results reveal that particle–fan interactions redirect particles from the core stream flow into the bypass stream tube, which leads to a significant particle mass concentration reduction inside the engine core section. The results also show that the particle–fan interactions increase with increasing fan rotation rates and VA particle size. Depending on ingested VA size distributions, the particle mass inside the engine core flow can be up to 30% reduced compared to the incoming particle mass flow. The presented results enable future calculations of effective core flow exposure or dosages based on simulated or observed atmospheric VA particle size distribution, which is required to quantify engine failure mechanisms after exposure to VA. As an example, we applied our methodology to a recent aircraft encounter during the Mt. Kelud 2014 eruption. Based on ambient VA concentrations simulated with an atmospheric particle dispersion model (FLEXPART), we calculated the effective particle mass concentration inside the core stream flow along the actual flight track and compared it with the whole engine exposure.

Concepts of Gas Turbine Engine Improvement Through the Use of Choked-Flow Turbines

The new conceptual design of simple gas turbine engines for different purposes is presented. A choked-flow low-pressure turbine is used that allows reducing the number of stages and gas temperature in the relative motion ahead of blades. Application of such turbines with uncooled blades to reduce the cost of power-limited engines is considered.