Turbine Blade Row Research Articles

Sound generation by entropy waves accelerated by blades

Entropy noise is the sound produced by the acceleration of convective temperature fluctuations (entropy waves) by non-uniform flows and is a major contributor to indirect combustion noise. Entropy waves are accelerated by nozzles in rocket engines or by the turbine blade rows in the combustors of gas turbines. While several analytical models have been developed for entropy noise generated by nozzles, not many efficient models exist for the case of blade rows. This is due to the more complex physical mechanisms involved: when entropy waves pass through cascades of blades, sound is produced due to the acceleration and turning of the entropy waves in the interblade region, but also unsteady forces are created in the blades and vorticity is shed from the trailing edge of the aerofoils. All these mechanisms are strongly coupled and affect the generated entropy noise. In this work, we explore numerically the importance of these mechanisms. To this end, we present time- and frequency-domain simulations of entropy waves interacting with several canonical aerofoils. Finally, a simple model is presented to estimate the entropy noise generated by each configuration.

The Interaction of Main Stream Flow and Cavity Flows in Turbine Center Frames and Turbine Vane Frames

Abstract This paper focuses on the interaction between the last high-pressure turbine (HPT) stage purge flows and the intermediate turbine duct (ITD) in modern turbofan engines. Two state-of-the-art ITD concepts are analyzed in this work: the turbine center frame (TCF), which is supported by symmetric aerodynamic strut fairings and generally adopted in conventional dual-spool engines; the turbine vane frame (TVF), which features turning struts and splitters and is typical of geared turbofan engines. The measurement campaigns for both setups are carried out in the transonic test turbine facility (TTTF) at Graz University of Technology. The test vehicles consist of an HPT stage, the ITD (TCF or TVF), and the first low-pressure turbine (LPT) vane or blade row. The same HPT stage is used for both ducts, to enable consistent, engine-representative inlet conditions between the two solutions. All the HPT stator–rotor cavities are supplied with purge flows by a secondary air system, with independent mass flow and temperature control for each purge line. Five-hole probe data are acquired at the inlet and outlet sections of the ITDs, to characterize the aerodynamic flow field entering and leaving the duct. In addition to the pneumatic probe tests, seed gas concentration measurements are performed in the same planes, to track down the cavity air in the mainstream and investigate its postegress behavior. Finally, detailed post-test computational fluid dynamics (CFD) results are presented to get additional insight into the flow phenomena developing through the strut passage.

Effects of recess depth and turbine reaction on the recessed tip thermal load in turbine blade cascades

Both tip leakage loss and thermal load of a turbine blade need to be kept as low as possible. In this investigation, detailed thermal load and near-wall flow distributions for recessed tips of different hst/s (recess depth-to-span ratio) in a turbine cascade of outlet-to-inlet velocity ratio (VR) = 2.01 are obtained with the naphthalene sublimation method and oil film technique, respectively. The results show that there is no repetitive local thermal load change on the recess floor, contrary to the flat tip case. As the recess deepens, the average thermal load on the recess floor decreases consistently, whereas the mass-averaged loss tends to decrease, to become minimal for hst/s = 1.88%, and then to increase. Present recessed tip flow models based on the flow visualizations clearly explain the reason for these tendencies. The deep recessed tip of hst/s = 5.63% has higher leakage outflow rate and lower near-wall momentum over the recess floor than the shallow recessed tip of hst/s = 1.88%. The comparisons of the previous data for VR = 1.2 and 2.4 with the present one show that as VR increases, the average thermal load on the recess floor as well as on the flat tip tends to increase, regardless of hst/s. VR (turbine reaction) is considered as the most important parameter that influences tip average thermal load in different turbine blade rows.

Effects of pressure side film cooling hole placement and condition on adiabatic film cooling effectiveness characteristics of a transonic turbine blade tip

The effects of film cooling hole placement location along the upper pressure side of a transonic squealer are considered. The thermal performance of four different film cooling configurations; B1, B2, B3 and B4, are considered using the University of Alabama in Huntsville's SS/TS/WT (supersonic/transonic/wind tunnel) experimental facility and a simulated turbine blade row using a linear cascade. Surface-varying results are provided for both the squealer blade tip surface, and for the upper pressure side of the squealer blade. These results are given for blowing ratios ranging from 0.42 to 3.20 in the form of spatially-resolved and spatially-averaged adiabatic film cooling effectiveness distributions. Because local static pressure variations and gradients vary in a significant manner at different blade locations, coolant accumulations near surfaces (and the associated surface thermal protection and film effectiveness distributions) produced by the different film cooling configurations are vastly different along the squealer blade tip surface. The B1 and B3 film cooling configurations show that substantial magnitudes of adiabatic film cooling effectiveness are present along the pressure side rim and the suction side rim. Also present is a considerable region of locally increased film cooling effectiveness and a wider region of surface protection within the squealer recess region for the B2 film cooling configuration. The contributions are unique because of a deficit of experimental data for film cooled squealer blades which operate with transonic flow, and because the present investigation is the first to consider the effects of hole placement location for upper pressure side film cooling arrangements with transonic flow conditions.

Efficiency of Pulsating Base Bleeding to Control Trailing Edge Flow Configurations

As high-pressure-turbines operate at extreme temperature conditions, base bleed can be applied at the trailing edge of the airfoils, enhancing the thermal protection along the trailing edge surface, but also disrupting the trailing edge flow and altering the overall aerodynamic pressure losses. The current work explores the potential use of base bleed as a flow control tool to modulate the flow between turbine blade rows. Through the numerical analysis of a symmetric airfoil immersed in a subsonic flow, the effects that trailing edge ejection has on the base region properties and the downstream flow are evaluated. In particular, previous research constrained to steady blowing is now extended to consider an unsteady pulsating base bleed injection. Three injection frequencies are investigated, covering a wide range of base bleed intensities. The results presented herein demonstrate that pulsating bleed flow is more efficient than its steady counterpart in terms of reducing pressure losses and controlling the primary frequency of the downstream oscillations for the same mass flow injection.

Aerodynamic performance of suction-side squealer tip in impulse/reaction turbine blade rows

In order to understand the impact of turbine reaction on the SS (suction-side squealer) tip loss, the loss map for the SS tip with the variations of both tip clearance and squealer height is completed in the blade row of medium turbine reaction (exit-to-inlet velocity ratio, VR = 2.0), and then the present SS tip losses are compared with the previous ones in the blade rows of weak and strong turbine reactions (VR = 1.2 and 2.4). The SS tip loss map shows that the loss change with squealer height is not so large and its maximum loss reduction is less than2.9% of flat tip loss. Regardless of tip clearance and squealer height, the present SS tip performs worse than the full squealer tip in the loss reduction for VR = 2.0, and this performance gap is more noticeable for wider tip clearance. The comparisons of the present SS tip losses for VR = 2.0 with those for VR = 1.2 and 2.4 show that the SS tip-to-flat tip loss ratio is influenced strongly by VR (reaction) of impulse/reaction turbine blade row. Irrespective of squealer height, the SS tip has a considerable loss reduction only in the blade row of VR = 2.4. Especially for low squealer heights without structural defect, the SS tip-to-flat tip loss ratio continues to decrease, as VR increases from 1.2 to 2.4. The present over-tip flow visualizations give a clear explanation about tip gap flow physics in the blade rows of different VR.

Effects of Near-Wall Streamwise Vortex on the Film Cooling Performance of a Fan-Shaped Hole

Abstract Fan-shaped hole is a typical shaped hole frequently used in gas turbine film cooling, which may operate in a complex vortical flow environment of turbine blade rows. It is critical to understand the destructive mechanism of the film cooling by near-wall vortex on the endwall surface. A vortex generator (VG) was installed in front of a discrete fan-shaped hole to produce a vortical environment on a flat plate. Heat transfer effects were experimentally studied for mainstream velocities of 20 m/s by measuring the adiabatic film cooling effectiveness at blowing ratios M = 0.5 to 2.5. Fluid dynamics effects of vortical upwash and downwash movements were numerically analyzed. Results show that near-wall streamwise vortex destroys the film flow and intensifies the mixing of hot gas and coolant flow, which leads to the deterioration of film cooling performance. Film coverage area varies remarkably with respect to the positions of streamwise vortex. An efficient blowing ratio has a strong ability to mitigate the influence of near-wall vortex. At M = 2.5, streamwise vortex can restrain liftoff tendency of coolant jet and retard the decrease of film cooling effects.

Modeling of a turbine bladerow with stagger angle variation using the multi-fidelity influence superposition method

Manufacturing variabilities can significantly affect a turbine bladerow's performance. With the push for turbine's efficiency and reliability more than ever, the understanding of manufacturing variability impacts is sought after. It has been shown that a multi-passage simulation domain is required to capture the interaction among varied blades. This is a challenge for conventional single-passage methods, largely due to the requirement of prohibitive computational resources. The present work introduces an attempt to solve this challenge by combining the multi-fidelity method and the influence superposition approach. The proposed methodology combines the accuracy of a high-fidelity simulation and the speed of a low-fidelity simulation. Therefore, the multi-fidelity predictions tend to be both accurate and fast. The key enabler is that only a small set of configurations is needed to pre-compute the source term. Two representative geometries of a subsonic low-pressure turbine and a transonic high-pressure turbine have been chosen as the test cases. Each test case is subject to two further stagger angle variation patterns, namely alternating and sinusoidal pattern. The low-fidelity method has been shown to be inadequate for the transonic high-pressure turbine test case, owning to its incapability to capture correctly the shockwave formation and the shockwave/wake interaction. On the other hand, the proposed multi-fidelity method has been successful to match qualitatively and quantitatively compared to the direct high-fidelity solution. More interestingly, the multi-fidelity method has a reduced computational overhead of one order of magnitude compared to the direct high-fidelity simulation. With an ability to accurately and efficiently predict the manufacturing variability effects, this method provides a tool for engineers to explore and optimize their blading designs.

Innovative Design, Structural Optimization, and Additive Manufacturing of New-Generation Turbine Blades

Abstract The need for high performances is pushing the complexity of mechanical design at very high levels, especially for turbomachinery components. In this field, structural topology optimization methods together with additive manufacturing techniques for high-resistant alloys are considered very promising tools, but their potentialities have not been deeply investigated yet for critical rotating components like new-generation turbine blades. In this context, this research work proposes a methodology for the design, the optimization, and the additive manufacturing of extremely stressed turbomachinery components like turbine blade rows. The presented procedure pays particular attention to important aspects as fluid–structure interactions (forced response and flutter phenomena) and fatigue behavior of materials, going beyond the standard structural optimization approaches. The new design strategy enables a substantial reduction of the component mass, limiting the maximum stress and improving the vibrational behavior of the system in terms of eigenfrequencies, modal shapes, and fatigue life. Furthermore, the numerical procedure shows robustness and efficiency, making the proposed methodology an appropriate tool for rapid design and prototyping and for reducing the design costs and the typical time to market of this type of mechanical elements. The procedure has been applied to a low-pressure turbine rotor to improve the aeromechanical characteristics while keeping the aerodynamic performance. From the original geometry, mode-shapes, forcing functions (due to rotor/stator interactions), and aerodynamic damping have been numerically evaluated and used as input data for the subsequent topological optimization. Finally, the optimized geometry has been numerically and experimentally verified to confirm the improved aeromechanical design. After the structural topology optimization, the final geometries provided by the procedure have been properly rendered to make them suitable for additive manufacturing. Some prototypes of the new optimized turbine blade have been manufactured from aluminum alloy to be tested mechanically and in terms of fatigue.

An Experimental Study of the Leakage Flow Effect on the Film Cooling Effectiveness of a Gas Turbine Shroud

Abstract In the vicinity of gas turbine blades, a complex flow field is formed due to the flow separation, reattachment, and secondary flows, and this results in a locally non-uniform and high heat transfer on the surfaces. The present study experimentally investigates the effects of leakage flow through the slot between the gas turbine vane and blade rows on the film cooling effectiveness of the forward region of the shroud ring segment. The experiment is carried out in a linear cascade with five blades. Instead of the vane, a row of rods at the location of the vane trailing edge is installed to consider the wake effect. The leakage flow is introduced through the slot between the vane and blade rows, and additional coolant air is injected from the cooling holes installed at the vane’s outer zone. The effects of the slot geometry, cooling hole configuration, and blowing ratio on the film cooling effectiveness are experimentally investigated using the pressure-sensitive paint (PSP) technique. CO2 gas and a mixture of SF6 and N2 (25% + 75%) are used to simulate the leakage flow to the mainstream density ratios of 1.5 and 2.0, respectively. The results indicate that the area-averaged film cooling effectiveness is affected more by the slot width than by the cooling hole configuration at the same injection conditions, and the lower density ratio cases show higher film cooling effectiveness than the higher density ratio case at the same cooling configuration.

Development of a fully-coupled harmonic balance method and a refined energy method for the computation of flutter-induced Limit Cycle Oscillations of bladed disks with nonlinear friction contacts

Flutter stability is a dominant design constraint of modern gas and steam turbines. To further increase the feasible design space, flutter-tolerant designs are currently explored, which may undergo Limit Cycle Oscillations (LCOs) of acceptable, yet not vanishing, level. Bounded self-excited oscillations are a priori a nonlinear phenomenon, and can thus only be explained by nonlinear interactions such as dry stick–slip friction in mechanical joints. The currently available simulation methods for blade flutter account for nonlinear interactions, at most, in only one domain, the structure or the fluid, and assume the behavior in the other domain as linear. In this work, we develop a fully-coupled nonlinear frequency domain method which is capable of resolving nonlinear flow and structural effects. We demonstrate the computational performance of this method for a state-of-the-art aeroelastic model of a shrouded turbine blade row. Besides simulating limit cycles, we predict, for the first time, the phenomenon of nonlinear instability, i.e., a situation where the equilibrium point is locally stable, but for sufficiently strong perturbation (caused e.g.by an impact), the dry frictional dissipation cannot bound the flutter vibrations. This implies that linearized theory does not necessary lead to a conservative design of turbine blades. We show that this phenomenon is due to the nonlinear contact interactions at the tip shrouds, which cause a change of the vibrational deflection shape and frequency, which in turn leads to a loss of aeroelastic stability. Finally, we extend the well-known energy method to capture these effects, and conclude that it provides a good approximation and is useful for initializing the fully-coupled solver.

Forced Response of a Low-Pressure Turbine Blade using Spectral/hp Element Method: Direct Numerical Simulation

The Spectral/HP element method has been applied to perform Direct Numerical Simulations (DNS) over a single T106A turbine blade-row using the open source software Nektar++. The main goal of the current study is to perform preliminary investigations at modest Reynolds and Mach numbers, 8000 and 0.1 respectively, for uniform, steady flow past the aerofoil by employing Nektar++’s solver for the 2D Navier-Stokes equations for incompressible flow. The mesh was firstly validated against results obtained using the same software and for a similar set of parameter values. One dimensional, pitch-wise harmonic vibrations were subsequently imposed on the blade by means of a coordinate transformation. A parametric study in terms of the frequency and amplitude of the vibrations was carried out. The effects of the vibrations on entire domain, along the blade surface and in its wake were assessed. The pressure on the blade surface and the wake loss were each decomposed into components arising due to the mean flow and due to the vibrations. In each case the dominant components were then identified for the values of frequency and amplitude considered here.

Analysis of Friction-Saturated Flutter Vibrations With a Fully Coupled Frequency Domain Method

Abstract Flutter stability is a dominant design constraint of modern turbines. Thus, flutter-tolerant designs are currently explored, where the resulting vibrations remain within acceptable bounds. In particular, friction damping has the potential to yield limit cycle oscillations (LCOs) in the presence of a flutter instability. To predict such LCOs, it is the current practice to model the aerodynamic forces in terms of aerodynamic influence coefficients for a linearized structural model with fixed oscillation frequency. This approach neglects that both the nonlinear contact interactions and the aerodynamic stiffness cause a change in the deflection shape and the frequency of the LCO. This, in turn, may have a substantial effect on the aerodynamic damping. The goal of this paper is to assess the importance of these neglected interactions. To this end, a state-of-the-art aero-elastic model of a low pressure turbine blade row is considered, undergoing nonlinear frictional contact interactions in the tip shroud interfaces. The LCOs are computed with a fully coupled harmonic balance method, which iteratively computes the Fourier coefficients of structural deformation and conservative flow variables, as well as the a priori unknown frequency. The coupled algorithm was found to provide excellent computational robustness and efficiency. Moreover, a refinement of the conventional energy method is developed and assessed, which accounts for both the nonlinear contact boundary conditions and the linearized aerodynamic influence. It is found that the conventional energy method may not predict a limit cycle oscillation at all while the novel approach presented here can.

Effect of Purge on the Secondary Flow-Field of a Gas Turbine Blade-Row

Abstract Turbine disc cooling is required to protect vulnerable components from exposure to the high temperatures found in the mainstream gas path. Purge air, bled from the latter stages of the compressor, is introduced to the turbine wheelspace at low radius before exiting through the rim-seal at the periphery of the discs. The unsteady, complex flowfield that arises from the interaction between the purge and mainstream gases modifies the structure of secondary flows within the blade passage. A computational study was conducted using an unsteady Reynolds-averaged Navir–Stokes (RANS) solver, modeling an engine-representative turbine stage. Preliminary results were validated using experimental data from a test rig. The baseline secondary flowfield was described, in the absence of purge flow, demonstrating the classical rollup of the horseshoe vortex and subsequent convection of the two legs downstream. The unsteady behavior of the model was investigated and addressed, resulting in recommendations for modeling interaction phenomena in turbines. A superposed purge flow, resulting in egress through the upstream rim-seal, was shown to modify the secondary flowfield in the turbine annulus. The most notable effect of egress was the formation of a large plume forming near the pressure minima associated with the blade suction surface. The egress was turned by the mainstream flow, creating a vortical structure consistent in rotational direction to the pressure-side leg of the horseshoe vortex; the pressure-side leg was subsequently strengthened and showed an increased radial migration relative to the unpurged case. The egress plume was also shown to overwhelm the suction-side leg of the horseshoe vortex, reducing its strength.

The Impact of Combustor Turbulence on Turbine Loss Mechanisms

Abstract A blade row that is located downstream of a combustor has an extremely high turbulence intensity at the inlet, typically above 10%. The peak turbulent length scale is also high, at around 20% of the chord of the downstream blade row. In a combustor, the turbulence is created by impinging jets in crossflow. This may result in the turbulence being anisotropic in nature. The aim of this paper is to investigate the effect of combustor turbulence on the loss mechanisms which occur in a turbine blade row. The paper has a number of important findings. The combustor turbulence is characterized and is shown to be isotropic in nature. It shows that, when no pressure gradient is present, combustor turbulence increases the loss of a turbulent boundary layer by 22%. The mechanism responsible for this change is shown to be a deep penetration of the turbulence into the boundary layer. It shows that the presence of combustor turbulence increases the profile loss and endwall loss in the turbine cascade studied by 37% and 47%, respectively. The presence of combustor turbulence also introduces a freestream loss resulting in the total loss of the turbine cascade rising by 47%. When these loss mechanisms were applied to the vane alone, of an engine representative high-pressure turbine, it was found to result in a 1.3% reduction in stage efficiency.

Multi-Wave Vibration Caused by Flutter Instability and Nonlinear Tip-Shroud Friction

Abstract Measurements revealed the contribution of multiple traveling waves to the flutter vibrations of bladed disks. Saturated flutter vibration, whether in this multi-wave or in its better-understood single-wave form, is a nonlinear phenomenon. However, it is still not understood of what physical origin the relevant nonlinearities are, and under what conditions single-wave or multi-wave flutter vibration occurs. Recent theoretical work suggests that multi-wave flutter vibration can be explained by strongly nonlinear frictional interblade coupling. The verity of this hypothesis is strictly limited by the simplicity of the considered model, namely, a cyclic chain of seven oscillators with frictional coupling and rather unrealistic aeroelastic behavior. In this work, it is demonstrated that nonlinear dynamical contact interactions at tip-shrouds are a likely cause for the observed multi-wave flutter vibration. To this end, a more realistic structural turbine blade row model with a more realistic aeroelastic behavior is considered. Some insight into its intriguing dynamics, dependence of limit states on initial conditions, and eigenvalue placement is provided. For instance, it is shown that there is an intimate relation between internal combination resonance conditions of certain traveling wave modes and the spectral content of single- and multi-wave flutter oscillations.

Constructal Design of Elliptical Conduits for Cooling of Gas Turbine Blades with External Thermal Barrier Coating

Gas turbines (GTs) are thermal machines used to transform the energy released in combustion with a hydrocarbon into mechanical power, in order to drive a machine or generate thrust in aircraft. The critical issue in the GT design are the parts exposed to extreme mechanical and thermal conditions, e.g., the first row of turbine blades. The GT thermal efficiency is limited by the maximum temperature the blade materials can withstand without softening or creeping. Currently, the maximum operating temperature is above the softening point of the blade material thanks to techniques of ceramic coatings of low thermal conductivity, called Thermal Barrier Coating (TBC), and techniques of blade cooling. The internal cooling of blades involves conduits inside them for air that comes from a bleed in an intermediate compressor stage. The air bleeding is around 3 to 5% of the main GT flow. This air and the heat flow that it receives are not used to generate power, so it is necessary to optimize the cooling techniques in order to control the temperature using the least amount of air and minimum heat flux evacuated, for holding the GT overall efficiency high. The present work studies the internal cooling of Elemental Gas Turbine Blade (EGTB) with a fixed thickness of the TBC and the optimization of the conduit shape and position over a cross section in 2D. The optimization is carried out by exhaustive searching method based on the Constructal Theory. The optimization of the position, size, and aspect ratio of EGTBs was done for two types of standard elliptical conduits of different geometries, uniformly distributed. Two different objective functions are analyzed: minimum maximum temperature on the metal and maximum heat evacuation efficiency. The outcome of this work establishes that the use of elliptical conduits of aspect ratio 2:5 leads to improvement in the thermal performance of cooled blades. As compared with circular conduits of the same area, elliptical conduits allow transfer of a greater amount of heat; with a correct design, they enable a lower maximum temperature on the metal. Besides, the constructal designs obtained in this study for the minimum maximum relative temperature $$\tilde{T}_{\rm{max}}$$ or maximum heat evacuation efficiency ξ were not identical.

Analysis of Nonstationary Loadings and Vibration Amplitudes of the Blades of the Last Turbine Cascade Tak-ing Into Consideration the Detuning

Nonstationary phenomena caused by the blade vibration under the action of disturbing forces that are characterized by the energy exchange between the gas flow and the vibrating blades and make the foundation for a physical mechanism of self-excited oscillations that can either damp (aerodamping) or develop in the stable mode of self-induced vibrations or in the unstable mode of flutter that can result in the construction failure. One of the approaches to an increase in the vibration stability of the blades is the detuning of natural modes and frequencies. Based on the developed mathematical model and the numerical method of aeroelastic behavior of the blade rim in the transonic gas flow (the coupled problem of nonstationary aerodynamics and elastic vibrations of the blades), we gave the numerical analysis of aeroelastic behavior of the turbine blade row in the three-dimensional flow of ideal gas through the turbine cascade taking into consideration the detuning of blade natural modes and frequencies. To solve the coupled problem we used the partially integral method that includes Euler integral equations and those of the dynamics of vibrating blades (modal approach) at each time step with the information exchange. This method of the solution of coupled aeroelastic problem enables the prediction of the amplitude-frequency spectrum of blade vibrations in the three dimensional flow of ideal gas, including forced vibrations, self-excited vibrations and self-induces vibrations in order to increase the efficiency and reliability of the blade rows of turbine machines. As a result of the investigation we obtained gas dynamic parameters in the form of nonstationary fields, nonstationary loads that have action on the blades and the amplitude and frequency spectra of blade vibrations.

Linear Acoustic Analysis of the Preburner of an Oxidizer-Rich Staged Combustion Engine

The linear acoustic behavior of an oxidizer-rich staged combustion engine preburner assembly that was similar to that of the RD-170 has been investigated by using real fluid speeds of sound to accommodate supercritical pressures. The speeds of sound, and hence the resulting frequencies, were approximately 20% higher than would be expected from ideal gas assumptions, which is a critical point of difference between the acoustic behavior of the preburner versus the main chamber. The results show that the mode shapes in the preburner–turbine inlet assembly are largely invariant with respect to the boundary condition used for the turbine blade row, but the frequencies are noticeably affected. The 2L mode is totally suppressed for both the acoustically open case and the case with the blade row impedance condition. The remaining longitudinal modes are damped. Both the frequency and the mode shape of the 1T mode remain totally invariant because the oscillations are confined to the injector regions and do not exhibit acoustic velocity with components normal to the impedance surface. The lack of damping of the 1T mode by the blade rows, combined with the relatively high magnitudes of acoustic velocity near the combustion zone, suggests that the 1T mode is a likely candidate to experience instability if it is suitably energized by the heat release.

Aeroelastic Behaviour of Turbine Blade Row in 3D Viscous Flow

Aeroelastic Behaviour of Turbine Blade Row in 3D Viscous Flow