Within an industrial setting, mesh adaptation has so far found very limited use. This is, in part, due to the complexity of the geometries and flow features that are to be dealt with. However, the successful utilisation of grid modification techniques, could help engineers achieve more accurate estimates of quantities of interest quickly and efficiently. For this reason, in this paper, adjoint error mesh adaptation technology is developed and applied to steady-3D turbo-machinery solutions. The grid modification strategy proposed comprises of a combined mesh movement and mesh refinement procedure, entirely based on errors related to the functional of interest. The node addition scheme makes use of the output to the flow adjoint solver and an interpolation to an embedded grid. The determined error is used in an edge-refinement approach developed in the in-house MeshPost software. The mesh relocation technique, instead, employs the sensitivity of the functional of interest with respect to the nodes’ coordinates to compute a Riemmannian metric. This parameter is then equi-distributed over the mesh by applying a spring-stiffness approach.

As one of the most important components in a rotor-bearing system, journal bearings provide proper support and damping to the rotor so that it can run both smoothly and efficiently and keep stable under different working conditions. As the rotating speed of the rotor growing faster and load getting heavier, the traditional cylinder journal bearing can no longer meet the demand of stabilizing the rotor, so different kinds of non-circular journal bearings were invented, such as elliptical bearing, multi-lobe bearing, wave bearing and etc., to provide better stability and greater load capacity. However, these kinds of non-circular bearings were mostly designed by experience of the engineers, and also the current hydrodynamic bearing design methodology still depends on empirical design. There lacks of corresponding theoretical foundation. In order to develop a theoretical method for bearing designing, an innovative analyzing approach needs to be carried out to explore the mechanism of the bearings and its performance. In this paper, a new approach is presented focusing on the profile of each bearing and their film thickness. A universal mathematical expression for different types of non-circular bearings has been put forward based on the Fourier series theory. The influence of periodic harmonics of film thickness on the static performance of non-circular bearings of finite length is studied for incompressible lubricant. The results show that the film thickness can always be expanded into a Fourier series, and the harmonic components of film pressure can be obtained by solving the Reynolds equation. Finally, the relation between the k -th order harmonic component H 0, k and the corresponding static pressure component P 0, k is established. This new investigation can be used to improve the non-circular bearing designing methodology with theoretical guidance.

Validation and tuning methods for analytical algorithms used in modeling heat transfer for engine parts, such as turbine blades, require the direct comparison of calculated and experimental metal temperature measurement results. It is important to ensure that the geometric characteristics as well as the boundary condition parameters of the gas flow are held in common for both the physical experiment and the calculation. Unfortunately, the task of obtaining knowledge of the gas parameters such as the radial temperature profile at the turbine inlet is far from trivial. The authors will briefly survey commonly used methods and the uncertainties related to their application based on publicly available data. Among these techniques, a relatively new method of using Uniform Crystal Temperature Sensors (UCTS) for gas temperature measurement, may offer some advantages and will be analyzed in greater detail. The authors use 3D Finite Element Analysis code to analyze sources of potential systematic error for gas temperature measurement using UCTS. A matrix of varying materials and configurations is evaluated. Recommendations drawn from this analysis will enable analytical designers and instrumentation engineers to optimize the application of this technique specific to their test requirements.

While reliability stays the most critical requirement for power generation in the Oil & Gas sector, energy effectiveness is becoming an increasingly important topic. Gas turbines are efficient and flexible tools both in electrical and mechanical drive applications as they cope with the multiple energy demand profiles of this sector. Aeroderivative gas turbines boast fast response and fast O&M procedures which fits with the exploitation of hydrocarbon fields. Heavy duty machines are robust and very popular in refinery utilities or as large mechanical drivers. However, it is sometimes desirable to boost GT performances: operators of the Oil & Gas sector may need to increase power output or wish to improve efficiency, especially during part load operation where the latter is degraded. While a number of commercial hardware, service and software products cover these needs, the operators may want to keep their own control on their production tools. To that end resorting to simple thermodynamic considerations enables rule-of-thumb evaluations of the impact of cycle changes on GT performances but this possibility is often unexploited. This joint paper is intended to provide the energy community, especially the engineers of the Oil & Gas sector, with some basic methods for the estimation/anticipation of gas turbine performance changes.

Gas turbine combustion engineers strive to obtain high flame stability at low power conditions (idle); whereas good emissions characteristics are desired at high power conditions. Although lean combustors provide good emissions characteristics, they suffer from the issue of poor flame stability. Often a pilot flame along with a lean main flame is used in order to improve flame stability issue. Positioning of an axial jet concentric to a swirl jet is one of commonly used configurations for a pilot flame. The interaction between both co-annular jets leads to distinct flow patterns in the combustor and subsequently the flame structure is impacted. If the axial jet is able to penetrate the inner recirculation zone (IRZ) generated by the swirled flow of the main jet, then a jet type of flame is obtained. The jet flame in vicinity of co-annular swirled flow is also denoted as type 1 flame, and is known to have good flame stability. On the other hand, if the pilot jet is not able to penetrate the IRZ, then the recirculating type of flame is obtained. A recirculating flame type has good emissions characteristics but suffers from poor flame stability. In this work the numerical predictions have been performed to gain more insight into occurrences of these two different flame structures which have been experimentally recorded. Using ANSYS Fluent CFD software, non-reactive steady state turbulent flow simulations have been performed to understand the flow and mixing field in a 3D combustor. Laser Doppler Anemometry (LDA) and hydroxyl radical (OH*) chemiluminescence have been used to measure the flow and characterize the flame structure, respectively.

In the near future, combustion engineers will shape the burner according to the flame, and not the opposite way anymore. In this contribution, this idea is explored with the help of additive manufacturing (AM). The focus is put on the design and the production of swirlers using advanced materials with the least possible efforts in terms of manufacturing. The material chosen for this study is Inconel 718. There are three motivations to this project. The first one is to design new shapes and assess these in comparison to conventional ones. The second motivation is to be able to manufacture them using additive manufacturing, and to gather know-how on selective laser melting. The third motivation is to elaborate a methodology involving engineering, research and education to promote — only if and when this is desirable — the production of series of premium parts such as high-end components of gas turbine combustor using AM. First-of-a-kind swirler shapes are explained and designed. These are unlikely to be produced using conventional manufacturing. They are then successfully produced in Inconel 718 using AM. The raw parts are directly submitted for testing with no surface post-processing. The paper states why at current state-of-the-art the raw surface quality still needs improvement, and highlights the benefits of the new swirler shape versus conventional.

Kinetic energy recovery is a key objective for low pressure exhaust hood design and optimization. Numerical simulation of the exhaust hood helps the engineers to explore and confirm the causes of the loss in the hood. Many studies have suggested that it is necessary for the simulation to include the last stage blade to get a realistic assessment. For the sole exhaust hood study, the inlet boundary condition is hard to set precisely like the downstream flow of the last stage blade. And the studies have also shown that the performances generated from the simulations may vary evidently between the sole exhaust hood and exhaust hood with last stage blade. It is obvious that the blade influences the exhaust hood, but the exact effect factors of the blade and the way they work are not thoroughly discussed. This paper has conducted many numerical tests to audit the influence of the common effect factors of the last stage blade. The internal flow field of the exhaust hood was numerically investigated using three-dimensional Reynolds-Averaged Navier-Stokes (RANS) solutions based on the ANSYS-CFX. In the first part of the paper, the tests are conducted by changing each effect factor of the inlet boundary condition for sole exhaust hood studies. These factors include the mass flow flux, the angle of the exit flow of the last stage, both the circumferential and the radial ones, and the speed and position of the jet-flow downstream of the seal over the shroud of the bucket. The tests show that each factor has its own distinctive style and extent for influence. Some of them may maximize the performance at some certain point, and some may deteriorate the performance rapidly beyond a threshold. And some factors may change the performance insignificantly within a wide range. However, these influences are not good enough to be consistent with the difference between the sole exhaust hood and the hood with blade simulations. In the second part of this paper, the focus locates on the direction of the jet-flow of the bucket seal. The tests prove that this direction is the prominent factor to influence the exhaust hood performance. Some extra tests for the seal have also been conducted to analyze this factor. The static pressure recovery for the simulation with labyrinth seal is about only half of the sole exhaust hood simulation. The discussion of these tests show that the seal jet is the main cause for this performance dive, and explain how the seal jet direction changes the flow field of the exhaust hood. It also suggests that the procedure to optimize the seal design is not mature yet, for some nature of the jet-flow remains unclear. It may need more detailed study in the future.

Effusion cooling is one of the most widespread system used to cool combustion chamber liners nowadays: it is efficient, cheap and light. Effusion cooling consists of drilling thousands of submillimetric holes into the combustor wall in order to cool it down from inside the holes and create a cooling film inside the combustor protecting it from the hot gases. Effusion cooling has long time been very challenging for combustor simulations because it involves lengthscales ranging from ½ millimeter (about the size of the effusion holes) to ½ meter (about the diameter of the combustor). That is one of the main reasons for which 3D simulations of effusion cooling has long been inaccessible, and has generated the studies presented in this paper. This study will focus on the effusion cooling holes discharge coefficient evaluation as a function of numerous aerothermal and design parameters. Many attempts to describe aerodynamically effusion have occurred, mainly based on experiments exploring few of the parameters mentioned above. But none of these studies tried to elaborate a model able to handle most of them. This is the purpose of this study which will set up a 3D model able to describe finely the physical phenomena involved in combustor effusion cooling holes and the influence of the design parameters available to combustor engineers on these phenomena. The strategy which prevails in the setup of the numerical 3D detailed model is to find a compromise between the reliability and the CPU cost of the simulation. Indeed the objective is to study the influence of a wide range of effusion cooling design parameters such as hole diameter, orientation, shape, etc. . . on the effective cross section. In addition, for a better understanding of the physical phenomena, all the simulations are performed at the same aerothermal conditions. These aerothermal conditions as blowing ratio, cooling temperature, pressure are not design parameters of effusion cooled walls. They are usually imposed by the gas turbine thermodynamic cycle very early in the development of a new engine. A preliminary study allowed to select the parameters which were both the least known and the most influential on the effusion hole mass flowrate according to literature, preexisting numerical and experimental databases. More than 60 new CFD simulations have been performed and show the influence of each effusion cooling design parameter taken separately: effusion holes inclination, orientation and taper angle. A mesh sensitivity study has been performed in order to validate the numerical approach. Then, the analysis of both the preexisting data and this new numerical database showed that some of the design parameters have strong effects and coupled influences on the mass flow rate through the holes. On the other hand, some other parameters could be easily described by simple models or even neglected. This study concludes by quantifying the improvement of a proprietary effective cross section correlation of effusion cooled walls, based on the analysis mentioned in this study.

This paper presents a description of Detached Eddy Simulations being carried out on a variable stator vane with a penny-cavity in order to determine the secondary flow phenomena in the main flowpath. Variable stator vanes are common in multi-stage compressors to prevent flow separations on rotor and stator blades at off-design operation points. The bearing of the stators at hub and tip generate unavoidable circular-shaped ring gaps, which are called penny-cavities. The aim of this paper is to determine secondary flow phenomena in variable stator vanes on an annular cascade testbed resulting from the throughflow of the penny-cavities. Reynolds-Averaged-Navier-Stokes simulations and scale resolving Detached-Eddy-Simulations of a variable stator vane with hub penny-cavity were therefore performed using Ansys CFX. The results of these simulations will be compared to corresponding simulations without penny-cavity. The study shows secondary flow phenomena, which are comparable to the interaction of a transverse jet in a free stream. Due to the low momentum ratio of R = 0.5 , the jet immediately veers in the direction of the main flow. The typical vortices which develop from a transverse jet in a free stream are identified. The steady RANS simulation shows an asymmetrical counter-rotating vortex pair. A lack of unsteady secondary flow interaction can be seen in the RANS simulations in contrast to the Detached-Eddy-Simulations, which resolve large turbulent scales. Hence an interaction between the counter-rotating vortex pair and the unsteady shear layer vortices in the stator is visible. In the Detached Eddy Simulations the counter-rotating vortex pair is superimposed by the unsteady shear-layer vortices. The vortices produce significant additional mixing losses, which will be shown in detail. By comparing simulations with and without penny-cavity, the penny-cavity losses are quantified. In conclusion, this paper will help design engineers become more aware of the significance of the penny-cavity with variable stator vanes.

There is evidence of a lack of knowledge in the design of the blade/disk attachment so that the strength of the materials is not fully exploited and the load capability of the attachment is underestimated. The aim of this work is to improve the engineers’ capability in designing the attachment so that higher loads can be carried with the same material. To this end, an optimization method has been applied to the attachment design. A dovetail blade root was chosen as case study and the objective function was the static equivalent stress in the blade and the disc. The dovetail was described by variable parameters under geometrical and physical constraints. Optimization was performed with a Genetic Algorithm (GA). The result of the optimization procedure is the optimal set of parameter values that minimizes the equivalent stress on the critical areas. Moreover, a surrogate function was utilized as a booster to the GA to save computational time. Stress analysis was performed with a commercial Finite Element (FE) software to provide the exact fitness value. An in-house code was developed to manage both the optimization process and the input/output interface with the FE software. The same code provides a decision-making core. This core checks for feasibility of the geometry of the current set of parameters. The expected result is an optimized profile in terms of Von-Misses equivalent stress.

Fluid film bearings for turbomachinery are designed to support the loads applied by the rotor system, often at high speeds when power loss in the bearing becomes significant and bearing temperatures can reach levels that can be detrimental to the long-term reliability of the support system. These requirements of supportive bearings require an intimate understanding of how bearing design variables affect their overall performance. Ideal bearings minimize power loss to increase machine efficiency and maintain low operating temperatures to ensure long-term reliability while meeting other design criteria such as minimum film thickness to provide proper support and avoiding high fluid pressures that can be harmful to the bearing structure. However, real world designs are often forced to sacrifice some of these design goals in order to preserve others. Therefore, further understanding of the relative opportunity costs associated with optimizing the bearing design with differently weighted performance metrics and their relationships to bearing design variables is invaluable to design engineers. This study explores the impact of eight bearing design variables on the performance of two tilting pad journal bearings supporting an eight-stage centrifugal compressor using design of experiments techniques applied to an established thermoelastohydrodynamic (TEHD) bearing model of tilting pad bearing performance. The bearing design variables analyzed include the radial clearance, pad arc spacing, pad axial length, pivot offset, preload, working fluid viscosity and viscosity index, and the number of pads. Each of the design variables — excluding the number of pads which was realistically constrained — were first varied over five levels each in a central composite design. These central composite designs were repeated for each of three values for number of pads. The responses obtained from the TEHD numerical simulations for each bearing design point were power loss, maximum pad temperature, minimum film thickness, and maximum fluid film pressure. The results from the central composite studies were fit with a multivariate least-squares regression model and a secondary series of experimental design studies were simulated around potential optimum design points to obtain a learning set to initialize direct optimization methods. Two direct multi-objective optimization methods, a sequential quadratic programming method and a multi-island genetic algorithm, were performed using Isight, a commercial software. A range of weighting parameters were selected for the optimization functions to find bearing designs that minimized power loss and pad temperature while maintaining pressure and film thickness criteria within acceptable design ranges for fluid film bearings. The resulting optimum design points allowed for a comparison between the design optimization approaches. The various strengths and weaknesses of the different methods are discussed. This study demonstrates how designers can use these approaches to view the relationships between design variables and important performance metrics to design better bearings for a wide range of applications.

Matching of various components (impeller exducer to impeller inducer, and vaneless diffuser and volute to the impeller) in a centrifugal compressor is critical for stage performance, but this is often neglected during compressor design and selection. This paper studies the importance of flow area matching for stage efficiency and how this matching process can be performed rapidly following a few simple principles. Methods for achieving optimum efficiency under different compressor operating conditions and size constraint are proposed and compared with experimental results. The purpose of this work is to draw attention to the key aspects of the matching, and provide easy-to-use design guidelines for engineers.

The present contribution proposes a Reduced Order Model based multi-fidelity optimization methodology for the design of highly loaded blades in low pressure compressors. Environmental, as well as, economical limitations applied to engine manufacturers make the design of modern turbofans an extremely complex task. A smart compromise has to be found to guarantee both a high efficiency and a high average stage loading imposed for mass reduction constraints, while satisfying stability requirements. The design of compressor blades, usually involves at the same time a dedicated parametrization set-up in highdimensional space and high-fidelity simulations capturing, at least, efficiency and stability as most impacting phenomena. Despite recent advances in the high-performance computing area, introducing high-fidelity simulations into automated optimization, or even surrogate assisted optimization, loops still stands as a endeavor for engineers. In this framework, the proposed methodology is based on multi-fidelity surrogate models capable of representing the physics at hand in reduced spaces inferred from both precise, albeit costly, high-fidelity simulations and abundant, yet less accurate lower-fidelity data. Finally, we investigate the coupling of the proposed hierarchised multi-fidelity non-intrusive Proper Orthogonal Decomposition based surrogates with an evolutionary algorithm to reduce the number of high-fidelity simulation calls towards the targeted optimum.

An optimal design of film cooling is a key factor in the effort of producing high efficiency gas turbine. Understanding of the fluid dynamics interaction between cooling holes can help engineers to improve overall thermal effectiveness. Modelling and correct prediction is a very complex problem, since the multiple phenomena involved, such as: mixing, turbulence and heat transfer. The present work performs an investigation of different cooling configurations ranging from single hole up to two rows. The main objective is to evaluate the double-rows interaction and the effect on film cooling. Strong nonlinear effects are underlined by different simulations, while varying blowing ratio and geometrical configuration of cooling holes. Meanwhile an initial analysis is performed using flat plate geometry, verification and validation is then extended to realistic stage of high pressure turbine. Multiple cooling holes configurations are embedded on the pressure and suction sides of the single stage. The main outcome is the verification of the thermal effectiveness improvement obtained by cooling jets interaction of multiple rows design. The effects of curvature surface and frame of reference rotation are also evaluated, underlying the differences with standard flat plate test cases.

In this paper we will numerically analyse flow mixing in multiple jets in a crossflow. The system comprises a row of six radially-distributed injectors around the main pipe. The configuration represents mixing zones in industrial systems where a counter jet can be formed in the injection plane. Flow mixing can be modified as a result of geometry and injection velocities. We propose a simple model to describe the counter jet length as a function of injection flow characteristics. We also develop empirical laws to help engineers design practical test facilities. We then vary the velocity ratio to obtain both impinging and non-impinging jets in the injection plane. The focus is mainly on flow characteristics around the radial injection plane in the case of impinging jets, examining the mixing quality and efficiency by introducing a passive scalar discharge in a nitrogen flow. The mean velocity and width of the counter jet are finally analyzed by changing the injection velocities. These results are compared to those of non-impinging jets. It is found that the non-impinging jet configurations are convenient for short length mixing chambers, while the impinging ones should be considered in the case of longer mixing chambers.

Many factors come into play in a successful turbomachinery design, but peak aerodynamic performance, maximum operating range, stress levels, and manufacturability are generally the dominant concerns. Depending on the application, other considerations such as acoustics, size, weight, and environmental issues can also come into the picture. The most effective designers take a holistic approach, which tries to incorporate all of these concerns at the beginning of the design process. One of the challenges facing aerodynamicists is a basic understanding of how these downstream issues can be accommodated in the initial design process. Typically, an aerodynamic engineer has minimal knowledge of manufacturing methods such as 5-axis machining. This paper will provide a basic understanding of the subsequent manufacturing impact of various decisions typically made during the preliminary and detailed aerodynamic design process. The paper will focus on radial turbomachinery, but much of the information provided is common for axial turbomachinery produced by 5-axis Computer Numerical Controlled (CNC) machining. In the initial phase of the design, basic layout decisions generally dictate the overarching manufacturing process. For components produced with 5-axis machining, these would typically be: flank milled or point milled, open or covered impeller, and for covered impellers, integral or welded shroud. The relative costs of these processes are considered in this work along with first-order estimates of their typical impact on aerodynamic performance and stress levels. Once a general layout (and accompanying manufacturing process) is determined, other geometric parameters of the components drive secondary costs in the manufacturing process. The secondary parameters include main blade count, the presence of splitter blades, the blade length, thickness, curvatures, leading and trailing edge shapes, and fillet radii. The use of Computational Fluid Dynamics (CFD) and numerical optimization has driven even more choices into blade shapes, such as splitter blades with different shapes and offsets from the main blade, bowed blades, irregular blade patterns, and non-axisymmetric hub shapes. In many cases, the aerodynamicists and CFD analysts have pushed geometry further than manufacturing capabilities are ready to accept. These secondary costs, and the aerodynamic compromises needed to reduce them, are also considered here. This paper attempts to lay out the basic principles of cost-effective manufacturing, and how these can be considered as early in the design process as possible. Specific examples are considered, and quantitative information is provided which can help guide the designer from the beginning and avoid expensive reworks resulting from downstream revisions. This paper will provide a framework for collaboration between aerodynamic engineers and the manufacturing teams assigned to produce the parts they have designed.

The measurement of unsteady total temperature is of great interest for the examination of loss mechanisms in turbomachinery with respect to the improvement of the efficiency. Since conventional thermocouples are limited in frequency response, several fast-response total temperature probes have been developed over the past years. To improve the spatial resolution compared to these existing probes and maintaining a high temporal resolution, a new fast-response total temperature probe has been developed at the Institute of Aircraft Propulsion Systems (ILA) in cooperation with Berns Engineers. The design of the probe allows a sensitive measuring surface below one square millimeter. A detailed insight into the design of the probe, the measurement principle, the calibration process, and an estimation of the measurement uncertainty is given in the present paper. Furthermore, to prove the functionality of the probe, first experimental results of a simple test bed and of area traverses downstream of the first rotor of a two-stage low pressure turbine are presented. It is shown, that the new probe is capable of detecting rotor characteristic effects as well as rotor-stator-interactions. In addition, a hot-spot is investigated downstream of the first rotor of the turbine and the findings are compared to effects known from literature.

Prior to the detailed design of components, turbomachinery engineers must guide a mean-line or throughflow design towards an optimum configuration. This process requires a combination of informed judgement and low-order correlations for the principle sources of loss. With these requirements in mind, this paper examines the impact of key design parameters on endwall loss in turbines, a problem which remains poorly understood. The analysis focuses on parametric computational studies of linear cascades, which represent a simplified model of real-engine flow. The designs are nominally representative of the Low Pressure Turbine blades of an aero engine. The paper examines the variations across the design space in a quantitative manner, examining the impact of flow angles, blade thickness and suction surface aerodynamics.

With the efforts of many researchers and engineers on the Supercritical CO 2 (S-CO 2 ) Brayton cycle technology development, the S-CO 2 Brayton cycle is now considered as one of the key power technologies for the future. Since S-CO 2 Brayton cycle has advantages in economics due to high efficiency and compactness of system, various industries have been trying to develop technologies on the design and analysis of S-CO 2 Brayton cycle components. Among various technical issues on the S-CO 2 Brayton cycle technology development, treatment of thermodynamic property near the critical point of S-CO 2 is very important since the property shows non-linear variation which causes large error on design and analysis results for ideal gas based methodologies. Due to the special behavior of thermodynamic property of CO 2 near the critical point, KAIST research team has been trying to develop a S-CO 2 compressor design and analysis tool to reflect real gas effect accurately for better design and performance prediction results. The main motivation for developing an in-house code is to establish turbomachinery design methodology based on general equations to improve accuracy of design and analysis results for various working fluids including S-CO 2 . One of the key improvements of KAIST_TMD which is an in-house tool for S-CO 2 turbomachinery design and analysis is the conversion process between stagnation condition and static condition. Since fluid is moving with high flow velocity in a compressor, the conversion process between stagnation and static condition is important and it can have an impact on the design and analysis results significantly. A common process for the conversion is based on the specific heat ratio which is typically a constant from ideal gas assumption. However, specific heat ratio cannot be assumed as a constant for the case of S-CO 2 compressor design and analysis because it varies dramatically near the critical point. Thus, in this paper, sensitivity study results on the state condition conversion between stagnation and static conditions with different approaches will be presented and further analysis on impact of the selected approaches on the final impeller design results will be discussed.

Driven by the increasing fossil fuel prices, global warming and climatic change, the world is currently witnessing an increasing development in renewable energy technologies, particularly those of wind energy. As such, engineers around the world are trying to optimize the design of wind turbines to maximize the captured energy while simultaneously minimizing the cost. This work aims to develop a mathematical tool to be used to compare different wind turbine designs and hence to reach the ultimate goal of an optimized wind turbine rotor designed specifically to operate in the Saharan regions of North Africa and the Middle East. As a case study, the main aerodynamic and structural parameters of the NREL 5MW virtual rotor have been optimized for the wind conditions prevalent at the Zaafarana site in Egypt. Specifically, the airfoil chord lengths and twist angles — smoothed using Bezier curves — as well as the layup sequence of the spar caps have been considered i n the optimization process which was carried out using a Genetic Algorithm (GA) developed i n MATLAB and coupled with NREL’s FAST Modularization Framework. The results showed that the NREL 5MW wind turbine design optimized for the site specific wind conditions of Zaafarana using airfoil families with low-sensitivity to dust accumulation, achieved a drop of 2.41% of the Levelized Cost of Energy of Energy (LCOE) over that of the baseline design. The developed turbine rotor design is tested for structural integrity commensurate with IEC standards.