Flow in a Turbine Cascade: Part 1—Losses and Leading-Edge Effects

1984 ◽  
Vol 106 (2) ◽  
pp. 400-407 ◽  
Author(s):  
J. Moore ◽  
A. Ransmayr
Keyword(s):  
Large Scale ◽  
Static Pressure ◽  
Flow Direction ◽  
Turbine Blades ◽  
Leading Edge ◽  
Turbine Cascade ◽  
Trailing Edge ◽  
Linear Cascade ◽  
High Loss ◽  
Downstream Flow

An experimental investigation was conducted to study the effect of the leading-edge shape on the overall losses in a large-scale linear cascade of turbine blades. The leading-edge shapes used were a cylinder and a wedge. The cascade was designed to be geometrically similar to the cascade used by Langston et al. at United Technologies Research Center, with the same span/chord and pitch/chord ratios. Measurements of wall static pressure on the blades and of total pressure and flow direction downstream of the cascade showed only minor changes due to the alteration of the leading-edge shape. The measurements of the flow and loss distributions downstream of the cascade complement the results of Langston et al., which showed the flow development only within the cascade. The downstream flow is important, however, as apppoximately 50 percent of the losses occur downstream of the trailing edge. Regions of high loss were found near midspan at an axial location 40 percent of the axial chord downstream of the trailing edge. The sources of fluid in these regions are determined in Part 2.

Influence of Tip Clearance on the Inter Blade and Exit Flow Field of a Turbine Rotor Cascade

1994 ◽  
Author(s):  
M. Govardhan ◽  
N. Venkatrayulu ◽  
V. S. Vishnubhotla
Keyword(s):  
Static Pressure ◽  
Leading Edge ◽  
Turbine Rotor ◽  
Tip Clearance ◽  
Turbine Cascade ◽  
Suction Surface ◽  
Static Pressure Distribution ◽  
Blade Tip ◽  
Flow Through ◽  
Downstream Flow

A detailed study of flow through the blade passage and downstream of a linear turbine cascade was carried out for four cases of tip clearance including zero clearance. Apart from inlet traverse, a total of eight stations were chosen for inter-blade flow traversing between 5% and 95% of axial chord from leading edge. Downstream flow surveys were made at distances of 106% of axial chord from the blade leading edge. Pitchwise and spanwise traverses were conducted for each tip clearance at these stations using a small five hole probe. Provision was also made for the measurement of static pressure distribution on the suction and pressure surfaces and also on the blade tip surface when clearance is present. At about 40% of axial chord from the leading edge, the presence of clearance vortex is identified inside the passage. The growth of the clearance vortex in size, its movement towards the suction surface and its increase in strength with the gap size were observed beyond 55% of axial chord till the trailing edge region. The rate of growth of the losses in the endwall region increased with clearance. Horse shoe vortex was not observed for the highest clearance. The overall losses increase rapidly with clearance in the rear half of the blade.

Effect of an Oscillating Flow Direction on Leading Edge Heat Transfer

1984 ◽  
Vol 106 (1) ◽  
pp. 222-228 ◽  
Author(s):  
M. L. Marziale ◽  
R. E. Mayle
Keyword(s):  
Heat Transfer ◽  
Strouhal Number ◽  
Large Scale ◽  
Flow Direction ◽  
Leading Edge ◽  
Periodic Variation ◽  
Scale Model ◽  
Transfer Rates ◽  
Large Scale Model ◽  
Turbulence Length

An experimental investigation was conducted to examine the effect of a periodic variation in the angle of attack on heat transfer at the leading edge of a gas turbine blade. A circular cylinder was used as a large-scale model of the leading edge region. The cylinder was placed in a wind tunnel and was oscillated rotationally about its axis. The incident flow Reynolds number and the Strouhal number of oscillation were chosen to model an actual turbine condition. Incident turbulence levels up to 4.9 percent were produced by grids placed upstream of the cylinder. The transfer rate was measured using a mass transfer technique and heat transfer rates inferred from the results. A direct comparison of the unsteady and steady results indicate that the effect is dependent on the Strouhal number, turbulence level, and the turbulence length scale, but that the largest observed effect was only a 10 percent augmentation at the nominal stagnation position.

Experimental Investigation on the Annular Sector Cascade of a High Endwall-Angle Turbine

2016 ◽  
Author(s):  
Weiliang Fu ◽  
Jie Gao ◽  
Chen Liang ◽  
Fukai Wang ◽  
Qun Zheng ◽  
...  
Keyword(s):  
Mach Number ◽  
Static Pressure ◽  
Incidence Angle ◽  
Pressure Sensors ◽  
Leading Edge ◽  
Control Program ◽  
Turbine Cascade ◽  
Cascade Flow ◽  
Annular Sector ◽  
The Way

The flow in high endwall-angle turbine is complex, and it is different from the ordinary turbine flow in characteristics. In order to study the flow field characteristics of high endwall-angle turbines, the annular sector cascade experimental study of high endwall-angle turbines is carried out. The blade is studied experimentally in the form of annular sector cascade. The cascade includes 7 blades, and makes up 6 flow passages, in order to simulate full cascade flow. The experimental Mach number is adjusted by the way of changing inlet total pressure, and the Mach number influence (0.7, 0.8 and 0.9) on annular sector cascade flow is studied. Based on it, the inlet incidence angle (−15°, −7.5°, 0°, 7.5° and 15° )is changed with the way of changing sector straight pipes upstream of the cascade, and its influence on turbine flow fields is studied at the Mach number of 0.8. Here, five-hole probes are used to measure aerodynamic parameters distributions downstream of the cascade, and static pressure taps are positioned on the blade surface to measure surface static pressure distribution. The auto-traversing system and pressure sensors were operated by a self-compiled program based control program. The results indicate that there are two passage vortices inside the turbine cascade flow passage under the high Mach number condition, and the passage vortex near the high endwall-angle region is bigger. As Mach number increases, the passage vortices inside turbine cascade passage will become strong, and moves towards the blade mid-span. Besides, it is shown that the way of changing sector straight pipes can achieve the variation of inlet incidence angles. And, the blade profile with big leading-edge radius has good design and off-design performance. Detailed results and analyses are presented in the paper.

A Turbulent Boundary Layer Flow Over an Open Shallow Cavity

2016 ◽  
Author(s):  
Khaled J. Hammad
Keyword(s):  
Boundary Layer ◽  
Large Scale ◽  
Vector Fields ◽  
Turbulence Models ◽  
Leading Edge ◽  
Field Measurements ◽  
Recirculation Region ◽  
Mean Flow ◽  
Trailing Edge ◽  
Cavity Depth

Particle Image Velocimetry (PIV) was used to study the flow structure and turbulence, upstream, over, and downstream a shallow open cavity. Three sets of PIV measurements, corresponding to a turbulent incoming boundary layer and a cavity length-to-depth ratio of four, are reported. The cavity depth based Reynolds numbers were 21,000; 42,000; and 54,000. The selected flow configuration and well characterized inflow conditions allow for straightforward assessment of turbulence models and numerical schemes. All mean flow field measurements display a large flow recirculation region, spanning most of the cavity and a smaller, counter-rotating, secondary vortex, immediately downstream of the cavity leading edge. The Galilean decomposed instantaneous velocity vector fields, clearly demonstrate two distinct modes of interaction between the free shear and the cavity trailing edge. The first corresponds to a cascade of vortical structures emanating from the tip of the leading edge of the cavity that grow in size as they travel downstream and directly interact with the trailing edge, i.e., impinging vortices. The second represents vortices that travel above the trailing edge of the cavity, i.e., non-impinging vortices. In the case of impinging vortices, a strong, large scale region of recirculation forms inside the cavity and carries the flow disturbances, arising from the impingement of vortices on the trailing edge of the cavity, upstream in a manner that interacts with and influences the flow as it separates from the cavity leading edge.

scholarly journals Interpreting Aerodynamics of a Transonic Impeller from Static Pressure Measurements

2018 ◽  
Vol 2018 ◽  
pp. 1-9
Author(s):  
Fangyuan Lou ◽  
John Charles Fabian ◽  
Nicole Leanne Key
Keyword(s):  
Flow Field ◽  
High Speed ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Pressure Measurements ◽  
High Loss ◽  
Supersonic Inlet ◽  
Mach Numbers ◽  
Inlet Conditions

This paper investigates the aerodynamics of a transonic impeller using static pressure measurements. The impeller is a high-speed, high-pressure-ratio wheel used in small gas turbine engines. The experiment was conducted on the single stage centrifugal compressor facility in the compressor research laboratory at Purdue University. Data were acquired from choke to near-surge at four different corrected speeds (Nc) from 80% to 100% design speed, which covers both subsonic and supersonic inlet conditions. Details of the impeller flow field are discussed using data acquired from both steady and time-resolved static pressure measurements along the impeller shroud. The flow field is compared at different loading conditions, from subsonic to supersonic inlet conditions. The impeller performance was strongly dependent on the inducer, where the majority of relative diffusion occurs. The inducer diffuses flow more efficiently for inlet tip relative Mach numbers close to unity, and the performance diminishes at other Mach numbers. Shock waves emerging upstream of the impeller leading edge were observed from 90% to 100% corrected speed, and they move towards the impeller trailing edge as the inlet tip relative Mach number increases. There is no shock wave present in the inducer at 80% corrected speed. However, a high-loss region near the inducer throat was observed at 80% corrected speed resulting in a lower impeller efficiency at subsonic inlet conditions.

Leading Edge Modification Effects on Turbine Cascade Endwall Loss

2003 ◽  
Author(s):  
S. Becz ◽  
M. S. Majewski ◽  
L. S. Langston
Keyword(s):  
Large Scale ◽  
Leading Edge ◽  
Total Loss ◽  
Optimum Shape ◽  
Loss Reduction ◽  
Turbine Cascade ◽  
Pressure Loss Coefficient ◽  
Contour Plots ◽  
Flow Loss ◽  
Total Pressure Loss Coefficient

Experimental results are presented which provide area-averaged total pressure loss coefficient measurements for four different turbine airfoil leading edge configurations. A baseline (Langston) configuration, two leading edge bulbs, and a leading edge fillet were tested in a large-scale, low aspect ratio, high turning linear cascade. Results show that both the small bulb and fillet geometries each reduced area averaged total loss by 8%, while the large bulb exhibited a slight increase in total loss. Contour plots for all geometries are presented and the major differences between each are discussed. Through investigation of pitch averaged loss profiles it is found that the area of greatest reduction differs between the small bulb and fillet, leading to the possibility that the mechanisms through which each is affecting the flow may be different. This provides hope that the best features of each may potentially be combined to determine an optimum shape for secondary flow loss reduction.

Unsteady Flow in Oscillating Turbine Cascade: Part 1 — Linear Cascade Experiment

1996 ◽  
Author(s):  
L. He
Keyword(s):  
Computational Study ◽  
Separation Bubble ◽  
Turbine Blades ◽  
Leading Edge ◽  
Computational Method ◽  
Reattachment Point ◽  
Suction Surface ◽  
Linear Cascade ◽  
Unsteady Pressure ◽  
Pressure Surface

An experimental and computational study has been carried out on a linear cascade of low pressure turbine blades with the middle blade oscillating in a torsion mode. The main objectives of the present work were to enhance understanding of the behaviour of bubble type of flow separation and to examine the predictive ability of a computational method. In addition, an attempt was made to address a general modelling issue: was the linear assumption adequately valid for such kind of flow? In Part 1 of this paper, the experimental work was described. Unsteady pressure was measured along blade surfaces using off-board mounted pressure transducers at realistic reduced frequency conditions. A short separation bubble on the suction surface near the trailing edge and a long leading-edge separation bubble on the pressure surface were identified. It was found that in the regions of separation bubbles, unsteady pressure was largely influenced by the movement of reattachment point, featured by an abrupt phase shift and an amplitude trough in the 1st harmonic distribution. The short bubble on the suction surface seemed to follow closely a laminar bubble transition model in a quasi-steady manner, and had a localized effect. The leading-edge long bubble on the pressure surface, on the other hand, was featured by a large movement of the reattachment point, which affected the surface unsteady pressure distribution substantially. As far as the aerodynamic damping was concerned, there was a destabilizing effect in the separated flow region, which was however largely balanced by the stabilizing effect downstream of the reattachment point due to the abrupt phase change.

Blade Thickness Effect on Impeller Slip Factor

2010 ◽  
Author(s):  
Donghui Zhang ◽  
Jean-Luc Di Liberti ◽  
Michael Cave
Keyword(s):  
Factor Model ◽  
Numerical Study ◽  
Flow Direction ◽  
Leading Edge ◽  
Thickness Effect ◽  
Centrifugal Impeller ◽  
Trailing Edge ◽  
Slip Factor ◽  
Blade Thickness ◽  
Blockage Factor

A numerical study of the effect of the blade thickness on centrifugal impeller slip factor is presented in this paper. The CFD results show that generally the slip factor decreases as the blade thickness increases. Changing the thickness at different locations has different effects on the slip factor. The shroud side blade thickness has more effect on the impeller slip factor than the hub side blade thickness. In the flow direction, the blade thickness at 50% meridional distance is the major factor affecting the slip factor. The leading edge thickness has little effect on slip factor. There is an optimum thickness at the trailing edge for the maximum slip factor. For this impeller, the hub side thickness ratio of 0.5 between the trailing edge and the middle of the impeller gives the highest value of the slip factor, while the ratio of 0.25 at shroud side gives the highest value of the slip factor. A blockage factor is added into the slip factor model to include the aerodynamic blockage effect on the slip factor. The model explains the phenomena observed in the CFD results and the test data very well.

Effect of the wing trailing-edge flaps and spoilers position on the jet-vortex wake behind an aircraft during takeoff and landing run

2019 ◽  
pp. 095441001988215
Author(s):  
Yu M Tsirkunov ◽  
MA Lobanova ◽  
AI Tsvetkov ◽  
BA Schepanyuk
Keyword(s):  
Large Scale ◽  
Stokes Equations ◽  
Computational Simulation ◽  
Leading Edge ◽  
Velocity Profiles ◽  
Essential Elements ◽  
Trailing Edge ◽  
Structure Of Flow ◽  
Trailing Edge Flaps ◽  
Calculated Velocity

The large-scale vortex structure of flow in the near wake behind an aircraft during its run on a runway is investigated numerically. The geometrical aircraft configuration was taken close to a mid-range commercial aircraft like Boeing 737-300. It included all essential elements: a body (fuselage), wings with winglets, horizontal and vertical stabilizers, engine nacelles, nacelle pylons, inboard flap track fairings, leading-edge and trailing-edge flaps, and spoilers. The position of flaps and spoilers corresponded to the takeoff and landing run conditions. Computational simulation was based on solving the Reynolds averaged Navier–Stokes equations closed with the Menter Shear Stress Transport turbulence model. Patterns of streamlines, fields of the axial vorticity and the turbulent intensity, vertical and horizontal velocity profiles in the wake are compared and discussed for both run regimes. The flow model was preliminary tested for validity by comparison of the calculated velocity profiles behind a reduced-scale aircraft model with those obtained in special wind tunnel experiments.

EXPERIMENTAL STUDY OF THE ENDWALL HEAT TRANSFER OF A TRANSONIC NOZZLE GUIDE VANE WITH UPSTREAM JET PURGE COOLING PART 1 - EFFECT OF DENSITY RATIO

2021 ◽  
pp. 1-36
Author(s):  
Shuo Mao ◽  
Ridge A. Sibold ◽  
Wing Ng ◽  
Zhigang LI ◽  
Bo Bai ◽  
...  
Keyword(s):  
Large Scale ◽  
Guide Vane ◽  
Flow Direction ◽  
Density Ratio ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Vital Role ◽  
Blowing Ratio ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract Nozzle guide vane platforms often employ complex cooling schemes to mitigate the ever-increasing thermal loads on endwall. This study analyzes, experimentally and numerically, and describes the effect of coolant to mainstream blowing ratio, momentum ratio and density ratio for a typical axisymmetric converging nozzle guide vane platform with an upstream doublet staggered, steep-injection, cylindrical hole purge cooling scheme. Nominal flow conditions were engine-representative and as follows: Maexit = 0.85, Reexit,Cax = 1.5×106 and an inlet large-scale freestream turbulence intensity of 16%. Two blowing ratios were investigated, each corresponding to the design condition and its upper extrema at M = 2.5 and 3.5, respectively. For each blowing ratio, the coolant to mainstream density ratio was varied between DR=1.2, representing typical experimental neglect of coolant density, and DR=1.95, representative of typical engine conditions. The results show that with a fixed coolant-to-mainstream blowing ratio, the density ratio plays a vital role in the coolant-mainstream mixing and the interaction between coolant and horseshoe vortex near the vane leading edge. A higher density ratio leads to a better coolant coverage immediately downstream of the cooling holes but exposes the in-passage endwall near the pressure side. It also causes the in-passage coolant coverage to decay at a higher rate in the flow direction. From the results gathered, both density ratio and blowing ratio should be considered for accurate testing, analysis, and prediction of purge jet cooling scheme performance.

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