Dragan Kožulović

Flow Characteristics of Axial Compressor Tandem Cascades at Large Off-Design Incidence Angles

In a number of recent and former publications, compressor tandem blade configurations show potential to outperform single blade configurations in terms of turning, loss and operating range at high aerodynamic loading levels. However, very little insight is given into the mechanisms of flow breakdown when comparing tandem blades to single blades at large off-design incidence angles.Single blade cascades tend to fail as a result of either pressure side flow separation for high negative incidence or suction side flow separation for high positive incidence, the latter being mostly accompanied by significant increase of underturning. Tandem blade cascades are expected to show a different behavior due to the aerodynamic interaction in the blade overlapping region.Two different tandem blade configurations are examined together with their respective reference single blades, one being a recently designed and optimized tandem blade for high subsonic inlet Mach numbers, which has also been investigated in cascade wind tunnel testing. The other one is a more generic tandem blade based on NACA65 family, designed for medium inlet Mach numbers using current state-of-the-art understanding of tandem design.The mechanisms of flow breakdown are examined using quasi two-dimensional RANS simulations which are validated with test data for one of the aforementioned tandem configurations. A detailed analysis of the flow structure at heavy off-design conditions gives insight into the characteristics of tandem flow breakdown. In particular, the ability of the tandem configuration to extend the operating range to larger positive incidence is described. The shortcomings of the tandem cascade at large negative incidence are also commented. These and further conclusions can be used to improve tandem blade performance at moderate off-design conditions.Copyright

Analysis of Laminar-Turbulent Transition of a Low-Loss Generic Low Pressure Turbine Distribution

The efficiency of modern Turbofan engines can be significantly increased by using a gearbox between compressor and turbine of the low pressure section. Rotational speed of the low pressure turbine (LPT) in a Geared Turbofan is much higher than in normal LPT’s which lead to necessary adjustments in blade design.This work has investigated the transition behavior of a modified profile geometry for low-loss at engine cruise conditions. Typical LPT conditions have thus been chosen as baseline for the experimental work. A pressure distribution has been created on a flat plate by means of contoured walls in a low speed wind tunnel. The paper will analyze the experimental results and show additionally the numerical predictions of the test case.The experimental part of this paper describe how the blade was Mach number scaled to obtain the geometry of the wind tunnel wall contour. The pressure distribution for the incompressible test case show a very good agreement to the compressible case. Boundary layer (BL) measurements with hot-wire-anemometry have been performed at high spatial resolution under a freestream turbulence of almost 8%. Different Reynolds numbers have been investigated and will be compared with special attention being paid to the transition on the suction side by contour plots (turbulence levels, turbulent intermittency) and integral BL parameters. It was found that the transition on the suction side is not completed for small Reynolds numbers but takes place at higher velocities.In the numerical part studies by means of steady RANS simulations with k-ω – SST turbulence model and γ-Reθ transition model have been conducted. The aim is to validate the RANS solver for the low-loss LPT application. Hence, comparison is made to the measured data and the transitional behavior of the BL. Furthermore, additional parameter variations have been conducted (turbulence intensity and Reynolds number).The numerical investigations show partially a good comparison for the BL development indicating the different transition modi with increasing Reynolds number and turbulence intensity.Copyright

Aerodynamic Optimization of Turboprop Turbine Blades Using a Response Surface Methodology Based Algorithm

Turbomachinery blade design improvement and optimization by CFD is a time-consuming engineering challenge. Such an optimization process, which requires advanced numerical simulations, uses a large amount of computational resources to provide the required solutions. This paper presents a turbine blade optimization process which uses an algorithm based on response surface methodology (RSM) to increase the simulation speed. The main idea of RSM is to start with a lower number of sample points to generate an analytical model that describes the relationship between the pre-defined numbers of design parameters. In this study, the Kriging approximation is used to generate the surface model. The global minimum on the surface is searched by applying a gradient method. The increase in the convergence speed is achieved by using an adaptive scheme, which creates additional points around the previous minimum while reducing the solution space at each iteration step, until convergence is achieved. Each iteration step is composed of several CFD simulation runs where each point represents different designed geometries inside the n-dimensional parameter space. The process combines a Bezier-spline based airfoil-generator with a parametric meshing tool —G3DMESH— and a CFD solver —TRACE—, both developed and provided by DLR, into a MATLAB script function. A particular characteristic of this optimization method is its lower evaluation number requirement to reach convergence, as well as its capability to run multiple simultaneous RANS. The optimizer process was initially tested by using basic functions to analyze its solution behavior and its performance in comparison to a genetic algorithm (GA) type optimizer. It is observed from this comparison that RSM optimization reaches the convergence faster and more stable than the GA method applied on the test case. Preliminary optimization results show an improvement in function evaluation requirements by up to 50%, which depends on the complexity of the respective surface model of the test case. As an application, a 4-stage low pressure turbine for a turboprop engine is designed by multi-streamline analysis. 2D mid-span cross-sections from both rotor and stator are produced by the Bezier-spline based airfoil-generator. The basis tool requires input parameters as leading and trailing edge blade angles and maximum thickness position. The blade generator is further improved by the additional ability to work with high values of deviation angles between the leading and trailing edges, up to 90°. 6 control points are used to define the two curves, for pressure and suction sides, which encompass the cross-section geometry. Optimization process runs to improve these airfoil parameters. The 2D airfoils of the first stage are optimized by an objective function based on total pressure loss coefficient at the engine on-design point. The same geometry is also optimized using the GA method as a comparison case.Copyright

Investigations on Aerodynamic Loading Limits of Subsonic Compressor Tandem Cascades: End Wall Flow

An optimized subsonic compressor tandem cascade was investigated experimentally and numerically. Since the design aims at applications under incompressible flow conditions, a low inlet Mach number of 0.175 was used. The experiments were carried out at the low speed cascade wind tunnel at the Technische Universitat Braunschweig. For the numerical simulations, the CFD-solver TRACE of DLR Cologne was used, together with a curvature corrected k-ω turbulence model and the γ-Reθ transition model. The aerodynamic loading was varied by incidence variation. Results are presented and discussed for different inlet angles: spanwise loss coefficient, turning, pressure rise coefficient and AVR together with contour plots of the wake plane, flow visualization and oil flow pictures. Experimental and numerical results were compared and found to be in good agreement. The secondary flow topology of the front blade is considerably altered by the aerodynamic loading variation, whereas the topology of the rear blade surface is almost unchanged. The effect of the nozzle between the tandem blades, was observable up to the end wall for all investigated incidences. In addition, a comparison is made to published results of previous experimental and numerical investigations of a transonic tandem compressor cascade [1] and its reference single compressor cascade [2]. The comparison of the tandem cascades revealed that the general structures of the secondary flow seem to be similar for similar loading.Copyright

Improved Turbulence and Transition Prediction for Turbomachinery Flows

A correlation based approach for estimation of the turbulence length scale lT at the inflow boundary is proposed and presented. This estimation yields reasonable turbulence decay, supporting the transition model in accurately predicting the laminar-turbulent transition location and development. As an additional element of the approach, the sensitivity of the turbulence model to free stream values is suppressed by limiting the eddy viscosity in non-viscous regions. Therefore a criterion to detect those regions, based only on local variables, is derived. The method is implemented in DLR’s turbomachinery flow solver TRACE in the framework of the k–ω turbulence model by Wilcox 1988 [1] and the γ–Reθ transition model by Langtry and Menter [2] in combination with a cross flow (CF) induced transition criterion after Muller [3]. The improved model is tested to the T161 turbine test case [4], [5] and validated at the Durham turbine Cascade [6] and an outlet guide vane for low pressure turbine configurations [7].Copyright

Investigations on Aerodynamic Loading Limits of Subsonic Compressor Tandem Cascades: Midspan Flow

An optimized subsonic compressor tandem cascade was investigated experimentally and numerically. Since the design aims at incompressible applications, a low inlet Mach number of 0.175 was used. The experiments were carried out at the low speed cascade wind tunnel at the Technische Universitat Braunschweig. For the numerical simulations, the CFD-solver TRACE of DLR Cologne was used, together with a curvature corrected k-ω turbulence model and the γ-Reθ transition model. Besides the incidence variation, the aerodynamic loading has also been varied by contracting endwalls. Results are presented and discussed for different inlet angles and endwall contractions: pressure distribution, loss coefficient, turning, pressure rise, AVDR and Mach number. The comparison of experimental and numerical results is always adequate for a large range of incidence. In addition, a comparison is made to an existing high subsonic tandem cascade and conventional cascades. For the latter the Lieblein diffusion factor has been employed as a measure of aerodynamic loading to complete the Lieblein Chart of McGlumphy [1].© 2013 ASME

Performance and Boundary Layer Development of a High Turning Compressor Cascade at Sub- and Supercritical Flow Conditions

Based on current numerical investigations, the present paper reports on new Q2D midspan-calculations and results for the well known high turning (Δβ = 50°) supercritical (Ma1 = 0.85) compressor cascade V2. A Q2D treatment of the problem was chosen in order to avoid the difficult modelling of the porous endwalls in a corresponding 3D approach. All simulations were done with the RANS solver TRACE of the DLR Cologne in combination with modified versions of the Wilcox turbulence model and Langtry/Menter transition model. Existing experimental Q2D midspan-results for the V2 compressor cascade were used to demonstrate the improved ability of the numerical code to determine performance characteristics, blade pressure and Mach number distributions as well as boundary layer parameter and velocity distributions. The loss characteristics show minimum loss regions when plotted against inlet angle or axial velocity density ratio. Within these regions, increasing with decreasing Mach number, the experimental results were adequately predicted. Outside these regions it turned out difficult to reproduce the experimental results due to increasing boundary layer separation. Furthermore, the prediction quality was very good for subsonic conditions (Ma1 = 0.60) and still reasonable for supercritical conditions (Ma1 = 0.85), where shock/boundary layer interaction made the prediction more difficult.Copyright