Despite significant advancements in computational power and various numerical modeling in past decades, flow simulation of a multi-stage axial-flow compressor is still one of the most active areas of research, for it is the critical component in engine performance and operability, and there are so many elements that need to be looked into to predicting correct matching of the stages and accurate flow distribution inside the machine. Modeling unsteadiness, both deterministic and random types, and real geometries are among the most important features to be considered in such prediction. The authors have conducted in their previous studies a series of unsteady RANS (URANS) simulations of a 6.5-stage high-speed highly-loaded axial-flow compressor, and explored many unsteady effects as well as effects of real geometries such as Variable Stator Vane (VSV) clearance and inter-stage seal leakage flow on the compressor performance. However, all the analyses failed to predict correct stage matching, total pressure and temperature radial profiles, or mass-flow with adequate accuracies. In the present study, an Improved Delayed Detached Eddy Simulation (IDDES) with SST k-omega model is applied to the simulation of the same compressor configuration at aerodynamic design point. Fifth-order WENO scheme is employed for improved spatial accuracy to suppress significant increase in mesh size. Total number of mesh points are over 400 million for 1/10th sector model. Computations are ensemble averaged for 20 sector passage. Computed overall performance and flow field are compared with the compressor rig test data. The predictions of inter-stage total temperature radial profiles are noticeably improved over the URANS with the same mesh, discretization scheme and eddy turbulence model. Good comparison with the rig data indicates the current simulation is properly capturing the span-wise mixing phenomena. Unsteady flow field are compared between IDDES and URANS to locate the cause for the enhanced mixing. It is shown that components of Reynolds stress responsible for radial diffusion and anisotropic features are intensified in the tip leakage vortex at the rotor exit for the IDDES.

This paper aims at quantifying the stabilization effect of mistuning in transonic fan flutter. The results are used to support the evaluation of flutter boundary and to clarify the reason for the mismatch observed in the numerical predictions reported in our previous study. Mistuning is modeled by the deviation of blade-mode frequency, and the stability analysis of vibrating blades is formulated as an eigenproblem of the equation of motion including self-excited aerodynamic force obtained by fluid-structure interaction simulations. Statistics about the modal properties are obtained by Monte Carlo simulation. The change in the averaged damping rate and flutter boundary is evaluated in a wide range of mistuning levels and operating conditions. Nominal levels of mistuning due to manufacturing tolerance have little effect to the flutter boundary because the decline in aerodynamic damping is very steep. Therefore, the accuracy associated with the computational fluid dynamics is likely to have caused the mismatch in the flutter boundary. Histograms of modal properties show that the inter-blade phase angle and blade amplitudes in flutter mode can be highly scattered, even if the level of mistuning is nominal. For largely mistuned cases, new crests which do not exist in nominal cases appear in the eigenvalue histogram. They were found to be highly-localized, single-blade dominant mode.

Flutter has been a very important and severe problem for gas turbines, and its importance is increasing since a modern jet engine has very thin blades to reduce weight. There have been a lot of researches on its mechanism and evaluation technique[1][2][3], however, almost all of these researches are done by CFD, forced excitation and post evaluation of engine test. Among these activities, it became clear that torsion axis plays an important role to suppress flutter onset, however, there are few data on direct measurement during flutter on an influence of torsion axis. In the present study, two blade cascades which has different torsion axis were prepared and evaluate flutter can truly suppressed or not. Test rig was designed not to disturb circumferential disturbance which is generated by flutter at the test cascade, and pressure fluctuation transduces are introduced to measure pressure field during flutter vibration. The first test campaign will be held February 2016, and the experimental data will be compared with design and CFD results. The data will help clarifying the present design criteria can be truly applied to actual flutter onset and suppression. The present paper reports rig and blade design, and these evaluations by CFD simulations.

The unsteady aerodynamic characteristics of oscillating thin turbine blades were studied both experimentally and numerically to obtain the comprehensive knowledge on the aerodynamic damping of the blades operating in transonic flows. The experiment was carried out in a linear cascade tunnel by use of the influence coefficient method. The two flow conditions were adopted, namely, a near-design condition and an off-design condition with a higher back pressure. In the results for the near-design case, a strong vibration instability was observed in the positive side of the interblade phase angle. In the off-design case, however, the instability did not appear for almost all the interblade phase angles. A drastic change was found in the phase angle of unsteady aerodynamic force between the two cases, which change was a governing factor for the oscillation instability. Numerical simulation based on 2-D Euler equation revealed that the phase change came from the change in phase of the unsteady surface pressure across the shock impingement point on the blade suction surface in the off-design case. The numerical results also showed that the aerodynamic damping increased with increasing reduced frequency, and that the oscillation instability disappeared.

This paper describes numerical investigation of fan transonic stall flutter, especially focused on flutter bite. A transonic stall flutter occurs in high loaded condition at part rotating speed. A region of the transonic stall flutter occasionally protrudes to an operating line at narrow rotational speed range. This protrusion of flutter boundary is called flutter bite. In that case, it is necessary to re-design the blade for securing sufficient operating range. The re-design process might require some compromise on performance and/or weight, and takes long time. So it is important to understand the mechanism of the flutter bite. Two types of fan blade, one has a flutter bite and another dose not, are numerically studied with a 3D Navier Stokes CFD code. Numerical results show agreement with rig test results for the fans in qualitative sense. Detailed flow fields reveal that a detached shock wave and separation due to the shock boundary layer interaction play significant role for the flutter stability.

This paper describes the calculation of transonic stall flutter of a fan. A new CFD code has been developed and validated. The code is an unsteady 3D multi-block flow solver. The Reynolds-Averaged Navier-Stokes equations are solved using a finite volume method with Spallart-Allmaras 1 equation turbulence model. A grid deforming system is applied, so the new code is capable of simulating an oscillating blade row. This grid deforming system produces less grid distortion and the code has robustness for a blade oscillating calculation. The code has validated on an IHI’s research transonic fan rig test, and the result was in good agreement with the test data in the prediction of the flutter boundary. In the rig test at part-speed condition, stall-side flutter was experienced. In that condition, the inlet relative Mach number in the tip region is about unity. The aerodynamic work by the CFD at the near flutter condition is positive, which means that the flutter characteristic is unstable, while at other conditions the aerodynamic work is negative. The aerodynamic work increases rapidly just before the zero damping point with the increase of the blade loading. From the detailed CFD result, the shock wave on the suction surface contributes to the excitement of the blade oscillation, and the aerodynamic work of the shock wave has large value at the flutter condition.

The unsteady aerodynamic characteristics of an oscillating compressor cascade composed of Double-Circular-Arc airfoil blades were both experimentally and numerically studied under transonic flow conditions. The study aimed at clarifying the role of shock waves and boundary layer separation due to the shock boundary layer interaction on the vibration characteristics of the blades. The measurement of the unsteady aerodynamic moment on the blades was conducted in a transonic linear cascade tunnel using an influence coefficient method. The cascade was composed of seven DCA blades, the central one of which was an oscillating blade in a pitching mode. The unsteady moment was measured on the central blade as well as the two neighboring blades. The behavior of the shock waves was visualized through a schlieren technique. A quasi-three dimensional Navier-Stokes code was developed for the present numerical simulation of the unsteady flow fields around the oscillating blades. A k-ε turbulence model was utilized to adequately simulate the flow separation phenomena caused by the shock-boundary layer interaction. The experimental and numerical results complemented each other and enabled a detailed understanding of the unsteady aerodynamic behavior of the cascade. It was found that the surface pressure fluctuations induced by the shock oscillation were the governing factor for the unsteady aerodynamic moment acting on the blades. Such pressure fluctuations were primarily induced by the movement of impingement point of the shock on the blade surface. During the shock oscillation the separated region caused by the shock boundary layer interaction also oscillated along the blade surface, and induced additional pressure fluctuations. The shock oscillation and the movement of the separated region were found to play the principal role in the unsteady aerodynamic and vibration characteristics of the transonic compressor cascade.

Unsteady aerodynamic characteristics of an oscillating cascade composed of DCA (Double Circular Arc airfoil) blades were studied both experimentally and numerically. The test cascade was operated in high subsonic flow fields with incidence angles up to 5 degrees. Above 3 degrees of the incidence, a separation bubble was produced at the leading edge. The principal concern of the present study was placed on the influence of the separated region on the vibration instability of the cascade blades. The experiment was conducted in a linear cascade wind tunnel in which seven DCA blades were equipped. The central one could be oscillated in a pitching mode. The influence coefficient method was adopted for the measurement, where the unsteady aerodynamic moments were measured on the central blade and neighboring ones. For the numerical analysis, a quasi 3-D N-S code with k–ε turbulence model was developed. The experimental and numerical results complemented each other to obtain detailed understanding of the unsteady aerodynamic behavior of the cascade. It was found that the separation bubble at the leading edge governed the vibration characteristics of blades through the oscillation of the separation bubble itself on the blade surfaces. From the results of parametric studies, the phase shift of the oscillation of the separation bubble was found to be a key factor for determining the unsteady aerodynamic characteristics of the oscillating blades.