Influence of a Novel Three-Dimensional Leading Edge Geometry on the Aerodynamic Performance of Transonic Cascade Vanes

2012 ◽  
Author(s):  
D. Bouchard ◽  
A. Asghar ◽  
J. Hardes ◽  
R. Edwards ◽  
W. D. E. Allan ◽  
...  
Keyword(s):  
Total Pressure ◽  
Three Dimensional ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Surface Flow ◽  
Pressure Losses ◽  
Low Pressure Turbine ◽  
Edge Geometry ◽  
Turbine Vane ◽  
Transonic Cascade

This paper addresses the issue of aerodynamic performance of a novel 3D leading edge modification to a reference vane. An analysis of tubercles found in nature and some engineering applications was used to synthesize new leading edge geometry. Three variations of the reference low pressure turbine vane were obtained by changing the characteristic parameters of the tubercles. Shock structure, surface flow visualization and total pressure measurements were made through experiments in a cascade rig, as well as through computational fluid dynamics. The tests were carried out at design zero incidence and off-design ±10-deg and ±5-deg incidences. The performance of the new 3D leading edge geometries was compared against the reference vane. Some leading edge tubercle configurations were effective at decreasing total pressure losses at positive inlet incidence angles. Numerical results supplemented experimental results.

Influence of a Novel 3D Leading Edge Geometry on the Aerodynamic Performance of Low Pressure Turbine Blade Cascade Vanes

2014 ◽  
Author(s):  
A. Asghar ◽  
W. D. E. Allan ◽  
M. LaViolette ◽  
R. Woodason
Keyword(s):  
Turbine Blade ◽  
Pressure Loss ◽  
Total Pressure ◽  
Separated Flow ◽  
Leading Edge ◽  
Low Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Low Pressure Turbine ◽  
Edge Geometry

This paper addresses the issue of aerodynamic performance of a novel 3D leading edge modification to a reference low pressure turbine blade. An analysis of tubercles found in nature and used in some engineering applications was employed to synthesize new leading edge geometry. A sinusoidal wave-like geometry characterized by wavelength and amplitude was used to modify the leading edge along the span of a 2D profile, rendering a 3D blade shape. The rationale behind using the sinusoidal leading edge was that they induce streamwise vortices at the leading edge which influence the separation behaviour downstream. Surface pressure and total pressure measurements were made in experiments on a cascade rig. These were complemented with computational fluid dynamics studies where flow visualization was also made from numerical results. The tests were carried out at low Reynolds number of 5.5 × 104 on a well-researched profile representative of conventional low pressure turbine profiles. The performance of the new 3D leading edge geometries was compared against the reference blade revealing a downstream shift in separated flow for the LE tubercle blades; however, total pressure loss reduction was not conclusively substantiated for the blade with leading edge tubercles when compared with the performance of the baseline blade. Factors contributing to the total pressure loss are discussed.

Investigation of Variated Unsteady Inflow Boundary Conditions on the Transition Behavior of a Low Pressure Turbine Cascade Family

2014 ◽  
Author(s):  
Franz F. Blaim ◽  
Roland E. Brachmanski ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Condition ◽  
Transport Equation ◽  
Flow Separation ◽  
Total Pressure ◽  
Low Pressure ◽  
Test Facility ◽  
Number Range ◽  
Pressure Losses ◽  
Low Pressure Turbine ◽  
Transition Behavior

The objective of this paper is to analyze the influence of incoming periodic wakes, considering the variable width, on the integral total pressure loss for two low pressure turbine (LPT) airfoils. In order to reduce the overall weight of a LPT, the pitch to chord ratio was continuously increased, during the past decades. However, this increase encourages the development of the transition phenomena or even flow separation on the suction side of the blade. At low Reynolds numbers, large separation bubbles can occur there, which are linked with high total pressure losses. The incoming wakes of the upstream blades are known to trigger early transition, leading to a reduced risk of flow separation and hence minor integral total pressure losses caused by separation. For the further investigation of these effects, different widths of the incoming wakes will be examined in detail, here. This variation is carried out by using the numerical Unsteady Reynolds Averaged Solver TRACE developed by the DLR Cologne in collaboration with MTU Aero Engines AG. For the variation of the width of the wakes, a variable boundary condition was modeled, which includes the wake vorticity parameters. The width of the incoming wakes was used as the relevant variable parameter. The implemented boundary condition models the unsteady behavior of the incoming wakes by the variation of the velocity profile, using a prescribed frequenc. TRACE can use two different transition models; the main focus here is set to the γ–Reθt transition model, which uses local variables in a transport equation, to trigger the transition within the turbulence transport equation system. The experimental results were conducted at the high speed cascade open loop test facility at the Institute for Jet Propulsion at the University of the German Federal Armed Forces in Munich. For the investigation presented here, two LPT profiles — which were designed with a similar inlet angle, turning, and pitch are analyzed. However, with a common exit Mach number and a similar Reynolds number range between 40k and 400k, one profile is front loaded and the other one is aft loaded. Numerical unsteady results are in good agreement with the conducted measurements. The influence of the width of the wake on the time resolved transition behavior, represented by friction coefficient plots and momentum loss thickness will be analyzed in this paper.

Influence of Pre-History and Leading Edge Contouring on Aero-Performance of a 3D Nozzle Guide Vane

2013 ◽  
Author(s):  
Ranjan Saha ◽  
Boris I. Mamaev ◽  
Jens Fridh ◽  
Björn Laumert ◽  
Torsten H. Fransson
Keyword(s):  
Guide Vane ◽  
Three Dimensional ◽  
Leading Edge ◽  
Stagnation Line ◽  
Wide Range ◽  
Nozzle Guide Vane ◽  
Turbine Vane ◽  
Loss Coefficients ◽  
Exit Mach Number ◽  
Nozzle Guide

Experiments are conducted to investigate the effect of the pre-history in the aerodynamic performance of a three-dimensional nozzle guide vane with a hub leading edge contouring. The performance is determined with two pneumatic probes (5 hole and 3 hole) concentrating mainly on the endwall. The investigated vane is a geometrically similar gas turbine vane for the first stage with a reference exit Mach number of 0.9. Results are compared for the baseline and filleted cases for a wide range of operating exit Mach numbers from 0.5 to 0.9. The presented data includes loading distributions, loss distributions, fields of exit flow angles, velocity vector and vorticity contour, as well as, mass-averaged loss coefficients. The results show an insignificant influence of the leading edge fillet on the performance of the vane. However, the pre-history (inlet condition) affects significantly in the secondary loss. Additionally, an oil visualization technique yields information about the streamlines on the solid vane surface which allows identifying the locations of secondary flow vortices, stagnation line and saddle point.

scholarly journals Cold Flow Computations for the Diffuser-Combustor Section of an Industrial Gas Turbine

1996 ◽  
Author(s):  
Dadong Zhou ◽  
Ting Wang ◽  
William R. Ryan
Keyword(s):  
Gas Turbine ◽  
Total Pressure ◽  
Reverse Flow ◽  
Three Dimensional ◽  
Complex Flow ◽  
Pressure Losses ◽  
Structured Grid ◽  
Computational Fluid Dynamics Cfd ◽  
Enhance Efficiency ◽  
Industrial Gas Turbine

In the first part of a multipart project to analyze and optimize the complex three-dimensional diffuser-combustor section of a highly advanced industrial gas turbine under development, a computational fluid dynamics (CFD) analysts has been conducted. The commercial FEA code I-DEAS was used to complete the three-dimensional solid modeling and the structured grid generation. The flow calculation was conducted using the commercial CFD code PHOENICS. The multiblock method was employed to enhance computational capabilities. The mechanisms of the total pressure losses and possible ways to enhance efficiency by reducing the total pressure losses were examined. Mechanisms that contribute to the nonuniform velocity distribution of flow entering the combustor were also identified. The CFD results were informative and provided insight to the complex flow patterns in the reverse flow dump diffuser, however, the results are qualitative and are useful primarily as guidelines for optimization as opposed to firm design configuration selections.

Adjoint Based Aerodynamic Optimization of a Multi-Splitter Turbine Vane Frame

2019 ◽  
Author(s):  
Vincenzo Russo ◽  
Simone Orsenigo ◽  
Lasse Mueller ◽  
Tom Verstraete ◽  
Sergio Lavagnoli
Keyword(s):  
Total Pressure ◽  
Low Pressure ◽  
Design Parameters ◽  
Structural Constraints ◽  
Aerodynamic Optimization ◽  
Pressure Losses ◽  
Turbine Vane ◽  
Gradient Based ◽  
Adjoint Approach ◽  
Multi Body

Abstract This work presents a 2D optimization of a multi-body turbine vane frame (TVF), a particular configuration that can lead to considerable shortening of the aero-engine shaft as well as weight reduction. Traditionally, the turbine vane frame is used to guide the flow from the high pressure (HP) turbine to the low pressure (LP) turbine. Current designs have a mid turbine frame equipped with non lifting bodies that have structural and servicing functions, while multi-body configurations are characterized by the fact that, in order to shorten the duct length, the mid turbine struts are merged with the LP stator vanes, traditionally located downstream. This design architecture consists therefore of a multi-body vane row, where lifting long-chord struts replace some of the low pressure vane airfoils. However, the bulky struts cause significant aerodynamics losses and penalize the aerodynamics of the small vanes. The objective of the present work is to numerically optimize a TVF geometry with multi-body architecture using a gradient based algorithm coupled with the adjoint approach, enabling the use of a rich design space. Steady-state CFD simulations have been used to this end. The aim of this study is to reduce the total pressure losses of the TVF, while imposing several aerodynamic and structural constraints. The parametrization of the TVF geometry represents the airfoil shapes and their relative pitch-wise positions. The outcome of the optimization is to evaluate the potential improvements introduced by the optimized TVF geometry and to quantify the influence of the different design parameters on the total pressure losses.

Shockwave and Noise Abatement of Transonic Fans

2004 ◽  
Author(s):  
Liping Xu
Keyword(s):  
Shock Wave ◽  
Noise Reduction ◽  
Three Dimensional ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Test Results ◽  
Shock Strength ◽  
Noise Levels ◽  
Noise Abatement ◽  
Blade Passage

The aerodynamic sources of the forward tone noise of transonic civil fans are analysed. The leading edge shockwave near the rotor tip section is identified as the main source of tone noise. By comparing the tone noise levels of the same fan operating at two different working lines, numerical calculations show that on the lower working line, the main passage shockwave is swallowed and locked into the blade passage, and the fan blades act as a shield to prevent the strong passage shock wave from propagating upstream. The calculations show that, by running the fan at a lower working line, up to 6 db abatement in the blade passing frequency (BPF) tone can be achieved through shielding the shockwave. With three dimensional CFD it is possible to design swept rotors which have desired shockwave structures near the tip region. Fan rotors with different swept leading edges have been designed to study this effect and comparisons in aerodynamics performances as well as the tone noise levels are made. It is predicted that in a swept rotor the leading edge shock strength can be further weakened and up to 5db further reduction in tone noise is possible. With a more secure shockwave shielding, a forward swept rotor has the combination of better aerodynamic performance and better noise abatement feature. The design and test results of a three dimensional fan rotor LNR2, featuring localised forward swept rotor are presented. Rig test results show that although the noise reduction through shock shielding has been demonstrated, the aerodynamics and noise are complicated by the problems specific to such localised forward swept fan.

Experimental Study on the Three-Dimensional Flow Within a Compressor Cascade With Tip Clearance: Part I—Velocity and Pressure Fields

1993 ◽  
Vol 115 (3) ◽  
pp. 435-443 ◽  
Author(s):  
S. Kang ◽  
C. Hirsch
Keyword(s):  
Three Dimensional ◽  
Leading Edge ◽  
Surface Flow ◽  
Tip Clearance ◽  
Tip Vortex ◽  
Compressor Cascade ◽  
Three Dimensional Flow ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Dimensional Flow

Experimental results from a study of the three-dimensional flow in a linear compressor cascade with stationary endwall at design conditions are presented for tip clearance levels of 1.0, 2.0, and 3.3 percent of chord, compared with the no-clearance case. In addition to five-hole probe measurements, extensive surface flow visualizations are conducted. It is observed that for the smaller clearance cases a weak horseshoe vortex forms in the front of the blade leading edge. At all the tip gap cases, a multiple tip vortex structure with three discrete vortices around the midchord is found. The tip leakage vortex core is well defined after the midchord but does not cover a significant area in traverse planes. The presence of the tip leakage vortex results in the passage vortex moving close to the endwall and the suction side.

Detailed Three-Dimensional Velocity Field Measurements of a Complex Internal Cooling Flow Within a Gas Turbine Vane

2019 ◽  
Author(s):  
Michael J. Benson ◽  
David Helmer ◽  
Bret P. Van Poppel ◽  
Benjamin Duhaime ◽  
David Bindon ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Three Dimensional ◽  
Leading Edge ◽  
Scale Model ◽  
Internal Cooling ◽  
Complex Flow ◽  
Copper Sulfate Solution ◽  
Flow Rates ◽  
Experimental Apparatus ◽  
Turbine Vane

Abstract A 6.67 scale model of the Advanced Recirculation Total Impingement Cooling (ARTIC) gas turbine vane insert’s leading edge was designed, built using stereolithography (SLA) fabrication methods, and tested using Magnetic Resonance Velocimetry (MRV), a non-invasive data acquisition technique that captures three-dimensional, three-component velocity fields of a copper sulfate solution over the course of several hours. The experimental apparatus supplied constant flow rates through a test section placed within a 3.0 Tesla MRI magnet. Tests were run at two fully turbulent flow rates corresponding to Reynolds numbers based on hydraulic diameter of 10,000 and 20,000 with the higher flow rate case achieving dynamic similarity with the full-scale ARTIC device. The experimental results elucidated key details and intricacies of the complex flow within the insert. Analysis of flow distribution between each of the three independent impingement zones revealed a degree of measurable jet to jet variability. Stagnation and recirculation zones were detected, informing design modifications and enabling assessment of inlet effects on impingement. Measurement uncertainty was assessed and estimated to be approximately 7.5% of the peak velocity at the inlet to the central feed cavity.

scholarly journals Effect of Wing Corrugation on the Aerodynamic Efficiency of Two-Dimensional Flapping Wings

2020 ◽  
Vol 10 (20) ◽  
pp. 7375
Author(s):  
Thanh Tien Dao ◽  
Thi Kim Loan Au ◽  
Soo Hyung Park ◽  
Hoon Cheol Park
Keyword(s):  
Three Dimensional ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Flapping Wing ◽  
Flapping Wings ◽  
Two Dimensional ◽  
Aerodynamic Efficiency ◽  
Insect Wing ◽  
Wing Motion ◽  
Computational Fluid Dynamics Cfd

Many previous studies have shown that wing corrugation of an insect wing is only structurally beneficial in enhancing the wing’s bending stiffness and does not much help to improve the aerodynamic performance of flapping wings. This study uses two-dimensional computational fluid dynamics (CFD) in aiming to identify a proper wing corrugation that can enhance the aerodynamic performance of the KUBeetle, an insect-like flapping-wing micro air vehicle (MAV), which operates at a Reynolds number of less than 13,000. For this purpose, various two-dimensional corrugated wings were numerically investigated. The two-dimensional flapping wing motion was extracted from the measured three-dimensional wing kinematics of the KUBeetle at spanwise locations of r = (0.375 and 0.75)R. The CFD analysis showed that at both spanwise locations, the corrugations placed over the entire wing were not beneficial for improving aerodynamic efficiency. However, for the two-dimensional flapping wing at the spanwise location of r = 0.375R, where the wing experiences relatively high angles of attack, three specially designed wings with leading-edge corrugation showed higher aerodynamic performance than that of the non-corrugated smooth wing. The improvement is closely related to the flow patterns formed around the wings. Therefore, the proposed leading-edge corrugation is suggested for the inboard wing of the KUBeetle to enhance aerodynamic performance. The corrugation in the inboard wing may also be structurally beneficial.

A study of fan-distortion interaction within the fan rotor

2020 ◽  
pp. 095765092092004
Author(s):  
Zhihui Li ◽  
Juan Du ◽  
Qianfeng Zhang ◽  
Guofeng Ji ◽  
Hongwu Zhang
Keyword(s):  
High Performance ◽  
Fluid Dynamic ◽  
Three Dimensional ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Stagnation Pressure ◽  
Detached Eddy Simulation ◽  
Axial Fan ◽  
Fan Performance ◽  
Inlet Distortion

Boundary-layer-ingesting fans and compressors in the next-generation turbofan engines require high-performance operations under distorted inflow. The aim of this work is to study the effects of inlet distortions including inlet stagnation pressure and temperature distortion, on the aerodynamic performance of a transonic axial fan. Firstly, the validated full-annulus, unsteady, three-dimensional computational fluid dynamic code in conjunction with detached Eddy simulation approach is used here to simulate the fan flows assembly with individual inlet stagnation pressure/temperature distortion. Then, the propagation process of the inlet distortion waves is analyzed to understand how the aerodynamic performance degradation is triggered. The simulation results show that the fan performance is remarkably degraded when the inlet distortion is introduced. The leading-edge spillage, the trailing edge back flow and the “tornado vortex” occur when parts of fan blades encounter the incoming distorted flows. Finally, the responses of fan to the combined inlet stagnation distortion effects are discussed in this paper. It is found that the combined distortion effects can be predicted based on the sum of the performance responses to the individual constituent distortions. Furthermore, the relative location of the constituent distortions shows a non-ignorable influence on the overall fan performance, especially for the intensified inlet distortion.

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