In this paper, a systematic CFD work is carried out with the aim to inspect the influence of different cascade parameters on the aerodynamic performance of a reversible fan blade profile. From the obtained results, we derive a meta-model for the aerodynamic properties of this profile. Through RANS simulations of different arrangements in cascades, the aerodynamic performance of airfoils are analyzed as Reynolds number, solidity, pitch angle and angle of attack are varied. The definition of a trial matrix allows the reduction of the minimum number of simulations required. The computed CFD values of lift and drag coefficients, stall margin and the zero-lift angle strongly depend on cascade configuration and differ significantly from standard panel method software predictions. In this work, X-Foil has been used as a benchmark. Particularly, the high influence of pitch angle and solidity is here highlighted, while a less marked dependence from the Reynolds number has been found. Meta-models for lift and drag coefficients have been later derived, and an analysis of variance has improved the models by reducing the number of significant factors. The application of the meta-models to a quasi-3D in-house software for fan performance prediction is also shown. The effectiveness of the derived meta-models is proven through a spanwise comparison of a reversible fan with the X-Foil based and meta-model based versions of the software and 3D fields from a standard CFD simulation. The meta-model improves the software prediction capability, leading to a very low global overestimation of the specific work of the fan.

The paper presents comparative aerodynamic and aeroacoustic studies on basic models of blade sections of low-speed, low-Reynolds-number axial fans. The wind tunnel experiments incorporated representative cambered plate and airfoil blade profiles. The aerodynamic measurements revealed that, for low Reynolds numbers, cambered plate blade sections may perform aerodynamically better than airfoil sections. A phased array microphone system, combined with a dipole beamforming and spatial filtering technique, offered a potential for localizing the noise sources in both streamwise and transversal direction. The acoustic studies focused on the profile vortex shedding noise. The results were qualitatively evaluated and compared with the semi-empirical noise prediction model developed by Brooks, Pope, and Marcolini. The measurements are considered as preparation of a dataset contributing to the background for designing high-efficiency, low-noise axial fans operating at low Reynolds number.

Typical turbine blade design systems are based on parametric studies and engineering experience. Best practices use feedback from manufactured hardware to additionally create producible designs and establish tolerances. Often there is a large gap in understanding of true hardware variation and the ability to describe and analyze it without over simplification. Previously, hardware inspections that feed analyses were limited based on Coordinate Measurement Machines (CMM) point inspection, optical measurements or gages designed to measure specific features. Those details were input into parametric models in an attempt to quantify the impact of hardware variation. Each measurement method has its limitations including accuracy and methodology which can affect the evaluation results. With the advent of 3D scanning, the complex 3D nature of airfoil designs can be described accurately with minimal simplification or assumptions. Actual hardware can be analyzed and assessed for design requirement compliance. In this study, 3D structured light inspection technology is used to assess aerodynamic performance and mechanical durability variation on a statistical sample of actual production hardware. A Honeywell designed auxiliary power unit second stage uncooled turbine rotor was used as a test case for this study. Thirty-five (35) blade castings randomly selected from three different casting mold lots were scanned using structured light to capture manufacturing process variation. The 3D scanning was completed using controls and robust scanning techniques certified for metrology use. Scans were assessed according to manufacturing requirements and correlated to aerodynamic and mechanical requirements using nominal machining geometry based on datum alignments. Analyses were performed using STAR-CCM+ for aerodynamic performance and ANSYS for mechanical durability. The designed nominal geometry was used as a baseline for comparison. HEEDS was used to automate the aerodynamic analysis process — replacing the baseline model with the scanned data, merging the new geometry with the base model, meshing, applying boundary conditions, and solving in remote Linux clusters. The automation process achieved a significant reduction in cycle time from several weeks to a few days. Similarly, the same structured light scan data was used to evaluate mechanical durability in ANSYS. Scan data was aligned to machining datums, boolean cut with nominal CAD machining geometry and directly converted to solid models using SpaceClaim. The solid model was meshed in ANSYS and merged onto the baseline FEA model for solving. Macros were used to automate the analysis process of each part scan and output stresses at critical locations for durability assessment. Using a combination of analytical tools and actual scan hardware, the impact of manufacturing variation could be understood. Integrating scan data directly to analysis reduces engineering assumptions and gives a better understanding of reality.

The flow leaking through the gap between rotor blade tips and casing surface in a turbine stage is an important source of energy loss. The current study uses a new concept named as T ip L eakage I nterrupters ( TLI ) to mitigate some of the adverse effects of the tip leakage flows and improve the efficiency of an axial turbine stage. The TLIs are a system of vortex generators attached onto the suction side of the turbine blade tip. The TLI design was developed in a proof of concept effort and they operate by inducing controlled vortical structures originating from strategically shaped/oriented multiple and sub-miniature vortex generators. These induced vortical structures, when properly interact with the tip leakage vortex reduce the damaging aerodynamic effects of the leakage flow. The TLIs in this investigation were mounted near the suction side corner of turbine blade tips rotating in a single-stage cold-flow turbine facility. In this investigation, three different parameters such as the mounting location of TLI on the airfoil tip region, the number of TLIs mounted and the specific orientation of TLI were varied. The TLI mounted near the minimum pressure point on the suction side of the blade generated the largest vortical structure that is counter rotating to the leakage vortex system and hence had the greatest effect in reducing the strength of the leakage vortex. Adding more TLIs on the blade suction surface was found to improve the tip leakage mitigation effort. The study showed that changing the specific orientation of the TLI with respect to the incoming flow drastically changes the rotational direction of the vortex it generates and its nature of interaction with the leakage vortex.

Gas turbine performance is highly dependent on the quality of the manufactured parts. Manufacturing variations in the parts can significantly alter the performance, especially efficiency and thus SFC. The legacy process is to accept variations within predefined profile tolerance limits and a few other qualitative parameters, mostly at a few, key two-dimensional aerofoil sections. With the widespread use of White light scans and other similar three-dimensional scans, this has improved to include the three-dimensional profile. The future however may lie with performance based quality assessment of manufactured parts, combined with quantitative surface quality assessment to implement an intelligent screening process for the parts. The adjoint method, typically used for shape optimization is adapted to provide a prediction of the impact on performance due to manufacturing variations. The work presented outlines a three stage quality assessment process for manufactured parts, involving three-dimensional profile tolerance based screening, followed by a surface curvature based screening and finally an Adjoint based performance prediction.

As engine development continues to advance toward increased efficiency and reduced fuel consumption, efficient use of compressor bypass flow, commonly used as cooling flow, becomes increasingly important. In particular, optimal use of compressor bypass flow yields an overall reduction of harmful emissions. The cooling flows used for cavity sealing between stages are critical to the engine and must be sufficiently maintained to prevent damaging ingestion from the hot gas path. To assess these cavity seals, the present study utilizes a one-stage turbine with true-scale engine hardware operated at engine-representative rotational Reynolds number and Mach number. Past experimental studies have made use of part-span rather than full-span blades to reduce flow rate requirements for the turbine test rig; however, such decisions raise questions about potential influences of the blade span on sealing effectiveness measurements in the rim cavity. For this study, a tracer gas facilitates measurements of sealing effectiveness in the rim cavity to compare measurements collected with full-span engine airfoils and simplified, part-span airfoils. The results from this study show sealing effectiveness does not scale as a function of relative purge flow with respect to main gas path flow rate when airfoil span is changed. However, scaling the sealing effectiveness for differing spans can be achieved if the fully-purged flow rate is known. Results also suggest reductions of purge flow may have a relatively small loss of seal performance if the design is already near a fully-purged condition. Rotor tip clearance is shown to have no effect on measured sealing effectiveness.

The flow in intake manifolds of internal combustion engine becomes more unsteady because of engine downsizing which is achieved by reducing cylinder number and increasing the boosting. Turbocharger compressor is thus exposed to the enhanced pulsating backpressure. This paper studies responses of a centrifugal compressor to the pulsating backpressure via experimentally validated numerical method. Firstly, CFD model with the volute and all impeller passages is established and validated by experimental measurements. Then the unsteady three-dimensional simulation is conducted on a single passage imposed by the pulsating backpressure conditions which are obtained by 1-D unsteady simulation. Results show that the performance of the passage evidently deviates from steady performance. Hysteresis loops of the performance appear at pulsating backpressure conditions, which encapsulate the performance at steady conditions. Moreover, the unsteadiness of the impeller performance is enhanced as the mass flow rate reduces. The performance and flow structures of the impeller near stall for the pulsating case are more favorable than that at corresponding constant backpressure. Furthermore, flow structures at points with the same instant mass flow rate are also notably different when they are located at different strokes of the pulse. The flow in the impeller is determined by not only the instant boundary condition but also the evolution history of flow field. The dynamic stall which is analogue to the phenomenon on pitching airfoil happens in the compressor and delays the instability of the device when it is exposed to pulsating backpressure. This study provides useful insights in the influence of pulsating backpressure on compressor performance in actual engine situations, from which better turbo-engine matching might be benefited.

The role of gas-to-wall temperature ratio on bypass transition along a highly loaded turbine guide vane is investigated through time-resolved heat flux measurements for different flow conditions. The tests are conducted in the von Karman Institute (VKI) isentropic compression tube (CT-2) facility. High response thin films mounted on the vane (VKI LS89 airfoil) coupled with analog circuits are used for the heat flux measurements. The first detectable wall heat flux fluctuations denote the onset of transition which is also evaluated by both the heat transfer coefficient and the intermittency factor distributions along the suction side. The exit Mach number is kept constant during each test by means of a downstream sonic throat while the gas-to-wall temperature ratio is varied from 1.14 to 1.51 by changing the inlet gas temperature. Four test cases are considered for different exit Mach numbers (0.52 and 0.86) and freestream turbulence intensity levels (0.8% and 5.3%) while the isentropic exit chord Reynolds number is maintained at 10 6 . In the present test campaign, the length and the evolution of the phenomenon indicate a measurable effect of the gas-to-wall temperature ratio for test cases with mild pressure gradients.

A detailed experimental and numerical investigation of the effect of endwall contouring on the secondary flow was performed in a highly loaded low pressure linear cascade. A comparison between a planar and three contoured configurations has been performed, and the three-dimensional endwall secondary flow structures and the secondary flow losses are here analyzed and discussed. For the contoured endwall, three locations of the same contouring were investigated, first one with the contouring starting upstream the airfoil and ending just before the leading edge of the airfoil, secondary one with the contouring starting upstream the airfoil and ending in the middle of the airfoil passage, the last one with the contouring starting just from the leading edge of the airfoil and continuing through the airfoil passage just ending at the trailing edge. The results suggest that the numerical results offer a reliable prediction for the endwall secondary flow structure together with suction surface separation bubble. It was found that all the three locations of contouring could reduce the secondary flow losses effectively. The location through the airfoil passage got the most benefit in the reduction of secondary flow losses whereas the all contouring upstream the airfoil location got the least. It was also found that the profile losses was affected by the contoured endwall.

A preliminary design optimization approach of axial flow compressors is developed. Loss correlations associated with airfoil geometry are introduced to relax the stringent requirement for the designer to prescribe the stage efficiency. In face of the preliminary design complexity resulted from the large number of design variables together with their stringent variation ranges and multiple design goals, the multi-objective optimization algorithm is incorporated. With such a developed preliminary design optimization method, the design space can be then explored extensively and the optimum designs of both high level overall efficiency and wide stall margin can be readily achieved. The preliminary design optimization method is validated in two steps. Firstly, an existing 5-stage compressor is redesigned without optimization. The obtained geometries and flow parameters are compared to the existing data and a good consistency is achieved. Then, the redesigned compressor is used as initial design and optimized by the developed multi-objective preliminary design optimization method, and significant performance gains are obtained, which demonstrates the effectiveness of the developed optimization methods.

This is the first part of a series of three papers on the simulation of turbulence and transition effects in a multistage low pressure turbine. In this first part, the extension, verification and validation of a Harmonic Balance (HB) method recently proposed by the authors to fully include established turbulence and transition models in the method is presented. As an alternating frequency/time-domain type method the implemented HB solver has the advantage of being able to utilize models (e.g. boundary conditions or residual functions) formulated in either the frequency- or time-domain. On the one hand this allows highly accurate nonreflecting boundary conditions formulated in the frequency-domain to be used along entry, exit or interface boundaries, and on the other hand complex nonlinear terms formulated in the time-domain to be used to describe nonlinear effects. Nevertheless, the wish to minimize the number of harmonics used to describe a given time-periodic unsteady flow, coupled with the often highly nonlinear nature of turbulence and transition models makes the full inclusion of such models in the HB method challenging. In this work the integration of Menter’s SST two-equation k–ω turbulence model along with Menter and Langtry’s two-equation γ– Re θ transition model in the context of a general framework for transport equations in the CFD solver TRACE is described in detail. Following the basic verification of the underlying transport equation framework, the implemented models are used to compute the well known high lift, low pressure turbine airfoil T106C and results are compared with the available experimental data as well as results from more conventional time-domain simulations. Alongside the basic validation of the models this testcase is furthermore used to investigate the importance of including higher harmonics, as opposed to only the zeroth harmonic, of the turbulence and transition models for the accurate prediction of the time-mean flow.

The accurate prediction of wall–roughness effects in turbomachinery is becoming critical as turbine designers address airfoil surface quality and degradation concerns arising from the shift to advanced ceramic matrix composite (CMC) or additively–manufactured airfoils operating in higher temperature environments. In this paper, a recently developed computational capability for accurate and efficient scale–resolving simulations of turbomachinery is extended to analyze the boundary–layer separation and transition characteristics in a rough–wall low–pressure turbine (LPT) cascade. The computational capability is based on an entropy–stable discontinuous–Galerkin spectral–element approach that extends to arbitrarily high orders of spatial and temporal accuracy, and is implemented in an efficient manner on a modern high–performance computer architecture. Results from the scale–resolving simulations of both smooth and rough airfoil cascades are presented and compared to previous experiments and numerical simulations. The results show that the suction–surface boundary layer undergoes laminar separation, transition, and turbulent reattachment for the smooth airfoil cascade, while in the presence of roughness the separation and transition behavior of the suction–surface boundary layer is substantially modified. The differences between the smooth and rough airfoil cascades are then highlighted by a detailed analysis of their respective turbulent flow fields.

Modern trend in installation design is moving towards very high-bypass ratio turbofans. Very high-bypass turbofans represent an effective way of improving the propulsive efficiency of civil aero-engines. Such engines require larger and heavier nacelles, which partially offset the gains in specific fuel consumption. The penalty associated with a larger installation can be mitigated by adopting thinner walls for the nacelle and by shortening the intake section. Such inlet sections are characterized by more restrictive operation condition because they are more prone to separation at high incidence flight conditions. Moreover, in short nacelle installations the by-pass guide vanes and pylon are closer to the fan blades and consequently the distortion due to potential effects induced by the presence of the pylon and non-axisymmetric OGV stage play a significant role in terms of unsteady interaction in the entire system. It is mandatory to consider the inlet, fan, bypass and pylon as a unique coupled system also at the design stage, for assessment of fan force. This kind of assessment is usually carried on by expensive URANS calculation. The factors leading to high computational demands are the spatial resolution required in the fan domain and the time resolution required to sample the fan blade passing frequency. Large savings are therefore possible if simplifications are introduced which relax the resolution requirements in the fan passages and change the nature of the computation into a steady-state computation for the ducts. The present contribution documents a simplified fan model for fan-intake computations based on the solution of the double linearization problem for unsteady, transonic flow past a cascade of thin aerofoils with finite mean load. The coupling with the intake flow and the bypass is performed by using the flow patterns at fan face and fan exit as boundary conditions for the fan model and computing circumferentially non-uniform boundary conditions for the intake and the bypass from the fan model. The computation of the flow in the intake, bypass and pylon is therefore reduced to a steady problem, whereas the computation of the flow in the fan is reduced to one steady problem and a set of linearised models in the frequency domain. The model is applied to a well-documented test case and compares favourably with experimental data and much more expensive three-dimensional, time domain computations.

In recent years, the demands of Maintenance, Repair and Overhaul (MRO) customers to provide resource-efficient after market services have grown increasingly. One way to meet these requirements is by making use of predictive maintenance methods. These are ideas that involve the derivation of workscoping guidance by assessing and processing previously unused or undocumented service data. In this context a novel approach on predictive maintenance is presented in form of a performance-based classification method for high pressure compressor (HPC) airfoils. The procedure features machine learning algorithms that establish a relation between the airfoil geometry and the associated aerodynamic behavior and is hereby able to divide individual operating characteristics into a finite number of distinct aero-classes. By this means the introduced method not only provides a fast and simple way to assess piece part performance through geometrical data, but also facilitates the consideration of stage matching (axial as well as circumferential) in a simplified manner. It thus serves as prerequisite for an improved customary HPC performance workscope as well as for an automated optimization process for compressor buildup with used or repaired material that would be applicable in an MRO environment. The methods of machine learning that are used in the present work enable the formation of distinct groups of similar aero-performance by unsupervised (step 1) and supervised learning (step 2). The application of the overall classification procedure is shown exemplary on an artificially generated dataset based on real characteristics of a front and a rear rotor of a 10-stage axial compressor that contains both geometry as well as aerodynamic information. In step 1 of the investigation only the aerodynamic quantities in terms of multivariate functional data are used in order to benchmark different clustering algorithms and generate a foundation for a geometry-based aero-classification. Corresponding classifiers are created in step 2 by means of both, the k Nearest Neighbor and the linear Support Vector Machine algorithms. The methods’ fidelities are brought to the test with the attempt to recover the aero-based similarity classes solely by using normalized and reduced geometry data. This results in high classification probabilities of up to 96 % which is proven by using stratified k-fold cross-validation.

Solid particles ingestion can deteriorate performance and stability of gas turbines through the build-up of deposits on the aerofoils. In some cases, the effect is sufficient to require an engine shut-down or to cause a failure. The fouling phenomenon can be characterized by two different phases: particle sticking and deposit evolution. The sticking process of an impinging particle has been deeply studied with several models available in the literature, both deterministic and stochastic. A less investigated phase is the the evolution of the deposit over time. A deposit is subjected to several forces that either tend to make it sticking to the surface or to detach from it. In the hot section of the gas turbine, the forces that act on the build-up tending to make it to adhere to the surface can be traced back to the van der Waals forces and, possibly capillary force if a certain amount of liquid phase is present. On the other hand, the detaching mechanism is related to the component investigated: if the particle is deposited on a vane, the drag and the shearing force are the only forces that tend to detach the particle, if a rotor blade is investigated, the centrifugal forces must be considered as well. On top of that, the deposit evolves over time in what is called sintering. During this process, the single particles deposited can melt together forming necks. If the temperature is sufficiently high, these necks increase in size until the former pores among particles are completely filled. This process is of paramount importance if the effect of the applied forces on the deposit needs to be investigated. The amount of material detached is indeed strongly dependent on the tension exchanged within various layers that constitute the build-up. This article focuses on the prediction of the build-up evolution on an HPT nozzle. The sintering process is modeled and related to the resisting strength of the deposit: an increasing sintering time reduces the deposit porosity and thus increase its strength. In order to monitor the stability of the deposit, the balance among detaching and attaching forces is carried out. The evolution of the vane shape is taken into account by using a moving mesh technique.

Ceramic matrix composite (CMC) have higher temperature durability and lower density property compared to nickel-based super-alloys which so far have been widely applied to hot section components of aero-engines / gas turbines. One of promising CMC systems, SiC-SiC CMC is able to sustain its mechanical property at higher temperature, though it inherently needs environmental barrier coating (EBC) to avoid oxidation. There are several requirements for EBC. One of such critical requirements is its resistance to particle erosion, whereas this subject hasn’t been well investigated in the past. The present work presents the results of a combined experimental and numerical research to evaluate the erosion characteristics of CMC+EBC material developed by IHI. First, experiments were carried out in an erosion test facility using 50 micron diameter silica as erosion media under typical engine conditions with velocity of 225 m/s, temperature of 1311 K, and impingement angles of 30, 60, and 80deg. The data displayed brittle erosion mode in that the erosion rate increased with impact angles. Next, multi-phase flow simulations were carried out for the experimental setup and the predicted erosion rates were compared with the measured data, and consequently, the applied numerical method was verified. The validated numerical method was then applied to compute three dimensional flow field, particle trajectories, and erosion profile around a generic turbine airfoil to assess the erosion characteristics of the proposed CMC+EBC material when applied to airfoil. The trajectories indicated that the particles primarily impacted the airfoil leading edge and the pressure surface. Surface erosion patterns were predicted based on the calculated trajectories and the experimentally-based erosion characteristics.

Transpiration Cooling is an effective cooling technology to protect hot section components such as gas turbine airfoils, rocket heads and space craft. This external cooling method has much higher efficiency than film cooling with holes when consuming the same amount of coolant, due to the uniformity of coolant distribution. However, pore blockage, which frequently occur during the operation of transpiration cooled components, prevented its application in turbine components which require long term stability. Dust deposition was one the main reasons causing blockage of pores for transpiration cooling. A lot of effort was devoted into dust deposition and erosion while optimization for the components themselves were generally difficult as the blockage caused by dusts was unpredictable for traditional sintered porous media. Additive manufacturing, with capability to precisely construct structures in small scales, is a considerable tool to enhance the controllability of porous media, and furthermore, to find a good solution to minimize the blockage disadvantage. Present study selected a cooling configurations containing perforate straight holes with an additive manufacturable diameter of 0.4 mm. Computational Fluid Dynamics (CFD) methods were utilized to model the pore blockage and its effect on heat transfer. A scripting code in addition to the ANSYS CFX solver was utilized to simulate the random blockage conditions of the holes. Two hundred numerical cases with four different blockage probabilities were calculated and statistically evaluated to quantify the disadvantage of pore blockage on the cooling effectiveness. Results obtained from the numerical analysis indicated that the overall blockage ratio was a dominating parameter for the cooling effectiveness. Upstream regions of the cooled surface were more sensitive to local blockage compared to downstream regions. Randomness of the cooling effectiveness increased with the increase of blockage probability. Present study provided a quantitative understanding of the random blockage disadvantage on the specific transpiration cooling configuration, and could benefit further optimization effort to reduce the blockage disadvantage of transpiration cooling using additive manufacturing.

A double-wall cooling scheme combined with effusion cooling offers a practical approximation to transpiration cooling which in turn presents the potential for very high cooling effectiveness. The use of the conventional conjugate CFD for the double-wall blade can be computationally expensive and this approach is therefore less than ideal in cases where only the preliminary results are required. This paper presents a computationally efficient numerical approach for analysing a double-wall effusion cooled gas turbine blade. An existing correlation from the literature was modified and used to represent the two-dimensional distribution of film cooling effectiveness. The internal heat transfer coefficient was calculated from a validated conjugate analysis of a wall element representing an element of the aerofoil wall and the conduction through the blade solved using a finite element code in ANSYS*. The numerical procedure developed has permitted a rapid evaluation of the critical parameters including; film cooling effectiveness, blade temperature distribution (and hence metal effectiveness) as well as coolant mass flow consumption. Good agreement was found between the results from this study and that from literature. This paper shows that a straightforward numerical approach that combines an existing correlation for film cooling from the literature with a conjugate analysis of a small wall element can be used to quickly predict the blade temperature distribution and other crucial blade performance parameters.

Gas turbine design has been characterized over the years by a continuous increase of the maximum cycle temperature, justified by a corresponding increase of cycle efficiency and power output. In such way turbine components heat load management has become a compulsory activity and then, a reliable procedure to evaluate the blades and vanes metal temperatures, is, nowadays, a crucial aspect for a safe components design. In the framework of the design and validation process of HPT (High Pressure Turbine) cooled components of the BHGE NovaLT™ 16 gas turbine, a decoupled methodology for conjugate heat transfer prediction has been applied and validated against measurement data. The procedure consists of a conjugate heat transfer analysis in which the internal cooling system (for both airfoils and platforms) is modeled by an in-house one-dimensional thermo-fluid network solver, the external heat loads and pressure distribution are evaluated through 3D CFD analysis and the heat conduction in the solid is carried out through a 3D FEM solution. Film cooling effect has been treated by means of a dedicated CFD analysis, implementing a source term approach. Predicted metal temperatures are finally compared with measurements from an extensive test campaign of the engine, in order to validate the presented procedure.

A fan blade is a complicated object and obviously it is subjected to geometrical uncertainties from manufacture tolerances and other production deviations. In spite of all uncertainties a fan blade should provide stable aerodynamic efficiency and strength properties. That is why it is considered to solve a multidimensional and multidisciplinary optimization task (aerodynamics, strength and flutter sensitivity) in robust statement under geometrical uncertainties. In the proposed test case geometrical uncertainties from the fan blade manufacture tolerances and deviations are considered. The probability density function (pdf) was obtained as a result of statistical operation upon the results of blade coordinate measurements. Approximately 2500 fan blades were measured by means of CMM process to reconstruct the pdf for more than 40 geometrical uncertainties (there are blade thicknesses for different airfoil locations in several cross-sections). CFD and FEM calculations were carried out in NUMECA FINE/Turbo and ANSYS software, correspondingly. The surrogate model technique (the response surface and the Monte-Carlo method implemented to RSM results) was applied for the uncertainty quantification and the robust optimization process for the task under consideration. APPROX software was used for surrogate model construction. The IOSO technology was employed as one of the robust optimization tools. This technology is also based on a widespread application of the response surface technique. As a result, robust optimal solutions (the Pareto set) for all 4 considered criteria (aerodynamic efficiency, structural properties, stall margin and flutter sensitivity) were obtained. The probabilistic criteria were assessed based on the results obtained. The robust optimization results were compared with the deterministic optimization results.