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.

Due to Intermediate Compressor Ducts (ICDs) potential to reduce overall engine weight and therefore specific fuel consumption, many aerospace industry participants have embarked on programs to develop shorter and more aerodynamically "aggressive" ICD. Unfortunately, most "aggressive" designs have tended to either separate or closely approach their separation limit. As a result, in this paper, the ability of active flow control techniques (boundary layer (BL) blowing and suction) were investigated and novel schemes were designed to suppress the flow separations and thus increase the aerodynamic robustness of such ducts. The test case used in this study is a state-of-the-art, highly aggressive, ICD developed by the Institute of Propulsion Technology, German Aerospace Center (DLR), located aerodynamically close to its separation limit operating at its nominal operating point. Numerical simulations were used to characterize the aerodynamic baseline of the duct as well as the modified design. Analysis of the baseline design results confirmed the strong tendency of the flow to separate at the hub-strut corner (corner separation), mid-strut (passage separation) and in the shroud (near duct exit) regions. Boundary layer energization schemes were then explored at different streamwise locations, to identify its optimum location. The results showed that the application of a BL energization scheme (BL blowing) on the hub resulted in a positive reduction in the passage and hub-strut corner separation while the application of BL suction (at the shroud) assisted in reducing the aerodynamic separations. Finally, when both schemes were employed simultaneously, an aggregated reduction of approximately 21% in the mass-averaged total pressure loss coefficient (compared to the baseline case) was predicted.

The paper addresses the aeromechanical design of lightweight fan blades for an electric fan thruster. It explores the use of carbon-fiber reinforced composites to reduce overall weight and enhance aeroelastic performance, particularly in terms of flutter stability. A comparison is made between the baseline metallic fan blade and different laminate stackups in the composite blade. The results indicate that when carefully tailored, composite fan blades can potentially provide higher aeroelastic stability than the reference metallic blade. At nodal diameters where disk motion dominates the mode shape, the aeroelastic response is less dependent on the choice of stackup. Additionally, the study demonstrates that choosing the appropriate laminate stackup can avoid potentially dangerous resonant crossings throughout the flight envelope. Several prototype blades were manufactured and tested in a vibration test rig to assess manufacturability and validate the numerical models used in blade design.

The aerospace industry has seen a significant push to develop shorter and more aerodynamically “aggressive” intermediate compressor ducts (ICDs) due to their potential to reduce engine weight and, consequently, overall specific fuel consumption. However, many aggressive designs tend to either experience flow separation or operate close to their separation limit. This study investigates the use of circumferential splitter vanes as a passive flow control device to suppress flow separations and enhance the aerodynamic robustness of these ducts. The test case examined is a state-of-the-art, highly aggressive ICD developed by the Institute of Propulsion Technology at the German Aerospace Center (DLR). This ICD operates near its separation limit at its nominal operating point. Numerical simulations were employed to assess the aerodynamic performance of both the baseline and augmented (splittered) ICD designs. Analysis of the baseline design revealed a strong tendency for flow separation at the hub-strut corner, mid-strut region, and in the shroud region near the duct’s exit plane. The introduction of a double splitter vane in the augmented designs was found to be effective in reducing flow separation in the hub and shroud regions. Nevertheless, the potential benefits must be balanced against the resulting increase in overall total pressure losses, indicating a need for further optimization of the splitter vane design.

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.

The present report contains the summary of project activities and main achievements of the ARIAS project. It starts with a brief overview of the project goals and objectives, breaking these down on a WP level. A description of each subtask in the work package with its specific objectives is included, providing the necessary background for the interpretation of the results. The main results and achievements of each project's tasks are thereafter highlighted.

The present work focuses on experimental and numerical investigations of the unsteady response in a transonic compressor cascade oscillating at high-reduced frequencies. An attempt is made to highlight the main features that drive the unsteady response on the cascade blades. The operating point set in the rig is representative for modern transonic compressors, with a supersonic inflow and oblique shock waves forming at the leading edge of the blades. The cascade consisting of five blades, has been operated in the influence coefficient domain, where the central blade 0 is excited to oscillate in its natural modes of vibration. The investigated modes are mode 3, 4 and 7, with a reduced frequency ranging from 2.08 to 4.55. Steady aerodynamics in the rig has been assessed by measuring the static pressure distributions on the blade and Laser-Two-Focus (L2F) flow field measurements within the central two passages of interest. The aeroelastic testing data indicated that the unsteady blade surface pressure response is mainly driven by the position and relative motion of the shock waves, caused by the blade oscillation. Unsteady response distributions are fairly similar for modes 3 and 4, with some observed differences when compared to mode 7 response. The preliminary results from unsteady simulations qualitatively correspond quite well to the test data, and the main trends are overall well captured both in amplitude and phase. The parametric studies on sensitivity of the numerical solution are ongoing.

Paper explores the potential of using carbon-fibre reinforced composites for designing low-pressure compressor blades with improved aeroelastic performance. Comparison between the blades with different laminate  stackups is made with respect to the modal behaviour and aerodynamic damping. It is found that if carefully designed, the composite blades can provide higher aeroelastic stability than the reference metallic blade. At the same time the results reveal that a laminate stackup with stabilizing behaviour in one mode could have a destabilizing effect for the other mode. The dependency on ply angle and arrangement of plies in laminates is observed to be complex and furtheri nvestigations and experimental validation is therefore deemed necessary.

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.

The paper presents experiences from an international student project course in turbomachinery aeromechanics, which has been established within the framework of the H2020 research project ARIAS (“Advanced Research Into Aeromechanical Solutions”). The main activities, challenges, and lessons learned from the students’ feedback and instructors’ experience from 2020 to 2022 are presented from an educational point of view. A key goal of the aeromechanic project course (APC) is to strengthen the link between education and research based on multinational collaborative work. The course focuses on gradually enabling the students to perform a forced response analysis within one academic semester. This is achieved by providing the necessary fundamentals as well as training and application of interdisciplinary tools to investigate structural and aerodynamic effects. The course is based on online lectures, tool-specific tutorials, literature reviews, and guest lectures. The students work via web-based communication in international teams conducting various numerical analyses to determine distinct influences on forced response phenomena at the ARIAS TUDa transonic compressor. Metrics regarding the course structure, quality, project rating, and students’ feedback are obtained via an anonymous survey. The shown collaborative results and conclusions presented by the students display that the gap between theoretical knowledge and application is reduced.