Flow Measurements in a Fishtail Diffuser With Strong Curvature

1999 ◽  
Vol 121 (2) ◽  
pp. 410-417 ◽  
Author(s):  
M. I. Yaras
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Cross Sectional ◽  
Inlet Flow ◽  
Inlet Conditions ◽  
Strong Curvature

The paper presents detailed measurements of the incompressible flow development in a large-scale 90 deg curved diffuser with strong curvature and significant streamwise variation in cross-sectional aspect ratio. The flow path approximates the so-called fishtail diffuser utilized on small gas turbine engines for the transition between the centrifugal impeller and the combustion chamber. Two variations of the inlet flow, differing in boundary layer thickness and turbulence intensity, are considered. Measurements consist of three components of velocity, static pressure and total pressure distributions at several cross-sectional planes throughout the diffusing bend. The development and mutual interaction of multiple pairs of streamwise vortices, redistribution of the streamwise flow under the influence of these vortices, the resultant streamwise variations in mass-averaged total-pressure and static pressure, and the effect of inlet conditions on these aspects of the flow are examined. The strengths of the vortical structures are found to be sensitive to the inlet flow conditions, with the inlet flow comprising a thinner boundary layer and lower turbulence intensity yielding stronger secondary flows. For both inlet conditions a pair of streamwise vortices develop rapidly within the bend, reaching their peak strength at about 30 deg into the bend. The development of a second pair of vortices commences downstream of this location and continues for the remainder of the bend. Little evidence of the first vortex pair remains at the exit of the diffusing bend. The mass-averaged total pressure loss is found to be insensitive to the range of inlet-flow variations considered herein. However, the rate of generation of this loss along the length of the diffusing bend differs between the two test cases. For the case with the thinner inlet boundary layer, stronger secondary flows result in larger distortion of the streamwise velocity field. Consequently, the static pressure recovery is somewhat lower for this test case. The difference between the streamwise distributions of measured and ideal static pressure is found to be primarily due to total pressure loss in the case of the thick inlet boundary layer. For the thin inlet boundary layer case, however, total pressure loss and flow distortion are observed to influence the pressure recovery by comparable amounts.

Secondary Flow Control Using Vortex Generator Jets

2004 ◽  
Vol 126 (4) ◽  
pp. 650-657 ◽  
Author(s):  
R. K. Sullerey ◽  
A. M. Pradeep
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Vortex Generator ◽  
Control Flow ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Vortex Generators ◽  
Vortex Generator Jets ◽  
Inflow Conditions

In this paper, results are presented of an experimental investigation into the effectiveness of vortex generator jets in controlling secondary flows in two-dimensional S-duct diffusers. The experiments were performed in uniform and distorted inflow conditions and the performance evaluation of the diffuser was carried out in terms of static pressure recovery and quality of the exit flow. In the case with inflow distortion, tapered fin vortex generators were employed in addition to vortex generator jets to control flow separation that was detected on the wall with inflow distortion. Detailed measurements including total pressure, velocity distribution, surface static pressure, skin friction, and boundary layer measurements were taken at a Reynolds number of 7.8×105. These results are presented in terms of static pressure rise, distortion coefficient, and total pressure loss coefficient at the duct exit. For uniform inflow, the use of vortex generator jets resulted in more than a 30 percent decrease in total pressure loss and flow distortion coefficients. In combination with passive device (tapered fin vortex generators), the vortex generator jets reduce total pressure losses by about 25 percent for distorted inflow conditions. A potential application of this method may include control of secondary flows in turbo machinery.

A Comprehensive Investigation of Blade Row Interaction Effects on Stator Loss Utilizing Vane Clocking

2017 ◽  
Author(s):  
Natalie R. Smith ◽  
Nicole L. Key
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Transition ◽  
Total Pressure Loss ◽  
Tip Leakage Flow ◽  
Inlet Flow ◽  
Blade Row ◽  
Vane Clocking

Blade row interactions drive the unsteady performance of high pressure compressors. Vane clocking is the relative circumferential positioning of consecutive stationary vane rows with the same vane count. By altering the upstream vane wake’s path with respect to the downstream vane, vane clocking changes the blade row interactions and results in a change in steady total pressure loss on the downstream vane. The open literature lacks a conclusive discussion of the flow physics governing these interactions in compressors. This paper presents the details of a comprehensive vane clocking study on the embedded stage of the Purdue 3-stage axial compressor. The steady loss results, including radial total pressure profiles and surface flow visualization, suggest a shift in the Stator 2 corner separations occurs between clocking configurations associated with the maximum and minimum total pressure loss. To better understand the flow mechanisms driving the vane clocking effects on the steady Stator 2 performance, time-resolved interrogations of the Stator 2 inlet flow field, surface pressure unsteadiness, and boundary layer response were conducted. The Stator 2 surface flows, both pressure unsteadiness and boundary layer transition, are influenced by vane clocking and interactions between Rotor 1 and Rotor 2, but neither of these results indicate a cause for the change in steady total pressure loss. Moreover, they are a result of upstream changes in the flow field: the interaction between the Stator 1 wake and Rotor 2 results in a circumferentially varying pattern which alters the inlet flow field for the downstream row, including the unsteadiness and frequency content in the tip and hub regions. Therefore, under different clocking configurations, Stator 2 experiences significantly different inlet blockage and unsteadiness from the Rotor 2 tip leakage flow and hub corner separation, which, in turn, shifts the radial blade loading distribution and subsequent loss development of Stator 2.

scholarly journals Numerical Investigation of flow through Annular diffusing duct

2011 ◽  
pp. 60-63
Author(s):  
Prasanta K. Sinha ◽  
Ananta Kumar Das ◽  
Bireswar Majumdar
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Line Length ◽  
Area Ratio ◽  
Experimental Results ◽  
Total Pressure Loss ◽  
Pressure Recovery ◽  
Mean Velocity ◽  
Inlet Conditions

In the present investigation the distribution of mean velocity, static pressure and total pressure are experimentally studied on an annular curved diffuser of 30° angle of turn with an area ratio of 1.283 and centerline length was chosen as three times of inlet diameter. The experimental results then were numerically validated with the help of Fluent and then a series of parametric investigations are conducted with same centre line length and inlet diameter but with different area ratios varying from 1.15 to 3.75. The measurements were taken at Reynolds number 2.25 x 105 based on inlet diameter and mass average inlet velocity. Predicted results of coefficient of mass averaged static pressure recovery (30%) and coefficient of mass averaged total pressure loss (21%) are in good agreement with the experimental results of coefficient of mass averaged static pressure recovery (26%) and coefficient of mass averaged total pressure loss (17%) respectively. Standard k-ε model in Fluent solver was chosen for validation. From the parametric investigation it is observed that static pressure recovery increases up to an area ratio of 2.86 and between the area ration 2.86 to 3.75, pressure recovery decreases steadily. The coefficient of total pressure loss almost remains constant with the change in area ratio for similar inlet conditions.

Numerical Investigation of Dump Diffuser Combustor Performance at Uniform and Non-Uniform Inlet Conditions

2020 ◽  
Author(s):  
Heyu Wang ◽  
Kai Hong Luo
Keyword(s):  
Boundary Layer ◽  
Numerical Investigation ◽  
Layer Thickness ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Thickness ◽  
Static Pressure ◽  
Aerodynamic Performance ◽  
Aero Engine ◽  
Inlet Conditions

Abstract A numerical investigation has been conducted for an axisymmetric dump diffuser combustor, which is a simplified geometry of a typical lean-burn combustor in a modern civil aero-engine gas turbine. The aerodynamic performance of the combustor is analyzed with an emphasis on two common performance parameters: static pressure recovery and total pressure loss. The former is essential in maintaining high-pressure air flow across the liner, whereas the latter involves the specific fuel consumption of the aero-engine. At first, the effects of geometrical parameters of the dump diffuser combustor are investigated. A high diffuser angle seems to be detrimental to both static pressure recovery and total pressure loss. On the other hand, a high dump gap ratio is beneficial from the aerodynamic performance point of view. However, all these desired characteristics are subject to mechanical constraints and their implications for specific consumption. Optimum values of those parameters should exist for a given desired aerodynamics performance. The majority of previous researches, including the first part of this study, have been carried out with uniform inlet conditions due to a typical independent design cycle of each component. The effects of compressor exit conditions are usually not considered in the early stage design process. In the second part of this study, various inlet conditions representing a more realistic compressor exit condition such as inlet symmetrical and asymmetrical boundary layer thickness are investigated. The performance of an asymmetrical configuration with a thin boundary layer thickness near the outer annulus is almost comparable to that of its uniform counterpart. Findings of this study provide useful input for combustor designers to improve the combustor’s performance based on the compressor exit conditions.

A Comprehensive Investigation of Blade Row Interaction Effects on Stator Loss Utilizing Vane Clocking

2018 ◽  
Vol 140 (7) ◽  
Author(s):  
Natalie R. Smith ◽  
Nicole L. Key
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Transition ◽  
Total Pressure Loss ◽  
Tip Leakage Flow ◽  
Inlet Flow ◽  
Blade Row ◽  
Vane Clocking

Blade row interactions drive the unsteady performance of high-pressure compressors. Vane clocking is the relative circumferential positioning of consecutive stationary vane rows with the same vane count. By altering the upstream vane wake's path with respect to the downstream vane, vane clocking changes the blade row interactions and results in a change in steady total pressure loss on the downstream vane. The open literature lacks a conclusive discussion of the flow physics governing these interactions in compressors. This paper presents the details of a comprehensive vane clocking study on the embedded stage of the Purdue three-stage axial compressor. The steady loss results, including radial total pressure profiles and surface flow visualization, suggest a shift in the stator 2 corner separations occurs between clocking configurations associated with the maximum and minimum total pressure loss. To better understand the flow mechanisms driving the vane clocking effects on the steady stator 2 performance, time-resolved interrogations of the stator 2 inlet flow field, surface pressure unsteadiness, and boundary layer response were conducted. The stator 2 surface flows, both pressure unsteadiness and boundary layer transition, are influenced by vane clocking and interactions between rotor 1 and rotor 2, but neither of these results indicate a cause for the change in steady total pressure loss. Moreover, they are a result of upstream changes in the flow field: the interaction between the stator 1 wake and rotor 2 results in a circumferentially varying pattern which alters the inlet flow field for the downstream row, including the unsteadiness and frequency content in the tip and hub regions. Therefore, under different clocking configurations, stator 2 experiences significantly different inlet blockage and unsteadiness from the rotor 2 tip leakage flow and hub corner separation, which, in turn, shifts the radial blade loading distribution and subsequent loss development of stator 2.

Second Vane Total Pressure Loss Due to Endwall Iceform Contouring

2008 ◽  
Author(s):  
Ronald S. LaFleur
Keyword(s):  
Heat Transfer ◽  
Secondary Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Drag ◽  
Pressure Coefficient ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Surface Shape ◽  
Exit Plane

The iceformation design method generates an endwall contour, altering the secondary flows that produce elevated endwall heat transfer load and total pressure losses. Iceformation is an analog to regions of metal melting where a hot fluid alters the isothermal surface shape of a part as it is maintained by a cooling fluid. The passage flow, heat transfer and geometry evolve together under the constraints of flow and thermal boundary conditions. The iceformation concept is not media dependent and can be used in analogous flows and materials to evolve novel boundary shapes. In the past, this method has been shown to reduce aerodynamic drag and total pressure loss in flows such as diffusers and cylinder/endwall junctures. A prior paper [1] showed that the Reynolds number matched iceform geometry had a 24% lower average endwall heat transfer than the rotationally symmetric endwall geometry of the Energy Efficiency Engine (E3). Comparisons were made between three endwall geometries: the ‘iceform’, the ‘E3’ and the ‘flat’ as a limiting case of the endwall design space. This paper adds to the iceformation design record by reporting the endwall aerodynamic performances. Second vane exit flow velocities and pressures were measured using an automated 2-D traverse of a 1.2 mm diameter five-hole probe. Exit plane maps for the three endwall geometries are presented showing the details of the total pressure coefficient contours and the velocity vectors. The formation of secondary flow vortices is shown in the exit plane and this results in an impact on exit plane total pressure loss distribution, off-design over- and under-turning of the exit flow. The exit plane contours are integrated to form overall measures of the total pressure loss. Relative to the E3 endwall, the iceform endwall has a slightly higher total pressure loss attributed to higher dissipation of the secondary flow within the passage. The iceform endwall has a closer-to-design exit flow pattern than the E3 endwall.

The Influence of Aero-Derivative Combustor Turbulence and Reynolds Number on Vane Aerodynamics Losses, Secondary Flows, and Wake Growth

2006 ◽  
Author(s):  
F. E. Ames ◽  
J. D. Johnson ◽  
N. J. Fiala
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Pressure Loss ◽  
Total Pressure ◽  
Reynolds Numbers ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Surface Boundary ◽  
High Turbulence ◽  
Turbulence Condition

Exit surveys detailing total pressure loss, turning angle, and secondary velocities have been acquired for a fully loaded vane profile in a large scale low speed cascade facility. Exit surveys have been taken over a four-to-one range in Reynolds numbers based on exit conditions and for both a low turbulence condition and a high turbulence condition. The high turbulence condition was generated using a mock aero-derivative combustor. Exit loss, angle, and secondary velocity measurements were acquired in the facility using a five-hole cone probe at two stations representing axial chord spacings of 0.25 and 0.50. Substantial differences in the level of losses, distribution of losses, and secondary flow vectors are seen with the different turbulence conditions and at the different Reynolds numbers. The higher turbulence condition produces a significantly broader wake than the low turbulence case and shows a measurable total pressure loss in the region outside the wakes. Generally, total pressure losses are about 0.02 greater for the high turbulence case compared with the low turbulence case primarily due to the state of the suction surface boundary layers. Losses decrease moderately with increasing Reynolds number. Cascade inlet velocity distributions have been previously documented in an endwall heat transfer study of this same geometry. These exit survey measurements support our understanding of the endwall heat transfer distributions, the secondary flows in the passage, and the origin of losses.

Location Effect of Boundary Layer Suction on Compressor Hub-Corner Separation

2014 ◽  
Author(s):  
Ping-Ping Chen ◽  
Wei-Yang Qiao ◽  
Karsten Liesner ◽  
Robert Meyer
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Base Flow ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Corner Separation ◽  
Boundary Layer Suction

The large secondary flow area in the compressor hub-corner region usually leads to three-dimensional separation in the passage with large amounts of total pressure loss. In this paper numerical simulations of a linear high-speed compressor cascade, consisting of five NACA 65-K48 stator profiles, were performed to analyze the flow mechanism of hub-corner separation for the base flow. Experimental validation is used to verify the numerical results. Active control of the hub-corner separation was investigated by using boundary layer suction. The influence of the selected locations of the endwall suction slot was investigated in an effort to quantify the gains of the compressor cascade performance. The results show that the optimal chordwise location should contain the development section of the three-dimensional corner separation downstream of the 3D corner separation onset. The best pitchwise location should be close enough to the vanes’ suction surface. Therefore the optimal endwall suction location is the MTE slot, the one from 50% to 75% chord at the hub, close to the blade suction surface. By use of the MTE slot with 1% suction flow ratio, the total-pressure loss is substantially decreased by about 15.2% in the CFD calculations and 9.7% in the measurement at the design operating condition.

Numerically Optimizing an Annular Diffuser Using a Genetic Algorithm With Three Objectives

2012 ◽  
Author(s):  
David J. Cerantola ◽  
A. M. Birk
Keyword(s):  
Genetic Algorithm ◽  
Pressure Loss ◽  
Total Pressure ◽  
Exhaust System ◽  
Outer Wall ◽  
Total Pressure Loss ◽  
Pressure Recovery ◽  
Annular Diffuser ◽  
Inlet Conditions ◽  
Wall Profile

A genetic algorithm was implemented to determine preferential solutions of a short annular diffuser exhaust system of length 1.5Do (outer annulus diameters). Five free variables defined the centre body shape and two variables determined the outer wall profile. Diffuser performance was evaluated using three objectives—(i) diffuser pressure recovery, (ii) outlet velocity uniformity, and (iii) total pressure loss—that were calculated from steady state solutions obtained using the computational fluid dynamics software FLUENT 13.0 with the realizable k-ε turbulence model and enhanced wall treatment. Inlet conditions were ReDh = 8.5 × 104 and M = 0.23. After thirty-five generations, a paraboloid-shaped centre body with length 0.74Do and initial annular expansion of approximately 14° produced preferential solutions. A configuration with a converging outer wall above the centre body developed greater outlet flow uniformity and lower total pressure loss whereas a straight outer wall followed by the solid diffuser generated more static pressure recovery.

Flow Investigation Through a 30° Turn Diffusing Duct in Subsonic Flow Regime

2009 ◽  
Author(s):  
Prasanta K. Sinha ◽  
Biswajit Haldar ◽  
Amar N. Mullick ◽  
Bireswar Majumdar
Keyword(s):  
Axial Velocity ◽  
Pressure Loss ◽  
Total Pressure ◽  
High Speed ◽  
Static Pressure ◽  
Transverse Velocity ◽  
Total Pressure Loss ◽  
Pressure Recovery ◽  
Pressure Probe ◽  
Mean Velocity

Curved diffusers are an integral component of the gas turbine engines of high-speed aircraft. These facilitate effective operation of the combustor by reducing the total pressure loss. The performance characteristics of these diffusers depend on their geometry and the inlet conditions. In the present investigation the distribution of axial velocity, transverse velocity, mean velocity, static and total pressures are experimentally studied on a curved diffuser of 30° angle of turn with an area ratio of 1.27. The centreline length was chosen as three times of inlet diameter. The experimental results then were numerically validated with the help of Fluent, the commercial CFD software. The measurements of axial velocity, transverse velocity, mean velocity, static pressure and total pressure distribution were taken at Reynolds number 1.9 × 105 based on inlet diameter and mass average inlet velocity. The mean velocity and all the three components of mean velocity were measured with the help of a pre-calibrated five-hole pressure probe. The velocity distribution shows that the flow is symmetrical and uniform at the inlet and exit sections and high velocity cores are accumulated at the top concave surface due to the combined effect of velocity diffusion and centrifugal action. It also indicates the possible development of secondary motions between the concave and convex walls of the test diffuser. The mass average static pressure recovery and total pressure loss within the curved diffuser increases continuously from inlet to exit and they attained maximum values of 35% and 14% respectively. A comparison between the experimental and predicated results shows a good qualitative agreement between the two. Standard k-ε model in Fluent solver was chosen for validation. It has been observed that coefficient of pressure recovery Cpr for the computational investigation was obtained as 38% compared to the experimental investigation which was 35% and the coefficient of pressure loss is obtained as 13% in computation investigation compared to the 14% in experimental study, which indicates a very good qualitative matching.

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