scholarly journals Three Dimensional Analysis of Advanced Swirl Vane/Nozzle Assemblies

1996 ◽  
Author(s):  
Gerald J. Micklow ◽  
Michael Benjamin
Keyword(s):  
Pressure Gradient ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Adverse Pressure Gradient ◽  
Total Pressure Loss ◽  
Linear Variation ◽  
Exit Velocity ◽  
Fuel Atomization ◽  
Vane Tip

The performance of high shear axial inflow/radial outflow airblast fuel injectors for advanced gas turbine combustors is highly dependent on the design of the swirl vanes. Curved vanes usually exhibit lower losses but straight vanes are also used due to lower cost and ease of manufacture. These type of vanes often operate under highly stalled conditions with high total pressure loss and a highly non-uniform exit velocity profile. This may produce poor fuel atomization with a non-uniform combustor fuel distribution resulting in lowered combustor efficiency and increased pollutant emissions. Properly designed vanes result in a greatly reduced total pressure loss. The exit velocity distribution is more uniform and higher in magnitude which can result in improved fuel atomization and distribution in the combustor. The present study investigates two curved swirler/nozzle shroud configurations operating at 1 and 10 atmospheres pressure for the same inlet temperature of 293°K. The first configuration was a twisted curved vane with thickness where the turning angle varied non-linearly from hub to tip with a maximum turning at the tip of 70 degrees. The second configuration was a curved vane with a linear variation of turning with 70 degrees turning at the tip. The results from a three dimensional viscous numerical flow simulation of these configurations shows similar performance for all cases investigated. The non-linear twisted vane however, had an approximately 3% higher mass flow rate than the vane with the linear variation in turning for the same exit static pressure at the hub. One problem which existed for all the conditions analyzed was a high loss region near the vane tip. This was due to the interaction with the shroud. As the flow exits the vane row and progresses along the nozzle outer lip, the flow area increases. This condition along with the streamline curvature effect of the outer nozzle lip causes an adverse pressure gradient to be formed in this region. This adverse pressure gradient causes the flow to separate from the vane suction surface. The problem initiated in the region of 70% span and increased in magnitude to the vane tip.

Three Dimensional Performance Prediction of Advanced Swirl Vanes for Gas Turbine Airblast Atomizers

1993 ◽  
Author(s):  
Gerald J. Micklow ◽  
Karthikeyan Shivaraman ◽  
H. Lee Nguyen
Keyword(s):  
Velocity Profile ◽  
Gas Turbine ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Total Pressure Loss ◽  
Pollutant Emissions ◽  
Exit Velocity ◽  
Fuel Atomization ◽  
Swirl Vane

The performance of high shear airblast fuel injectors for advanced gas turbine combustors is highly dependent on the design of the swirl vanes. The vanes may be of the straight or curved type. Curved vanes usually exhibit lower losses but straight vanes are also used due to lower cost and ease of manufacture. These type of vanes often operate under highly stalled conditions with high total pressure loss and a highly non-uniform exit velocity profile. This can produce poor fuel atomization with a non-uniform combustor fuel distribution resulting in lowered combustor efficiency and increased pollutant emissions. Constant turning curved vanes can also be easily manufactured. Properly designed vanes result in a greatly reduced total pressure loss. The exit velocity distribution is more uniform and higher in magnitude resulting in improved fuel atomization and distribution in the combustor. The present study compares standard helical flat vane performance with a low loss curved vane designed by the author for idle, takeoff and cruise conditions. The results from a three dimensional viscous numerical flow simulation show the curved swirl vane to be clearly superior to the standard flat helical swirl vane. The curved vane has a much lower total pressure loss with a more uniform exit velocity profile. This will result in improved combustor and engine performance and reduced pollutant emissions.

Shroud Influence on Gas Turbine Airblast Atomizer Swirler Flowfields

1995 ◽  
Author(s):  
Gerald J. Micklow ◽  
Karthikeyan Shivaraman ◽  
Insoo Cho
Keyword(s):  
Velocity Profile ◽  
Velocity Component ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Pollutant Emissions ◽  
Swirl Number ◽  
Exit Velocity ◽  
Fuel Atomization ◽  
Swirl Vane

The performance of high shear airblast fuel injectors for advanced gas turbine combustors is highly dependent on the design of the swirl vanes. The vanes may be of the straight or curved type. Curved vanes usually exhibit lower losses but straight vanes are also used due to lower cost and ease of manufacture. These type of vanes often operate under highly stalled conditions with high total pressure loss and a highly non-uniform exit velocity profile. This may produce poor fuel atomization with a non-uniform combustor fuel distribution resulting in lowered combustor efficiency and increased pollutant emissions. Constant turning curved vanes can also be easily manufactured. Properly designed vanes result in a greatly reduced total pressure loss. The exit velocity distribution is more uniform and higher in magnitude which can result in improved fuel atomization and distribution in the combustor. Further, the presence of a shroud is seen to have a major effect on the downstream flowfield. The present study compares standard helical flat vane performance with a low loss curved vane designed by the author for idle, and cruise conditions both with and without a shroud. The results from a three dimensional viscous numerical flow simulation show the curved swirl vane to be clearly superior to the standard flat helical swirl vane. The curved vane has a much lower total pressure loss with a more uniform exit velocity profile. This may result in improved combustor and engine performance and reduced pollutant emissions. The effect of the shroud was seen to reduce the size of the stall cell found in the vane passage for the helical vane. This resulted in a decrease in the magnitude of the axial velocity component in the outer vane passage and a decrease in the circumferential velocity component. This may result in a decrease in the swirl number. For the curved vane however, an increase in the magnitude of all velocity components was found which will result in a higher swirl number and better nozzle performance.

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.

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.

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).

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.

Performance Evaluation of Linear Turbine Cascades Using Three-Dimensional Viscous Flow Calculations

1985 ◽  
Vol 107 (4) ◽  
pp. 969-975 ◽  
Author(s):  
J. Moore ◽  
J. G. Moore
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Surface Flow ◽  
Total Pressure Loss ◽  
Suction Surface ◽  
Cross Sectional ◽  
High Loss ◽  
Flow Development ◽  
Turbine Cascades

The overall performance of two geometrically similar linear turbine cascades is calculated using an elliptic flow program. The increase in the mass-averaged total pressure loss is calculated within and downstream of the cascades and the results show good agreement with the measured values. The buildup and decay of the secondary kinetic energy are also shown; measurements are available for one of the cascades near and downstream of the trailing edge and these are in close agreement with the calculated values. Details of the flow development are also compared with measurements. Calculated velocity vectors near the endwall show the overturning revealed by surface flow visualization and similarly near the suction surface the strong spanwise flow is well calculated. Calculated contours of total pressure loss in cross-sectional planes confirm the important interaction of the passage vortex with the profile boundary layer at midspan. Regions of high loss near midspan are calculated downstream of both cascades; this three-dimensional flow development is followed in the calculations.

Non-Axisymmetric Endwall Contouring in a Compressor Cascade With Tip Gap

2014 ◽  
Author(s):  
Mahesh K. Varpe ◽  
A. M. Pradeep
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Axial Flow ◽  
Tip Clearance ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Flow Solver ◽  
Pressure Loss Coefficient ◽  
Flow Compressor

This paper describes the design of a non-axisymmetric hub contouring in a shroudless axial flow compressor cascade operating at near stall condition. Although, an optimum tip clearance reduces the total pressure loss, further minimization of the losses using hub contouring was achieved. The design methodology presented here combines an evolutionary principle with a three-dimensional CFD flow solver to generate different geometric profiles of the hub systematically. The total pressure loss coefficient was used as a single objective function to guide the search process for the optimum hub geometry. The resulting three dimensionally complex hub promises considerable benefits discussed in detail in this paper. A reduction of 15.2% and 16.23% in the total pressure loss and secondary kinetic energy, respectively, was achieved in the wake. The blade loading was observed to improve by about 4.53%. Other complementary benefits are also listed in the paper. The results confirm that non-axisymmetric contouring is an effective method for reducing the losses and thereby improving the performance of the cascade.

Characterization of a Supersonic Turbine Downstream of a Rotating Detonation Combustor

2018 ◽  
Author(s):  
Zhe Liu ◽  
James Braun ◽  
Guillermo Paniagua
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Oblique Shock ◽  
Total Pressure Loss ◽  
Flow Angle ◽  
Pressure Losses ◽  
Power Extraction ◽  
Inlet Conditions ◽  
Rotating Detonation

Rotating detonation combustors offer theoretically a significant total pressure increase, which may result in enhanced cycle efficiency. The fluctuating exhaust of rotating detonation combustors, however, induces low supersonic flow and large flow angle fluctuations at several kHz which affects the performance of the downstream turbine. For such flows, power extraction can be achieved by either integrating a diffuser with a conventional subsonic turbine or a nozzle with a supersonic turbine. In this paper, a numerical methodology is proposed to characterize a supersonic turbine exposed to fluctuations from rotating detonation combustors without any dilution. The inlet conditions of the turbine were extracted from a three dimensional unsteady Reynolds-Averaged Navier-Stokes simulation of a nozzle attached to a rotating detonation combustor, optimized for minimum flow fluctuations and a mass-flow averaged Mach number of 2 at the nozzle outlet. In a first step, a supersonic turbine able to handle steady Mach 2 inflow was designed based on a method of characteristics solver and total pressure loss was assessed. Afterwards unsteady simulations of eight stator passages exposed to periodic oblique shocks were performed. Total pressure loss was evaluated for several oblique shock frequencies and amplitudes. The unsteady stator outlet profile was extracted and used as inlet condition for the unsteady rotor simulations. Finally, a full stage unsteady simulation was performed to characterize the flow field across the entire turbine stage. Power extraction, airfoil base pressure, and total pressure losses were were assessed, which enabled the estimation of the loss mechanisms in supersonic turbine exposed to large unsteady inlet conditions. Frequency analysis of the pressure field across the turbine rows was used to evaluate the damping of the oblique shock waves.

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