scholarly journals Effects of Nozzle Helmholtz Number on Indirect Combustion Noise by Compositional Perturbations

2017 ◽  
Vol 140 (3) ◽  
Author(s):  
Luca Magri ◽  
Jeffrey O'Brien ◽  
Matthias Ihme
Keyword(s):  
Velocity Profile ◽  
Combustion Chamber ◽  
Transfer Functions ◽  
Mixture Fraction ◽  
Supersonic Nozzle ◽  
Noise Pollution ◽  
Combustion Noise ◽  
Dipole Source ◽  
Multicomponent Gas ◽  
Thermoacoustic Oscillations

By modeling a multicomponent gas, a new source of indirect combustion noise is identified, which is named compositional indirect noise. The advection of mixture inhomogeneities exiting the gas-turbine combustion chamber through subsonic and supersonic nozzles is shown to be an acoustic dipole source of sound. The level of mixture inhomogeneity is described by a difference in composition with the mixture fraction. An n-dodecane mixture, which is a kerosene fuel relevant to aeronautics, is used to evaluate the level of compositional noise. By relaxing the compact-nozzle assumption, the indirect noise is numerically calculated for Helmholtz numbers up to 2 in nozzles with linear velocity profile. The compact-nozzle limit is discussed. Only in this limit, it is possible to derive analytical transfer functions for (i) the noise emitted by the nozzle and (ii) the acoustics traveling back to the combustion chamber generated by accelerated compositional inhomogeneities. The former contributes to noise pollution, whereas the latter has the potential to induce thermoacoustic oscillations. It is shown that the compositional indirect noise can be at least as large as the direct noise and entropy noise in choked nozzles and lean mixtures. As the frequency with which the compositional inhomogeneities enter the nozzle increases, or as the nozzle spatial length increases, the level of compositional noise decreases, with a similar, but not equal, trend to the entropy noise. The noisiest configuration is found to be a compact supersonic nozzle.

Effects of Nozzle Helmholtz Number on Indirect Combustion Noise by Compositional Perturbations

2017 ◽  
Author(s):  
Luca Magri ◽  
Jeffrey O’Brien ◽  
Matthias Ihme
Keyword(s):  
Velocity Profile ◽  
Combustion Chamber ◽  
Gas Turbine ◽  
Transfer Functions ◽  
Mixture Fraction ◽  
Supersonic Nozzle ◽  
Noise Pollution ◽  
Combustion Noise ◽  
Dipole Source ◽  
Thermoacoustic Oscillations

By modeling a multi-component gas, a new source of indirect combustion noise is identified, which is named compositional indirect noise. The advection of mixture inhomogeneities exiting the gas-turbine combustion chamber through subsonic and supersonic nozzles is shown to be an acoustic dipole source of sound. The level of mixture inhomogeneity is described by a difference in composition with the mixture fraction. An n-dodecane mixture, which is a kerosene fuel relevant to aeronautics, is used to evaluate the level of compositional noise. By relaxing the compact-nozzle assumption, the indirect noise is numerically calculated for Helmholtz numbers up to 2 in nozzles with linear velocity profile. The compact-nozzle limit is discussed. Only in this limit, it is possible to derive analytical transfer functions for (i) the noise emitted by the nozzle and (ii) the acoustics travelling back to the combustion chamber generated by accelerated compositional inhomogeneities. The former contributes to noise pollution, whereas the latter has the potential to induce thermoacoustic oscillations. It is shown that the compositional indirect noise can be at least as large as the direct noise and entropy noise in chocked nozzles and lean mixtures. As the frequency with which the compositional inhomogeneities enter the nozzle increases, or as the nozzle spatial length increases, the level of compositional noise decreases, with a similar, but not equal, trend to the entropy noise. The noisiest configuration is found to be a compact supersonic nozzle.

Design Analysis of a Micro Gas Turbine Combustion Chamber Burning Natural Gas

2012 ◽  
Author(s):  
André Perpignan V. de Campos ◽  
Fernando L. Sacomano Filho ◽  
Guenther C. Krieger Filho
Keyword(s):  
Experimental Data ◽  
Natural Gas ◽  
Combustion Chamber ◽  
Gas Turbines ◽  
Low Cost ◽  
Mixture Fraction ◽  
Combustion Model ◽  
Inlet Temperature ◽  
Gas Combustion ◽  
Micro Turbine

Gas turbines are reliable energy conversion systems since they are able to operate with variable fuels and independently from seasonal natural changes. Within that reality, micro gas turbines have been increasing the importance of its usage on the onsite generation. Comparatively, less research has been done, leaving more room for improvements in this class of gas turbines. Focusing on the study of a flexible micro turbine set, this work is part of the development of a low cost electric generation micro turbine, which is capable of burning natural gas, LPG and ethanol. It is composed of an originally automotive turbocompressor, a combustion chamber specifically designed for this application, as well as a single stage axial power turbine. The combustion chamber is a reversed flow type and has a swirl stabilized combustor. This paper is dedicated to the diagnosis of the natural gas combustion in this chamber using computational fluid dynamics techniques compared to measured experimental data of temperature inside the combustion chamber. The study emphasizes the near inner wall temperature, turbine inlet temperature and dilution holes effectiveness. The calculation was conducted with the Reynolds Stress turbulence model coupled with the conventional β-PDF equilibrium along with mixture fraction transport combustion model. Thermal radiation was also considered. Reasonable agreement between experimental data and computational simulations was achieved, providing confidence on the phenomena observed on the simulations, which enabled the design improvement suggestions and analysis included in this work.

scholarly journals Flow Inhomogeneities in a Realistic Aeronautical Gas-Turbine Combustor: Formation, Evolution and Indirect Noise

2018 ◽  
Author(s):  
Andrea Giusti ◽  
Luca Magri ◽  
Marco Zedda
Keyword(s):  
Gas Turbines ◽  
Transfer Functions ◽  
Guide Vane ◽  
Flame Structure ◽  
Combustion Noise ◽  
Moment Closure ◽  
Combustion Model ◽  
Conditional Moment ◽  
Eddy Simulation ◽  
Nozzle Guide

Indirect noise generated by the acceleration of combustion inhomogeneities is an important aspect in the design of aeroengines because of its impact on the overall noise emitted by an aircraft and the possible contribution to combustion instabilities. In this study, a realistic rich-quench-lean combustor is numerically investigated, with the objective of quantitatively analyzing the formation and evolution of flow inhomogeneities and determine the level of indirect combustion noise in the nozzle guide vane (NGV). Both entropy and compositional noise are calculated in this work. A high-fidelity numerical simulation of the combustion chamber, based on the Large-Eddy Simulation (LES) approach with the Conditional Moment Closure (CMC) combustion model, is performed. The contributions of the different air streams to the formation of flow inhomogeneities are pinned down and separated with seven dedicated passive scalars. LES-CMC results are then used to determine the acoustic sources to feed an NGV aeroacoustic model, which outputs the noise generated by entropy and compositional inhomogeneities. Results show that non-negligible fluctuations of temperature and composition reach the combustor’s exit. Combustion inhomogeneities originate both from finite-rate chemistry effects and incomplete mixing. In particular, the role of mixing with dilution and liner air flows on the level of combustion inhomogeneities at the combustor’s exit is highlighted. The species that most contribute to indirect noise are identified and the transfer functions of a realistic NGV are computed. The noise level indicates that indirect noise generated by temperature fluctuations is larger that the indirect noise generated by compositional inhomogeneities, although the latter is not negligible and is expected to become louder in supersonic nozzles. It is also shown that relatively small fluctuations of the local flame structure can lead to significant variations of the nozzle transfer function, whose gain increases with the Mach number. This highlights the necessity of an on-line solution of the local flame structure, which is performed in this paper by CMC, for an accurate prediction of the level of compositional noise. This study opens new possibilities for the identification, separation and calculation of the sources of indirect combustion noise in realistic aeronautical gas turbines.

Influence of Turbulent Combustion Noise on the Decrement Estimates of Gas Oscillations

2018 ◽  
Vol 140 (10) ◽  
Author(s):  
V. Karmalita
Keyword(s):  
Combustion Chamber ◽  
Linear Model ◽  
Natural Frequency ◽  
Turbulent Combustion ◽  
Logarithmic Decrement ◽  
Combustion Noise ◽  
Vibrating Combustion ◽  
The Impact ◽  
Oscillation Decrement

This paper deals with the autoregression method to determine the logarithmic decrement and natural frequency of gas oscillations in a combustion chamber. The proposed approach to quantify the impact of combustion noise on the estimates of oscillation decrement is based on a linear model of the vibrating combustion phenomena. An application of the proposed approach for afterburner tests is presented as well.

scholarly journals Flow Inhomogeneities in a Realistic Aeronautical Gas-Turbine Combustor: Formation, Evolution, and Indirect Noise

2018 ◽  
Vol 141 (1) ◽  
Author(s):  
Andrea Giusti ◽  
Luca Magri ◽  
Marco Zedda
Keyword(s):  
Gas Turbines ◽  
Transfer Functions ◽  
Guide Vane ◽  
Flame Structure ◽  
Combustion Noise ◽  
Moment Closure ◽  
Combustion Model ◽  
Conditional Moment ◽  
Eddy Simulation ◽  
Aero Engines

Indirect noise generated by the acceleration of combustion inhomogeneities is an important aspect in the design of aero-engines because of its impact on the overall noise emitted by an aircraft and the possible contribution to combustion instabilities. In this study, a realistic rich-quench-lean (RQL) combustor is numerically investigated, with the objective of quantitatively analyzing the formation and evolution of flow inhomogeneities and determining the level of indirect combustion noise in the nozzle guide vane (NGV). Both entropy and compositional noise are calculated in this work. A high-fidelity numerical simulation of the combustion chamber, based on the large-eddy simulation (LES) approach with the conditional moment closure (CMC) combustion model, is performed. The contributions of the different air streams to the formation of flow inhomogeneities are pinned down and separated with seven dedicated passive scalars. LES-CMC results are then used to determine the acoustic sources to feed an NGV aeroacoustic model, which outputs the noise generated by entropy and compositional inhomogeneities. Results show that non-negligible fluctuations of temperature and composition reach the combustor's exit. Combustion inhomogeneities originate both from finite-rate chemistry effects and incomplete mixing. In particular, the role of mixing with dilution and liner air flows on the level of combustion inhomogeneities at the combustor's exit is highlighted. The species that most contribute to indirect noise are identified and the transfer functions of a realistic NGV are computed. The noise level indicates that indirect noise generated by temperature fluctuations is larger than the indirect noise generated by compositional inhomogeneities, although the latter is not negligible and is expected to become louder in supersonic nozzles. It is also shown that relatively small fluctuations of the local flame structure can lead to significant variations of the nozzle transfer function, whose gain increases with the Mach number. This highlights the necessity of an on-line solution of the local flame structure, which is performed in this paper by CMC, for an accurate prediction of the level of compositional noise. This study opens new possibilities for the identification, separation, and calculation of the sources of indirect combustion noise in realistic aeronautical gas turbines.

Effects of Pilot Fuel and Liner Cooling on the Flame Structure in a Full Scale Swirl-Stabilized Combustion Setup

2009 ◽  
Author(s):  
Jens Fa¨rber ◽  
Rainer Koch ◽  
Hans-Jo¨rg Bauer ◽  
Matthias Hase ◽  
Werner Krebs
Keyword(s):  
Combustion Chamber ◽  
Fuel Injection ◽  
Flame Structure ◽  
Cone Angle ◽  
Flow Patterns ◽  
Flame Stabilization ◽  
Operating Conditions ◽  
Combustion Noise ◽  
Pilot Fuel ◽  
Combustor Liner

The flame structure and the limits of operation of a lean premixed swirl flame were experimentally investigated under piloted and non-piloted conditions. Flame stabilization and blow out limits are discussed with respect to pilot fuel injection and combustor liner cooling for lean operating conditions. Two distinctly different flow patterns are found to develop depending on piloting and liner cooling parameters. These flow patterns are characterized with respect to flame stability, blow out limits, combustion noise and emissions. The combustion system explored consists of a single burner similar to the burners used in Siemens annular combustion systems. The burner feeds a distinctively non-adiabatic combustion chamber operated with natural gas under atmospheric pressure. Liner cooling is mimicked by purely convective cooling and an additional flow of ‘leakage air’ injected into the combustion chamber. Both, this additional air flow and the pilot fuel ratio were found to have a strong influence on the flow structure and stability of the flame close to the lean blow off limit (LBO). It is shown by Laser Doppler Velocimetry (LDV) that the angle of the swirl cone is strongly affected by pilot fuel injection. Two distinct types of flow patterns are observed close to LBO in this large scale setup: While non-piloted flames exhibit tight cone angles and small inner recirculation zones (IRZ), sufficient piloting results in a wide cone angle and a large IRZ. Only in the latter case, the main flow becomes attached to the combustor liner. Flame structures deduced from flow fields and CH-Chemiluminescence images depend on both the pilot fuel injection and liner cooling.

scholarly journals Numerical Prediction of Far-Field Combustion Noise from Aeronautical Engines

Acoustics ◽  
2019 ◽  
Vol 1 (1) ◽  
pp. 174-198 ◽  
Author(s):  
Mélissa Férand ◽  
Thomas Livebardon ◽  
Stéphane Moreau ◽  
Marlène Sanjosé
Keyword(s):  
Temperature Field ◽  
Low Power ◽  
Combustion Chamber ◽  
High Power ◽  
Acoustic Propagation ◽  
Combustion Noise ◽  
Far Field ◽  
Low Frequencies ◽  
Turboshaft Engine ◽  
Good Agreement

A hybrid methodology combining a detailed Large Eddy Simulation of a combustion chamber sector, an analytical propagation model of the extracted acoustic and entropy waves at the combustor exit through the turbine stages, and a far-field acoustic propagation through a variable exhaust temperature field was shown to predict far-field combustion noise from helicopter and aircraft propulsion systems accurately for the first time. For the single-stream turboshaft engine, the validation was achieved from engine core to the turbine exit. Propagation to the far field was then performed through a modeled axisymmetric jet. Its temperature modified the acoustic propagation of combustion noise significantly and a simple analytical model based on the Snell–Descarte law was shown to predict the directivity for axisymmetric single jet exhaust accurately. Good agreement with measured far-field spectra for all turboshaft-engine regimes below 2 kHz stresses that combustion noise is most likely the dominant noise source at low frequencies in such engines. For the more complex dual-stream turbofan engine, two regime computations showed that direct noise is mostly generated by the unsteady flame dynamics and the indirect combustion noise by the temperature stratification induced by the dilution holes in the combustion chamber, as found previously in the turboshaft case. However, in the turboengine, direct noise was found dominant at the combustor exit for the low power case and equivalent contributions of both combustion noise sources for the high power case. The propagation to the far-field was achieved through the temperature field provided by a Reynolds-Averaged Navier–Stokes simulation. Good agreement with measured spectra was also found at low frequencies for the low power turboengine case. At high power, however, turboengine jet noise overcomes combustion noise at low frequencies.

Thermoacoustic Modeling and Transfer Functions Determination for a Matrix Burner Using Unsteady CFD

2008 ◽  
Author(s):  
Dieter Bohn ◽  
James F. Willie ◽  
Nils Ohlendorf
Keyword(s):  
Velocity Field ◽  
Combustion Chamber ◽  
Frequency Response ◽  
Transfer Functions ◽  
Time Lag ◽  
Heat Rate ◽  
Network Code ◽  
Flame Holder ◽  
Air Stream ◽  
Flow Variables

Turbulent combustion of a lean premixed methane-air mixture is simulated numerically using unsteady CFD. The configuration is a matrix burner suitable for stationary gas turbine applications. The geometry consists of the following: seven slots which constitute the flame holder for stabilizing the flame, a diffuser which serves the purpose of lowering the pressure loss across the burner and a combustion chamber. A contraction ensures that a recirculation zone is created close to the exit of the flame holder for anchoring and stabilizing the flame. Fuel is injected in 112 holes, 8 along each end of the 7 slots. The injected fuel meets the in-coming high velocity air stream for mixing to begin in the premixed ducts before finally entering the combustion chamber. This paper validates the cold flow velocity field and the steady flame results from CFD with measurements and investigates combustion instability in a matrix burner, the onset of which can be attributed to changes in flow variables using URANS. Particularly, the effect of the mixture strength variation caused by fluctuations in the velocity field on the unsteady heat rate inside the combustor is investigated. The fuel inlet is assumed to be choked due to the high pressure drop across it. The time lag between the time fuel is injected and the time it reaches the flame front is estimated. Quantifying this time delay (or “flight time”) helps to characterize the burner with respect to thermo-acoustic instabilities. The flame frequency response to a white noise forcing at the air inlet is determined. This is followed by the determination of the acoustic transfer matrix linking the pressure and velocity downstream and upstream of the burner/flame. This is done by using system identification that is common in control theory. The determined flame frequency response and the time lag are used in a 1D acoustic network code for determining the longitudinal eigenmodes of the combustor of the matrix burner.

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