Influence of Combustor Flow With Swirl on Integrated Combustor Vane Concept Full-Stage Performance

2017 ◽  
Vol 140 (1) ◽  
Author(s):  
Simon Jacobi ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Power Generation ◽  
Gas Turbines ◽  
Aerodynamic Loss ◽  
Inlet Swirl ◽  
Stage Performance ◽  
Performance Penalty ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Stage Efficiency

The integrated combustor vane concept for power generation gas turbines with can combustors has been shown to have significant benefits compared to conventional nozzle guide vanes (NGV). Aerodynamic loss, heat transfer levels, and cooling requirements are reduced while stage efficiency is improved by approximately 1.5% (for a no-swirl scenario). Engine realistic combustor flow with swirl, however, leads to increased turning nonuniformity downstream of the integrated vanes. This paper thus illustrates the altered integrated vane stage performance caused by inlet swirl. The study shows a distinct performance penalty for the integrated vane rotor as a result of increased rotor incidence and the rotor's interaction with the residual swirl core. The stage efficiency advantage of the integrated combustor vane concept compared to the conventional design is thus reduced to 0.7%. It is furthermore illustrated how integrated vane profiling is suitable to reduce the turning variation across the span downstream of the vane, further improve stage efficiency (in this case by 0.23%) and thus mitigate the distinct impact of inlet swirl on integrated vane stage performance.

Influence of Combustor Flow With Swirl on Integrated Combustor Vane Concept Full-Stage Performance

2017 ◽  
Author(s):  
Simon Jacobi ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Power Generation ◽  
Gas Turbines ◽  
Aerodynamic Loss ◽  
Inlet Swirl ◽  
Stage Performance ◽  
Performance Penalty ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Stage Efficiency

The integrated combustor vane concept for power generation gas turbines with can combustors has been shown to have significant benefits compared to conventional nozzle guide vanes. Aerodynamic loss, heat transfer levels and cooling requirements are reduced while stage efficiency is improved by approximately 1.5% (for a no-swirl scenario). Engine realistic combustor flow with swirl however leads to increased turning non-uniformity downstream of the integrated vanes. This paper thus illustrates the altered integrated vane stage performance caused by inlet swirl. The study shows a distinct performance penalty for the integrated vane rotor as a result of increased rotor incidence and the rotor’s interaction with the residual swirl core. The stage efficiency advantage of the integrated combustor vane concept compared to the conventional design is thus reduced to 0.7%. It is furthermore illustrated how integrated vane profiling is suitable to reduce the turning variation across the span downstream of the vane, further improve stage efficiency (in this case by 0.23%) and thus mitigate the distinct impact of inlet swirl on integrated vane stage performance.

scholarly journals Development and Aerothermal Investigation of Integrated Combustor Vane Concept

2015 ◽  
Vol 138 (1) ◽  
Author(s):  
Simon Jacobi ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Gas Turbines ◽  
High Speed ◽  
Leading Edge ◽  
Cfd Simulations ◽  
Pressure Loss Coefficient ◽  
Novel Concept ◽  
Nozzle Guide ◽  
Stage Efficiency

This paper presents the development and aerothermal investigation of the integrated combustor vane concept for power generation gas turbines with individual can combustors. In this novel concept, first introduced in 2010, the conventional nozzle guide vanes (NGVs) are removed and flow turning is achieved by vanes that extend the combustor walls. The concept was developed using the in-house computational fluid dyanamics (CFD) code TBLOCK. Aerothermal experiments were conducted using a modular high-speed linear cascade, designed to model the flow at the combustor–vane interface. The facility is comprised of two can combustor transition ducts and either four conventional vanes (CVs) or two integrated vanes (IVs). The experimental study validates the linear CFD simulations of the IV development. Annular full-stage CFD simulations, used to evaluate aerodynamics, heat transfer, and stage efficiency, confirm the trends of the linear numerical and experimental results, and thus demonstrate the concept's potential for real gas turbine applications. Results show a reduction of the total pressure loss coefficient at the exit of the stator vanes by more than 25% due to a reduction in profile and endwall loss. Combined with an improved rotor performance demonstrated by unsteady stage simulations, these aerodynamic benefits result in a gain in stage efficiency of above 1%. A distinct reduction in heat transfer coefficient (HTC) levels on vane surfaces, on the order of 25–50%, and endwalls is observed and attributed to an altered state of boundary layer (BL) thickness. The development of IV's endwall- and leading edge (LE)-cooling geometry shows a superior surface coverage of cooling effectiveness, and the cooling requirements for the first vane are expected to be halved. Moreover, by halving the number of vanes, simplifying the design and eliminating the need for vane LE film cooling, manufacturing and development costs can be significantly reduced.

Improving the Performance of a Turbine With Low Aspect Ratio Stators by Aft-Loading

2004 ◽  
Vol 128 (3) ◽  
pp. 492-499 ◽  
Author(s):  
Graham Pullan ◽  
John Denton ◽  
Eric Curtis
Keyword(s):  
Aspect Ratio ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Loss Reduction ◽  
Low Aspect Ratio ◽  
Surface Pressure Distribution ◽  
Loaded Surface ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Stage Efficiency

Experimental data and numerical simulations are presented from a research turbine with low aspect ratio nozzle guide vanes (NGVs). The combined effects of mechanical and aerodynamic constraints on the NGV create very strong secondary flows. This paper describes three designs of NGV that have been tested in the turbine, using the same rotor row in each case. NGV 2 used three-dimensional design techniques in an attempt to improve the performance of the datum NGV 1 blade, but succeeded only in creating an intense vortex shed from the trailing edge (as previously reported) and lowering the measured stage efficiency by 1.1% points. NGV 3 was produced to avoid the “shed vortex” while adopting a highly aft-loaded surface pressure distribution to reduce the influence of the secondary flows. The stage with NGV 3 had an efficiency 0.5% points greater than that with NGV 1. Detailed comparisons between experiment and computations, including predicted entropy generation rates, are used to highlight the areas where the loss reduction has occurred and hence to quantify the effects of employing highly aft-loaded NGVs.

Nozzle Guide Vane Static Strip Seals

2006 ◽  
Author(s):  
Arash Farahani ◽  
Peter Childs
Keyword(s):  
Gas Turbines ◽  
Guide Vane ◽  
Main Stream ◽  
Guide Vanes ◽  
Advantages And Disadvantages ◽  
Air System ◽  
Improved Performance ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Assembly Constraints

Sealing of components where there is no relative motion between the elements concerned remains a significant challenge in many gas turbine engine applications. Loss of sealing and cooling air from the internal air system through seals impacts on specific fuel consumption and can lead to undesirable flow interactions with resultant cost implications. For gas turbines, various strip seal types have been developed for use between Nozzle Guide Vanes in order to limit the flow of gas between the main stream annulus and the internal air system. Many different types of design have been proposed for overcoming strip seal problems such as misalignment of the grooves due to manufacturing and assembly constraints. In this paper various methods, with a particular focus on patents, for minimising the amount of leakage caused by such problems for strip seals between nozzle guide vanes are reviewed. By considering the advantages and disadvantages of each technique it is concluded that although apparently new strip seal designs for NGVs have improved performance, none of them can be considered to be ideal. This paper reviews the techniques and makes recommendations for future designs.

Thermal Investigation of Integrated Combustor Vane Concept Under Engine-Realistic Conditions

2016 ◽  
Author(s):  
Simon Jacobi ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbines ◽  
High Speed ◽  
Secondary Flows ◽  
Experimental Measurements ◽  
Thermal Investigation ◽  
Nozzle Guide ◽  
The Heat Transfer Coefficient

This paper presents a thermal investigation of the Integrated Combustor Vane concept for power generation gas turbines with individual can combustors. This concept has the potential to replace the high-pressure turbine’s first vanes by prolonged combustor walls. Experimental measurements are performed on a linear high-speed cascade consisting of two can combustors and two integrated vanes. The modularity of the facility allows for the testing at engine-realistic high turbulence levels, as well as swirl strengths with opposing swirl directions. The heat transfer characteristics of the integrated vanes are compared to conventional nozzle guide vanes. The experimental measurements are supported by detailed numerical simulations using the inhouse CFD code TBLOCK. Experimental as well as numerical results congruently indicate a considerable reduction of the heat transfer coefficient (HTC) on the integrated vanes surfaces and endwalls caused by a differing state of boundary layer thickness. The studies furthermore depict a slight, non-detrimental shift in the heat transfer coefficient distributions and the strength of the integrated vanes secondary flows as a result of engine-realistic combustor swirl.

Controlling Secondary-Flow Structure by Leading-Edge Airfoil Fillet and Inlet Swirl to Reduce Aerodynamic Loss and Surface Heat Transfer

2002 ◽  
Author(s):  
T. I.-P. Shih ◽  
Y.-L. Lin
Keyword(s):  
Heat Transfer ◽  
Stokes Equations ◽  
Leading Edge ◽  
Surface Heat ◽  
Secondary Flows ◽  
Maximum Height ◽  
Surface Heat Transfer ◽  
Aerodynamic Loss ◽  
Swirl Angle ◽  
Inlet Swirl

Computations, based on the ensemble-averaged compressible Navier-Stokes equations closed by the shear-stress transport (SST) turbulence model, were performed to investigate the effects of leading-edge airfoil fillet and inlet-swirl angle on the flow and heat transfer in a turbine-nozzle guide vane. Three fillet configurations were simulated: no fillet (baseline), a fillet whose thickness fades on the airfoil, and a fillet whose thickness fades on the endwall. For both fillets, the maximum height above the endwall is positioned along the stagnation zone/line on the airfoil under the condition of no swirl. For each configuration, three inlet swirls were investigated: no swirl (baseline) and two linearly varying swirl angle from one endwall to the other (+30° to −30° and −30° to +30°). Results obtained show that both leading-edge fillet and inlet swirl can reduce aerodynamic loss and surface heat transfer. For the conditions of this study, the difference in stagnation pressure from the nozzle’s inlet to its exit were reduced by more than 40% with swirl or with fillet without swirl. Surface heat transfer was reduced by more than 10% on the airfoil and by more than 30% on the endwalls. When there is swirl, leading-edge fillets became less effective in reducing aerodynamic loss and surface heat transfer, because the fillets were not optimized for swirl angles imposed. Since the intensity and size of the cross flow were found to increase instead of decrease by inlet swirl and by the type of fillet geometries investigated, the results of this study indicate that the mechanisms responsible for aerodynamic loss and surface heat transfer are more complex than just the intensity and the magnitude of the secondary flows. This study shows their location and interaction with the main flow to be more important, and this could be exploited for positive results.

scholarly journals Development and Aerothermal Investigation of Integrated Combustor Vane Concept

2015 ◽  
Author(s):  
Simon Jacobi ◽  
Budimir Rosic
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
High Speed ◽  
Cfd Simulations ◽  
Pressure Loss Coefficient ◽  
Development Costs ◽  
Superior Surface ◽  
Novel Concept ◽  
Nozzle Guide ◽  
Stage Efficiency

This paper presents the development and aerothermal investigation of the Integrated Combustor Vane concept for power generation gas turbines with individual can combustors. In this novel concept, first introduced in 2010 [1], the conventional Nozzle Guide Vanes (NGVs) are removed and flow turning is achieved by vanes that extend the combustor walls. The concept is developed using the inhouse CFD code TBLOCK. Aerothermal experiments are conducted using a modular high speed linear cascade, designed to model the flow at the combustor-vane interface. The facility comprises two can combustor transition ducts and either four Conventional Vanes (CVs) or two Integrated Vanes (IVs). The experimental study validates the linear CFD-simulations of the IV development. Annular full stage CFD-simulations, used to evaluate aerodynamics, heat transfer and stage efficiency, confirm the trends of the linear numerical and experimental results and thus demonstrate the concept’s potential for real gas turbine applications. Results show a reduction of the total pressure loss coefficient at the exit of the stator vanes by more than 25% due to a reduction in profile- and endwall-loss. Combined with an improved rotor performance these aerodynamic benefits result in a gain in stage efficiency of above 1%, illustrated by unsteady stage simulations. A distinct reduction in HTC levels on vane surfaces, in the order of 25%–50%, and endwalls is observed and attributed to an altered state of boundary layer thickness. The development of IV’s endwall- and LE-cooling geometry shows a superior surface coverage of cooling effectiveness, and the cooling requirements for the first vane are expected to be halved. Moreover, by halving the number of vanes, simplifying the design and eliminating the need for vane LE film cooling, manufacturing and development costs can be significantly reduced.

Film Cooling Research on the Endwall of a Turbine Nozzle Guide Vane in a Short Duration Annular Cascade: Part 1—Experimental Technique and Results

1992 ◽  
Vol 114 (4) ◽  
pp. 734-740 ◽  
Author(s):  
S. P. Harasgama ◽  
C. D. Burton
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Suction Side ◽  
Coolant Temperature ◽  
Guide Vanes ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
Nozzle Guide Vanes

Heat transfer and aerodynamic measurements have been made on the endwalls of an annular cascade of turbine nozzle guide vanes in the presence of film cooling. The results indicate that high levels of cooling effectiveness can be achieved on the endwalls of turbine nozzle guide vanes (NGV). The NGV were operated at the correct engine nondimensional conditions of Reynolds number, Mach number, gas-to-wall temperature ratio, and gas-to-coolant density ratio. The results show that the secondary flow and horseshoe vortex act on the coolant, which is convected toward the suction side of the NG V endwall passage. Consequently the coolant does not quite reach the pressure side/casing trailing edge, leading to diminished cooling in this region. Increasing the blowing rate from 0.52 to 1.1 results in significant reductions in heat transfer to the endwall. Similar trends are evident when the coolant temperature is reduced. Measured heat transfer rates indicate that over most of the endwall region the film cooling reduces the Nusselt number by 50 to 75 percent.

Aerothermodynamics of Micro-Turbomachinery

2004 ◽  
Author(s):  
Y. Gong ◽  
B. T. Sirakov ◽  
A. H. Epstein ◽  
C. S. Tan
Keyword(s):  
Reynolds Number ◽  
Gas Turbine ◽  
Preliminary Design ◽  
Micro Gas Turbine ◽  
Heat Addition ◽  
Order Of Magnitude ◽  
Performance Penalty ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Engineering foundation for micro-turbomachinery aerothermal design, as an enabling element of the MIT micro-gas turbine technology, is developed. Fundamental differences between conventional, large scale and micro turbomachinery operation are delineated and the implications on design are discussed. These differences are largely a consequence of low operating Reynolds number, hence a relatively higher skin friction and heat transfer rate. While the size of the micro-gas turbine engine is ∼ a few mm, several order of magnitude smaller than conventional gas turbine, the required compressor stage pressure ratio (∼3–4) and impeller tip Mach number (∼1 and greater) are comparable; however, the disparity in the size implies that the operating Reynolds number of the micro-turbomachiery components is correspondingly several order of magnitudes smaller. Thus the design and operating requirements for micro-turbomachinery are distinctly different from those of conventional turbomachinery used for propulsion and power generation. Important distinctions are summarized in the following. 1. The high surface-to-flow rate ratio has the consequence that the flow in micro-compressor flow path can no longer be taken as adiabatic; the performance penalty associated with heat addition to compressor flow path from turbine is a primary performance limiting factor. 2. Endwall torque on the flow can be significant compared to that from the impeller blade surfaces so that direct use of Euler Turbine Equation is no longer appropriate. 3. Losses in turbine nozzle guide vanes (NGVs) can be one order of magnitude higher than those in conventional sized nozzle guide vanes. 4. The high level of kinetic energy in the flow exiting the turbine rotor is a source of performance penalty, largely a consequence of geometrical constraints. It can be inferred from these distinctions that standard preliminary design procedures based on the Euler equation, the adiabatic assumption, the loss correlations for large Reynolds numbers, and the three-dimensional geometry, are inapplicable to micro-turbomachinery. The preliminary design procedure, therefore, must account for these important differences. Characterization of the effects of heat addition on compressor performance, modification of Euler turbine equation for casing torque, characterization of turbine NGV performance and turbine exhaust effects are presented.

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