scholarly journals Cavitation in Turbopumps—Part 1

1962 ◽  
Vol 84 (3) ◽  
pp. 326-338 ◽  
Author(s):  
L. B. Stripling ◽  
A. J. Acosta
Keyword(s):  
Leading Edge ◽  
Simplified Model ◽  
Flat Plates ◽  
Streamline Flow ◽  
Cavitation Process ◽  
Flow Through ◽  
Loss Coefficients ◽  
Blade Geometry

A free-streamline flow through a cascade of semi-infinite flat plates is taken as a simplified model of the cavitation process in a helical inducer pump. The length and thickness of the resulting cavity is determined as a function of blade geometry and cavitation parameter. Loss coefficients resulting from the cavitation are estimated and representative cavity shapes are calculated to aid in designing the leading edge shape of the blades.

The Prediction of Flow Through Leading Edge Holes in a Film Cooled Airfoil With and Without Inserts

1985 ◽  
Vol 107 (1) ◽  
pp. 92-98 ◽  
Author(s):  
E. S. Tillman ◽  
E. O. Hartel ◽  
H. F. Jen
Keyword(s):  
Leading Edge ◽  
Flow Rates ◽  
Cooling Flow ◽  
Internal Loss ◽  
Measured Flow ◽  
Flow Through ◽  
Loss Coefficients ◽  
Cooling Air Flow ◽  
Cooling Air ◽  
Good Agreement

A method for predicting cooling air flow rates using tests on cylindrical models of typical turbine blade leading edges has been extended to include blades with inserts and blades with reversed-angled holes. When an insert is used, the pressure loss across the insert can be determined from flow tests and added to other losses in the flow path to determine cooling flow rates. Calculated and experimentally determined flow rates are compared with good agreement. The second experiment was performed to determine internal loss coefficients for reverse-angled holes oriented so the flow makes a reverse turn to enter the holes. The reversed flow case produced significantly greater internal loss coefficients than when the same holes were oriented in the direction of flow. These results were used to predict flow from arrays of reverse- angled holes and from a cylinder containing both reverse-angled holes and nonreversed holes. In all cases, good agreement was found between predicted and measured flow rates.

On sound generation by the interaction between turbulence and a cascade of airfoils with non-uniform mean flow

2002 ◽  
Vol 463 ◽  
pp. 25-52 ◽  
Author(s):  
I. EVERS ◽  
N. PEAKE
Keyword(s):  
Flat Plate ◽  
Steady Flow ◽  
Power Level ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Sound Generation ◽  
Mean Flow ◽  
Noise Levels ◽  
Flat Plates ◽  
Blade Geometry

The sound generated by the interaction between a turbulent rotor wake and a stator is modelled by considering the gust response of a cascade of blades in non-uniform, subsonic mean flow. Previous work by Hanson & Horan (1998) that considers a cascade of flat plates at zero incidence is extended to take into account blade geometry and angle of attack. Our approach is based on the work of Peake & Kerschen (1997), who calculate the forward radiation due to the interaction between a single vortical gust and a cascade of flat plates at non-zero angle of attack. The extensions completed in this present paper are two-fold: first we include the effects of small but non-zero camber and thickness; and second we produce uniformly valid approximations which predict the upstream radiation near modal cut-off. The thin-airfoil singularity in the steady flow at each leading edge is crucial in our model of the sound generation. A new analytical expression for the coefficient of this singularity is derived via a sequence of conformal mappings, and it turns out that in our asymptotic limit this is the only quantity which needs to be calculated from the steady flow in order to predict time-averaged noise levels. Once the response to a single gust has been completed, we use Hanson & Horan (1998)'s approach to determine the response to an incident turbulent spectrum, and find that as well as the noise corresponding to the auto-correlation of the gust velocity component normal to the blade, there is also a contribution from the cross-correlation of the normal and tangential velocities. Predictions are made of the effects of blade geometry on the upstream acoustic power level. The blade geometry can have a very significant effect on the noise generated by interaction with a single gust, with changes of up to 10 dB from the flat-plate noise levels. However, once these gust results have been integrated over a full incident turbulence spectrum the effects of the geometry are rather smaller, although still potentially significant, leading to changes of up to about 2 dB from the flat-plate results. The implication of all this is that the blade geometry can have a significant effect on the tonal noise components generated by rotor–stator interaction (i.e. by single harmonic gusts), but that the broadband part of the noise spectrum is relatively unaffected.

Some Unresolved Features in the Physics of Quasi Two-Dimensional Flows Over Turbine Blading

2015 ◽  
Author(s):  
J. P. Gostelow ◽  
W. D. E. Allan ◽  
A. Mahallati
Keyword(s):  
Plate Boundary ◽  
Leading Edge ◽  
Circular Cylinders ◽  
Pressure Gradients ◽  
Two Dimensional ◽  
Flat Plates ◽  
Two Dimensional Flow ◽  
Flow Through ◽  
Turbine Blading ◽  
Two Dimensional Flows

Even for the ostensibly two-dimensional flow through cascades of blades, many details of the flow physics are neither well-understood nor well-predicted. Gaps in knowledge are identified that cover entire blade and nozzle vane surfaces from leading edge to trailing edge and beyond. To give improved prediction capability these gaps require improved understanding. The goal of this work is to draw attention to five most significant internal aerodynamic phenomena that affect turbomachinery blade performance and design. By drawing together important experimental results awareness can be raised of these features in blading aerodynamics that are not yet clearly understood. The emphasis is on quasi two-dimensional flows. As well as work on blade cascades this research draws on fundamental investigations over flat plates and circular cylinders. Similar behavior was observed between tests under strong adverse pressure gradients on triggered spots, wake-disturbed flat plate boundary layers, and on turbine blading.

Study of the Nonsteady Flow in a Multipulse Converter

1991 ◽  
Vol 113 (3) ◽  
pp. 413-418
Author(s):  
P. Flamang ◽  
R. Sierens
Keyword(s):  
Exhaust System ◽  
Simplified Model ◽  
Flow Conditions ◽  
Velocity Measurements ◽  
Test Rig ◽  
Steady State Flow ◽  
Cyclic Flow ◽  
Flow Through ◽  
Loss Coefficients ◽  
Pressure And Velocity Measurements

In a previous paper [1] a simplified model has been proposed to calculate the pressure loss coefficients of a multipulse converter under steady-state flow conditions. Therefore a special test rig has been built, which simulates the nonsteady but cyclic flow in the exhaust system of a real engine. Pressure and velocity measurements (with LDA) are compared with the results of the numerical simulation for the flow through the multipulse converter of the test rig. Finally, a comparison is made between measurements and calculations of the pressure history in the exhaust system of a real engine. This paper proves that this simplified model accurately predicts the behavior of the multipulse converter under nonstationary flow conditions.

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 An analytical and experimental investigation of aerofoil–turbulence interaction noise for plates with spanwise-varying leading edges

2019 ◽  
Vol 865 ◽  
pp. 137-168 ◽  
Author(s):  
Lorna J. Ayton ◽  
Paruchuri Chaitanya
Keyword(s):  
Noise Reduction ◽  
Analytic Solution ◽  
Piecewise Linear ◽  
Separation Of Variables ◽  
Leading Edge ◽  
Test Case ◽  
Far Field ◽  
Mean Flow ◽  
Flat Plates ◽  
Leading Edges

This paper presents an analytic solution for gust–aerofoil interaction noise for flat plates with spanwise-varying periodic leading edges in uniform mean flow. The solution is obtained by solving the linear inviscid equations via separation of variables and the Wiener–Hopf technique, and is suitable for calculating the far-field noise generated by any leading edge with a single-valued piecewise linear periodic spanwise geometry. Acoustic results for homogeneous isotropic turbulent flow are calculated by integrating the single-gust solution over a wavenumber spectrum. The far-sound pressure level is calculated for five test-case geometries; sawtooth serration, slitted $v$-root, slitted $u$-root, chopped peak and square wave, and compared to experimental measurements. Good agreement is seen over a range of frequencies and tip-to-root ratios (varying the sharpness of the serration). The analytic solution is then used to calculate the propagating pressure along the leading edge of the serration for fixed spanwise wavenumbers, i.e. only the contribution to the surface pressure which propagates to the far field. Using these results, two primary mechanisms for noise reduction are discussed; tip and root interference, and a redistribution of energy from cuton modes to cutoff modes. A secondary noise-reduction mechanism due to nonlinear features is also discussed and seen to be particularly important for leading edges with very narrow slits.

The Reynolds Analogy Applied to a Cylinder Rotating in Air

1954 ◽  
Vol 58 (519) ◽  
pp. 205-208 ◽  
Author(s):  
Y. R. Mayhew
Keyword(s):  
Heat Transfer ◽  
Shear Stress ◽  
Heat Flow ◽  
Solid Surface ◽  
Fluid Flows ◽  
Flat Plates ◽  
Reynolds Analogy ◽  
Transfer Data ◽  
Turbulent Fluid ◽  
Flow Through

When a turbulent fluid flows past a solid surface whose temperature differs from that of the fluid, the shear stress at the surface and the heat flow from it can be related by means of the Reynolds analogy. This analogy has been improved by Prandtl, Taylor, von Kármán and others, and its validity has been tested for flow through tubes and past flat plates by several investigators. In this note the analogy is checked against shear stress data and heat transfer data for a cylinder rotating in “still” air, when the flow is turbulent.

An Investigation of the Flow Characteristics and of Losses in Radial Nozzle Cascades

1984 ◽  
Vol 106 (2) ◽  
pp. 502-509 ◽  
Author(s):  
S. G. R. Hashemi ◽  
R. J. Lemak ◽  
J. A. Owczarek
Keyword(s):  
Flow Characteristics ◽  
Leading Edge ◽  
Injection Technique ◽  
Test Results ◽  
Test Rig ◽  
Water Test ◽  
Edge Vortex ◽  
Loss Coefficients ◽  
Radial Nozzle ◽  
Nozzle Blade

A study was made of the flow in radial nozzle cascades using an air test rig and a water test rig. In the air test rig, three cobra probes were used in circumferential and spanwise traverses to determine the total pressure variations in the flow field at three radii downstream of the nozzles at which static pressure was also measured. The tests were made on two sets of nozzle blades having heights of 0.148 in. (0.376 cm) and 0.200 in. (0.508 cm), at trailing edge angles (measured from circumferential direction) of 15, 20, and 25 deg, and at two flow Mach numbers of approximately 0.2 and 0.35. The test results presented in this paper, in the form of loss coefficients and flow angles, were flow-weighted and averaged. Flow visualization in the air test rig was made on the walls bounding the nozzle blades using the graphite power-oil mixture technique. Additional tests were made on the water test rig using dye injection technique. Photographs were obtained showing clearly formation of secondary flow around each nozzle blade in the form of the leading edge vortex. The test results confirm the existence of the leading edge vortices reported peviously, and extend their study to the radial nozzle cascades.

Flow Instabilities in Gas Turbine Chute Seals

2020 ◽  
Vol 142 (2) ◽  
Author(s):  
Joshua T. M. Horwood ◽  
Fabian P. Hualca ◽  
Mike Wilson ◽  
James A. Scobie ◽  
Carl M. Sangan ◽  
...  
Keyword(s):  
Large Scale ◽  
Computational Study ◽  
Leading Edge ◽  
Large Scale Structures ◽  
Strong Shear ◽  
Time Marching ◽  
Flow Through ◽  
Scale Structures ◽  
Large Scale Flow ◽  
Secondary Flow Field

Abstract The ingress of hot annulus gas into stator–rotor cavities is an important topic to engine designers. Rim-seals reduce the pressurized purge required to protect highly stressed components. This paper describes an experimental and computational study of flow through a turbine chute seal. The computations—which include a 360 deg domain—were undertaken using dlrtrace's time-marching solver. The experiments used a low Reynolds number turbine rig operating with an engine-representative flow structure. The simulations provide an excellent prediction of cavity pressure and swirl, and good overall agreement of sealing effectiveness when compared to experiment. Computation of flow within the chute seal showed strong shear gradients which influence the pressure distribution and secondary-flow field near the blade leading edge. High levels of shear across the rim-seal promote the formation of large-scale structures at the wheel-space periphery; the number and speed of which were measured experimentally and captured, qualitatively and quantitatively, by computations. A comparison of computational domains ranging from 30 deg to 360 deg indicates that steady features of the flow are largely unaffected by sector size. However, differences in large-scale flow structures were pronounced with a 60 deg sector and suggest that modeling an even number of blades in small sector simulations should be avoided.

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