Gas Temperature Profile Attenuation Through a Multistage Axial-Flow Turbine

Nowadays 2D through-flow models are widespread for designing and analysis of a turbine. Unlike 2D calculations, the measurements show that a non-uniform inlet gas temperature profile is strongly attenuated to the outlet of a turbine. This attenuation can be taken into account in through-flow models only using some corrective coefficients. The objective of this work was to find such an empirical coefficient. The results of full-scale tests of several power GTUs and aviation GTEs were employed to obtain values of the temperature profile attenuation coefficient in the through-flow model for various span locations of airfoil rows. During the tests detailed radial-circumferential distributions of the gas temperature upstream of each of rows and downstream of the turbines were measured (in absolute motion for stator and in relative motion for rotor). The values of the attenuation coefficient for airfoil rows of the three front stages were obtained by means comparison of experimental and computed results. The experience shows that the attenuation coefficient is easily incorporated into the 2D gas-dynamic codes. This incorporation allows spanwise distributions of flow parameters to be predicted and the airfoil geometry and cooling mass flow to be chosen more correctly.

Multistage Compressor and Turbine Modeling for the Prediction of the Maximum Turbine Speed Resulting From Shaft Breakage

A mathematical model for the prediction of the maximum speed of a high pressure turbine following a shaft failure event was developed. The model predicts the high pressure compressor and ducting system pre- and post-stall behavior like rotating stall and surge after the shaft breakage. The corresponding time-dependent high pressure turbine inlet conditions are used to calculate the turbine maximum speed, taking into account friction and blade&vane tip clearance variations as a result of the rearward movement of the turbine and destruction of the turbine blading. The compressor and ducting system is modeled by a 1-dimensional, stage-by-stage approach. The approach uses a finite-difference numerical technique to solve the nonlinear system of equations for continuity, momentum and energy including source terms for the compressible flow through inlet ducting, compressor and combustor. The compressor blade forces and shaft work are provided by a set of quasi steady state stage characteristics being valid for pre-stall and post-stall operations. The maximum turbine speed is calculated from a thermodynamic turbine stand-alone model, derived from a performance synthesis program. Friction and blade&vane tip clearance variations are determined iteratively from graphical data depending on the axial rearward movement of the turbine. The compressor and ducting system model was validated in pre-stall and post-stall operation mode with measured high pressure compressor data of a modern 2-shaft engine. The turbine model was validated with measured intermediate pressure shaft failure data of a 3-shaft engine. The shaft failure model was applied on a modern 2-shaft engine. The model was used to carry out a sensitivity study to demonstrate the impact of control system reactions on the resulting maximum high pressure turbine speed following a shaft failure event.

Effects of Endwall Profiling on Axial Flow Compressor Stage

In recent years endwall profiling has been well validated as a major new engineering design tool for the reduction of secondary loss in turbines. However, its application on compressors have been rarely performed and reported. This paper documents the findings of the analysis for diminishing compressor stator corner separation using endwall profiling; In the study, novel profiled endwalls were designed and numerically studied on a subsonic axial-flow compressor stage. The compressor stator endwalls were profiled on both axial and azimuthal directions. The results showed, the stator corner separation was significantly suppressed under all the operating conditions by implementing this profiled endwall. Significant improvements on stage pressure ratios and stage efficiency were observed. Detailed flow field changes, as well as endwall profiling methods are provided in the paper, so that the results of this research can be referenced to other compressor designs.

The Application of Non-Axisymmetric Endwall Contouring in a Single Stage, Rotating Turbine

As turbine manufacturers strive to develop machines that are more efficient, one area of focus has been the control of secondary flows. To a large extent these methods have been developed through the use of computational fluid dynamics and detailed measurements in linear and annular cascades and proven in full scale engine tests. This study utilises 5-hole probe measurements in a low speed, model turbine in conjunction with computational fluid dynamics to gain a more detailed understanding of the influence of a generic endwall design on the structure of secondary flows within the rotor. This work is aimed at understanding the influence of such endwalls on the structure of secondary flows in the presence of inlet skew, unsteadiness and rotational forces. Results indicate a 0.4% improvement in rotor efficiency as a result of the application of the generic non-axisymmetric endwall contouring. CFD results indicate a clear weakening of the cross passage pressure gradient, but there are also indications that custom endwalls could further improve the gains. Evidence of the influence of endwall contouring on tip clearance flows is also presented.

A New Concept of a Two-Dimensional Supersonic Relative Inlet Mach Number Compressor Cascade

In this study, a numerical shape optimization method based on evolutionary algorithms coupled with a verified CFD solver has been applied to the ambitious target of a shock free 2-D supersonic inlet Mach number compressor cascade. The study is based on the DLR-PAV-1.5 supersonic compressor cascade designed by the pre-compression blading concept. The DLR cascade airfoil has been optimized using a verified CFD code. A superior performance of the optimized supersonic cascade with about 24% reduction of the total pressure loss coefficient compared to the original cascade has been realized. The flow mechanisms observable around the blade with improved performance and the resulting design concept are discussed in this paper.

Influence of Labyrinth Seal Leakage on Centrifugal Compressor Performance

The focus of this paper lies on the leakage flow across the shroud of a centrifugal compressor impeller. It is common practice to use shrouded impellers in multi stage compressors featuring a single shaft. The rotating impeller then has to be sealed against the higher pressure in the downstream diffuser by means of labyrinths. The relative amount of leakage is higher for stages designed for low flow, meaning that the associated losses gain in relevance. In addition to this loss source, the injection of the leakage flow has a serious influence on the main flow in a region where it is prone to separation, i.e. at the suction side of the impeller blades close to the shroud, where the highest relative velocities are found. The present paper discusses the numerical results of several geometrical arrangements where the leakage flow was mixed with the main flow in different ways. The distance between the location of injection and the leading edge of the impeller as well as the orientation of the injected flow showed a distinct influence on the performance of the entire stage, mainly on stability.

Validating Numerical CFD Simulations With Experimental Data for Turbulence Phenomena in Axial Flow Gas Turbine Diffusers

Scale models give engineers an excellent understanding of the aerodynamic behavior behind their design; nevertheless, scale models are time consuming and expensive. Therefore computer simulations such as Computational Fluid Dynamics (CFD) are an excellent alternative to scale models. One must ask the question, how close are the CFD results to the actual fluid behavior of the scale model? In order to answer this question the engineering team investigated the performance of a large industrial Gas Turbine (GT) exhaust diffuser scale model with performance predicted by commercially available CFD software. The experimental results were obtained from a 1:12 scale model of a GT exhaust diffuser with a fixed row of blades to simulate the swirl generated by the last row of turbine blades five blade configurations. This work is to validate the effect of the turbulent inlet conditions on an axial diffuser, both on the experimental front and on the numerical analysis approach. The object of this work is to bring forward a better understanding of velocity and static pressure profiles along the gas turbine diffusers and to provide an accurate experimental data set to validate the CFD prediction. For the CFD aspect, ANSYS CFX software was chosen as the solver. Two different types of mesh (hexagonal and tetrahedral) will be compared to the experimental results. It is understood that hexagonal (HEX) meshes are more time consuming and more computationally demanding, they are less prone to mesh sensitivity and have the tendancy to converge at a faster rate than the tetrahedral (TET) mesh. It was found that the HEX mesh was able to generate more consistent results and had less error than TET mesh.

Fully Parametric High-Fidelity CFD Model for the Design Optimisation of the Cyclic Stagger Pattern of a Set of Fan Outlet Guide Vanes

The present paper deals with the redesign of cyclic variation of a set of fan outlet guide vanes by means of high-fidelity full-annulus CFD. The necessity for the aerodynamic redesign originated from a change to the original project requirement, when the customer requested an increase in specific thrust above the original engine specification. The main objectives of this paper are: 1) make use of 3D CFD simulations to accurately model the flow field and identify high-loss regions; 2) elaborate an effective optimisation strategy using engineering judgement in order to define realistic objectives, constraints and design variables; 3) emphasise the importance of parametric geometry modelling and meshing for automatic design optimisation of complex turbomachinery configurations; 4) illustrate that the combination of advanced optimisation algorithms and aerodynamic expertise can lead to successful optimisations of complex turbomachinery components within practical time and costs constrains. The current design optimisation exercise was carried out using an in-house set of software tools to mesh, resolve, analyse and optimise turbomachinery components by means of Reynolds-averaged Navier-Stokes simulations. The original configuration was analysed using the 3D CFD model and thereafter assessed against experimental data and flow visualisations. The main objective of this phase was to acquire a deep insight of the aerodynamics and the loss mechanisms. This was important to appropriately limit the design scope and to drive the optimisation in the desirable direction with a limited number of design variables. A mesh sensitivity study was performed in order to minimise computational costs. Partially converged CFD solutions with restart and response surface models were used to speed up the optimisation loop. Finally, the single-point optimised circumferential stagger pattern was manually adjusted to increase the robustness of the design at other flight operating conditions. Overall, the optimisation resulted in a major loss reduction and increased operating range. Most important, it provided the project with an alternative and improved design within the time schedule requested and demonstrated that CFD tools can be used effectively not only for the analysis but also to provide new design solutions as a matter of routine even for very complex geometry configurations.

Effects of Tip Clearance on Unsteady Flow Characteristics in an Axial Turbine Stage

The clearance between the rotor blade tip and casing wall in turbomachinery passages induces leakage flow loss and thus degrades aerodynamic performance of the machine. The flow field in turbomachinery is significantly influenced by the rotor blade tip clearance size. To investigate the effects of tip clearance size on the rotor-stator interaction, the turbine stage profile from Matsunuma’s experimental tests was adopted, and the unsteady flow fields with two tip clearance sizes of 0.67% and 2.00% of blade span was numerical simulated based on Harmonic method using NUMECA software. By comparing with the domain scaling method, the accuracy of the harmonic method was verified. The interaction mechanism between the stator wake and the leakage flow was investigated. It is found that the recirculation induced by the stator wake is separated by a significant “interaction line” from the flow field close to the suction side in the clearance region. The trend of the pressure fluctuation is contrary on both sides of the line. When the stator wakes pass by the suction side, the pressure field fluctuates and the intensity of the tip leakage flow varies. With the clearance size increasing, the “interaction line” is more far away from the suction side and the intensity of tip leakage flow also fluctuates more strongly.

Investigation of the Flow Field on a Transonic Turbine Nozzle Guide Vane With Rim Seal Cavity Flow Ejection

Ensuring an adequate life of high pressure turbines requires efficient cooling methods, such as rim seal flow ejection from the stator-rotor wheel space cavity interface, which prevents hot gas ingress into the rotor disk. The present work addresses the potential to improve the efficiency in transonic turbines at certain rim seal ejection rates. To understand this process a numerical study was carried out combining computational fluid dynamic simulations (CFD) and experiments on a single stage axial test turbine. The three dimensional steady CFD analysis was performed modeling the purge cavity flow ejected downstream of the stator blade row, at three flow regimes, subsonic M2 = 0.73, transonic M2 = 1.12 and supersonic M2 = 1.33. Experimental static pressure measurements were used to calibrate the computational model. The main flow field-purge flow interaction is found to be governed by the vane shock structures at the stator hub. The interaction between the vane shocks at the hub and the purge flow has been studied and quantitatively characterized as function of the purge ejection rate. The ejection of 1% of the core flow from the rim seal cavity leads to an increase of the hub static pressure of approximately 7% at the vane trailing edge. This local reduction of the stator exit Mach number decreases the trailing edge losses in the transonic regime. Finally, a numerically predicted loss breakdown is presented, focusing on the relative importance of the trailing edge losses, boundary layer losses, shock losses and mixing losses, as a function of the purge rate ejected. Contrary to the experience in subsonic turbines, results in a transonic model demonstrate that ejecting purge flow improves the vane efficiency due to the shock structures modification downstream of the stator.