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|Title: ||Combustion Instability Screech In Gas Turbine Afterburner|
|Authors: ||Ashirvadam, Kampa|
|Advisors: ||Paul, P J|
|Keywords: ||Gas Turbine|
Gas Turbines - Combustion
Gas Turbine Afterburners
Aircraft Gas Turbines - Combustion
Gas Turbines - Combustion Chambers
Gas Turbine Afterburners - Combustion Instabiltiy
|Submitted Date: ||Jul-2007|
|Series/Report no.: ||G21108|
|Abstract: ||Gas turbine reheat thrust augmenters known as afterburners are used to provide additional thrust during emergencies, take oﬀ, combat, and in supersonic ﬂight of high-performance aircrafts. During the course of reheat development, the most persistent trouble has been the onset of high frequency combustion instability, also known as screech, invariably followed by rapid mechanical failure. The coupling of acoustic pressure upstream of the ﬂame stabilizer with in-phase heat-release downstream, results in combustion instability by which the amplitude at various resonant modes — longitudinal (buzz — low frequency), tangential or radial (screech — high frequency) – ampliﬁes leading to deterioration of the afterburner components.
Various researchers in early 1950s have performed extensive testing on straight jet afterburners, to identify screech frequencies. Theoretical and experimental work at test rig level has been reported in the case of buzz to validate the heat release combustion models. In this work, focus is given to study the high frequency tangential combustion instability by vibro-acoustic software and the tests are conducted on the scaled bypass ﬂow afterburner for conﬁrmation of predicted screech frequencies.
The wave equation for the afterburner is solved taking the appropriate geometry of the afterburner and taking into account the factors aﬀecting the stability. Nozzle of the afterburner is taken into account by using the nozzle admittance condition derived for a choked nozzle. Screech liner admittance boundary condition is imposed and the eﬀect on acoustic attenuation is studied. A new combustion model has been proposed for obtaining the heat release rate response function to acoustic oscillations. Acoustic wave – ﬂame interactions involve unsteady kinetic, ﬂuid mechanic and acoustic processes over a large range of time scales. Three types of ﬂow disturbances exist such as : vortical, entropy, and acoustic. In a homogeneous, uniform ﬂow, these three disturbance modes propagate independently in the linear approximation. Unsteady heat release also generates entropy and vorticity disturbances. Since ﬂow is not accelerated in the region of uniform area duct, vortical and entropy disturbances are treated as in signiﬁcant, as these disturbances are convected out into atmosphere like an open-ended tube, but these are considered in deriving the nozzle admittance condition. Heat release ﬂuctuations that arise due to ﬂuctuating pressure and temperature are taken into consideration. The aim is to provide results on how ﬂames respond to pressure disturbances of diﬀerent amplitudes and characterised by diﬀerent length scales. The development of the theory is based on large activation energy asymptotics. One-dimensional conservation equations are used for obtaining the response function for the heat release rate assuming the laminar ﬂamelet model to be valid. The estimates are compared with the published data and deviations are discussed.
The normalized acoustic pressure variation in the afterburner is predicted using the models discussed earlier to provide an indication of the resonant modes of the pressure oscillations and the ampliﬁcation and attenuation of oscillations caused by the various processes. Similar frequency spectrum is also obtained experimentally using a test rig for a range of inlet mean pressures and temperatures with combustion and core and bypass ﬂows simulated, for conﬁrmation of predicted results.
Without the heat source only longitudinal acoustic modes are found to be excited in the afterburner test section. With heat release, three additional tangential modes are excited. By the use of eight probes in the circumferential cross section of afterburner it was possible to identify the tangential modes by their respective phase shift in the experiments.
Comparison of normalized acoustic pressure and phase with and without the incorporation of perforate liner is made to study the eﬀectiveness of the screech liner in attenuating the amplitude of screech modes. By the analysis, conclusion is drawn about modes that get eﬀectively attenuated with the presence of perforate liner. Parametric study of screech liner porosity factor of 1.5 % has not shown appreciable attenuation. Whereas with 2.5 % porosity signiﬁcant attenuation is noticed, but with 4 % porosity, the gain is very minimal. Hence, the perforate screech liner with the porosity of 2.5 % is ﬁnalized.
From the rig runs, ﬁrst pure screech tangential mode and second screech coupled tangential modes are captured. The theoretical frequencies for ﬁrst and second tangential modes with their phases are comparable with experimental results. Though third tangential mode is predicted, it was not excited in the experiments. There was certain level of deviation in the prediction of these frequencies, when compared to the experimentally obtained values. For this test section of length to diameter ratio of 5, no radial modes are encountered both in the analysis and experiments in the frequency range of interest.
In summary, an acoustic model has been developed for the afterburner combustor, taking into account the combustion response, the screech liner and the nozzle to study the acoustic instability of the afterburner. The model has been validated experimentally for screech frequencies using a model test rig and the results have given suﬃcient conﬁdence to apply the model for full scale afterburners as a predictive design tool.|
|Appears in Collections:||Aerospace Engineering (aero)|
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