A Critical Evaluation of the Classical k−εk-\varepsilon Model and Wall-Laws for Unsteady Flows over Bluff Bodies

Projet M3NThe classical $k-\varepsilon$ model and a new implementation of wall-laws have been evaluated for unsteady turbulent flows past square and circular cylinders. We show that good numerics can lead to a serious improvement of the results. In the case of the flow past a square cylinder (Lyn's test case), results agree surprisingly well with the experimental data. In the analysis of the vortex shedding flow past a circular cylinder also the agreement is quite good for the sub-critical regime but not entirely satisfactory after the critical Reynolds number (i.e. the drag crisis phenomeno

NSIKE - An Incompressible Navier-Stokes Solver for Unstructured Meshes

Projet M3NA new solver for 3D incompressible Navier-Stokes equations is developed. The standard k-epsilon turbulence model is implemented for turbulent flow simulation. The incompressible Navier-Stokes solver is based on Chorin's projection method. with finite element discretization. The computation of intermediate velocity and turbulence variables is done explicitely and the stabilization of convection terms is based on PSI (positive streamwise invariant) residual distribution scheme. The Poisson problem for pressure is solved using conjugate gradient technique. In the case of turbulent flow simulations the wall-laws are used for solid boundaries. Test cases presented are 2D flat plate, cavity, backward facing step, circular cylinder and 3D cavity, backward facing step, both in laminar and turbulent regime

GT2006-91234 ON NEAR-WALL DYNAMIC COUPLING OF LES WITH RANS TURBULENCE MODELS

ABSTRACT In this paper, the RANS/LES coupling formulation proposed i

Aerodynamic Performance of a Compact, High Work-Factor Centrifugal Compressor at the Stage and Subcomponent Level

The measured aerodynamic performance of a compact, high work-factor, single-stage centrifugal compressor, comprising an impeller, diffuser, 90-bend, and exit guide vane is reported. Performance levels are based on steady-state total-pressure and total-temperature rake and angularity-probe data acquired at key machine rating planes during recent testing at NASA Glenn Research Center. Aerodynamic performance at the stage level is reported for operation between 70 to 105% of design corrected speed, with subcomponent (impeller, diffuser, and exit-guide-vane) flow field measurements presented and discussed at the 100% design-speed condition. Individual component losses from measurements are compared with pre-test CFD predictions on a limited basis

Evaluation of alternatives for two-dimensional linear cascade facilities

This paper presents two low-cost alternatives for turbine blade surface heat transfer and fluid dynamics measurements. These models embody careful compromises between typical academic and full-scale turbomachinery experiments and represent a comprehensive strategy to develop experiments that can directly test shortcomings in current turbomachinery simulation tools. A full contextual history of the wide range of approaches to simulate turbine flow conditions is presented, along with a discussion of their deficiencies. Both models are simplifications of a linear cascade: the current standard for simulating two-dimensional turbine blade geometries. A single passage model is presented as a curved duct consisting of two half-blade geometries, carefully designed inlet and exit walls and inlet suction. This facility was determined to be best suited for heat transfer measurements where minimal surface conduction losses are necessary to allow accurate numerical model replication. A double passage model is defined as a single blade with two precisely designed outer walls, which is most appropriate for flow measurements. The design procedures necessary to achieve a desired flow condition are discussed

Aerodynamic Performance of a Compact, High Work-Factor Centrifugal Compressor at the Stage and Subcomponent Level

The measured aerodynamic performance of a compact, high work factor, single-stage centrifugal compressor, comprising an impeller, diffuser, 90-bend, and exit guide vane (EGV), is reported. Performance levels are based on steady-state total-pressure and total-temperature rake and angularity-probe data acquired at key machine rating planes during recent testing at NASA Glenn Research Center. Aerodynamic performance at the stage level are reported for operation between 70 to 105 of design corrected speed, with subcomponent (impeller, diffuser, and exitguide-vane) detailed flow field measurements presented and discussed at the 100 design-speed condition. Individual component losses from measurements are compared with pre-test predictions on a limited basis

High Efficiency Centrifugal Compressor for Rotorcraft Applications

The report "High Efficiency Centrifugal Compressor for Rotorcraft Applications" documents the work conducted at UTRC under the NRA Contract NNC08CB03C, with cost share 2/3 NASA, and 1/3 UTRC, that has been extended to 4.5 years. The purpose of this effort was to identify key technical barriers to advancing the state-of-the-art of small centrifugal compressor stages; to delineate the measurements required to provide insight into the flow physics of the technical barriers; to design, fabricate, install, and test a state-of-the-art research compressor that is representative of the rear stage of an axial-centrifugal aero-engine; and to acquire detailed aerodynamic performance and research quality data to clarify flow physics and to establish detailed data sets for future application. The design activity centered on meeting the goal set outlined in the NASA solicitation-the design target was to increase efficiency at higher work factor, while also reducing the maximum diameter of the stage. To fit within the existing Small Engine Components Test Facility at NASA Glenn Research Center (GRC) and to facilitate component re-use, certain key design parameters were fixed by UTRC, including impeller tip diameter, impeller rotational speed, and impeller inlet hub and shroud radii. This report describes the design effort of the High Efficiency Centrifugal Compressor stage (HECC) and delineation of measurements, fabrication of the compressor, and the initial tests that were performed. A new High-Efficiency Centrifugal Compressor stage with a very challenging reduction in radius ratio was successfully designed, fabricated and installed at GRC. The testing was successful, with no mechanical problems and the running clearances were achieved without impeller rubs. Overall, measured pressure ratio of 4.68, work factor of 0.81, and at design exit corrected flow rate of 3 lbm/s met the target requirements. Polytropic efficiency of 85.5 percent and stall margin of 7.5 percent were measured at design flow rate and speed. The measured efficiency and stall margin were lower than pre-test CFD predictions by 2.4 percentage points (pt) and 4.5 pt, respectively. Initial impressions from the experimental data indicated that the loss in the efficiency and stall margin can be attributed to a design shortfall in the impeller. However, detailed investigation of experimental data and post-test CFD simulations of higher fidelity than pre-test CFD, and in particular the unsteady CFD simulations and the assessment with a wider range of turbulence models, have indicated that the loss in efficiency is most likely due to the impact of unfavorable unsteady impeller/diffuser interactions induced by diffuser vanes, an impeller/diffuser corrected flow-rate mismatch (and associated incidence levels), and, potentially, flow separation in the radial-to-axial bend. An experimental program with a vaneless diffuser is recommended to evaluate this observation. A subsequent redesign of the diffuser (and the radial-to-axial bend) is also recommended. The diffuser needs to be redesigned to eliminate the mismatching of the impeller and the diffuser, targeting a slightly higher flow capacity. Furthermore, diffuser vanes need to be adjusted to align the incidence angles, to optimize the splitter vane location (both radially and circumferentially), and to minimize the unsteady interactions with the impeller. The radial-to-axial bend needs to be redesigned to eliminate, or at least minimize, the flow separation at the inner wall, and its impact on the flow in the diffuser upstream. Lessons were also learned in terms of CFD methodology and the importance of unsteady CFD simulations for centrifugal compressors was highlighted. Inconsistencies in the implementation of a widely used two-equation turbulence model were identified and corrections are recommended. It was also observed that unsteady simulations for centrifugal compressors require significantly longer integration times than what is current practice in industry

Proceedings of Turbo Expo 2007 ASME Turbo Expo 2007: Power for Land, Sea and Air ON COUPLING OF RANS AND LES FOR INTEGRATED COMPUTATIONS OF JET ENGINES

ABSTRACT Large scale integrated computations of jet engines can be performed by using the unsteady RANS framework to compute the flow in turbomachinery components while using the LES framework to compute the flow in the combustor. This requires a proper coupling of the flow variables at the interfaces between the RANS and LES solvers. In this paper, a novel approach to turbulence coupling is proposed. It is based on the observation that in full operating conditions the mean flow at the interfaces is highly non-uniform and local turbulence production dominates convection effects in regions of large velocity gradients. This observation has lead to the concept of using auxilliary ducts to compute turbulence based on the mean velocity at the interface. In the case of the RANS/LES interface, turbulent fluctuations are reconstructed from an LES computation in an auxiliary three-dimensional duct using a recycling technique. For the LES/RANS interface, the turbulence variables for the RANS model are computed from an auxilliary solution of the RANS turbulence model in a quasi-2D duct. We have demonstrated the feasibility of this approach for the integrated flow simulation of a 20 o sector of an entire jet engine