Accurate prediction of aerodynamic damping is essential for flutter and forced response analysis of turbomachinery components. Reaching a high level of confidence in numerical simulations requires that the models have been validated against the experiments. Even though a number of test cases have been established over the past decades, there is still a lack of suitable detailed test data that can be used for validation purposes in particular when it comes to aero damping at high reduced frequencies which is more relevant in the context of forced response analysis. A new transonic cascade test rig, currently undergoing commissioning at KTH, has been designed with the goal to provide detailed blade surface unsteady pressure data for compressor blades profiles oscillating at high reduced frequencies. The paper provides an overview of the blade actuation system employed in the test rig and presents the result of a series of bench tests characterizing the blade vibration amplitudes achieved with this actuation system.

A new set of steady-state aerodynamics data from the recently build KTH transonic linear cascade (TLC) is presented. The operational point is representative of an open source virtual compressor (VINK) operating near stall at 70% speed line. The inlet Mach number (M = 0.81) is above its airfoil critical value with high incidence angle, producing a leading edge separation bubble. The test object in the TLC has a 2% tip gap, and the reference measurement plane is at 85% span. Results show repeatability, and periodicity. Upstream flow field was mapped by a non-intrusive, Laser-2-Focus (L2F) method. The blade loading was recovered by a set of pressure taps. Flow field properties downstream the test section were recovered by a five-hole probe. A flow structures comparison was performed between numerical and experimental data. Numerical results were obtained from the commercial software Ansys CFX, where different turbulence models were tested: SST, SST with reattachment method (RM) and SST with turbulent transition (γ - θ). The closest match to the experimental data was recovered by SST-γ - θ RM. Separation onset and flow properties downstream are strongly affected by the selected turbulence transition model. The results are aimed to work as step towards validation for CFD in separated regions. These data is the first of a series of test campaigns for experimental validation for aerodynamic damping in separated flows. 

The steady-state aerodynamics and the aeroelastic response have been analyzed in an oscillating linear transonic cascade at the KTH Royal Institute of Technology. The investigated operating points (Π=1.29 and 1.25) represent an open-source virtual compressor (VINK) operating at a part speed line. At these conditions, a shock-induced separation mechanism is present on the suction side. In the cascade, the central blade vibrates in its first natural modeshape with a 0.69 reduced frequency, and the reference measurement span is 85%. The numerical results are computed from the commercial software Ansys CFX with an SST turbulence model, including a reattachment modification (RM). Steady-state results consist of a Laser-2-Focus anemometer (L2F), pressure taps, and flow visualization. Steady-state numerical results indicate good agreement with experimental data, including the reattachment line length, at both operating points, while discrepancies are observed at low-momentum regions within the passage. Experimental unsteady pressure coefficients at the oscillating blade display a fast amplitude decrease downstream, while numerical results overpredict the amplitude response. Numerical results indicate that, at the measurement plane, for both operating points, the harmonic amplitude is dominated by the shock location. At midspan, there is an interaction between the shock and the separation onset, where large pressure gradients are located. Experimental and numerical responses at blades adjacent to the oscillating blade are in good agreement at both operating points.

In a turbomachine, with time, wear and depositions modify the surface from smooth to rougher characteristics. These effects are also prevalent in emerging manufacturing technologies such as additive manufacturing (AM). Surface roughness characterization is based on a correlation between a physical length scale or surface statistic moments, and a non-physical equivalent sand-grain roughness (ks). Depending on the flow characteristics and ks, three wall regimes can be considered: hydraulically smooth, transitional or fully rough. In compressors, an increase in surface roughness translates into a reduction in efficiency and pressure ratio. While steady-state roughness effects over airfoils and stage performance are well documented, its consequences over the aerodynamic damping are not. This paper aims to investigate numerically the effect of surface roughness on the aerodamping for the three wall regimes. The test geometry is the first stage of an open-source transonic axial compressor. Operating points considered are peak efficiency and near-stall conditions at nominal speed (N100) and part-speed (N70). The presented numerical simulations are obtained using the commercial software Ansys CFX with SST as the turbulence model with a reattachment modification. At near-stall part-speed, there are non-physical separation regions that are mesh independent but assumed to come from the numerical roughness implementation. Amplitude fluctuations in the aerodynamic damping per unit area, s∗, are driven by the tip gap presence as well as normal shocks at N100 whereas the presence of separated flow regions appear to have a negligible effect at N70. The phase between the pressure and blade motion appears to remain almost constant regardless of the wall regimes, implying that only the amplitude distribution drives the aerodynamic damping stability. This numerical observation is aimed to be experimentally tested at the transonic linear cascade at KTH Royal Institute of Technology.