The performance of the compressor blade is considerably influenced by secondary flow effects, like the cross flow on the end wall as well as corner flow separation between the wall and the blade. The present work is focused on the studying the effects of Vortex Generator (VG) on NASA Rotor 37 test case using Computational Fluid Dynamics (CFD). VG helps in controlling the inception of the stall by generating vortices and energizes the low momentum boundary layer flow which enhances the rotor performance. Three design configuration namely, Counter-rotating, Co-rotating and Plow configuration VG are selected based on the improved aerodynamic performance discussed in reference [1]. These VG are located at 90% span and 42% chord on suction side surface of the blade. Among the three configurations, the first configuration has greater impact on the end wall cross flow and flow deflection which resulted in enhanced numerical stall margin of 5.4% from baseline. The reasons for this numerical stall margin improvement are discussed in detail.

The effect of blade row interaction and hub leakage flow on the performance of moderately loaded NASA transonic hybrid compressor stage (Rotor 35 / Stator 37) is investigated through three-dimensional steady state and time-accurate, Navier Stokes calculations of the stage using the ANSYS CFX code at peak efficiency and near stall operating conditions. Understanding unsteady flow phenomena in compressor stages requires the use of time-accurate CFD simulations. Due to the inherent differences in blade counts between adjacent blade rows, the flow conditions at any given instant in adjacent blade rows differ. Depending on the blade counts, it may be necessary to model the entire annulus of the stage; however, this requires considerable computational time and memory resources. Several methods for modeling the transient flow in turbo machinery stages which require a minimal number of blade passages per row, and therefore reduced computational demands, have been presented in the literature. Recently, some of these methods have become available in commercial CFD solvers. The paper describes the steady and the unsteady CFD approaches used for investigating the ability to predict the measured performance of the NASA transonic axial stage design known as the hybrid stage, which consists of the axial Rotor35 and the axial stator 37. The steady approach employs the mixing-plane while the unsteady approaches are URANS with one based on full annulus simulation for the stage and the second enables simulations for the stage using reduced computational model, with a single passage from each blade row based on the time-tilting or the time-transformation technique. The above methods are evaluated and compared in terms of computational efficiency and comparison is made to steady stage simulations. Comparisons to overall performance data and two-dimensional Laser Doppler Velocimeter measurements of the velocity field are used to assess the predictive capabilities of the methods. Computed flow features are examined, and compared with reported measurements. This paper presents validation and calibration of methods used for determining blade row interactions and the respective predictive capabilities against the full annulus and the experimental test data.

This paper summarizes the results of a validation and calibration study for two modern Computational Fluid Dynamics programs that are capable of modeling multistage axial compressors in a multi-blade row environment. The validation test case is a modern 4-stage high pressure ratio axial compressor designed and tested by Honeywell Aerospace in the late 90’s. The two CFD programs employ two different techniques for simulating the steady three-dimensional viscous flow field in a multistage/multiblade row turbo-machine. The first code, APNASA, was developed by NASA Glenn Research Center “GRC” and applies the approach by Adamczyk [1] for solving the average-passage equations which is a time and passage-averaged version of the Reynolds Averaged Navier Stokes (RANS) equations. The second CFD code is commercially marketed by ANSYS-CFX and applies a much simpler approach, known as the mixing-plane model, for combining the relative and the stationary frames of reference in a single steady 3D viscous simulation. Results from the two CFD programs are compared against the tested compressor’s overall performance data and against measured flow profiles at the leading edge of the fourth stator. The paper also presents a turbulence modeling sensitivity study aimed at documenting the sensitivity of the prediction of the flow field of such compressors to use of different turbulence closures such as the standard K-ε model, the Wilcox K-ω model and the Shear-Stress-Transport K-ω/SST turbulence model. The paper also presents results that demonstrate the CFD prediction sensitivity to modeling the compressor’s hub leakages from the inner-banded stator cavities. Comparison to the test data, using the K-ε turbulence closure, show that APNASA provides better accuracy in predicting the absolute levels of the performance characteristics. The presented results also show that better predictions by CFX can be obtained using the K-ω and the SST turbulence models. Modeling of the hub leakage flow was found to have significant and more than expected impact on the compressor predicted overall performance. The authors recommend further validation and evaluation for the modeling of the hub leakage flow to ensure realistic predictions for turbo-machinery performance.

The need of increased stall margin is very high for aero gas turbine engines, as they operate under varied operating conditions. A number of different options are being used to increase the stall margin of gas turbine engines. Circumferential casing groove, in the compressor section of a gas turbine engine, is one of such methods. Incorporation of the grooves on the shroud increases the stall margin of the compressor, but this generally gives rise to loss of performance, such as efficiency and pressure ratio. By employing 3D blading techniques for rotor blades as well as stator vanes, performance of a compressor can be increased. 3D blading helps in reducing secondary flow losses and hence increased performance. Sweep and lean are examples of 3D blading, which is very common in any modern gas turbine compressor. A number of literatures are available in public domain, giving detailed understanding of effect of circumferential casing grooves and 3D blade features, but the interaction effect of sweep and casing grooves are not well published in public domain literature. In this work, an effort is made to understand, numerically, the interaction effect of sweep with circumferential grooves, using Computational Fluid Dynamics (CFD). Any numerical tool needs thorough validation before the results of numerical analyses can be used for analyzing the underlying physics. NASA Rotor37 is used to validate current CFD methodology. Mesh sensitivity is carried out to get mesh independence solution. Different turbulence models are used to get the best turbulence model for the problem in hand. 1D averaged performance data as well as hub to shroud variation of various flow parameters are compared to have full confidence on the CFD methodology. A baseline axial compressor rotor, without sweep and lean is generated, as the first step of this study. This rotor is created by using hub and tip profiles of NASA Rotor37. The profiles are stacked along a radial line through their center of gravities, which has resulted in rotor geometry without any sweep and lean. Modifications are done to the tip profile of the baseline rotor, in terms of stagger angle, to get comparable performance w.r.t. NASA Rotor37. Casing of the NASA Roto37 is used as the redesigned compressor casing. Circumferential casing grooves, with five grooves between leading edge to trailing edge, are created as per industry standards. Meshing and modeling are done according to the best practices developed while validating CFD methodology. It is to be noted that the casing grooves and the main flow domain are meshed with one to one mesh connectivity, in order to avoid any numerical losses due to interface interpolations. This is considered very critical in this work, as the vortices from the tip is expected to have a strong interaction with grooves. This interaction is expected to create high gradients of flow variables in this region. Valuable flow information might be lost, if flow variables are interpolated in this region. Baseline rotor is analyzed with and without casing grooves from choke to stall at 100% corrected speed. As expected, introduction of casing grooves has resulted in increased stall margin. A number of rotor geometries are created with different amount of sweeps. In the current study, blades are swept in the direction of chord, in order to avoid introduction of any sweep induced lean. The span location, where sweep starts, is also changed to understand the localized and global effect of this blade design features. Results obtained from numerical simulations of these geometries are presented in this paper. The performance and flow features are compared with respect to baseline rotor, with and without circumferential grooves, in an attempt to understand the underlying flow physics.

This paper presents a thorough assessment for two of the contemporary CFD programs available for modeling and predicting nonfilm-cooled surface heat transfer distributions on turbine airfoil surfaces. The CFD programs are capable of predicting laminar-turbulent transition and have been evaluated and validated against five test cases with experimental data. The suite of test cases considered for this study consists of two flat plat cases at zero and non-zero pressure gradient and three linear-turbine-cascade test cases that are representative of modern high pressure turbine designs. The flat plate test cases are the ERCOFTAC T3A and T3C2, while the linear turbine cascade cases are the MARKII, the Virginia Polytechnic Institute (VPI), and the Von Karman Institute (VKI) turbine cascades. The numerical tools assessed in this study are 3D viscous Reynolds Averaged-Navier-Stokes (RANS) equations programs that employ a variety of one-equation and two-equation models for turbulence closure. The assessment study focuses on the one-equation Spalart and Allmaras and the two-equation shear stress transport K-ω turbulence models with the ability of modeling and predicting laminar-turbulent transition. The RANS 3D viscous codes are Numeca’s Fine Turbo and ANSYS-CFX’ CFX5. Numerical results for skin friction, surface temperature distribution and heat transfer coefficient from the CFD programs are compared to measured experimental data. Sensitivity of the predictions to free stream turbulence and to inlet turbulence boundary conditions is also presented. The results of the study clearly illustrate the superiority of using the laminar-turbulent transition prediction in improving the accuracy of predicting the heat transfer coefficient on the surfaces of high pressure turbine airfoils.