The Effect of Inlet Boundary Layer Thickness on the Flow Within an Annular S-Shaped Duct

1999 ◽  
Vol 121 (3) ◽  
pp. 626-634 ◽  
Author(s):  
T. Sonoda ◽  
T. Arima ◽  
M. Oana
Keyword(s):  
Boundary Layer ◽  
Flow Pattern ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Vortex Pair ◽  
Separated Flow ◽  
Total Pressure Loss ◽  
Centrifugal Impeller

Experimental and numerical investigations were carried out to gain a better understanding of the flow characteristics within an annular S-shaped duct, including the effect of the inlet boundary layer (IBL) on the flow. A duct with six struts and the same geometry as that used to connect compressor spools on our experimental small two-spool turbofan engine was investigated. A curved downstream annular passage with a similar meridional flow path geometry to that of the centrifugal compressor has been fitted at the exit of S-shaped duct. Two types of the IBL (i.e., thin and thick IBL) were used. Results showed that large differences of flow pattern were observed at the S-shaped duct exit between two types of the IBL, though the value of “net” total pressure loss has not been remarkably changed. According to “overall” total pressure loss, which includes the IBL loss, the total pressure loss was greatly increased near the hub as compared to that for a thin one. For the thick IBL, a vortex pair related to the hub-side horseshoe vortex and the separated flow found at the strut trailing edge has been clearly captured in the form of the total pressure loss contours and secondary flow vectors, experimentally and numerically. The high-pressure loss regions on either side of the strut wake near the hub may act on a downstream compressor as a large inlet distortion, and strongly affect the downstream compressor performance. There is a much-distorted three-dimensional flow pattern at the exit of S-shaped duct. This means that the aerodynamic sensitivity of S-shaped duct to the IBL thickness is very high. Therefore, sufficient care is needed to design not only downstream aerodynamic components (for example, centrifugal impeller) but also upstream aerodynamic components (LPC OGV).

scholarly journals The Effect of Inlet Boundary Layer Thickness on the Flow Within an Annular S-Shaped Duct

1998 ◽  
Author(s):  
Toyotaka Sonoda ◽  
Toshiyuki Arima ◽  
Mineyasu Oana
Keyword(s):  
Boundary Layer ◽  
Flow Pattern ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Vortex Pair ◽  
Separated Flow ◽  
Total Pressure Loss ◽  
Centrifugal Impeller

Experimental and numerical investigations were carried out to gain a better understanding of the flow characteristics within an annular S-shaped duct, including the effect of the inlet boundary layer (IBL) on the flow. A duct with six struts and the same geometry as that used to connect compressor spools on our experimental small two-spool turbofan engine was investigated. A curved downstream annular passage with a similar meridional flow path geometry to that of the centrifugal compressor has been fitted at the exit of S-shaped duct. Two types of the IBL (i.e. thin and thick IBL) were used. Results showed that large differences of flow pattern were observed at the S-Shaped duct exit between two types of the IBL, though the value of “net” total pressure loss has not been remarkably changed. According to “overall” total pressure loss, which includes the IBL loss, the total pressure loss was greatly increased near the hub as compared to that for a thin one. For the thick IBL, a vortex pair related to the hub-side horseshoe vortex and the separated flow found at the strut trailing edge has been clearly captured in the form of the total pressure loss contours and secondary flow vectors, experimentally and numerically. The high-pressure loss regions on either side of the strut wake near the hub may act on a downstream compressor as a large inlet distortion, and strongly affect the downstream compressor performance. There is a much-distorted three-dimensional flow pattern at the exit of S-Shaped duct. This means that the aerodynamic sensitivity of S-Shaped duct to the IBL thickness is very high. Therefore, sufficient carefulness is needed to design not only downstream aerodynamic component (for example centrifugal impeller) but also upstream aerodynamic component (LPC OGV).

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.

scholarly journals The Influence of Downstream Passage on the Flow Within an Annular S-Shaped Duct

1997 ◽  
Author(s):  
Toyotaka Sonoda ◽  
Toshiyuki Arima ◽  
Mineyasu Oana
Keyword(s):  
High Pressure ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Horseshoe Vortex ◽  
Meridional Flow ◽  
Total Pressure Loss ◽  
Navier Stokes ◽  
Compressor Performance

Experimental and numerical investigations were carried out to gain a better understanding or the flow characteristics within an annular S-shaped duct, including the influence of the shape of the downstream passage located at the exit of the duct on the flow. A duct with six struts and the same geometry as that used to connect the compressor spools on our new experimental small two-spool turbofan engine was investigated. Two types of downstream passage were used. One type had a straight annular passage and the other a curved annular passage with a similar meridional flow path geometry to that of the centrifugal compressor. Results showed that the total pressure loss near the hub is large due to instability of the flow, as compared with that near the casing. Also, a vortex related to the horseshoe vortex was observed near the casing, in the case of the curved annular passage, the total pressure loss near the hub was greatly increased compared with the case of the straight annular passage, and the spatial position of the above vortex depends on the passage core pressure gradient. Furthermore, results of calculation using an in-house-developed three-dimensional Navier-Stokes code with a low Reynolds number k-ε turbulence model were in good qualitative agreement with experimental results. According to the simulation results, a region of very high pressure loss is observed near the hub at the duct exit with the increase of inlet boundary layer thickness. Such regions of high pressure loss may act on the downstream compressor as a large inlet distortion, and strongly affect downstream compressor performance.

The Influence of Downstream Passage on the Flow Within an Annular S-Shaped Duct

1998 ◽  
Vol 120 (4) ◽  
pp. 714-722 ◽  
Author(s):  
T. Sonoda ◽  
T. Arima ◽  
M. Oana
Keyword(s):  
High Pressure ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Horseshoe Vortex ◽  
Meridional Flow ◽  
Total Pressure Loss ◽  
Navier Stokes ◽  
Compressor Performance

Experimental and numerical investigations were carried out to gain a better understanding of the flow characteristics within an annular S-shaped duct, including the influence of the shape of the downstream passage located at the exit of the duct on the flow. A duct with six struts and the same geometry as that used to connect the compressor spools on our new experimental small two-spool turbofan engine was investigated. Two types of downstream passage were used. One type had a straight annular passage and the other a curved annular passage with a meridional flow path geometry similar to that of the centrifugal compressor. Results showed that the total pressure loss near the hub is large due to instability of the flow, as compared with that near the casing. Also, a vortex related to the horseshoe vortex was observed near the casing. In the case of the curved annular passage, the total pressure loss near the hub was greatly increased compared with the case of the straight annular passage, and the spatial position of this vortex depends on the passage core pressure gradient. Furthermore, results of calculation using an in-house-developed three-dimensional Navier–Stokes code with a low-Reynolds-number k–ε turbulence model were in good qualitative agreement with experimental results. According to the simulation results, a region of very high pressure loss is observed near the hub at the duct exit with the increase of inlet boundary layer thickness. Such regions of high pressure loss may act on the downstream compressor as a large inlet distortion, and strongly affect downstream compressor performance.

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.

Effects of axial length and integrated design on the aggressive intermediate turbine duct

2018 ◽  
Vol 233 (4) ◽  
pp. 443-456 ◽  
Author(s):  
Qingzong Xu ◽  
Pei Wang ◽  
Qiang Du ◽  
Jun Liu ◽  
Guang Liu
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Secondary Flow ◽  
Axial Length ◽  
Pressure Loss ◽  
Total Pressure ◽  
Guide Vane ◽  
Three Dimensional ◽  
Intermediate Turbine Duct ◽  
Aero Engine

With the increasing demand of high bypass ratio and thrust-to-weight ratio in civil aero-engine, the intermediate turbine duct between the high pressure and low pressure turbines of a modern gas turbine tends to shorter axial length, larger outlet-to-inlet area ratio and high pressure-to-low pressure radial offset. This paper experimentally and numerically investigated the three-dimensional flow characteristics of traditional (ITD1) and aggressive intermediate turbine duct (ITD2) at low Reynolds number. The baseline case of ITD1 is representative of a traditional intermediate turbine duct of aero-engine design with non-dimensional length of L/dR = 2.79 and middle angle of 20.12°. The detailed flow fields inside ITD1 and flow visualization were measured. Results showed the migration of boundary layer and a pair of counter-rotating vortexes were formed due to the radial migration of low momentum fluid. With the decreasing axial length of intermediate turbine duct, the radial and streamwise reverse pressure gradient in aggressive intermediate turbine duct (ITD2) were increased resulting in severe three-dimensional separation of boundary layer near casing surface and higher total pressure loss. The secondary flow and separation of boundary layer near the endwall were deeply analyzed to figure out the main source of high total pressure loss in the aggressive intermediate turbine duct (ITD2). Based on that, employing wide-chord guide vane to substitute “strut + guide vane”, this paper designed the super-aggressive intermediate turbine duct and realized the suppression of the three-dimensional separation and secondary flow.

Understanding of Secondary Loss Structures in a Multi-Stage Transonic Compressor Using DES

2016 ◽  
Author(s):  
Donghyun Kim ◽  
Changmin Son ◽  
Kuisoon Kim
Keyword(s):  
Flow Separation ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Total Pressure Loss ◽  
Suction Surface ◽  
Corner Region ◽  
Transonic Compressor ◽  
Multi Stage ◽  
Rans Analysis

In the present study, a multi-stage transonic compressor has been analyzed to investigate secondary loss structures and flow interactions in the corner region where the hub endwall and blade suction surface meet. The Detached Eddy Simulation (DES) approach is used successfully with the Shear Stress Transport (SST) turbulence model to directly resolve the eddy structure in the separated region. The SST-DES results for a transonic three stage axial compressor are compared with a RANS analysis obtained using ANSYS CFX. The present analysis indicates that the DES is better in simulating secondary losses and vortex structures than the RANS. With the DES, a large three-dimensional separation is predicted in the stator suction surface and hub endwall compared to the RANS prediction. The flow separation affects adversely the loss characteristics such as increases in the entropy and total pressure loss. The DES analysis indicates that the secondary flow phenomenon of the stator rows is apparent in all stages. It is observed to predict two distinct vortices induced by a three dimensional flow separation in the region adjacent to the suction surface and trailing edge of the last stage stator near the hub endwall. For the front two stages, the DES also predicts strong vortices and flow separation in the same corner region while the RANS analysis fails to predict them clearly. The total pressure loss prediction is concerned, the DES analysis predicts significantly larger than the RANS analysis in the region where the hub corner separation occurs. The DES is also found to predict a periodic fluctuations in the entropy, leading to the instantaneous efficiency variations with maximum differences of about 10% compared with the RANS solutions.

Numerical Investigation of the Influence of Real World Blade Profile Variations on the Aerodynamic Performance of Transonic Nozzle Guide Vanes

2011 ◽  
Vol 134 (2) ◽  
Author(s):  
R. Edwards ◽  
A. Asghar ◽  
R. Woodason ◽  
M. LaViolette ◽  
K. Goni Boulama ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Flow Pattern ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Guide Vanes ◽  
Trailing Edges ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

This paper addresses the issue of aerodynamic consequences of small variations in airfoil profile. A numerical comparison of flow field and cascade pressure losses for two representative repaired profiles and a reference new vane were made. Coordinates for the three airfoil profiles were obtained from the nozzle guide vanes of refurbished turboshaft engines using 3D optical scanning and digital modeling. The repaired profiles showed differences in geometry in comparison with the new vane, particularly near the leading and trailing edges. A numerical simulation was conducted using a commercial CFD code, which uses the finite volume approach for solving the governing equations. The computational predictions of the aerodynamic performance were compared with experimental results obtained from a cascade consisting of blades with the same airfoil profiles. The CFD analysis was performed for the cascade at subsonic inlet and transonic exit conditions. Boundary layer growth, wake formation, and shock boundary layer interactions were observed in the two-dimensional computations. The flow field showed the presence of shock waves downstream of the passage throat and near the trailing edges of the blades. A conspicuous change in flow pattern due to subtle variation in airfoil profile was observed. The calculated flow field was compared with the flow pattern visualized in the experimental test rig using the schlieren method. The total pressure calculation for the cascade exit showed an increase in pressure loss for one of the off-design profiles. The pressure loss calculations were also compared with the multihole total pressure probe measurement in the transonic cascade rig.

Effect of Casing Geometry on Flow Characteristics in a Model Can-Combustor

2012 ◽  
Author(s):  
Abdur Rahim ◽  
Dhirgham Alkhafagiy ◽  
Prabal Talukdar
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Numerical Study ◽  
Flow Characteristics ◽  
Total Pressure Loss ◽  
Stable Operation ◽  
Flow Uniformity ◽  
Flow Through ◽  
Combustor Liner

In a gas turbine combustor, it is necessary to use a diffuser to decelerate the high velocity air stream delivered by the compressor and thus avoid high total pressure loss. The interaction between the diffuser and combustor external flows plays a key role in controlling the pressure loss, air flow distribution around the combustor liner. Flow through casing-liner annulus is crucial as it feeds air to the primary, secondary and dilution holes. It is important that the annulus flow has sufficient static pressure to achieve adequate penetration of the jets. Moreover, the correct proportion of air enters the combustor liner through the dome and the various ports to maintain stable operation and good quality outlet condition. Length of combustor can be reduced if a provision is made for sufficient diffusion in the dump region. In the present numerical study, three can-combustor models of different geometry with a constant dump-gap have been analyzed with emphasis on the flow through annulus. A comparison has been made amongst the three models in terms of flow uniformity, static pressure recovery and total pressure loss. It is observed that flow uniformity in the annulus region is improved if a small divergence in length and a curved shape step height casing is made.

Accurate Estimation of Profile Losses and Analysis of Loss Generation Mechanisms in a Turbine Cascade

2017 ◽  
Vol 139 (12) ◽  
Author(s):  
D. Lengani ◽  
D. Simoni ◽  
M. Ubaldi ◽  
P. Zunino ◽  
F. Bertini ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Accurate Estimation ◽  
Loss Mechanism ◽  
Pressure Loss Coefficient ◽  
Generation Mechanisms ◽  
Total Pressure Loss Coefficient

The paper analyzes losses and the loss generation mechanisms in a low-pressure turbine (LPT) cascade by proper orthogonal decomposition (POD) applied to measurements. Total pressure probes and time-resolved particle image velocimetry (TR-PIV) are used to determine the flow field and performance of the blade with steady and unsteady inflow conditions varying the flow incidence. The total pressure loss coefficient is computed by traversing two Kiel probes upstream and downstream of the cascade simultaneously. This procedure allows a very accurate estimation of the total pressure loss coefficient also in the potential flow region affected by incoming wake migration. The TR-PIV investigation concentrates on the aft portion of the suction side boundary layer downstream of peak suction. In this adverse pressure gradient region, the interaction between the wake and the boundary layer is the strongest, and it leads to the largest deviation from a steady loss mechanism. POD applied to this portion of the domain provides a statistical representation of the flow oscillations by splitting the effects induced by the different dynamics. The paper also describes how POD can dissect the loss generation mechanisms by separating the contributions to the Reynolds stress tensor from the different modes. The steady condition loss generation, driven by boundary layer streaks and separation, is augmented in the presence of incoming wakes by the wake–boundary layer interaction and by the wake dilation mechanism. Wake migration losses have been found to be almost insensitive to incidence variation between nominal and negative (up to −9 deg) while at positive incidence, the losses have a steep increase due to the alteration of the wake path induced by the different loading distribution.

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