Short Helical Combustor: Concept Study of an Innovative Gas Turbine Combustor With Angular Air Supply

2015 ◽  
Vol 138 (3) ◽  
Author(s):  
B. Ariatabar ◽  
R. Koch ◽  
H.-J. Bauer ◽  
D.-A. Negulescu
Keyword(s):  
Gas Turbine ◽  
Scaling Laws ◽  
Generic Model ◽  
Reacting Flow ◽  
Air Supply ◽  
Tilting Angle ◽  
Annular Combustor ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Cooling Air

An innovative design of a gas turbine annular combustor is investigated analytically and numerically. Its principal feature is the helical arrangement of the burners around the turbine shaft. Hence, a shorter combustor with lower aerodynamic losses and cooling air demand might be realized. A generic model of the combustor is developed and analyzed by means of a parametric study. Scaling laws for the geometry of the flame tube and the burners are derived. Thereby, the relevant similarity parameters for fluid flow, combustion, and heat transfer are maintained constant. Subsequently, nonreacting and reacting flow regimes of selected design variants are numerically investigated. It is shown that a double annular (DA) configuration with a tilting angle of β = 45 deg, where circumferentially adjacent swirls are corotating and radially are counter-rotating, is the superior design in terms of (1) maintaining the relevant similarity rules, (2) size and location of the recirculation zones and swirl flames, and (3) flow pattern at the combustor exit. The deflection angle of the nozzle guide vanes (NGV) as well as the axial length of such a short helical combustor (SHC) could be reduced by approximately 30%.

Short Helical Combustor: Concept Study of an Innovative Gas Turbine Combustor

2015 ◽  
Author(s):  
B. Ariatabar ◽  
R. Koch ◽  
H.-J. Bauer
Keyword(s):  
Gas Turbine ◽  
Scaling Laws ◽  
Generic Model ◽  
Reacting Flow ◽  
Innovative Design ◽  
Tilting Angle ◽  
Annular Combustor ◽  
Recirculation Zones ◽  
Turbine Shaft ◽  
Cooling Air

An innovative design of a gas turbine annular combustor is investigated analytically and numerically. Its principal feature is the helical arrangement of the burners around the turbine shaft. Hence, a shorter combustor with lower aerodynamical losses and cooling air demand might be realized. A generic model of the combustor is developed and analyzed by means of a parametric study. Scaling laws for the geometry of the flame tube and the burners are derived. Thereby, the relevant similarity parameters for fluid flow, combustion and heat transfer are maintained constant. Subsequently, non-reacting and reacting flow regimes of selected design variants are numerically investigated. It is shown that a double annular configuration with a tilting angle of β = 45°, where circumferentially adjacent swirls are co-rotating and radially are counter-rotating, is the superior design in terms of 1. Maintaining the relevant similarity rules 2. Size and location of the recirculation zones and swirl flames 3. Flow pattern at the combustor exit The deflection angle of the NGV as well as the axial length of such a Short Helical Combustor could be reduced by approx. 30%.

A Calculation Tool of a Turbine Cooling Air Schedule for General Gas Turbine Simulation Algorithms

2011 ◽  
Vol 134 (4) ◽  
Author(s):  
Valerio Lallini ◽  
Jan Janikovic ◽  
Pericles Pilidis ◽  
Riti Singh ◽  
Panagiotis Laskaridis
Keyword(s):  
Gas Turbine ◽  
Operating Cost ◽  
Flight Path ◽  
Turbine Engine ◽  
Turbine Cooling ◽  
Cooling Flow ◽  
Conventional Cooling ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Cooling Air

A tool to evaluate the turbine cooling air schedule during the flight path is presented. The method is suitable for complex gas turbine engine models where accurate calculation of thermodynamic parameters of a cooling flow used for studies such as specific fuel consumption (SFC), lifing, and operating cost analysis. The method is also applicable for land-based gas turbine engines where the flight path is replaced by load pattern. The method can be employed even if only a little information about the engine is known. The calculation method has been investigated on a model of a high bypass ratio turbofan for long haul application. Using it with an aircraft model and after comparing to conventional cooling air simulations, the reduction of SFC during the flight path has been observed. A comparison has been performed on engine manufacturer’s public domain data where the method shows comparable value of total cooling air needed. The tool returns similar results but estimates a higher proportion of cooling air for the high pressure turbine nozzle guide vanes (NGV) compared with Young and Wilcock’s method.

Characteristics of Volcanic Ash in a Gas Turbine Combustor and Nozzle Guide Vanes

2018 ◽  
Vol 140 (7) ◽  
Author(s):  
Lei-Yong Jiang ◽  
Yinghua Han ◽  
Prakash Patnaik
Keyword(s):  
Gas Turbine ◽  
Volcanic Ash ◽  
Inlet Temperature ◽  
Gas Turbine Combustor ◽  
Flow Passage ◽  
Guide Vanes ◽  
Cooling Flow ◽  
Cooling Passage ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

To understand the physics of volcanic ash impact on gas turbine hot-components and develop much-needed tools for engine design and fleet management, the behaviors of volcanic ash in a gas turbine combustor and nozzle guide vanes (NGV) have been numerically investigated. High-fidelity numerical models are generated, and volcanic ash sample, physical, and thermal properties are identified. A simple critical particle viscosity—critical wall temperature model is proposed and implemented in all simulations to account for ash particles bouncing off or sticking on metal walls. The results indicate that due to the particle inertia and combustor geometry, the volcanic ash concentration in the NGV cooling passage generally increases with ash size and density, and is less sensitive to inlet velocity. It can reach three times as high as that at the air inlet for the engine conditions and ash properties investigated. More importantly, a large number of the ash particles entering the NGV cooling chamber are trapped in the cooling flow passage for all four turbine inlet temperature conditions. This may reveal another volcanic ash damage mechanism originated from engine cooling flow passage. Finally, some suggestions are recommended for further research and development in this challenging field. To the best of our knowledge, it is the first study on detailed ash behaviors inside practical gas turbine hot-components in the open literature.

Flow Field Simulations of a Gas Turbine Combustor

2002 ◽  
Vol 124 (3) ◽  
pp. 508-516 ◽  
Author(s):  
M. D. Barringer ◽  
O. T. Richard ◽  
J. P. Walter ◽  
S. M. Stitzel ◽  
K. A. Thole
Keyword(s):  
Flow Field ◽  
Gas Turbine ◽  
Large Scale ◽  
Guide Vane ◽  
Wall Region ◽  
Turbine Engine ◽  
Tunnel Section ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Combustor Liner

The flow field exiting the combustor in a gas turbine engine is quite complex considering the presence of large dilution jets and complicated cooling schemes for the combustor liner. For the most part, however, there has been a disconnect between the combustor and turbine when simulating the flow field that enters the nozzle guide vanes. To determine the effects of a representative combustor flow field on the nozzle guide vane, a large-scale wind tunnel section has been developed to simulate the flow conditions of a prototypical combustor. This paper presents experimental results of a combustor simulation with no downstream turbine section as a baseline for comparison to the case with a turbine vane. Results indicate that the dilution jets generate turbulence levels of 15–18% at the exit of the combustor with a length scale that closely matches that of the dilution hole diameter. The total pressure exiting the combustor in the near-wall region neither resembles a turbulent boundary layer nor is it completely uniform putting both of these commonly made assumptions into question.

Validation and Comparison of Strip Seal Designs for Gas Turbine Engine Nozzle Guide Vanes

2008 ◽  
Author(s):  
Arash Farahani ◽  
Peter Childs
Keyword(s):  
Gas Turbine ◽  
Guide Vane ◽  
Gas Turbine Engine ◽  
Experimental Comparison ◽  
Cfd Modelling ◽  
Guide Vanes ◽  
Turbine Engine ◽  
Seal Design ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Strip seals are commonly used to prevent or limit leakage flows between nozzle guide vanes (NGV) and other gas turbine engine components that are assembled from individual segments. Leakage flow across, for example, a nozzle guide vane platform, leads to increased demands on the gas turbine engine internal flow system and a rise in specific fuel consumption (SFC). Careful attention to the flow characteristics of strip seals is therefore necessary. The very tight tolerances associated with strip seals provides a particular challenge to their characterisation. This paper reports the validation of CFD modelling for the case of a strip seal under very carefully controlled conditions. In addition, experimental comparison of three types of strip seal design, straight, arcuate, and cloth, is presented. These seals are typical of those used by competing manufacturers of gas turbine engines. The results show that the straight seal provides the best flow sealing performance for the controlled configuration tested, although each design has its specific merits for a particular application.

Failure Analysis of Heavy Industrial Gas Turbine Engine First Stage Nozzel Guide Vane

2012 ◽  
Vol 445 ◽  
pp. 1047-1052
Author(s):  
Alaaeldin H. Mustafa
Keyword(s):  
Failure Analysis ◽  
Gas Turbine ◽  
Guide Vane ◽  
Xrd Analysis ◽  
Optical Microscope ◽  
Guide Vanes ◽  
X Ray ◽  
Nozzle Guide ◽  
Industrial Gas Turbine ◽  
Nozzle Guide Vanes

Failure analysis investigation was conducted on 70 MW set of 1st stage turbine nozzle guide vanes (NGVs) of heavy industrial gas turbine. The failure was investigated using the light optical microscope (LOM), X-ray diffraction analysis (XRD) and energy dispersive X-ray spectroscopy (EDS) in an environmental scanning electron microscope (ESEM). The results of the analysis indicate that the NGVs which were made of Co base superalloy FSX-414 had been operated above the recommended operating hours under different fuel types in addition to inadequate repair process in previous repair removal. The XRD analysis of the fractured areas sample shows presence ofwhich might indicate the prolonged operation at high temperature. Keywords: cobalt-base; nozzle guide vanes, gas turbine.

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.

An Impact Assessment of Erosion of Nozzle Guide Vanes and Rotor Blades on Aerodynamic Performance of a Gas Turbine by CFD

2019 ◽  
Author(s):  
Koichi Yonezawa ◽  
Masahiro Takayasu ◽  
Genki Nakai ◽  
Kazuyasu Sugiyama ◽  
Katsuhiko Sugita ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Rotor Blade ◽  
Cfd Simulation ◽  
Rotor Blades ◽  
Total Power ◽  
Guide Vanes ◽  
Turbine Performance ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Abstract Nozzle guide vanes (NGVs) and rotor blades deteriorate due to erosion, and this may affect the aerodynamic characteristics of gas turbines. According to previous studies, the erosion of first-stage NGVs significantly affected the blade loading of the first-stage rotor. An increase in the tip gap also may significantly affect the gas turbine performance. In the present study, numerical investigations have been carried out using a real eroded nozzle and blade geometries for two purposes. One purpose was to clarify the influences underlying the deterioration of the nozzle and the rotor blade, such as the effects on the erosion of NGVs in the first stage and the effects of the tip gap on the gas turbine performance. The other was to develop a method to estimate the total gas turbine performance using a CFD simulation and a heat balance analysis. The results show that the erosion of NGV leads to an increased flow rate and affects the operating condition of the gas turbine cycle. This, in turn, can decrease the total thermal efficiency. The experimental results suggest that an increase in the tip gap width decreases rotor output almost linearly, and the numerical results showed the same tendency. The influence of the tip gap in the real gas turbine condition was also examined, revealing that an increase in the tip gap leads to an increase in the pressure loss in the nozzle downstream as well as around the rotor blade itself. Consequently, the total power output and the isentropic efficiency of the turbine decreased.

Study over thermal state of gas turbine engine metal-ceramic rotor blades and nozzle guide vanes under thermal shock and thermal-cyclic loading conditions

2004 ◽  
Vol 13 (2) ◽  
pp. 163-166
Author(s):  
A. V. Soudarev ◽  
A. A. Souryaninov ◽  
V. Ya. Podgorets ◽  
V. V. Grishaev ◽  
V.Yu Tikhoplav ◽  
...  
Keyword(s):  
Cyclic Loading ◽  
Thermal Shock ◽  
Gas Turbine ◽  
Thermal State ◽  
Rotor Blades ◽  
Guide Vanes ◽  
Turbine Engine ◽  
Metal Ceramic ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes
Sign in / Sign up

Export Citation Format

Share Document