An empirical riblet model for manufactured V-shaped and trapezoidal riblets which is suitable for turbomachinery application is presented. The implementation of the riblet effect employs a correlation-based correction for the damping of the specific dissipation rate ω in the vicinity of the wall which has been previously presented by other researchers. In the current paper, the correlations are extended into the drag-increasing regime and are extended to account for the effect of misalignment of the riblets relative to the flow and for the effect of adverse pressure gradients. In order to account for the latter in modern, massive parallel Reynolds-averaged Navier–Stokes (RANS) codes, a local Clauser parameter has been newly derived. The model is implemented in a three-dimensional (3D) turbomachinery design code and validated with flat plate measurement data and a NACA6510 compressor cascade. The predictions of the experimental values are in very good agreement with the experimental data, showing the capability of the model for designing riblet structured turbomachinery blading.

The state-of-the-art design of turbomachinery components is based on Reynolds-averaged Navier–Stokes (RANS) solutions. RANS solvers model the effects of turbulence and boundary layer transition and therefore allow for a rapid prediction of the aerodynamic behavior. The only drawback is that modeling errors are introduced to the solution. Researchers and computational fluid dynamics developers are working on reducing these errors by improved model calibrations which are based on experimental data. These experiments do not typically, however, offer detailed insight into three-dimensional flow fields and the evolution of model quantities in an actual machine. This can be achieved through a direct step-by-step comparison of model quantities between RANS and direct numerical simulation (DNS). In the present work, the experimentally obtained model correlations are recomputed based on DNS of the same turbine profile simulated by RANS. The actual local values are compared to the modeled RANS results, providing information about the source of model deficits. The focus is on the transition process on the blade suction side (SS) and on evaluating the development of turbulent flow structures in the blade's wake. It is shown that the source of disagreement between RANS and DNS can be traced back to three major deficiencies that should be the focus of further model improvements.

A novel theoretical model of the internal flow field in multistage axial compressors based on an eigenvalue approach is developed, in order to predict the onset of acoustic resonance in aircraft engines. Using an example high-speed four-stage compressor, it is shown that one of the resultant frequencies is in excellent agreement with the experimental data in terms of acoustic resonance. On the basis of the computed natural frequency of the whole compression system and the measured spanwise distribution of static pressure, the location of the acoustic excitation source can be found in the third stage. Unsteady flow simulations of the full annulus of this stage reveal two criteria for acoustic excitation at the rotor-blade tip, reversed flow near the suction surface and flow impingement on the pressure surface. Additionally, a fast Fourier transform of the unsteady pressure field at the upper rotor-blade span verifies the existence of the computed unstable frequency of the oscillating tip leakage flow. Using this novel theory, which combines a theoretical calculation of flow-instability frequency of the global system with the computational simulation of a single stage, the onset mechanism and location of the excitation source of acoustic resonance in multistage turbomachinery can be explained at acceptable computational cost.

The current design of low-pressure turbines (LPTs) with steady-blowing vortex generating jets (VGJs) uses steady computational fluid dynamics (CFD). The present work aims to support this design approach by proposing a new semiempirical transition model for injection-induced laminar-turbulent boundary layer transition. It is based on the detection of cross-flow vortices in the boundary layer which cause inflectional cross-flow velocity profiles. The model is implemented in the CFD code TRACE within the framework of the $γ$-$Reθ$ transition model and is a reformulated, recalibrated, and extended version of a previously presented model. It is extensively validated by means of VGJ as well as non-VGJ test cases capturing the local transition process in a physically reasonable way. Quantitative aerodynamic design parameters of several VGJ configurations including steady and periodic-unsteady inflow conditions are predicted in good accordance with experimental values. Furthermore, the quantitative prediction of end-wall flows of LPTs is improved by detecting typical secondary flow structures. For the first time, the newly derived model allows the quantitative design and optimization of LPTs with VGJs.

This paper examines the diffuser flow with consideration to turbine outflow conditions. The setup consists of a low-speed axial diffuser test rig, that represents a 1/10 scaled heavy-duty exhaust diffuser with an annular and a conical diffuser part. In part A of this paper it was shown through experimental investigation that the turbulent kinetic energy as well as the Reynolds shear stresses are the relevant physical parameters that correlate with diffuser pressure recovery. To complement the experimental investigations, unsteady scale-resolving CFD simulations are performed, applying the SST-SAS turbulence model. As a first step, the numerical approach is validated by means of the experimental data with regards to the diffuser’s integral parameters as well as the prediction of local flow characteristics. In a second step, the interaction of coherent vortices generated by the rotor and the diffuser’s boundary layer are analyzed by means of the validated SST-SAS results. These vortices are found to have a major impact on the boundary layer separation in the region immediately downstream of the rotor and at the diffuser inlet.