© 2024 by the authors.The ratio between blade height and chord, named the aspect ratio (AR), plays an important role in compressor aerodynamic design. Once selected, it influences stage performance, blade losses and the stage stability margin. The choice of the design AR involves both aerodynamic and mechanical considerations, and an aim is frequently to achieve the desired operating range while maximizing efficiency. For a fixed set of aerodynamic and geometric parameters, there will be an optimal choice of AR that achieves a maximum efficiency. However, for a state-of-the-art aero-engine design, optimality means multi-objective optimality, that is, reaching the highest possible efficiency for a number of operating points while achieving a sufficient stability margin. To this end, the influence of the AR on the performance of the first rotor row of a multistage, multi-objective, high-speed compressor design is analyzed. A careful setup of the high-speed aerodynamic design problem allows the effect of the AR to be isolated. Close to the optimal AR, only a modest efficiency variation is observed, but a considerable change in compressor stability margin (SM) is noted. Decreasing the AR allows for increasing efficiency, but at the expense of a reduced surge margin. This allows the designer to trade efficiency for stability. Increasing the AR, however, is shown to reduce both the surge margin and efficiency; hence, a distinct optimality in stability is observed for the analyzed rotor blade row. In this work, optimality in the surge margin with respect to the AR is observed, whereas there is a close to optimal efficiency. The predicted range from AR = 1.10 to AR = 1.64 is only indicative, considering that the definition of multi-objective optimality requires balancing efficiency and the surge margin and that the choice of balancing these two criteria requires making a design choice along a pareto optimal front.

© 2024 by ASME.This work aims to provide a methodology for defining the design point for industrial gas turbine considering the economic, environmental, and engine performance aspects. The definition of the design point is a key step in the development project of a gas turbine since this definition involves the analysis of several operational points to verify if the desired performance can be obtained. Thus, to extend the methodology presented in the literature developed for micro-turbines to consider industrial gas turbines a computer program was developed in MATLAB®. This program is capable of performing thermodynamic calculations for design point definition and of performing single- and multi-objective thermoeconomic and thermodynamic optimizations using genetic algorithms. For the optimization process, total cost minimization, yield maximization, and gas turbine-specific work maximization were chosen as objective functions. The decision variables chosen were compressor pressure ratio, compressor polytropic efficiency, turbine polytropic efficiency, and maximum cycle temperature. For the calculation of economic aspects, fixed costs (equipment, installations, land acquisition cost, etc.) and variable costs (fuel, emissions, and operation and maintenance costs) were considered. The emission cost of NOx, CO, and UHC was considered for the environmental cost calculations. The thermodynamic calculations were based on enthalpy and entropy. The developed computer program was validated by simulating a commercial gas turbine and comparing the results obtained, also using a commercial program, GASTURB®. The presented optimization process shows results for a single objective, two objectives, and three objectives, where the results show a comparison between different design points obtained. The software developed will be of great assistance in the learning of engineering students.

© 2024 by ASME.The air conditioning system of executive, commercial, or military aircraft heavily relies on air cycle machines due to the availability of engine bleed air and their lightness and reliability compared to vapor cycle systems. The type of application, weight, refrigeration capacity, financial aspects, size, performance, and other specific design requirements drive the selection of suitable equipment for a particular aircraft. The motivation of this paper has been based on summarize the main concepts of the aeronautical environmental control system, as well as the mathematical aspects underlying the modeling of a simple/bootstrap air cycle unit in a software. The main aim is to develop software that can generate high level requirements that would be refined during the development phase of an aeronautical air conditioning system. It will be greatly benefit for engineers and students in the design of aeronautical air conditioning systems to better understand and to meet the design requirements. The results demonstrate the influence of the water-sprayer and chilled-recirculation system on air cycle performance and cabin inlet temperature, respectively. They also show changes in certain parameters of interest such as a function of altitude, power consumed by the secondary compressor, and air cycle machine fan. The computational model has proven to be a useful tool for performing parametric studies and evaluating critical points in designing and selecting an air conditioning unit based on a simple/bootstrap air cycle with humid air (any quantity of moist) as the working fluid.

© 2024 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.Decarbonization of the aviation sector is a key factor for current and future systems. Waste Heat Recovery (WHR) may be used to convert waste energy to electric power by using a bottoming cycle, which can reduce the overall fuel requirement of the airplane. One of the potential bottoming cycles for aircraft application is a Supercritical CO2 (sCO2) power system. The sCO2 power system has advantages because of the component compactness, which is crucial for aircraft integration. However, the main challenge for aircraft integration is the size and weight of the heat exchangers. The present work focuses on the performance of the Supercritical CO2 power system in both current and next-generation aircraft engines considering an innovative and advanced design of the sCO2 heat exchangers (cooler and primary heat exchanger). The first part of the work is focused on the analysis of the sCO2 WHR system for an aircraft engine. The second part of the work is focused on a detailed heat exchanger selection, design and optimization based on the aircraft engine parameters. The results show the potential of WHR utilization, which may generate an additional 100 - 200 kW. However, the heat exchangers may increase overall weight of the aircraft. For this reason, an advanced design is necessary.

© 2023, The Author(s), under exclusive licence to The Brazilian Society of Mechanical Sciences and Engineering.The Turbofan engine represents the type of propulsive technology mostly used in commercial aircrafts, and until that the new disruptive technologies take place, researches to optimize this propulsive system shall be continued to reduce the environmental impacts. The aim of this paper is to propose a methodology for the low-pressure system preliminary design (fan/low-pressure turbine), based on aircraft cruise thrust adjustment from commercial off-the-shelf turbofan engine, focusing on reducing specific fuel consumption for the individual aircraft mission. This work is carried out according to the following steps: (i) model development with calculation methodology for velocity diagram flow angles applied to the low-pressure system; (ii) estimation of baseline low-pressure system design parameters from limited engine data (an integrated engine aircraft model developed in the Gasturb and MATLAB commercial softwares are applied); (iii) evaluation of the strategies to increase the low-pressure system component efficiencies and their implementation by computer simulation; (iv) reapplication of the calculation methodology for estimation of the velocity diagram flow angles considering the adjusted low-pressure system components; and (v) analysis of the adjustment proposal results considering the matching between the lowest specific fuel consumption and the net thrust required for the cruise flight phase of the aircraft. As a final result, it demonstrates that the proposed strategies are promising for the adjustment of the low-pressure system in the preliminary design scope, and this approach may be considered feasibility from the standpoint of the engine manufacturer implementation, since the engine core and its external sizing do not affected.

© 2023 by ASME.Evaporative cooling systems are commonly used in thermoelectric plants to cool the air at gas turbines inlet, improving the performance of these engines. Normally, the evaporative cooling is modeled as adiabatic saturation and, in this case, the water-air equilibrium temperature depends only on the atmospheric air properties. However, other factors such as the water temperature that supplies the equipment and the ratio between the mass flow rates of water and air, also affect the equilibrium conditions of these systems. This work presents three methodologies to calculate the air temperature in equilibrium state, considering all the factors mentioned. The methodologies were implemented in a computer program written in FORTRAN. In all cases tested, the results obtained by the three models showed high convergence. As an example, for 70 different sets of inputs, the absolute and relative differences of the results were below 0.3236°C and 1.2480%, respectively. A statistical study, also on this sample of results, revealed that, for a confidence level of 99%, the hypothesis of the equivalence between the methods cannot be rejected.

© 2023, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.Correction Notice Reference 5 should be: L. Vesely, J. S. Kapat, C. Bringhenti, J. T. Tomita, M. F. Stoia, and K. Jui, “sCO2 Waste Heat Recovery System for Aircraft Engines,” AIAA 2022-1407. AIAA SCITECH 2022 Forum. January 2022. doi: https://doi.org/10.2514/6.2022-1407.

© 2023, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.The present work investigates the effect of reordering the nodes and elements of a grid according to the Hilbert curves on the cache utilization in an in-house parallel CFD code. A sorting algorithm is proposed based on domain decomposition techniques and the execution times are compared to those obtained by the structured grid format.

© 2023, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.Waste Heat Recovery is one of the key pathways to achieving reduced emissions and improving system efficiency. The Waste Heat Recovery (WHR) may be used to convert the waste energy to electric power by using a bottoming cycle. One of the potential bottoming cycles for aircraft application is a Supercritical CO2 (sCO2) power system. The sCO2 power system has advantages because of the component compactness, which is a key factor for aircraft integration. The present work focuses on the performance of the Supercritical CO2 power system in both the current and the next-generation aircraft engines considering the techno-economic evaluation of the bottoming cycle. The techno-economic evaluation needs to consider bottoming cycle integration and potential fuels, such as hydrogen, ammonia, or sustainable aviation fuel (SAF). The first part of the work is focused on the analysis of the sCO2 WHR system for an aircraft engine. The second part of the work is focused on a detailed techno-economic evaluation, including the capital, operation, and maintenance costs. The simulation was done using in-house computer programs for gas turbine performance and the sCO2 cycle. The results show the potential utilization of WHR in different operational regimes: idling on the ground, cruise, landing, and takeoff. The results show that the Waste Heat Recovery unit may generate an additional 100 - 200 kW. However, the additional power will require an additional cost for the system, approximately $ 2 Million.

Copyright © 2023 by ASME.The industry and the academy are continuously developing new approaches, technologies, and models for gas turbine design. However, there was not enough time to cover all the relevant subjects for undergraduate or graduate students in one or two-semester courses. So, in previous works, the authors described a developed interactive platform for the preliminary design of multistage axial flow turbines for uncooled blades and improved it based on the student’s feedback, so it could be as didactic as possible. Its application in the courses offered by the Turbomachines Department at Aeronautics Institute of Technology (ITA) successfully accelerated the learning process of the basics. In the graduate courses, the use of the program granted time to more complex topics, e.g., blade cooling, off-design performance, CFD simulations, manufacture, and machine learning applied to turbomachine design, which were not covered in previous years. The program initiates with the data from thermodynamic cycle calculation and the definition of the main design parameters. Then, it computes the aerothermodynamic properties of the flow stage-by-stage, from hub to tip, and the geometry of the blades. Finally, it estimates the losses by source, iteratively, through the models of Ainley and Mathieson [1], Dunham and Came [2], or Kacker and Okapuu [3]. This work presents some studies performed by the students using the platform. Firstly, it was varied some design key parameters such as loading and flow coefficients, the aspect ratio and the pitch-to-chord ratio of the blades, the airfoil section geometry, and the tip clearance, once at a time while maintaining the others. Then, it was possible to observe how these modifications affected the number of stages required, the stress levels, the machine size, and the isentropic efficiency, tracking the primary sources of loss. After, the students implemented other loss models, such as the one by Craig and Cox [4], aiming to analyze the effect of surface roughness on the losses. Finally, they compared the platform results with CFD simulations and experimental data from turbines developed at the Department. The paper concludes with the students’ insights through the project and comments on how the employed methodology improved their learning process.

Copyright © 2023 by ASME.The inducer is an axial pump that is part of the propellant injection system of Liquid Propellant Rocket Engines (LPRE). It is located at the inlet of the turbopump assembly and is critical for designing high performance LPREs. Its geometric and operational characteristics allow it to operate at low inlet pressures, delaying the appearance of cavitation and allowing the propellant tanks to operate at lower pressures. This allows the tanks to be lighter due to a reduced wall thickness requirement. The inducer also needs to operate harmoniously with the other components of the turbopump, especially with the main impeller which is located just downstream in the system. Therefore, it is important that the flow conditions at the inducer inlet and outlet are known and integrated with the turbopump and tank design. The present work aims to develop a methodology for inducer design based on literature established methods in order to obtain geometry and evaluate the flow conditions in liquid-propelled rocket engine inducer pumps. This work will assess outlet flow and pressure conditions in a way that it is possible to match them with the main impeller inlet. Performance criteria are evaluated in terms of the outlet pressure coefficient, flow coefficient and efficiency focusing exclusively on non-cavitating conditions. Two established analytical methods were implemented, one to provide inducer geometry in terms of system operational requirements and another, from National Aeronautics and Space Administration (NASA), for performance prediction based on geometrical and operational parameters. Further analysis is complemented by simulating the generated geometry in a CFD software. The methods were validated using published experimental data and the performances of the analytical, numerical and experimental results were compared. Results showed that the 3D turbulent CFD simulations provided very good agreement of efficiency. Satisfactory results were obtained for the general trends of characteristic curves over a range of flow rates and the spanwise distribution of key performance parameters near design point. The pressure coefficient was significantly overestimated. The results of the analytical models showed good agreement with simulated CFD results, indicating appropriate calibration of loss coefficients.

Copyright © 2023 by ASME.The use of inducers in turbopumps for liquid propellant rocket engines allowed operation at high rotational speeds, contributing to global vehicle performance improvement. Methods for designing inducers have been explored using analytical methods, experimental data and numerical solutions. The use of CFD for simulating and designing rocket turbopump inducers is a relevant practice because it can rapidly explore scenarios untested in the experimental endeavors for determining empirical functions. This technique also captures more problem details than reduced order analytical solutions. Turbomachines have a different accuracy in terms of solution prediction for different turbulence models and application. The flow specificity of the turbomachinery changes the adequate turbulence model to obtain a more accurate solution. For this reason, it is of interest to investigate how different turbulence modeling predicts the fluid flow behavior and their accuracy to calculate the inducers' performance. The present work aims to investigate the capability of different turbulence models on the performance and flowfield obtained via CFD simulations of an inducer pump. The CFD simulations were performed using a commercial software for k-ε, RNG k-ε and Shear Stress Transport turbulence models. The simulations were performed on a known inducer geometry for which published experimental performance data as a function of operational conditions is available in literature. The present paper discusses the differences in the performance prediction and the flow field calculated for the turbulence models simulated. For the cases studied, the k-ε standard model shows better predictions for the efficiency and head coefficient compared with the experimental data demonstrating their accuracy in solving rotational flows.

Copyright © 2023 by ASME.The boundary layer ingestion (BLI) concept has emerged as a novel technology for reducing aircraft fuel consumption. Several studies designed BLI-fans for aircraft. BLI-propellers, although, have still received little attention, and the choice of open-rotors or ducted propellers is still an open question regarding the best performance. The blade design is also challenging because the BLI-propulsors ingest a nonuniform flow. These aspects emphasize further investigation of unducted and ducted BLI-propulsors and the use of optimization frameworks, coupled with computational fluid dynamics simulations, to design the propeller to adapt to the incoming flow. This paper uses a multi-objective NSGA-II optimization framework, coupled with three-dimensional RANS simulations and radial basis function (RBF) metamodeling, used for the design and optimization of three propeller configurations at cruise conditions: (a) conventional propeller operating in the freestream, (b) unducted BLI-propeller, and (c) ducted BLI-propeller, both ingesting the airframe boundary layer. The optimization results showed a significant increase in chord and a decrease in the blade angles in the BLI configurations, emphasizing that these geometric parameters optimization highly affects the BLI-blade design. The unducted BLI-propeller needs approximately 40% less shaft power than the conventional propeller to generate the same amount of propeller force. The ducted BLI-propeller needs even less power, 47%. The duct contributes to the tip vortex weakening, recovering the swirl, and turning into propeller force, as noticed from 80% of the blade span to the tip. However, the unducted and ducted BLI-configurations presented a higher backward force, 26% and 46%, respectively, compared to the conventional propeller, which can be detrimental and narrow the use of these configurations.

© 2022, The Author(s), under exclusive licence to The Brazilian Society of Mechanical Sciences and Engineering.The costs involved in the design, manufacture, certification and maintenance of a helicopter have grown over the past few years. In the certification phase of embedded systems, their safety levels and their performance requirements are verified. The helicopter engine is a system that must be reliable and capable of providing the necessary power to produce lift and controllability for the aircraft. In this work was developed a computer model to evaluate the helicopter engine’s performance under any flight conditions and the pilot’s inputs. The developed software was incorporated as a module in a flight test simulator at the Flight Tests and Research Institute (IPEV) which belongs to the Brazilian Air Force. This simulation tool allows foreseeing and investigating possible situations that may occur during actual flight tests, improving safety and reducing costs. Using MATLAB® Simulink, it was possible to run at the same time: an iterative and a non-iterative methodology, a control system to set the fuel flow schedule, based on several inputs generated from the thermodynamic model. Based on classic thermodynamics laws and differential equations, the particularities due to the helicopter application were adjusted: the influence of the pilot’s commands; performance requirements; running line control; and the fuel flow control system. The simulation results were compared with commercial gas turbine performance simulation software and with the data provided by the IPEV in five real flight tests. These data were also used for obtaining engine output power requirements according to collective stick position.

© 2022, The Author(s), under exclusive licence to The Brazilian Society of Mechanical Sciences and Engineering.The tip clearance is the gap between the rotor blade row and its casing. In this region, a leakage flow on the rotor blade tip is induced by pressure differences from rotor blade pressure side to suction side, resulting in a loss in the turbomachine efficiency and drop in performance. High pressure turbines (HPT) operate in the limit of the energy transfer process with low-aspect ratio blades and high-pressure loading. The tip clearance loss is significant when compared with other loss sources. To minimize the performance drop, different desensitization techniques were tested for turbulent flow in steady state. First, the HPT developed by NASA in the Energy Efficient Engine (E3) program was studied with its original configuration of rotor tip, also called flat-tip. Then, the winglet was implemented on rotor tip geometry, for both suction and pressure sides. Numerical simulations using the computational fluid dynamics were performed, and the results are compared with experimental data for both cases. The results show that in general, for the same HPT pressure ratio, the use of winglet on the rotor tip pressure side achieved the best results showing an increase in efficiency of 1.025 % for 3.7 of pressure ratio. Even the winglet on the rotor tip suction side presented an efficiency increase of 0.625 % for 3.7 of pressure ratio compared with flat-tip rotor configuration. Overall, both winglet configurations obtained results better than the common rotor blade flat-tip geometry, for the same pressure ratio operational condition.

© 2022, The Author(s), under exclusive licence to The Brazilian Society of Mechanical Sciences and Engineering.For a long time, thermal power plants play important roles in world electricity and are expected to continue, at least, in the next decades. However, the finitude of fossil fuel sources leads to the crucial need for improving the existing power generation systems. In this study, an in-house computational code was developed and validated to evaluate the energy, exergy and economic performance for thermal power plants applications. Based on operating data of an actual lignite coal-fired steam power plant, two cycles were designed and compared. In the cycle in which more components were added, the fuel consumption was 9.44% lower to produce the same amount of power, making more effective use of the fuel resource. This substantial reduction in fuel consumption reflected lower electricity average costs for this plant. Comparing to the electricity price of a country using the same type of fuel, it was found that it could be lower by 1.62 percentage points for household consumers. Although the higher costs with capital investment and operational and maintenance (O&M) due to the addition of these components, the attractive economic performance of the cycle reduces the annual fuel costs and offsets the increase in capital and O&M costs.

© 2022 Elsevier Masson SASAxial turbines are machines widely used in different engineering applications. Due to their constructive characteristics, they must have a space between the rotor blades and the turbine casing, called tip clearance. Unfortunately, this gap allows a part of the fluid to leak from the pressure side to the suction side of the rotor blades. This leakage is undesirable and represents an energy loss. A way to avoid part of this loss is through the use of desensitization techniques. Although the use of these techniques is widely known, no studies in the open literature have evaluated these techniques in hydraulic turbines. This work presents a numerical analysis of squealer desensitization techniques applied in a hydraulic axial turbine. The turbomachine under study is the first stage of the hydraulic axial turbine used in the Low Pressure Oxidizer Turbopump (LPOTP) of the Space Shuttle Main Engine (SSME). Numerical simulations were performed using CFX v.19.2 software, and computational meshes were generated in ICEM v.19.2 software. Initially, the computational model was validated, using the experimental results published by the National Aeronautics and Space Administration (NASA). A parametric analysis was performed considering the variation in squealer cavity depth and rim thickness. The study found that the squealer cavity depth has a greater influence on the stage performance than its rim thickness. The tendency is that the greater the cavity depth, the greater the stage efficiency. One of the squealer geometries analyzed allowed an average increased efficiency of 1.43%, over the entire turbine operational range. The results obtained also show that the application of the proposed geometries would enable the reduction in cavitation close to the trailing edge of the rotor blades. This result is extremely valuable, as it can impact the life cycle of the turbine.

© 2021 Elsevier Inc.The present work describes the implementation of an implicit time-integration numerical scheme to solve viscous flows in an in-house CFD solver. The scheme is developed to calculate engineering problems involving compressible flows. This work extends the defect-correction technique for the 3D flow calculations, and all mathematical formulations are described. The CFD solver is based on the finite-volume method (FVM) to calculate the three-dimensional flow and can be applied to solve unstructured meshes. The current implementation uses the Flux-Difference Splitting method (FDS) developed by Roe combined with the MUSCL method and the Venkatakrishnan flux limiters to provide better accuracy of the numerical solutions. The implicit time-integration scheme was linearized applying the backward Euler method on the left-hand side (LHS) and a Newton-type linearization on the right-hand side (RHS) of the governing equations. The Jacobian matrix was computed analytically for the inviscid fluxes using the Roe fluxes, and for the viscous fluxes differentiating the conservative vector. Earlier work by Cavalca et al. (2018) showed the robustness and accuracy of this implicit solver to predict inviscid flows over the airfoil and into the supersonic nozzle. Finally, the Gauss-Seidel (GS) iterative method was applied to solve the resultant sparse and large system of equations. These numerical schemes and methods were applied to solve the laminar flow over a flat plate. Afterwards, the numerical solution was validated and verified with the exact Blasius solution. From the results, the numerical simulations exhibited superior robustness of the implicit-defect correction scheme when compared with the explicit scheme for compressible flows. All numerical particularities and their implementations are detailed in this paper.

© (2022) by International Council of Aeronautical Sciences (ICAS) All rights reserved.Variable cycle engines promise to enable adaptive cycles that give close to optimal performance over a wide range of conflicting mission requirements, such as low altitude high speed flight and supercruise still providing excellent range. Modelling such engines pose challenges for general purpose software since variable geometry gas paths modify the underlying set of equations being solved. It is possible to use multiple engine models transferring design data between the models. This, however, creates a high risk for inconsistency and modelling error. It is more attractive if the solutions obtained could be determined using the same model. In this work an in-house software was developed to model an Adaptive Cycle Engine (ACE). This development was used to show how variable cycle mode switches can be integrated into general purpose performance tools. The variable cycle engine uses a FLADE, which is a "fan on blade" component, to extend its range and to provide improved subsonic performance. The individual impact of the components, its effect on propulsion performance parameters and in the engine installation were analyzed as the main results. The contribution from this paper is thus two-fold, firstly the paper goes ahead and proposes new methods for the simulation of mode switching in generic performance tools by introducing dynamic equation systems. Secondly, the paper then studies the FLADE component and its potential performance benefits if added to a conventional turbofan architecture.

© (2022) by International Council of Aeronautical Sciences (ICAS) All rights reserved.An Axial Turbine Blade Tip has a great influence on its flow behavior and performance. Due to the clearance between the turbine casing and the rotor blades tips, part of the flow leaks from the pressure side to the suction side. This leakage reduces the turbomachine efficiency, and therefore must be minimized. Over the years, the use of desensitization techniques has proven to be an excellent strategy for reducing this unwanted flow. These techniques, however, has only been studied in machines that operate with compressible fluids. The objective of this work is to verify the effects of two Winglet geometries in the first stage of the Liquid Oxygen (LOX) Turbine used as booster in the Space Shuttle Main Engine (SSME). The two Winglet geometries evaluated have identical thickness and width, being differentiated by their trailing edge region configuration. In this region, the first geometry (W1) connects to the trailing edge with an angle close to 90°, while the second geometry (W2) presents a smooth connection. The results obtained show that it is possible to improve the stage efficiency depending on the geometry adopted, as well as to analyze the cavitation phenomenon. The mesh generation and simulations were done using a commercial software and the 3D flow calculations were based on the Reynolds Averaged Navier-Stokes (RANS) equations.

Copyright © 2022 by ASME.The gas turbine industry requires extensive knowledge in several areas of engineering, and since both industry and academy continuously develop new approaches, technologies, and models, usually, there is not enough time to cover all the relevant subjects in one or two-semester courses for undergraduate or graduate students. In previous work, the authors have presented an interactive platform for the preliminary design of single-stage axial turbines with uncooled blades, for use at the undergraduate courses offered by the Turbomachine Department at Aeronautics Institute of Technology to accelerate the learning process. The present work aims to present an expansion of this interactive learning platform, with the inclusion of a module for the thermodynamic cycle study, a module for off-design calculations, and the generation of a PDF file containing the step-by-step solution memorial with all the equations and values used in the design. The work also presents a structure for the conduction of a graduate course in turbomachines focused on the design of axial turbines. It comprehends theory and exercise classes, oriented study with the interactive learning platform, and a project in which the students have to implement some of the modules and run test cases. The authors observed more interest of the students and higher quality questions in the classes while using the interactive platform or programming, developing a better understanding of the design process until the end of the course. Also, while, in previous semesters, the preliminary design occupied almost half of the 48-hour course, it took only 12-hour to cover the same subject, granting time to more advanced topics, such as blade cooling, off-design performance and computational fluid dynamics simulations.

Copyright © 2022 by ASME.Passive wall treatments with circumferential grooves in axial compressors proved to be effective in increasing the compressor stall margin in previous researches by creating a resistance to the flow that leaks in the tip clearance region of the compressor, from the rotor blade pressure side to the suction side. In the present work, a passive wall treatment with circumferential grooves was implemented in a multi-stage axial compressor. Different configurations of circumferential grooves were created at the casing of the first rotor row used in a four-stage axial flow compressor. 3D CFD flow simulations were performed in order to evaluate all the specified configurations aiming to find improvements on compressor stall margin. Investigations on the compressor flow characteristics were realized and the stall margin variations were determined. The numerical simulations were performed based on the Reynolds-Averaged Navier Stokes equations and the turbulence model was the k-ω SST. After the simulations, several rotational speeds of the compressor map characteristics, including the design-point rotational speed, were obtained for the case without casing treatment (smooth wall case) and for the case with circumferential grooves. In the results, passive wall treatment with circumferential grooves demonstrated an improvement in the compressor stall margin, especially for N=0.60 and N=0.90 rotational speeds.

© 2022, American Institute of Aeronautics and Astronautics Inc, AIAA., All rights reserved.This work investigates the use of solid and permeable surfaces in the Ffowcs WilliamsHawkings (FW-H) analogy for predicting high-speed propeller noise. The CFD/CAA methodology encompasses unsteady Reynolds-Averaged Navier-Stokes simulations to compute the flowfield on the acoustic surface applied in the FWH analogy to obtain the noise signatures in the far-field. Furthermore, this manuscript also investigates the effects of the downstream end-cap position, on the propeller noise prediction, by using two permeable surfaces with different lengths to assess the propeller noise levels in each case. The former is a short SFW-H surface placed near the rotor, and the latter, namely the LFW-H, is a surface larger in length where the end-cap grid is placed farther downstream from the rotor. The results showed the capability of the permeable surface technique for predicting the noise with higher accuracy than the solid formulation, especially at the first blade passing frequency. Also, the larger LFW-H surface performed better than the SFW-H surface. A reason that could justify this is that the LFW-H end-cap surface is placed at a suitable distance downstream from the propeller. Therefore, the LFW-H surface can include more of the contributions of the non-linear effects or quadrupole sources enclosed within the permeable source surface region.

© 2022, American Institute of Aeronautics and Astronautics Inc.. All rights reserved.Waste heat recovery is a key pathway to achieving reduced emissions and improved system efficiency. Waste heat can potentially be converted to electric power by several methods. One of the most effective methods is based on using a supercritical CO2 waste heat recovery power system. The sCO2 power system has advantages because of component compactness, which is an important consideration for aircraft integration. The present work focuses on implementing the supercritical CO2 power system into both current and next-generation aircraft engines that may use different fuels, such as hydrogen, ammonia, or sustainable aviation fuel (SAF). The first part of the work is focused on detailed optimization of the sCO2 waste heat system for a real aircraft engine with two sCO2 cycle configurations. The second part of the work is focused on detailed design of the heat exchangers, including weight and pressure drop calculation. The simulation was done using an in-house computer program for gas turbine performance and for the sCO2 cycle. The results show the potential utilization of waste heat in different operational regimes: idling on the ground, cruise, landing, and takeoff. One engine (nominal thrust of 9kN) with two different waste recovery units are investigated. The results demonstrated that the waste heat unit could generate an additional 100-200 kW for the 9-kN-engine (under cruise operation), which may reduce fuel consumption, even if the sCO2 system weight is around 800 lbm / 364 kg.

© 2021 Elsevier LtdThe requirements of Liquid Propellant Rocket Engine (LPRE) are high for thrust, specific impulse, and flow rate; thus, its components also have strict requirements. For the turbopumps (TPs), this means high flow rate, high rotational speed, and high pressure ratio, which makes their operations susceptible to the cavitation phenomenon, as observed in two previous works. In the first, cavitation regions were observed in the first stage of the Space Shuttle Main Engine (SSME) Liquid Oxygen (LOX) booster turbine, for 3.0, 5.5, and 8.0% tip clearances (relative to rotor blade height), using monophase flow (Lindquist Whitacker et al., 2017). In the second, the simulations were performed with multiphase flow, producing results more physically coherent for the 3.0% gap configuration (Whitacker et al., 2018). The characteristics of both types of simulations in space propulsion applications still require better understanding. Therefore, to compare monophase and multiphase results at various operating points and turbine configurations, steady-state turbulent 3-D Computational Fluid Dynamics (CFD) simulations were performed, based on Reynolds-Averaged Navier-Stokes (RANS) formulation. The same three tip configurations for the turbine first stage were simulated, and the calculations were validated using experimental results from the National Aeronautics and Space Administration (NASA) (Boynton and Rohlik, 1976). This made it possible to verify the effect of the tip clearance on the machine performance and internal flowfield. When the gap increased, the pressure loading decreased in a large region of the blade tip, the interaction was greater between the Tip Clearance Vortex (TCV) and a vortex generated around the shroud cavitation region (Cavitation Vortex - CV), and this interaction moved towards the middle of the blade-to-blade passage. Thus, the losses increased and the efficiency decreased. Various comparative aspects between the simulations using both mono and multiphase numerical schemes are also discussed.

© 2021 by GE Research.The aim of this work is an evaluation of different turbulence models applied in Computational Fluid Dynamics (CFD) techniques in the turbomachinery area, in this case, in an axial turbine stage used in turbopump (TP) application. The tip clearance region was considered in this study because it has a high influence in turbomachinery performance. In this region, due to its geometry and the relative movement between the rotor row and casing, there are losses associated with vortices and secondary flow making the flowfield even more turbulent and complex. Moreover, the flow that leaks in the tip region does not participate in the energy transfer between the fluid and rotor blades, degradating the machine efficiency and performance. In this work, the usual flat tip rotor blade geometry was considered. The modeling of turbulent flow based on Reynolds Averaged Navier-Stokes (RANS) equations predicts the variation of turbine operational characteristics that is sufficient for the present turbomachine and flow analysis. Therefore, the appropriate choice of the turbulence model for the study of a given flow is essential to obtain adequate results using numerical approximations. This comparison become important due to the fact that there is no general turbulence model for all engineering applications that has fluid and flow. The turbomachine considered in the present work, is the first stage of the hydraulic axial turbine used in the Low Pressure Oxidizer Turbopump (LPOTP) of the Space Shuttle Main Engine (SSME), considering the 3.0% tip clearance configuration relative to rotor blade height. The turbulence models evaluated in this work were the SST (Shear Stress Transport), the k-e Standard and the k-e RNG. The computational domain was discretized in several control volumes based on unstructured mesh. All the simulations were performed using the commercial software developed by ANSYS, CFX v15.0 (ANSYS). All numerical settings and how the boundary conditions were imposed at different surfaces are explained in the work. The boundary conditions settings follow the same rule used in the test facility and needs some attention during the simulations to vary the Blade-Jet-Speed ratio parameter adequately. The results from numerical simulations, were synthesized and compared with the experimental data published by National Aeronautics and Space Administration (NASA), in which the turbine efficiency and its jet velocity parameter are analyzed for each turbulence model result. The work fluid considered in this work was water, the same fluid used in the NASA test facility.

© 2020The present article aims to implement and investigate a different preconditioning method based in a three-dimensional in-house compressible CFD code that ensures the robustness and numerical stability to determine the flowfield considering low Mach number flow. The present preconditioning method involve two different methodologies developed by references [8, 32]. The CFD solver was developed to calculate the Euler and Navier-Stokes equations, numerically, for steady-state regime based on the cell-centered finite volume method (FVM) using Reynolds Averaged Navier-Stokes equations (RANS). The centered second-order scheme was used for the discretization of convective terms from momentum equations. The explicit second-order five-step Runge-Kutta scheme was employed for the time-marching procedure, using an implicit residual smoothing technique to enhance the numerical stability. A local preconditioning method was implemented due to its robustness in predicting low Mach number flows in a compressible CFD code environment. However, for low Mach number flows, near stagnation points, numerical perturbations were amplified generating a stiffness in the convergence rate, which provided an inaccurate solution and numerical stability degradation. Aiming to improve the preconditioning robustness, a flux function and a new limiter were applied to operate with the preconditioning technique based on a pressure sensor. Those corrections re-scale the eigenvectors and ensure the locality of the algorithm, which improves the numerical stability and guarantees the convergence for low-speed flows. The inviscid flow over a NACA 0012 airfoil for compressible and incompressible cases shown accurate and robust solutions. For a viscous flow over a flat plate case in the compressible and incompressible cases, the preconditioning technique purposed in this work supplied good numerical solution in agreement with the analytical solution.

© 2020 Elsevier LtdGrowing environmental concerns are driving the energy market toward the development of thermodynamic cycles to harness renewable energy and waste heat. This manuscript introduces the novel organic Rankine flash cycle, which combines the organic Rankine cycle with the trilateral cycle, merging their advantages in terms of high specific power output and low heat transfer irreversibility, respectively. By comparing the organic Rankine flash cycle to the organic flash cycle, it was found that the proposed architecture reaches a peak exergy efficiency at a more realistic value of two-phase expansion volume flow ratio, consistently achieves higher energy and exergy efficiencies, presents a lower cost, and is not constrained to operate close to the working fluid saturation temperature, promising easier operability. Considering pentane as working fluid, the exergy efficiency of the organic Rankine flash cycle is 18%p higher for a heat source temperature of 150 °C, 12%p for 175 °C, and 4%p for 200 °C. The attractive thermoeconomic performance of the proposed organic Rankine flash cycle highlights the potential of such a cycle as a new paradigm in the ORC panorama, encouraging further investigation towards practical demonstration.

© Royal Aeronautical Society 2019.This paper presents a review of the various aeronautical air conditioning systems that are currently available and discusses possible system configurations in the context of the aeronautical environmental control systems. Descriptions of the standard vapor compression cycle and air cycles are provided. The latter includes, simple-cycle, bootstrap-cycle, simple-bootstrap cycle (3-wheel) and condensing cycle (4-wheel). Water separation and air recirculation systems are also explored. A comparison between vapor compression cycles and air cycles is provided, as well as a comparison between different air cycles. Air cycle units are far less efficient than vapor compression cycle units, but they are lighter and more reliable for an equivalent cooling capacity. Details regarding the aircraft conceptual design phase along with general criteria for the selection of an air conditioning system are provided. Additionally, industry trends and technological advances are examined. Conclusions are compiled to guide the systems engineer in the search for the most appropriate design for a particular application.

© 2019 Elsevier LtdIn an increasingly decentralized energy market, micro gas turbines are seen with great potential due to their low emissions and fuel flexibility, which aligns with growing environmental concerns. Although presenting a relatively low efficiency, these machines could be improved by coupling it with an organic Rankine cycle. This manuscript covers the thermoeconomic design and optimization of such bottoming cycle for a 100 kWe micro gas turbine. The tool employed for such calculations is extensively described and was developed using solely open resources. The results shown that the saturation temperature at ambient pressure was an important variable when the minimum pressure is constrained above ambient and that a high degree of superheating was favored when the recuperated cycle is heated directly by the microturbine flue gases. Pentane was flagged as the best working fluid, generating 14.1 kWe of additional power and increasing the overall electric efficiency from 30 to 34.2%. The Authors show that at the current state of the art an efficiency of around 35% is the upper practical limit for such microturbine organic Rankine cycle combination.

Copyright © 2020 ASMEThis work presents the implementation of an interactive learning platform for turbine design in an engineering teaching environment. Due to the abundance of strategies and problems encountered in a multidisciplinary iterative design process, presenting the student to the multitude of scenarios can be a laborious and time-consuming task, often not possible in one-semester courses for undergraduate students. The developed computational program breaks down the preliminary design methodology into a step-by-step analysis of a single-stage axial turbine for aeronautical application. In it, the student is guided through velocity diagram construction, performance prediction, tridimensional and compressible effects considerations, blade designing as well as accounting for losses. In this interactive learning tool, it is possible to explore the sensitivity and effects of each design choice at various design steps, generating insight and hopefully a more intimate understanding. This exploration generates real-time changes in the output interface, for example the velocity diagrams and major geometrical features, in which the student is able through different trials to observe and compare the impact of different approaches, choices and assumptions. The program is written in Python language and the loss models chosen were Kacker and Okapuu; Dunham and Came; and Ainley and Mathieson. As the same set of design requirements can lead to different - yet optimal - configurations, the student will be given guidelines based on established design methodologies with the aid of graphs and the usual ranges of the calculated parameters found in practice. At the end of this process, the student is able to harvest a final design from which it is possible to generate discussions among a class or examine the suitability of a final product in regards to a proposed assignment, objective or application.

Copyright © 2020 ASMEFilm cooling is an important technique to ensure safe operation and performance fulfillment of turbines. Its ultimate goal is to protect the axial turbine blades from high gas temperatures. An appropriate study is necessary in order to obtain a reliable representation of the flow characteristics involved in such phenomena. Because of the high computational cost of high-fidelity simulations, the low-fidelity simulation method Reynolds Averaged Navier Stokes (RANS) is commonly used in practical configurations. However, the majority of the current turbulent heat flux models fail to accurately predict heat transfer in film cooling flows. Recent work suggests the use of machine learning models to improve turbulent closure in these flows. In the present work, a machine learning model for spatially varying turbulent Prandtl number previously described in the literature is applied to a transverse film cooling flow consisting of a jet square channel. The results obtained in the present work were compared to adiabatic effectiveness experimental data available in the literature to assess the performance of the machine learning model. The results shown that for low blowing ratios (BR =0.2 and BR = 0.4) the proposed machine learning model has poor performance. However, for the case with the highest blowing ratio (BR = 0.8), the proposed model presented better results. These results are then explained in terms of the resulting turbulent Prandtl number field and suggest that the training set is not appropriate for capturing the turbulent heat flux in fully attached jets in crossflow.

Copyright © 2020 ASMEThe aim of this work is the evaluation of different mesh types applied in turbomachines area, in this case in an axial turbine stage used in turbopumps (TP) applications. The tip clearance region was considered in this study because it has high influence in turbomachines performance. Due to the complexity of the tip clearance region, structured mesh generation is not always feasible, therefore it is necessary to generate unstructured meshes that allow flow calculation through Computational Fluid Dynamics (CFD) techniques. The use of different mesh type is an interesting topic when different rotor tip geometries are evaluated, in which the desensitization methods are applied. In this work, only the common flat-tip was consider. Thus, as a first step, unstructured tetrahedral meshes (with prismatic layers close to the surfaces) with different y+ values were generated. After this, turbulent 3-D flow calculations were performed at design and off design conditions, based con Reynolds Averaged Navier-Stokes (RANS) equations. The methodology used is to present in a didactic way, for under and graduate students, the advantages and disadvantages of the unstructured mesh in relation to the structured one, already used in previous research. Unstructured meshes were generated using ICEM software (ANSYS), while structured ones were generated using AxCent software developed by CONCEPTS NREC. The machine under study is the first stage of the hydraulic axial turbine used in the Low Pressure Oxidizer Turbopump (LPOTP) of the Space Shuttle Main Engine (SSME), considering 3.0% tip clearance configuration relative to blade height. All simulations were done using CFX program (ANSYS). The result shows the comparison between the two mesh types considering the difficulty and time generation, discretization quality, effect of y+ parameter variation on flowfield, simulation time, and stage performance parameters calculation for different operating points.

© 2020 ASMENumerical simulations were carried out with the purpose of investigating the effect of applying circumferential grooves at axial compressor casing passive wall treatment to enhance the stall margin and change the tip leakage flow. The tip leakage flow is pointed out as one of the main contributors to stall inception in axial compressors. Hence, it is of major importance to treat appropriately the flow in this region. Circumferential grooves have shown a good performance in enhancing the stall margin in previous researches by changing the flow path in the tip clearance region. In this work, a passive wall treatment with four circumferential grooves was applied in the transonic axial compressor NASA Rotor 37. Its effect on the axial compressor performance and the flow in the tip clearance region was analyzed and set against the results attained for the smooth wall case. A 2.63% increase in the operational range of the axial compressor running at 100%N, was achieved, when compared with the original smooth wall casing configuration. The grooves installed at compressor casing, causes an increase in the flow entropy generation due to the high viscous effects in this gap region, between the rotor tip surface and casing with grooves. These viscous effects cause a drop in the turbomachine efficiency. For the grooves configurations used in this work, an efficiency drop of 0.7% was observed, compared with the original smooth wall. All the simulations were performed based on 3D turbulent flow calculations using Reynolds Averaged Navier-Stokes equations, and the flow eddy viscosity was determined using the two-equation SST turbulence model. The details of the grooves geometrical dimensions and its implementation are described in the paper.

© The Authors, published by EDP Sciences, 2019.Current energy conversion machines such as the micro gas turbine can be improved by harvesting the low-grade energy of the exhaust. A prominent option for such is the organic Rankine cycle due to its relatively efficient and reliable design. This manuscript presents a review on the subject and is the first step toward the design of an organic Rankine cycle bottoming a 100 kWe recuperated gas turbine. After introducing and covering the historical development of the technology, appropriate guidelines for defining the cycle arrangement and selecting the fluid are presented. At last, the viability of the cycle is assessed by assuming an appropriate efficiency value and general cost functions. The organic Rankine is expected to generate an additional 16.6 kWe of power, increasing the electrical efficiency from 30 to 35%. However, the capital cost increase was estimated in 48%.

© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.This work presents and analyzes the suitability of using differential GPS equipment (DGPS) for the determination of air data system errors by means of flight testing of the H-50 (AS 350) helicopter. Since the Pitot-error cannot be undervalued, two existing methodologies that use the DGPS as a data acquisition tool have been adapted for the usage in helicopters. Results of flight tests using a validated technique (tower flyby and ground speed course) are presented for the purpose of comparison with data obtained from both proposed techniques. Data are processed through mathematical equations and specific code in MatLab® platform, which has been adapted specifically for this research. Two methodologies using DGPS are set: the first needs three stabilized legs for each speed within the desired flight envelope whose trajectory describes a path similar to a clover leaf, which suggests the method's name: "cloverleaf"; The second uses acceleration and deceleration in 90 degrees alternated legs forming a windbox-like design. The data reduction is performed by iterations of the estimated parameters until a convergence criterion is reached using the analysis of the outputs. Results show the suitability of both proposed methods from a quantitative point of view. Qualitatively, the windbox maneuver with parameters estimation from the analysis of output-error allows reducing material resources expenditures without quality degradation.

© 2019, Avestia Publishing.In an effort to ensure the robustness and numerical stability of a three-dimensional explicit compressible code for all speed flows, a preconditioning technique was implemented. The code solves Euler steady-state equations into a three-dimensional flow. Local preconditioning was implemented due to their accuracy in predicting lift and drag forces on mixed flows. However, for low speed flows near stagnation points numerical perturbations are amplified, generating a loss in the convergence rate, code accuracy and robustness. Aiming to improve the preconditioning accuracy and convergence rate suggested a new limit to the preconditioning sensor based on the flow pressure. Numerical simulations of a subsonic flow over a cylinder showed a faster convergence rate when the preconditioning technique was implemented.

© 2019, Avestia Publishing.In this paper are presented a preconditioning technique to be implemented in a three-dimensional explicit compressible code to solve a turbulence flow to steady state regime. A local preconditioning technique with accurate predictions of mixed speed regimes is implemented in the original code, however, for low flow Mach numbers in the boundary layer region the numerical accuracy is lost to the preconditioning code. To improve the numerical solution are suggested a new limit to the preconditioning sensor based on a pressure sensor and is established an explicit flux function to evaluate the preconditioning sensor in the cell fluxes. The preconditioning code is validated for a supersonic case in nozzle and then to a subsonic case is studied the convergence rate for a low Mach number flow. Numerical solutions demonstrated that the changes applied in the original code improves the accuracy and robustness of the code for low speed flows.

© 2019, Avestia Publishing.A computational program used to calculate the preliminary design of axial turbines which uses the Kacker and Okapuu's loss model was modified to improve the losses predictions implementing Tournier and El-Genk's loss model. The in-house program was written in FORTRAN 90 and is based on the meanline technique to calculate the axial turbine. As the losses interfere in the geometrical calculations applying more accurate loss models, the predictions of the preliminary design are more reliable. The program was applied to design a single-stage turbine and the results are compared with the commercial turbomachine design software AXIALTM®. The improvements obtained by applying a more recent loss model have been discussed as the future works to improve the in-house program.

© 2019 Inderscience Enterprises Ltd.This manuscript provides an exergy-based parallel between combined- A nd steam-cycle power plant configurations burning blast furnace gas (BFG). The combined cycle (CC) was based on a currently operational power plant located in Rio de Janeiro, Brazil. The steam cycle (SC) was created by replacing the gas turbines (GTs) for steam generators (SGs) that handled the same amount of fuel. The results show that the combined cycle achieved 21.25% higher exergy efficiency, although emitting twice as much nitrogen oxide. The combined cycle generated 52.08% less steam while wasting 78.86% less exergy, which indicated that steam generators benefit from a higher amount of excess air. The gas turbine combustion chamber high exergy efficiency indicates that burning low-grade fuels is beneficial for reducing the intrinsic waste of chemical reactions. However, the compression process required prior to combustion undermines this benefit. Ultimately, this manuscript provides a comparison between two options to avail blast furnace gas.

Copyright © 2018 by ASME.This work describes a methodology used for counter-rotating (CR) propellers performance estimation. The method is implemented in an in-house program for gas turbine performance prediction, making possible the simulation of the counter-rotating open rotor (CROR) architecture. The methodology is used together with a variable geometry compressor control strategy to avoid surge conditions. Two cases are simulated under transient operation for both fixed and variable geometry compressor. The influence of the variable geometry control on the transient performance of CROR engines is evaluated and a comprehensive understanding on the transient behavior of this type of engine could be obtained. It is shown that the use of the variable geometry compressor control does not significantly affect the overall engine performance, while avoiding the surge conditions, thus ensuring the engine operation safety.

© 2018 Elsevier Inc.This paper reports the development and convergence analysis in steady-state of an effective and robust implicit finite-volume solver for compressible Euler equations on three-dimensional unstructured grids. A second-order upwind scheme (MUSCL) was employed based on Roe's approximate Flux-Difference Scheme (FDS) by using Venkatrakishnan flux limiters. The construction of the linear system for the implicit scheme was performed by applying the backward Euler on the left-hand side of the conservation equation and Newton-type linearization on the right-hand side. The Jacobian matrix that resulted from the linearization process was computed analytically using Roe flux terms. In this phase, the defect-correction technique was employed allowing effective time-dependent computations by an implicit time-integration scheme. In this approach, the flux integral on the right-hand side is computed based on a high-order of accuracy whilst the left-hand side the Jacobian is performed based on the low-order. The resulting sparse and large system of linear equations is solved by a sequential Gauss–Seidel iterative method. Simulations were performed and the developed implicit defect-correction solver was validated and verified. In addition, convergence analysis comparing the implicit solver and the explicit Runge–Kutta of 5-steps using Implicit Residual Smoothing (IRS) were performed showing the significant speed-up of the implicit solver over the explicit one. Simulations were performed for case studies to demonstrate the robustness of the developed implicit defect-correction solver in solving typical problems of aerodynamic involving transonic condition and shock wave captures for internal and external flows. Finally, the main particularities of the implicit scheme were investigated and discussed considering the simulation results, showing also its capacity to serve as an effective preconditioner (start-up method) to other implicit techniques.

© 2018 by ASME.This paper describes a methodology used for propeller performance estimation, which was implemented in an in-house modular program for gas turbine performance prediction. A model based on subsonic generic propeller maps and corrected for compressibility effects, under high subsonic speeds, was proposed and implemented. Considering this methodology, it is possible to simulate conventional turboprop architectures and counter-rotating open rotor (CROR) engines in both steady-state and transient operating conditions. Two simulation scenarios are available: variable pitch angle propeller with constant speed; or variable speed propeller with constant pitch angle. The simulations results were compared with test bench data and two gas turbine performance commercial software packages were used to fulfill the model validation for conventional turboprop configurations. Furthermore, a direct drive CROR engine was simulated using a variable inlet guide vanes (VIGV) control strategy during transient operation. The model has shown to be able to provide several information about propeller-based engine performance using few input data, and a comprehensive understanding on steady-state and transient performance behavior was achieved in the obtained results.

© 2018, The Brazilian Society of Mechanical Sciences and Engineering.The association of turbochargers with piston engines is widespread since both the efficiency and the power output of an engine could be improved. However, a piston engine operational range is wide and highly variable. This characteristic imposes challenges for the project and application of a turbocharger that should perform properly within the operational range. An important tool used to evaluate the performance of both turbine and compressor, which compose a turbocharger, is the characteristic map. The map condenses the main performance parameters into a single graphic that allow the evaluation of the machine characteristics, such as the operational width. A typical characteristic map relates the pressure ratio, mass flow rate, rotation and efficiency for each operational condition. The present work provides a technique to obtain the characteristic map of a turbocharger centrifugal compressor with reduced time consumption through steady state simulation using a fully unstructured mesh. Evaluation of the results indicated good accuracy for the predicted mass flow rate and pressure ratio. However, the resulting efficiency presented considerable discrepancy, which was aggravated when simulating extreme operational conditions or when the mass flow was used as a boundary condition. At last, the porter shroud and volute were evaluated within the entire range to provide an insight into the compressor operation.

© 2018, The Brazilian Society of Mechanical Sciences and Engineering.Advances in turbofan engine technology have led to engines with growing bypass ratios and lower fan pressure ratios, increasing the complexity of the in-flight thrust determination. Thrust values cannot be directly measured in flight; therefore, ground-level test are carried out, and the results calculated from thermodynamic properties of the gas are compared to the force exerted by the engine on the test bench. The result of this comparison is a scalar that is applied to the fan pressure ratio, fan pressure correlation, which attempts to minimize the error between the measured and calculated values. After the thermodynamic properties of the gas are measured during in-flight tests and together with the fan pressure correlation are used to calculate the in-flight thrust. The calculation procedure is implemented through VISUAL BASIC scripts, in the MICROSOFT EXCEL® environment. These scripts are used to calculate the generated thrust and the mass flow that go through the engine from the thermodynamic properties of the gas obtained from a high-fidelity numerical simulation of this engine. These results are then validated against the thrust and mass flow values calculated by this model. An analysis of the free-stream suppression effects on thrust is carried making use of these scripts.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.The pursuit of lower fuel consumption for aircraft is promoting a departure from contemporary arrangements. One example is the development of more synergetic airframe and propulsion system designs, which are expected to increase significantly aircraft efficiency mainly by means of boundary layer ingestion. By integrating propulsion and airframe, both systems will significantly impact each other. This mutual interference requires the development of novel performance evaluation methods that consider such effects. This manuscript introduces a propulsive efficiency equation for boundary layer ingestion propellers based on the power balance method. Two formulations are presented for numerical and analytical evaluations. The equation is bounded between 0 and 1 and allows a meaningful evaluation of shaft to propulsive powers conversion, which results in an accurate determination of thrust and drag. This manuscript is the first advance of a project that will develop an optimizing tool for boundary layer ingestion propellers based on computational fluid dynamic simulations. The results will be presented in subsequent manuscripts.

© Copyright 2018 ASME.Boosters are commonly used in liquid propellant rocket engines (LPRE) to allow lower propellant pressures in their storage tanks and, thus, smaller structural masses, contributing to cavitation free operation in the subsequent main turbopumps (TP). Boosters can be identified as key components for the overall performance of large engines, and if their operating requirements are stringent, they can operate under cavitation. Thus, effective design and performance tools are fundamental to design the components of these boosters considering this phenomenon. The simulation techniques based on turbulent and multiphase 3-D Computational Fluid Dynamics (CFD) were used in this work at steady state regime. The simulations were done using the commercial software CFX from ANSYS® Workbench. The study was conducted analyzing the performance of the first stage of the hydraulic axial turbine of the liquid oxygen (LOX) booster of the Space Shuttle Main Engine (SSME), at various operation points under cavitation, considering 3.0% tip clearance relative to blade height. The results obtained for, the performance parameters of this stage were compared with those obtained through monophase simulation, and the multiphase technique showed results closer to the experimental ones around the design point (DP), with increased simulation times acceptable for the computational resources currently available. Moreover, the results from the current work show the importance of considering the effects of cavitation through multiphase flow in hydraulic turbines.

Copyright © 2018 ASME.In modern gas turbine engines, many sophisticated cooling schemes are used to maintain the turbine blade temperature in acceptable levels. These schemes, such as convective cooling, film cooling, impingement cooling and the use of pin fins, can be combined to increase the cooling effectiveness. Jet impingement cooling, pin fins and convective cooling are internal cooling techniques, in which the cooling is achieved based on coolant flow through internal blade channels decreasing the blade metal temperature. Film cooling is an external cooling technique, in which the cold fluid (air) is injected into the hot gas flow through discrete holes providing a coolant film at blade surface, protecting the blade metal. In this way, the present work refers to the numerical investigation of internal and external cooling strategies applied in gas turbines. The methodology developed to analyze such strategies is based on the flat-plate approach with laboratory length scales and Computational Fluid Dynamics (CFD) techniques, being the flow, in the study domain, considered viscous, turbulent and compressible. A commercial CFD program is used to solve the general equations of fluid mechanics with Reynolds Average Navier-Stokes (RANS) technique for steady state regime and Shear Stress Transport (SST) turbulence model to determine the flow eddy viscosity. The combined effects of internal and external cooling is studied through a highly sophisticated scheme, called louver, which combines the effects of impingement and film cooling. Pin fins and ribs turbulator geometries applied in the channel between the impingement and the film cooling have the purpose of evaluating the impact of these geometries on the film cooling effectiveness over the flat surface in comparison to the louver scheme without turbulator. This study concluded that, pin fins proved to be the most promise solution because they increased in 7% the film cooling effectiveness. Ribs also have a good potential to increase the effectiveness, because an increase of 4% in film cooling effectiveness was observed. In addition, the effects of the turbulator are dependent on their location, since the turbulator positioned near the film cooling hole exit showed improvements in the film cooling effectiveness in relation to the turbulator near of the impingement cooling jet.

© 2018, Journal of Aerospace Technology and Management. All rights reserved.This article describes the run time characteristics of a gas turbine performance simulation using different solvers and components off-design performance database formats. Two different nonlinear systems of equation solvers, Newton-Raphson’s and Broyden’s, and two different formats of compressor and turbine off-design performance database (maps), tabulated values and fitted surface equations, were compared. Based on the results it is then possible to trade off and select the most appropriate combination of solver and component map type for the gas turbine performance simulation for real-time application.

© 2017 Elsevier Masson SASLarge launch vehicles use their propulsion systems based on Liquid Rocket Engines (LRE) equipped with turbopumps. Turbopumps are complex rotary machines that supply high power, mass flow, and pressures in the engine system to reach the thrust requirements as determined in the rocket engine thermodynamic cycle. Strong engines need a secondary turbopump system called a booster. These boosters have pumps and turbines smaller than those of the main engine turbopumps, and their important function is to increase the fluid pressure at the inlet of the main turbopumps, mainly to avoid cavitation. In the present work, the influence of the tip clearance issues in an axial turbine installed to operate as oxidizer booster in the Space Shuttle Main Engine (SSME) were evaluated numerically. The results are compared with experimental data from National Aeronautics and Space Administration (NASA). The flow characteristics and the variation in the turbine efficiency for different jet velocities were determined for three different tip clearance values associated with the percentage of turbine blade height: 3.0%, 5.5%, and 8.0%. The turbine design, numerical issues, mesh generation and results are described and discussed. The methodology and numerical simulations used in the present work was consistent with the experimental data and can be extended for other correlated numerical simulations related to axial hydraulic turbines.

© 2016 Elsevier Masson SASThe choice of the most appropriate rotor tip configuration is important, because it helps to avoid high blade tip losses due to the leakage flow that are responsible for efficiency and pressure ratio drops, mainly in High Pressure Turbine (HPT). This subject has been investigated to improve the axial turbines performance. The HPT used in this work is the turbine designed during the Energy Efficient Engine Program (E3 Program). This HPT was evaluated with different rotor tip geometry configurations: without tip clearance (hypothetical condition), with standard tip clearance geometry (flat-tip), with squealer, with winglet and squealer with winglet. Results were obtained based on the three-dimensional turbulent flow calculations making the use of a commercial CFD RANS equation-based solver with the addition of a two-equation turbulence model, in which the numerical solutions were compared with data available in the open literature for a HPT design-point operation. It was determined that for the HPT studied in this work, the machine efficiency can be improved using the rotor tip geometry equipped with winglet tip configuration. However, the rotor tip geometry equipped with squealer–winglet tip configuration presented a better pressure ratio compromise.

Copyright © 2017 ASME.Due to the critical importance of the turbopump applied in Liquid-Propellant Rocket Engines (LPRE) and the importance in the use of specific engineering software to design and analyze turbomachines, a Project-Based Learning (PBL) methodology was implemented in the undergraduate Turbopumps (TP) discipline at the Aeronautics Institute of Technology (ITA), taught for aerospace engineering students. This methodology was applied, using as a class example, the Liquid Oxygen (LOX) booster turbine of the Space Shuttle Main Engine (SSME), aiming at an enhancement in the discipline's syllabus, to become the theory and practice closer to the real engineering, and to increase the discipline's attractiveness. The results obtained with this methodology showed that the students have more interest and attention in the classes in which an engineering problem is evaluated and discussed with details using appropriate examples and engineering software that are used by the academia and industry. Several turbomachines issues as velocity triangles, power, blade geometrical aspects, flow quality, losses and in this case, the importance of tip clearance, could be better understood by the students. About the numerical results, the aim is that the students, after the preliminary project ends, evaluate the results and compare them with experimental data from National Aeronautics and Space Administration (NASA). One of the most important experience in this project is the results evaluation by the students and the discussion around it, as lessons learned, given suggestions to improve the project, if the results are not in the right way what can be done to correct them and understanding all physical phenomena involved. The learning experience was fascinating and effective, as noticed by students and noted by Professors.

Copyright © 2017 ASME.The increasing fuel prices and stringent environmental legislation compel industries worldwide to pursue means to increase their processes efficiency. A higher efficiency relates to a reduction in fuel consumption, which results in a lower operational cost and emissions. When considering a steel mill, processes encountered in the blast furnace and in the coke oven, for example, generate gases that can be availed as low-grade fuels to return some sort of energy back to the process. This practice reduces the amount of high-grade fuel required and increases the global efficiency of the industrial site; however, demands higher investments and increase the management complexity. A thorough evaluation of such power cycles is important to assess their application. This paper is based on a currently operational combinedcycle power plant composed by two gas turbines that are adapted to use blast furnace gas as main fuel and one steam turbine with a total power rating of 490 MWe. This power plant configuration is compared to another one in which the topping cycle - composed by two gas turbines - is eliminated, and the same amount of blast furnace gas is burnt in a conventional steam generator, operating as a Rankine-cycle. The software Gate Cycle™ was used to model and simulate both cycles and provide the main parameters to analyze their performance. Parameters such as power rating, efficiency, emissions, and expected capital expenditure provided means to assess both options and evaluate their application. The combined-cycle provided higher efficiency and power rating when compared with the Rankine-cycle. However, the expected values for capital expenditure showed to be also higher. A major difference between both cycles is the higher flexibility of the combined-cycle power plant, which is essential to guarantee an electric energy source within the industrial site. As a counterpart, the operational complexity is significantly higher when compared with the Rankine-cycle. Overall, the present work provides valuable information to assess both solutions.

Copyright © 2017 ASME.Turbodrill is a type of hydraulic axial turbomachine that rotates a bit by the action of the drilling fluid on turbine blades, which converts the hydraulic power provided by the high pressure from drilling fluid into mechanical power through turbine stages. The evaluation of hydraulic turbine performance characteristics are important to define feasible rotational speed and mass flow to attend the bit torque requirements during drilling through the post-salt and salt layers. As a result, optimum operational parameters are proposed for gaining the required rotational speed and torque for post-salt environments. The turbine motor presented in this study was established by design methods based on classical aeronautical turbomachinery blade profile to supply 30k Newton-meters (Nm) of torque requested by a polycrystalline diamond compact (PDC) bit to power the complex heterogeneous layer of rock. The performance evaluation of this innovative hydraulic turbine with 200 stages was carried out using computational fluid dynamics (CFD). The simulation considers two different drilling fluid types, sea water and brine. Besides, different flow rates were considered to investigate how velocity vectors, pressure profile, output power and other performance parameters are affected. Due the large amount of data, the first and second stages of the turbine have been used to predict the performance characteristics. This assumption gives interesting results and avoids too heavy computational costs. A commercial CFD solver (ANSYS CFX 15.0®) was used to calculate the governing equations based on Reynolds-Averaged Navier-Stokes (RANS equations) with the addition of turbulence model. The two-equation Shear-Stress Transport (SST) turbulence model was used to account the effects of flow eddy viscosity.

© 2017, Journal of Aerospace Technology and Management. All rights reserved.Industry and universities around the world invest time and money to develop digital computer programs to predict gas turbine performance. This study aims to demonstrate a brand new digital model developed with the ability to simulate gas turbine real time high fidelity performance. The model herein described run faster than 30ms per point, which is compatible with a high-definition video refresh rate: 30 frames per second. This user-friendly model, built in Visual Basic in modular structure, can be easily configured to simulate almost all the existing gas turbine architectures (single, 2 or 3 shaft engines mixed or unmixed flows). In addition, its real time capability enables simulations with the pilot in the loop at earlier design phases when their feedback may lead to design changes for improvements or corrections. In this paper, besides the model description, it is presented the model run time capability as well as a comparison of the simulated performance with a commercial gas turbine tool for single, 2 and 3 shaft engine architecture.

© 2016, Journal of Aerospace Technology and Management. All rights reserved.Power plants operating in combined cycle present higher thermal efficiency (over 60%) and increased power generation when compared to traditional simple cycles, such as gas or steam turbines operating alone. Considering that the power plant evaluated in this paper is already operational, a further development concerning to the power plant control system is required in order to evaluate disturbances and frequency variations, generated by the electrical grid during normal operation, as the loads applied to the turbines are intrinsically associated to the grid frequency. A computer program able to simulate the control system was developed to cope with these instabilities and to guarantee the necessary protection to the power plant operation. The develop program was made using MATLAB Simulink®. The main components of the power plant consists of 2 gas turbines of 90 MW each and a steam turbine of 320 MW, totalizing 500 MW. Firstly, the power plant main components were constructed separately. Once obtained stable models, the exhaust from the gas turbine was connected to the water-steam cycle through the heat recovery steam generator. The main parameters necessary to adjust the model such as gains, limits and constants were obtained from the power plant operational data. The simulation results allowed the evaluation of some key parameters; others are possible but not shown, such as power, exhaust gas temperature, fuel flow and variable stator angles during grid instabilities. The studies were conducted by testing the robustness, response time, transient analysis, steady state analysis and reliability of the proposed model.

© 2016 Elsevier Ltd. All rights reserved.The increase in electricity demand in Brazil and the frequent interruptions in its supply forced the industry and commerce use stationary generator sets, mainly in large urban centers like São Paulo. This city established a decree in order that these devices use cleaner fuels than diesel oil, or adopt post-gas treatment systems, since there are no national regulations for generator sets emissions. To meet this decree, ethanol appears as a good option because it is an environmentally friendly fuel, does not affect the ozone layer, since it is obtained from sugarcane, which helps to reduce the carbon dioxide emissions to the atmosphere through photosynthesis in sugarcane field. Within recommended specifications ethanol can be blended with diesel and gasoline, but can also be used without additives, without damaging the engine. The ethanol availability in Brazil and its consolidation in the automotive market make it an alternative for use in generator sets. The aim of this work was to characterize the performance of a heavy duty Otto cycle engine, developed based on 12 l diesel longblock, prepared to run on ethanol without additives, as a prototype generator set. The performance tests of the generator set were conducted at three different altitudes. A load bank was used to simulate the real electrical load in five different power settings. The performance characteristics obtained experimentally were compared with results obtained with a one-dimensional model using the commercial software GT-Power®. The results obtained during the development phase showed that the engine achieve up to 39.6% of the brake efficiency and a peak power of the 326 kW. The maximum electrical power achieved by the generator set was 302 kW at sea level and 278 kW at 1640 m, according to the employed methodology. The results showed that is feasible to use ethanol without additives for the energy generation, replacing diesel oil in heavy-duty engines, operating in the evaluated steady state condition.

© 2015, Journal of Aerospace Technology and Management. All rights reserved.Design and development of gas turbine components are a complex multidisciplinary process. At the beginning of the power class definition and engine configuration it is necessary to conduct a market study. The results obtained are used in gas turbine thermodynamic cycle calculations and analysis in order to define the gas turbine design point. Several possible design points are evaluated during this procedure. After this step, the gas turbine components are designed, including: compressor, combustion chamber and turbine. For industrial gas turbine purposes, it is common to use a free turbine after the gas generator, also commonly named power turbine. In this work, a power turbine was initially designed by meanline techniques, considering internal loss mechanisms, to obtain the main dimensions. The geometries of the components were generated in a 3-D environment to make possible the mesh generation, process to discretize the physical domain into a computational domain and use a 3-D Computational Fluid Dynamics tool. The results from the meanline approach and from the 3-D turbulent flow numerical simulations were compared to verify the turbine operational conditions and its predictions at design and off-design conditions. The gas turbine under study is a project, derived from a low thrust turbojet previously developed by Instituto de Aeronáutica e Espaço. The power turbine project uses the same turbojet gas generator, already designed and currently under tests.

Copyright © 2015 by ASME.This work describes the continuous study that is being done in a small gas turbine that can be used for power generation purposes. Previous studies were conducted aiming to develop a gas generator able to be used in both applications, as a turbojet or as a turboshaft. The gas generator was designed, manufactured and is still under test. The thermodynamic cycle calculation was evaluated as a project-based class, hence, a power turbine was specified and its requirements were determined. The outlet conditions from the gas generator were used to perform the preliminary size of the power turbine. At this phase, the students must use 1D design models considering loss modeling to improve the machine design prediction. The meanline technique was used and the calculations at leading and trailing edges were extrapolated from hub-to-tip, using vortex design methods. With the airfoil stacking for each blade row was possible to determine the 3D geometry of the single stage axial flow turbine. This geometry was assembled in a CAD software to start the mesh generation procedure. After this step, a commercial CFD software was used to calculate the continuity, momentum and energy equations from fluid mechanics. The flow was considered fully turbulent and the two-equation SST turbulence model was set to determine the flow eddy viscosity. The results from preliminary design and 3D techniques were compared and evaluated to complete the first round of the design phase. In this work, experiences from the project-based class on turbomachinery design are described together with the challenges and difficulties that appeared during the project.

Copyright © 2015 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.Nowadays the environmental and economic aspects emerge as essential alternatives in the design phase of microturbines. In design point definition for micro gas turbine cycles, not only engine performance requirements are necessary to have competitive microturbines but also external requirements, as cost and environmental issues must be in agreement simultaneously. To support engineers in defining the engine design point considering thermodynamic, economic, and environmental aspects, a gas turbine code was developed. The code uses a methodology that includes the sum of microturbine costs as power plant, fuel, and environmental emissions. The developed computer program was written in MATLAB® and is able to simulate the economic and thermodynamic performance of a given micro gas turbine cycle through an optimization process using genetic algorithm. The code is capable of calculating the suitable design point for a specific application. In this work, a 200 kW micro gas turbine recuperated cycle was chosen to study. As initial analysis, a parametric study was made to investigate the behavior of the main decision variables, considering costs and emissions. Afterward, single-objective and multiobjective optimizations were carried out using the objective function according to the proposed methodology. In sequence, a comparison was presented between the design point of a reference available microturbine and the same optimized by the code. The results reveal the importance of the cost optimization, showing how much savings can be achieved in choosing an appropriate design point for microturbines using the methodology implemented in the present work.

© 2015, Journal of Aerospace Technology and Management. All rights reserved.The Instituto Tecnológico de Aeronáutica (ITA) is an Engineering school maintained by the Air Force Command, Ministry of Defense. The aim of the Turbomachines Department at ITA is the human resources training for design and development of aeronautical and industrial gas turbines, necessary for the Gas Turbine Program of the Departamento de Ciência e Tecnologia Aeroespacial (DCTA). The human resources training is carried out in undergraduate and graduate courses at ITA, where topics in gas turbine and turbomachinery are taught. The gas turbine topic is taught in the undergraduate degree, in Mechanical-Aeronautical Engineering course, and focuses on gas turbines’ performance for different configurations (turboshaft, turbojet and turbofan). Lecture notes containing the essential elements of the course are made available for the students, addressing the basic theory of the gas turbines required for the performance study at the design and off-design point. The technological aspects are presented and discussed during detailed studies of the actual cycle. Simple gas turbines and more sophisticated ones are studied, for both aeronautical and industrial application. Performance calculations at design and off-design point of the main engine’s components and the cycle are done manually, encouraging students to develop spreadsheets. The theory is complemented with laboratory classes and technical visits, when the practicalities involving gas turbines operation and tests are presented to the students. As an activity laboratory class, the students perform disassembly-assembly of a small industrial gas turbine.

© 2015 Taylor & Francis Group, London.The application of optimization techniques in engineering designs is a fundamental parameter for a definition of the best design schedule. High performance compressors operate close to stall line for higher pressure ratio. When the compressor operates at off-design condition, it causes an efficiency degradation and pressure ratio drop due to detachment of the fluid on vanes. A technique to realign the flow is the use of variable geometry stator, but it is necessary a trade-off to define the amount of VSVs and angle of each grid. Therefore this will perform a match between all compressor stages in order to keep the efficiency optimal. In this work a hybridization technique is implemented by using an evolutionary algorithm together with a deterministic algorithm. The commercial software ModeFrontier® was coupled with the axial compressor computational program. The results show high efficiency using this technique in runs with multi-objective problems with low computational cost, and also it shows a comparison between the original compressor performance map and the map obtained after the optimization process.

Using "Non-Dimensional" parameters for measuring and presenting the level-flight performance of Turbine-Engined Helicopters is very efficient when conducting a flight test performance program. In this study, the level-flight performance of a helicopter-example (AS 550 Fennec) is determined using the collective pitch position as main parameter and the methodology described by the Engineering Sciences Data Unit ESDU No. 74042. The purpose of this study is threefold. The first is to present graphics of level-flight performance for a helicopter-example (AS 550 Fennec) using collective pitch position. The second is to present the relationship between the torque curve and collective pitch. The last is to exemplify some cases where this methodology can be applied to produce level flight performance graphics for operating data manual of a helicopter.

Copyright © 2014 by ASME.At Technological Institute of Aeronautics-ITA, the Flow Machines for Aerospace Applications course deals with turbopumps. It is offered to students of the second professional year in the undergraduate program. The objective of the course is to present enough information for the students to learn about machine preliminary design. The theory involved in flow machines is multidisciplinary, so that the students must fully understand the principles of fluid mechanics, heat transfer, gas dynamics and thermodynamics, whose complete understanding is vital to start the design of such machine. Lack of understanding the basic theory due to the problems like difficulty of associating the velocity triangles to blade angles, blade camber, incidence and deviation angles, makes the subject more complex than it really is. This problem is mitigated with hands-on activities during practical classes. In this work, the syllabus of flow machines design course at ITA is presented, details of how the subjects are taught and the procedures used during the classes are reported, based on real life engineering project. After a period of theory and lab classes the students are requested to do a complete preliminary design of a turbomachine, using adequate numerical tools based on specs of a turbomachine. The instructor follows the class with discussions towards adjustments that could result in operational characteristics improvement. The project starts with meanline calculations to produce the base geometry and proceeds up to 3D CFD calculations to check the performance of the designed turbomachine and to subsidize design modifications. Experience exchanged with students are reported and commented aiming at course improvement.

Copyright © 2014 by ASME.In high performance turbomachines the tip region is a key point to improve aiming at high pressure ratios without high penalties. In the case of HPT, several techniques are still in development by academic research laboratories and industry. Some geometrical configurations were created at the rotor tip region, as winglets and squealers geometries. In the case of squealers, the depth of their cavity is an important parameter to evaluate, because its values can cause different flow behavior on this region. Changing the heat transfer. In this work, the rotor blade of a HPT developed in the E3 program was changed, the aim is to study the influence of the squealer cavity depth variation on its performance. The flow within the turbine was calculated using a commercial CFD package. The details of the rotor geometrical changes, the differences between a simple flat rotor tip surface and squealer configurations are discussed and presented.

Copyright © 2014 by ASME.The paper aim is to study the transient performance using fuel flow schedule, variable geometry compressor control and combustion emissions for a simple turbojet engine in a thrust class of 5kN. This engine is under development, it was designed, manufactured and are being tested in test bench, it is composed by a 5-stage axial flow compressor, an annular combustor, an uncooled turbine and a convergent nozzle. The engine was originally designed to run on kerosene but other types of fuels, as biofuels, are intended to be used, having in mind a turboshaft in a class of 1.2 MW for power generation purpose. PID control is being studied to determine the appropriate setting of the VIGV in conjunction with a prescribed fuel flow injection necessary to accelerate, or decelerate, the engine from 80% to full thrust in a prescribed time interval. The engine control system is being studied during this engine design phase, so that all the components characteristics needed are being synthesized using in-house developed computer codes: compressor design and performance; combustion chamber design and performance; turbine design and performance; whole engine performance. The engine is required to accelerate from 80% to full thrust in a short time interval, which is also a limitation imposed to the control system. Compressor surge margin is controlled during accelerations using controlled positioning of the VIGV at each engine speed. The engine running lines, for accelerations and decelerations are shown and commented. They served as basis for the design of the engine control logic and hardware. The combustor was designed for kerosene, but other types of fuels can be burned, with the lower heating value and all the necessary parameters recalculated using reaction mechanisms, reactor network and stability loops approach.

© 2014 Acta Scientiarum. Technology. All rights reserved.An efficient combustion depends on many factors, such as injection, turbulence and ignition characteristics. With the improvement of internal combustion engines the turbulence intensity and internal pressure have risen, demanding more efficient and powerful ignition systems. In direct injection engines, the stratified charge resultant from the wall/air-guided or spray-guided system requires even more energy. The Paschen’s law shows that spark plug gap and mixture density are proportional to the dielectric rupture voltage. It is known that larger spark gaps promote higher efficiency in the internal combustion engines, since the mixture reaction rate rises proportionally. However, the ignition system must be adequate to the imposed gap, not only on energy, but also on voltage and spark duration. For the reported study in this work two test benches were built: a standard inductive ignition system and a capacitive discharge high energy ignition system, with variable voltage and capacitance. The influence of the important parameters energy and ignition voltage on the spark duration, as well as the electrode gap and shape were analyzed. It was also investigated the utilization of a coil with lower resistance and inductance values, as well as spark plugs with and without internal resistances.

Undergraduate courses at Technological Institute of Aeronautics (ITA) are 5-years course, divided into Fundamental (2 years) and Professional (3 years). The Flow Machines, in the Mechanical-Aeronautical Engineering Course, is offered by the Turbomachines Department and is taught in the first semester of the fourth year (2nd professional year). In the course, the basic theory, unified for all machines, is presented in details for the students, emphasizing the physics of all processes involved in the fluid-machine energy transfer. Incompressible and compressible fluids are treated accordingly. The flow machines types are individually studied, focusing attention to their performance characteristics and range of applications. The preliminary design and off-design operation issues are discussed in details with the students, with emphasis on relevant aspects of each machine, like cavitation, stall and surge. The students are taught on how to choose the flow properties at the blade edges for the sake of preliminary design and off-design performance estimations. Loss models are introduced during the theory classes and popular models are presented. At this point, in-house computer codes and commercial software are presented to the students, who are asked to solve simple problems. The installation, operation and basic performance calculations are also presented for the students during the lab classes for several hydraulic machines installed at ITA laboratories. All course material is transferred for the students in pdf format before classes. In this work, the experience with the teaching process in flow machines at ITA, theory and laboratory, is described. Copyright © 2013 by ASME.

During a gas turbine development phase an important engineer task is to find the appropriate engine design point that meet the required specifications. This task can be very arduous because all possible operating points in the gas turbine operational envelope need to be analyzed, for the sake of verification of whether or not the established performance might be achieved. In order to support engineers to best define the engine design point that meet required performance a methodology was developed in this work. To accomplish that a computer program was written in MatlabR. In this program was incorporated the thermoeconomic and thermodynamic optimization. The thermodynamic calculation process was done based in enthalpy and entropy function and then validated using a commercial program. The methodology uses genetic algorithm with single and multi-objective optimization. The micro gas turbine cycle chosen to study was the recuperated. The cycle efficiency, total cost and specific work were chosen as objective functions, while the pressure ratio, compressor and turbine polytropic efficiencies, turbine inlet temperature and heat exchange effectiveness were chosen as decision variables. For total cost were considered the fixed costs (equipment, installation, taxes, etc.) and variable costs (fuel, environmental and O&M). For emissions were taken into account the NOx, CO and UHC. An economic analysis was done for a recuperated cycle showing the costs behavior for different optimized design points. The optimization process was made for: single-objective, where each objective was optimized separately; two-objectives, where they were optimized in pairs; three-objectives, where it was optimized in trio. After, the results were compared each other showing the possible design points. Copyright © 2012 by ASME.

A small 5-kN thrust gas turbine, designed and manufactured having in mind a thorough source of validation data, serves as basis for the study. The engine is an uncooled turbine, 5:1 pressure ratio axial flow compressor, delivering 8.1 kg/s air mass flow, whose control is made by a FADEC. Cold runs of the jet engine version have already been completed. The engine characteristics are being developed using the technology indicated in the paper. Accelerations and decelerations from idle to full power in a prescribed time interval and positive surge margin are the limitations imposed to the control system. In order to accomplish such requirements, a proportional, integral and derivative (PID) has been implemented to control the variable geometry transients, which proved to drive the engine to the required operating points. Compressor surge is avoided during accelerations or decelerations, imposing operation limits to the surge margin. In order to simulate a jet engine under transient operation, use was made of high-fidelity in-house developed software. The results presented in the paper are related to the compressor inlet guide vane (VIGV) transients. The engine transient calculations were predicted with the IGV settings varying with time, and the results are being used for the initial calibration of the transfer functions for the real time control. Copyright © 2012 by ASME.

The design of a small gas turbine in the range of 5 kN thrust / 1.2 MW shaft power is being made in association with industry, aiming at distributed power generation and cogeneration. The gas turbine was constructed and its gas generator is being prepared for development tests. The results will be used for the final specification of the power section. The gas turbine design has been carried out using indigenous software, developed specially to fulfill the requirements of the engines design, as well as the support for validation of research ork. The work reported in this paper deals with the design methodology of a 5:1 pressure ratio, 5-stage axial flow compressor with VIGV and a single stage axial flow turbine. These components were designed and their maps synthesized and fed to the gas turbine performance simulation program. The engine performance results were analyzed and verified. The calculated behavior compares with similar engines', indicating they are qualitatively correct. Copyright © 2012 by ASME.

During the design of a gas turbine it is required the analysis of all possible operating points in the gas turbine operational envelope, for the sake of verification of whether or not the established performance might be achieved. In order to achieve the design requirements and to improve the engine off-design operation, a number of specific analyses must be carried out. This paper deals with the characterization of a small gas turbine under development with assistance from ITA (Technological Institute of Aeronautics), concerning the compressor variable geometry and its transient operation during accelerations and decelerations. The gas turbine is being prepared for the transient tests with the gas generator, whose results will be used for the final specification of the turboshaft power section. The gas turbine design has been carried out using indigenous software, developed specially to fulfill the requirements of the design of engines, as well as the support for validation of research work. The engine under construction is a small gas turbine in the range of 5 kN thrust / 1:2 MW shaft power, aiming at distributed power generation using combined cycle. The work reported in this paper deals with the variable inlet guide vane (VIGV) transients and the engine transients. A five stage 5:1 pressure ratio axial-flow compressor, delivering 8:1 kg=s air mass flow at design-point, is the basis for the study. The compressor was designed using computer programs developed at ITA for the preliminary design (meanline), for the axisymmetric analysis to calculate the full blade geometry (streamline curvature) and for the final compressor geometry definition (3-D RANS and turbulence models). The programs have been used interatively. After the final channel and blade geometry definition, the compressor map was generated and fed to the gas turbine performance simulation program. The transient study was carried out for a number of blade settings, using different VIGV geometry scheduling, giving indication that simulations needed to study the control strategy can be easily achieved. The results could not be validated yet, but are in agreement with the expected engine response when such configuration is used. Copyright © 2011 by ASME.

This work presents the performance study of a 1 MW gas turbine including the effects of blade cooling and compressor variable geometry. The axial flow compressor, with Variable Inlet Guide Vane (VIGV), was designed for this application and its performance maps synthesized using own high technological contents computer programs. The performance study was performed using a specially developed computer program, which is able to numerically simulate gas turbine engines performance with high confidence, in all possible operating conditions. The effects of turbine blades cooling were calculated for different turbine inlet temperatures (TIT) and the influence of the amount of compressor-bled cooling air was studied, aiming at efficiency maximization, for a specified blade life and cooling technology. Details of compressor maps generation, cycle analysis and blade cooling are discussed. Copyright © 2010 by ASME.

New health monitoring strategies were developed in the last decade aiming at improvement of gas turbines safety and reliability. Real time methodologies have been considered of major concern for safe operation at least cost. This paper describes a hybrid system approach for turboshaft faults diagnosis, using data obtained from a tuned high fidelity gas turbine simulator program, including those for multiple faults deteriorated performance. Kohonen neural network was used to analyze similarity together with an optimization strategy to reduce the volume of data used in the diagnostics phase. A Multi-Layer Perceptron (MLP) was used for training and validation. The MLP and Kohonen networks were tested for several configurations, in order to improve diagnosis. The hybrid system was also tested with noise-contaminated data and it was verified the capability of the neural approach to detect and isolate multiple faults better than the MLP alone. The results showed that the optimization strategy reduced significantly the database patterns and improved the learning process, demonstrating high precision to diagnose gas turbine operation problems. The reliability of the proposed system is explained both qualitatively and quantitatively. Copyright © 2010 by ASME.

This paper describes a procedure to measure the performance of detection and isolation of multiple faults in gas turbines using artificial neural network and optimization techniques. It is on a particular form of artificial neural networks, the traditional multi-layer perceptron (MLP). Error back-propagation and different activation functions are used. The main goal is to recognize single, double and triple faults in a turboshaft engine, whose performance data were output from a gas turbine simulator program, tuned to represent the engine running at an existing power station. MLP network is a nonlinear interpolation function usually made of input layer, hidden-layer and output-layer, with different neuronal units, but in this work, only one hidden-layer was used. Weights were altered by error back-propagation from the initial values established from a seed fixed between 0 and 1. The activation function in the MLP algorithm is the sigmoid function. The best moment to stop the training process and avoid the over fitting problem was chosen by cross-validation. Optimization of convergence error was achieved using the momentum criteria and reducing me oscillation problem in all nets trained. Several configurations of the neural network have been compared and evaluated, using several noise graduations incorporated to the data, aiming at finding the network most suitable to detect and isolate multiple faults in gas turbines. Based on the results obtained it is inferred that me procedure reported herein may be applied to actual systems in order to assist in maintenance programs, at least. Copyright © 2009 by ASME.

There are many different sources of loss in gas turbines. The turbine tip clearance loss is the focus of this work. In gas turbine components such as compressor and turbine the presence of rotating blades necessitates a small annular tip clearance between the rotor blade tip and the outer casing. This clearance, although mechanically necessary, may represent a source of large loss in a turbine. The gap height can be a fraction of a millimeter but can have a disproportionately high influence on the stage efficiency. A large space between the blades and the outer casing results in detrimental leakages, while contact between them can damage the blades. Therefore, the evaluation of the sources of the performance degradation independently presents useful information that can aid in the maintenance action. As part of the overall blade loss the turbine tip clearance loss arises because at the blade tip the gas does not follow the intended path and therefore does not contribute to the turbine power output and interacts with the outer wall boundary layer. Increasing turbine tip clearance causes performance deterioration of the gas turbine and therefore increases fuel consumption. The increase in turbine tip clearance may as a result of rubs during engine transients and the interaction between the blades and the outer casing. This work deals with the study of the influence of the turbine tip clearance on a gas turbine engine, using a turbine tip clearance model incorporated to an engine deck. Actual data of an existing engine were used to check the validity of the procedure. This paper refers to a single shaft turbojet engine under development, operating under steady state condition. Different compressor maps were used to study the influence of the curve shapes on the engine performance. Two cases were considered for the performance simulation: constant corrected speed and constant maximum cycle temperature. Copyright © 2008 by ASME.

This work deals with a nonlinear model, based on a particular form of artificial neural networks, ANN, for application to gas turbines fault diagnosis. The traditional multi-layer perceptron (MLP) is used, with error backpropagation and different activation functions. The application of the model is illustrated using test data from a gas turbine simulation computer program. A specially developed computer program is used to simulate the engine in operation, generating all needed engine data for both baseline and deteriorated engine. A test case using a turboshaft engine is used to demonstrate the capacity of this ANN to identify faults that may occur during engine operation. Copyright © 2007 by ASME.

Gas turbines need to operate efficiently due to the high specific fuel consumption. In order to reach the best possible efficiency the main gas turbine components, such as compressor and turbine, need to be optimized. This work reports the use of two specially developed computer programs: AFCC [1, 2] and GTAnalysis [3, 4] for such purpose. An axial flow compressor has been designed, using the AFCC computer program based on the stage-stacking technique. Major compressor design parameters are optimized at design point, searching for best efficiency and surge margin. Operation points are calculated and its characteristics maps are generated. The calculated compressor maps are incorporated to the GTAnalysis computer program for the engine performance calculation. Restrictions, like engine complexity, manufacture difficulties and control problems, are not taken into account. Copyright © 2006 by ASME.

Nacelles are responsible for good engine performance and considerable percentage of total aircraft drag, thus fuel consumption. Energy conservation and cost of fuel, among others, require good nacelle design. CFD calculations of the flow around it are a major design tool to predict shock waves, internal boundary layer in the nacelle forebody, high velocity zones and wake. Commercially available software may be used to calculate and visualize the flow at the most critical parts of the nacelle, allowing design modifications aiming at optimizations. This paper overviews the literature on nacelles, the methodologies involved in the design. A case study is presented for a long duct nacelle design, using an axissymmetric model. Performance characteristics at important operating conditions are also presented. Copyright © 2006 by ASME.

Poor part-load performance is a well-known undesirable characteristic of gas turbines. Running off-design, both compressor and turbine lose performance. Flow misalignment at the various rows causes losses to increase sharply, thereby decreasing net output faster than decreasing fuel consumption. To bring the flow to alignment with the blade passages, it is required to restagger the blades both at the compressor and at the turbine. To avoid mechanical complexities, it is generally accepted to restagger only the stators. This work deals with a numerical approach to the simulation of a gas turbine equipped with variable stators at the compressor and at the turbine, enabling the search for better-performance operation. A computer program has been developed to simulate virtually any gas turbine having variable stators at the compressor stages and turbine nozzle guide vanes. Variable-inlet guide vanes (VIGVs), variable-stator vanes (VSVs), variable-nozzle guide vanes (VNGVs), variable-geometry compressors (VGCs) and variable-geometry turbines (VGTs) are the focus in this work, which analyses a one-shaft free power turbine for power generation in the search for performance improvement at part load. © IMechE 2004.

For distributed power generation, sometimes the available gas turbines cannot match the power demands. It has been usual to uprate an existing gas turbine in the lower power range by increasing the firing temperature and speeding it up. The development costs are high and the time to make it operational is large. In the other hand, de-rating an existing gas turbine in the upper power range may be more convenient since it is expected to cut significantly the time for development and costs. In addition, the experience achieved with this engine may be easily extrapolated to the new engine. This paper deals with the performance analysis of an existing gas turbine, in the range of 25 MW, de-rated to the range of 18 MW, concerning the compressor modifications that could be more easily implemented. Analysis is performed for the base engine, running at part-load of 18 MW. A variable geometry compressor is derived from the existing one. Search for optimized performance is carried out for new firing temperatures. A variable geometry turbine analysis is performed for new NGV settings, aiming at better cycle performance.