scholarly journals Experimental and Computational Study of Hybrid Diffusers for Gas Turbine Combustors

2003 ◽  
Author(s):  
A. Duncan Walker ◽  
Paul A. Denman ◽  
James J. McGuirk
Keyword(s):  
Gas Turbine ◽  
Computational Study ◽  
Detailed Examination ◽  
Momentum Equation ◽  
Finite Limit ◽  
Recovery Coefficient ◽  
Pressure Loss Coefficient ◽  
Radial Depth ◽  
Flow Mechanisms ◽  
Total Pressure Loss Coefficient

The increasing radial depth of modern combustors poses a particularly difficult aerodynamic challenge for the prediffuser. Conventional diffuser systems have a finite limit to the diffusion that can be achieved in a given length and it is, therefore, necessary for designers to consider more radical and unconventional diffuser configurations. This paper will report on one such unconventional diffuser; the hybrid diffuser which, under the action of bleed, has been shown to achieve high rates of diffusion in relatively short lengths. However, previous studies have not been conducted under representative conditions and have failed to provide a complete description of the relevant flow mechanisms making optimisation difficult. Utilising an isothermal representation of a modern gas turbine combustor an experimental investigation was undertaken to study the performance of a hybrid diffuser compared to that of a conventional, single passage, dump diffuser system. The hybrid diffuser achieved a 53% increase in area ratio within the same axial length generating a 13% increase in the pre-diffuser static pressure recovery coefficient which, in turn, produced a 25% reduction in the combustor feed annulus total pressure loss coefficient. A computational investigation was also undertaken in order to investigate the governing flow mechanisms. A detailed examination of the flow field, including an analysis of the terms within the momentum equation, demonstrated that the controlling flow mechanisms were not simply a boundary layer bleed but involve a more complex interaction between the accelerating bleed flow and the diffusing mainstream flow. A greater understanding of these mechanisms enabled a more practical design of hybrid diffuser to be developed that not only simplified the geometry but also improved the quality of the bleed air making it more attractive for use in component cooling.

scholarly journals Experimental and Computational Study of Hybrid Diffusers for Gas Turbine Combustors

2004 ◽  
Vol 126 (4) ◽  
pp. 717-725 ◽  
Author(s):  
A. Duncan Walker ◽  
Paul A. Denman ◽  
James J. McGuirk
Keyword(s):  
Gas Turbine ◽  
Computational Study ◽  
Detailed Examination ◽  
Momentum Equation ◽  
Finite Limit ◽  
Recovery Coefficient ◽  
Pressure Loss Coefficient ◽  
Radial Depth ◽  
Flow Mechanisms ◽  
Total Pressure Loss Coefficient

The increasing radial depth of modern combustors poses a particularly difficult aerodynamic challenge for the pre-diffuser. Conventional diffuser systems have a finite limit to the diffusion that can be achieved in a given length and it is, therefore, necessary for designers to consider more radical and unconventional diffuser configurations. This paper will report on one such unconventional diffuser; the hybrid diffuser which, under the action of bleed, has been shown to achieve high rates of diffusion in relatively short lengths. However, previous studies have not been conducted under representative conditions and have failed to provide a complete description of the relevant flow mechanisms making optimization difficult. Utilizing an isothermal representation of a modern gas turbine combustor an experimental investigation was undertaken to study the performance of a hybrid diffuser compared to that of a conventional, single-passage, dump diffuser system. The hybrid diffuser achieved a 53% increase in area ratio within the same axial length generating a 13% increase in the pre-diffuser static pressure recovery coefficient which, in turn, produced a 25% reduction in the combustor feed annulus total pressure loss coefficient. A computational investigation was also undertaken in order to investigate the governing flow mechanisms. A detailed examination of the flow field, including an analysis of the terms within the momentum equation, demonstrated that the controlling flow mechanisms were not simply a boundary layer bleed but involve a more complex interaction between the accelerating bleed flow and the diffusing mainstream flow. A greater understanding of these mechanisms enabled a more practical design of hybrid diffuser to be developed that not only simplified the geometry but also improved the quality of the bleed air making it more attractive for use in component cooling.

Film Cooling and Aerodynamic Performance on Multi-Cavity Squealer Tip of a Turbine Blade

2021 ◽  
Author(s):  
Feng Li ◽  
Zhao Liu ◽  
Zhenping Feng
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Cooling Performance ◽  
Pressure Loss Coefficient ◽  
Blade Tip ◽  
Total Pressure Loss Coefficient

Abstract The blade tip region of the shroud-less high-pressure gas turbine is exposed to an extremely operating condition with combined high temperature and high heat transfer coefficient. It is critical to design new tip structures and apply effective cooling method to protect the blade tip. Multi-cavity squealer tip has the potential to reduce the huge thermal loads and improve the aerodynamic performance of the blade tip region. In this paper, numerical simulations were performed to predict the aerothermal performance of the multi-cavity squealer tip in a heavy-duty gas turbine cascade. Different turbulence models were validated by comparing to the experimental data. It was found that results predicted by the shear-stress transport with the γ-Reθ transition model have the best precision. Then, the film cooling performance, the flow field in the tip gap and the leakage losses were presented with several different multi-cavity squealer tip structures, under various coolant to mainstream mass flow ratios (MFR) from 0.05% to 0.15%. The results show that the ribs in the multi-cavity squealer tip could change the flow structure in the tip gap for that they would block the coolant and the leakage flow. In this study, the case with one-cavity (1C) achieves the best film cooling performance under a lower MFR. However, the cases with multi-cavity (2C, 3C, 4C) show higher film cooling effectiveness under a higher MFR of 0.15%, which are 32.6%%, 34.2%% and 41.0% higher than that of the 1C case. For the aerodynamic performance, the case with single-cavity has the largest total pressure loss coefficient in all MFR studied, whereas the case with two-cavity obtains the smallest total pressure loss coefficient, which is 7.6% lower than that of the 1C case.

A New Concept of a Two-Dimensional Supersonic Relative Inlet Mach Number Compressor Cascade

2009 ◽  
Author(s):  
Toyotaka Sonoda ◽  
Markus Olhofer ◽  
Toshiyuki Arima ◽  
Bernhard Sendhoff
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Optimization Method ◽  
Superior Performance ◽  
Compressor Cascade ◽  
Pressure Loss Coefficient ◽  
Supersonic Inlet ◽  
Flow Mechanisms ◽  
Improved Performance ◽  
Total Pressure Loss Coefficient

In this study, a numerical shape optimization method based on evolutionary algorithms coupled with a verified CFD solver has been applied to the ambitious target of a shock free 2-D supersonic inlet Mach number compressor cascade. The study is based on the DLR-PAV-1.5 supersonic compressor cascade designed by the pre-compression blading concept. The DLR cascade airfoil has been optimized using a verified CFD code. A superior performance of the optimized supersonic cascade with about 24% reduction of the total pressure loss coefficient compared to the original cascade has been realized. The flow mechanisms observable around the blade with improved performance and the resulting design concept are discussed in this paper.

Effects of casing angle on the performance of parallel hub axial annular diffuser

2020 ◽  
Vol 0 (0) ◽  
Author(s):  
Hardial Singh ◽  
B.B. Arora
Keyword(s):  
Static Pressure ◽  
Swirling Flow ◽  
Pressure Recovery ◽  
Recovery Coefficient ◽  
Pressure Loss Coefficient ◽  
Swirl Angle ◽  
Annular Diffuser ◽  
Inlet Swirl ◽  
Total Pressure Loss Coefficient ◽  
Pressure Recovery Coefficient

Abstract In this paper, the effects of non-swirling and swirling flow on the performance of parallel hub axial annular diffuser has been investigated. The study was conducted on a fully developed swirling flow and non-swirling flow to predict the separation of the flow from the wall. Three different annular diffusers were used with casing wall angles of 3°, 6°, and 9°. Furthermore, various swirl angles (0–25°) at the inlet of diffusers have been investigated to analyze the performance across the length. It was found that parallel hub axial annular diffuser performance increases up to a certain length as the inlet swirl angle increases. However, the performance also improves as the diffuser area ratio (AR) increases. The performance is evaluated based on the static pressure recovery coefficient (Cp) and the total pressure loss coefficient (CTL). The highest possible pressure recovery is achieved by the 12° swirl angle with a casing angle of 6°.

scholarly journals Effects of casing angle on the performance of parallel hub axial annular diffuser

2020 ◽  
Vol 0 (0) ◽  
Author(s):  
Hardial Singh ◽  
B.B. Arora
Keyword(s):  
Static Pressure ◽  
Swirling Flow ◽  
Pressure Recovery ◽  
Recovery Coefficient ◽  
Pressure Loss Coefficient ◽  
Swirl Angle ◽  
Annular Diffuser ◽  
Inlet Swirl ◽  
Total Pressure Loss Coefficient ◽  
Pressure Recovery Coefficient

AbstractIn this paper, the effects of non-swirling and swirling flow on the performance of parallel hub axial annular diffuser has been investigated. The study was conducted on a fully developed swirling flow and non-swirling flow to predict the separation of the flow from the wall. Three different annular diffusers were used with casing wall angles of 3°, 6°, and 9°. Furthermore, various swirl angles (0–25°) at the inlet of diffusers have been investigated to analyze the performance across the length. It was found that parallel hub axial annular diffuser performance increases up to a certain length as the inlet swirl angle increases. However, the performance also improves as the diffuser area ratio (AR) increases. The performance is evaluated based on the static pressure recovery coefficient (Cp) and the total pressure loss coefficient (CTL). The highest possible pressure recovery is achieved by the 12° swirl angle with a casing angle of 6°.

Influence of Inlet Shape on Twin Intake Duct Performance

2005 ◽  
Author(s):  
K. Saha ◽  
S. N. Singh ◽  
V. Seshadri
Keyword(s):  
Static Pressure ◽  
Cross Flow ◽  
Geometrical Parameters ◽  
Recovery Coefficient ◽  
Pressure Loss Coefficient ◽  
Secondary Motion ◽  
Vortical Motion ◽  
Total Pressure Loss Coefficient ◽  
Pressure Recovery Coefficient ◽  
Oval Shape

Performance of twin intake ducts with different inlet shapes has been analyzed using a commercial CFD (Computational Fluid Dynamics) code. The performance has been evaluated for incompressible flow at a fixed Reynolds number (1.4×105). The shapes studied are rectangular (Aspect ratio = 2), square, semicircular, elliptic-1, elliptic-2 (inverse-elliptic) and oval shape with all the other geometrical parameters remaining same. The performance of the twin intake ducts in terms of the static pressure recovery coefficient, total pressure loss coefficient and distortion coefficient, and the secondary motion at the merging plane and the downstream planes show that the inverse elliptic shape is the best followed by semi-circular inlet. The vectors plots of secondary motion at the merging plane and downstream have shown the presence of twin pairs of vortical motion possibly caused by the change in centerline curvature. The cross flow vector plots also show four distinct vortices after merger.

Numerical Simulation of Burner for Micro Gas Turbine

2012 ◽  
Vol 569 ◽  
pp. 51-55
Author(s):  
Lei Jia ◽  
Shi Liu ◽  
Yao Song Huang ◽  
Neng Wang ◽  
Fu Zhen Wang ◽  
...  
Keyword(s):  
Numerical Simulation ◽  
Natural Gas ◽  
Gas Turbine ◽  
Flue Gas ◽  
Outlet Temperature ◽  
Mixing Process ◽  
Micro Gas Turbine ◽  
Pressure Loss Coefficient ◽  
The Stability ◽  
Total Pressure Loss Coefficient

In order to study affects of oxy fuel combustion on micro gas turbine ,three axial swirl burners with different installation angles for micro gas turbine were designed, flue gas recycle/oxy fuel was used to burn natural gas. Numerical simulation was used to study flow field and combustion conditions. The result shows that application of axial swirl burner promotes mixing process of natural gas and oxygen and recirculation brought about to promote the stability of fire, uniformity of outlet temperature was reduced. With the increase of swirl installation angle, backflow becomes more intense, and uniformity of outlet temperature becomes smaller, however, total pressure loss coefficient increased. These results will have a great significance in the design of better burners.

Performance characteristics of flow in annular diffuser using CFD

2021 ◽  
Vol 0 (0) ◽  
Author(s):  
Hardial Singh ◽  
Bharat Bhushan Arora
Keyword(s):  
Swirling Flow ◽  
Flow Behavior ◽  
Pressure Recovery ◽  
Ansys Fluent ◽  
Pressure Loss Coefficient ◽  
Swirl Angle ◽  
Swirl Intensity ◽  
Annular Diffuser ◽  
Inlet Swirl ◽  
Total Pressure Loss Coefficient

Abstract An annular diffuser is a critical component of the turbomachinery, and its prime function is to reduce the flow velocity. The current work is carried to study the effect of four different geometrical designs of an annular diffuser using the ANSYS Fluent. The numerical simulations were carried out to examine the effect of fully developed turbulent swirling and non-swirling flow. The flow behavior of the annular diffuser is analyzed at Reynolds number 2.5 × 105. The simulated results reveal pressure recovery improvement at the casing wall with adequate swirl intensity at the diffuser inlet. Swirl intensity suppresses the flow separation on the casing and moves the flow from the hub wall to the casing wall of the annulus region. The results also show that the Equal Hub and Diverging Casing (EHDC) annular diffuser in comparison to other diffusers has a higher static pressure recovery (C p  = 0.76) and a lower total pressure loss coefficient of (C L  = 0.12) at a 17° swirl angle.

Numerical Modeling of Heat Transfer and Pressure Losses for Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2007 ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

scholarly journals Detailed Examination of a Modified Two-Stage Micro Gas Turbine Combustor

2017 ◽  
Vol 140 (2) ◽  
Author(s):  
Andreas Schwärzle ◽  
Thomas O. Monz ◽  
Andreas Huber ◽  
Manfred Aigner
Keyword(s):  
Gas Turbine ◽  
Detailed Examination ◽  
Flame Stabilization ◽  
Nox Emissions ◽  
Gas Turbine Combustor ◽  
Two Stage ◽  
Micro Gas Turbine ◽  
Previous Design ◽  
Recirculation Zones ◽  
Asme Paper

Jet-stabilized combustion is a promising technology for fuel flexible, reliable, highly efficient combustion systems. The aim of this work is a reduction of NOx emissions of a previously published two-stage micro gas turbine (MGT) combustor (Zanger et al., 2015, “Experimental Investigation of the Combustion Characteristics of a Double-Staged FLOX-Based Combustor on an Atmospheric and a Micro Gas Turbine Test Rig,” ASME Paper No. GT2015-42313 and Schwärzle et al., 2016, “Detailed Examination of Two-Stage Micro Gas Turbine Combustor,” ASME Paper No. GT2016-57730), where the pilot stage (PS) of the combustor was identified as the main contributor to NOx emissions. The geometry optimization was carried out regarding the shape of the pilot dome and the interface between PS and main stage (MS) in order to prevent the formation of high-temperature recirculation zones. Both stages have been run separately to allow a detailed understanding of the flame stabilization within the combustor, its range of stable combustion, the interaction between both stages, and the influence of the modified geometry. All experiments were conducted at atmospheric pressure and an air preheat temperature of 650  °C. The flame was analyzed in terms of shape, length, and lift-off height, using OH* chemiluminescence (OH-CL) images. Emission measurements for NOx, CO, and unburned hydrocarbons (UHC) emissions were carried out. At a global air number of λ = 2, a fuel split variation was carried out from 0 (only PS) to 1 (only MS). The modification of the geometry leads to a decrease in NOx and CO emissions throughout the fuel split variation in comparison with the previous design. Regarding CO emissions, the PS operations are beneficial for a fuel split above 0.8. The local maximum in NOx emissions observed for the previous combustor design at a fuel split of 0.78 was not apparent for the modified design. NOx emissions were increasing, when the local air number of the PS was below the global air number. In order to evaluate the influence of the modified design on the flow field and identify the origin of the emission reduction compared to the previous design, unsteady Reynolds-averaged Navier–Stokes simulations were carried out for both geometries at fuel splits of 0.93 and 0.78, respectively, using the DLR (German Aerospace Center) in-house code turbulent heat release extension of the tau code (theta) with the k–ω shear stress transport turbulence model and the DRM22 (Kazakov and Frenklach, 1995, “DRM22,” University of California at Berkeley, Berkeley, CA, accessed Sept. 21, 2017, http://www.me.berkeley.edu/drm/) detailed reaction mechanism. The numerical results showed a strong influence of the recirculation zones on the PS reaction zone.

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