Three-Dimensional Pressure Controlled Vortex Design of a Turbine Stage

A three-dimensional (3D) Pressure Controlled Vortex Design (PCVD) method for turbine stage design is proposed and discussed in this paper. The concept is developed from conventional Controlled Vortex Design (CVD) via pressure control approach and CVD technology. By specifying the static pressure and axial velocity distributions, the spanwise pressure gradient incorporated with pressure gradient in streamwise and azimuthal directions is moderated. Not only can profile loss profit from pressure control approach, but also secondary flow can be managed. The reasons for CVD are derived from stream surface thickness and stream surface twist. Through modifying stream surface thickness and inducing large stream surface twist, the secondary flow migrations are controlled properly and orderly. The relations of pressure control approach and CVD technology complement one another and finally lead to a well-posed flow pattern in turbine stage. The first stage redesign of a well-designed low pressure turbine demonstrates this technique application. A significant reduction of secondary flow losses and a corresponding increase of stage efficiency have achieved.

Novel Performance Prediction of a Transonic 4.5 Stage Compressor

Computing capacities have grown exponentially in recent years and 3D-Navier-Stokes methods were developed widely. However it is still not feasible to design a multi-stage compressor directly in three dimensions. Instead, compressor design starts with 1D-design. In accordance with this approach, basic parameters such as the number of stages and stage pressure ratios are determined. In the following 2D-design, the geometry of the flow channel and the main parameters of the blade geometries can be determined. Afterwards in the 3D-design, unsteady and 3D-flow-effects are considered and the design optimized accordingly. Therefore, it is virtually impossible to correct conceptual faults in the 3D-design phase. Thus a robust and reliable 2D-Throughflow-solver including a performance prediction for modern airfoil geometries is necessary. So far there is no efficient methodology known which predicts the performance for all kinds of airfoil geometries, as it would be necessary in a 2D-Throughflow optimization process. In [1, 2] a novel methodology was presented, which is able to predict the performance for a large number of airfoil geometries accurately. This method is based on a large airfoil database which is used to train a surrogate model for airfoil performance prediction. The scope of this work is to validate and to document the progress of this new approach. In Schmitz et al. [1] it was validated on rotor 1 of the 4.5 stage transonic test compressor DLR-RIG250 of the Institute of Propulsion Technology. In this work all 4.5 stages were calculated at different speedlines and different vane positions. The results of the S2-solver are compared to experimental data and 3D-CFD calculations, obtained using the DLR in-house solver TRACE.

Effects of Non-Uniform Combustor Exit Flow on Turbine Aerodynamics

Very-low NOx combustion concepts require a high swirl number of the flow in the combustion chamber to allow for lean burn combustion. This article deals with the influence of the resulting combustor exit swirl on the turbine aerodynamics of the first stage. This investigation is based on numerical simulations. According to the literature research additional insight into combustor-turbine interaction is achieved by taking into account a fully two dimensional inlet boundary condition. Up to now published results on combustor-turbine interaction were mostly restricted to the inhomogeneous temperature distribution at the turbine inlet. The investigations are carried out on a real engine geometry — the E3E Core 3/2 — a research project of Rolls-Royce Deutschland on lean combustion. Calculations are conducted by means of the Rolls-Royce plc code Hydra. The swirled inlet boundary condition is further scaled to test rig conditions to check for the transferability between the test rig and the real engine geometry. The results show a significant impact of the inhomogeneous turbine inflow on the stage efficiency and the thermal load. The optimization potential due to the clocking position of the combustor swirl is analyzed. The impact on the secondary flow mechanisms is analyzed with a novel visualization technique. A frequency spectrum analysis is carried out to investigate the effects of the 2D inlet boundary condition on the rotor row.

Aerodynamic Effects of an Unshrouded Low Pressure Turbine on a Low Aspect Ratio Exit Guide Vane

This paper presents the effects of an unshrouded low pressure turbine (LPT) onto the following exit guide vane row (EGV). The measurement results were obtained in the subsonic test turbine facility at Graz University of Technology by means of a fast response pressure probe in planes downstream of the rotor as well as oil flow visualisation. The test rig was designed in cooperation with MTU Aero Engines and represents the last 1.5 stages of a commercial aero engine. Considerable efforts were put into the adjustment of all relevant model parameters to reproduce the full scale LPT situation. Different tip clearances were evaluated by means of CFD obtained using a commercial Navier-Stokes code and validated with experimental results. The goal is to evaluate the effect of the varying leakage flow on the flow in the low aspect ratio EGV. Special attention is given to the impact on the development of secondary flows as well as the flow structures downstream of the EGV. The effect of the leakage flow causes a change of the flow structure of the EGV, particularly losses. Considering the largest investigated tip-clearance, the losses increased by 71% when compared to a zero-leakage case.

Use of Transition Modeling to Enable the Computation of Losses for Variable-Speed Power Turbine

To investigate the penalties associated with using a variable speed power turbine (VSPT) in a rotorcraft capable of vertical takeoff and landing, various analysis tools are required. Such analysis tools must be able to model the flow accurately within the operating envelope of VSPT. For power turbines low Reynolds numbers and a wide range of the incidence angles, positive and negative, due to the variation in the shaft speed at relatively fixed corrected flows, characterize this envelope. The flow in the turbine passage is expected to be transitional and separated at high incidence. The turbulence model of Walters and Leylek was implemented in the NASA Glenn-HT code to enable a more accurate analysis of such flows. Two-dimensional heat transfer predictions of flat plate flow and two and three dimensional heat transfer predictions on a turbine blade were performed and reported herein. Heat transfer computations were performed because it is a good marker for transition. The final goal is to be able to compute the aerodynamic losses. Armed with the new transition model, total pressure losses for three-dimensional flow of an Energy Efficient Engine (E3) tip section cascade for a range of incidence angles were computed in anticipation of the experimental data. The results obtained form a loss bucket for the chosen blade.

On LES Methods Applied to Seal Geometries

Nine Large Eddy Simulation (LES) methods are used to simulate flow through two labyrinth seal geometries and are compared with a wide range of Reynolds-Averaged Navier-Stokes (RANS) solutions. These involve one-equation, two-equation and Reynolds Stress RANS models. Also applied are linear and nonlinear pure LES models, hybrid RANS-Numerical-LES (RANS-NLES) and Numerical-LES (NLES). RANS is found to have a maximum error and a scatter of 20%. A similar level of scatter is also found among the same turbulence model implemented in different codes. In a design context, this makes RANS unusable as a final solution. Results show that LES and RANS-NLES is capable of accurately predicting flow behaviour of two seals with a scatter of less than 5%. The complex flow physics gives rise to both laminar and turbulent zones making most LES models inappropriate. Nonetheless, this is found to have minimal tangible results impact. In accord with experimental observations, the ability of LES to find multiple solutions due to solution non-uniqueness is also observed.

Modeling of Discrete Tip Injection in a Two-Dimensional Streamline Curvature Method

Fluid injection at the tip of highly loaded compressor rotors is known to be effective in suppressing the onset of rotating stall and eventually compressor instability. However, using such stability enhancement methods in a multistage compressor might not only stabilize certain stages but has also an impact on radial and axial matching. In order to account for tip injection during the early stages of compressor design, this paper focuses on the development of a method to model the physical effects underlying tip injection within a streamline curvature method. With the help of system identification it could be shown that a rotor subject to the discrete jets of tip injection adapts to the varying flow conditions according to a first order model. This information was used to generate a time-dependent input for the steady equations used with a streamline curvature method and eventually to model the unsteady response of the rotor to tip injection. Comparing the results obtained with the enhanced streamline curvature model to measurement results, good agreement could be shown which raised confidence that the influence of tip injection on axial and radial matching was sufficiently captured.

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.

Experimental Studies of Leading Edge Contouring Influence on Secondary Losses in Transonic Turbines

An experimental study of the hub leading edge contouring using fillets is performed in an annular sector cascade to observe the influence of secondary flows and aerodynamic losses. The investigated vane is a three dimensional gas turbine guide vane (geometrically similar) with a mid-span aspect ratio of 0.46. The measurements are carried out on the leading edge fillet and baseline cases using pneumatic probes. Significant precautions have been taken to increase the accuracy of the measurements. The investigations are performed for a wide range of operating exit Mach numbers from 0.5 to 0.9 at a design inlet flow angle of 90°. Data presented include the loading, fields of total pressures, exit flow angles, radial flow angles, as well as profile and secondary losses. The vane has a small profile loss of approximately 2.5% and secondary loss of about 1.1%. Contour plots of vorticity distributions and velocity vectors indicate there is a small influence of the vortex-structure in endwall regions when the leading edge fillet is used. Compared to the baseline case the loss for the filleted case is lower up to 13% of span and higher from 13% to 20% of the span for a reference condition with Mach no. of 0.9. For the filleted case, there is a small increase of turning up to 15% of the span and then a small decrease up to 35% of the span. Hence, there are no significant influences on the losses and turning for the filleted case. Results lead to the conclusion that one cannot expect a noticeable effect of leading edge contouring on the aerodynamic efficiency for the investigated 1st stage vane of a modern gas turbine.

The Effect of Four Part Gap Geometry Configurations for Variable Stator Vanes in a Compressor Cascade

The article describes numerical investigations on the influence of four different endwall clearance topologies for variable stator vanes to secondary flow field development and the performance of high pressure compressors. The aim of this work is to quantify the characteristics of different clearance configurations depending on the penny-axis position and the penny diameter for a typical operating range. All clearance configurations were implemented to a linear cascade of modern stator profiles. The analysis was introduced using a relative clearance size of 1.3% chord at three stagger angles and two characteristic Reynolds numbers to model the operating range on aircraft engines. 3D numerical calculations were carried out to gain information about the flow field inside the cascade. They were compared with measurements of a 5-hole-probe as well as pressure tappings on the airfoil and the endwall. The CFD shows the clearance characteristics in good agreement with the measurements for the lower and the nominal stagger angle. Small gaps in the rear part of the vane have a beneficial effect on the flow field. In contrast, a clearance in the higher loaded front part of the vane always resulted in increased losses. Otherwise, the significant enhanced performance of a rear part gap, which was measured at the higher stagger angle, was not reflected by the CFD. The reduced mixing losses and the higher averaged flow turning even compared to a configuration without a clearance are not verified with the calculations. Large flow separations at the high stagger angle result in a two to four times higher underturning of the CFD in comparison to the experiments. The clearance effects to the characteristic radial loss distribution up to 40 % bladeheight also deviate from the measurements due to heavy mixing of clearance and reversed separated flow.