In the past decades, several feedback mechanisms for longitudinal acoustic modes in gas turbine combustors have been investigated. These mechanisms are successfully used in predictive tools like acoustic network models to analyze low-frequency instabilities in combustion systems. In contrast, little is known about high-frequency oscillations — fluctuations at several kHz. Most theories are derived from experimental investigations of afterburners in the 1950s and 1960s, indicating an interaction of vortex shedding, fluctuating vorticity and heat release. In this work a different feedback mechanism for high-frequency oscillations in cylindrical flame tubes related to transverse acoustic modes is suggested and analysed: Transverse acoustic pressure fluctuations are linked to an oscillating velocity field. A time-dependent but periodic displacement field can be derived from these velocity fluctuations. The model assumes that the zone of heat release is displaced by the velocity fluctuations. Pressure oscillations and periodically deflected heat release lead to a contribution to the Rayleigh criterion without fluctuations in the global heat release. This effect is studied in a circular cross section presuming a circular zone of heat release. Expressions for the displacement of the flame front are derived from the analytical solution of the wave equation in cylindrical geometries assuming a quiescent medium, constant density and speed of sound. The Rayleigh criterion is integrated and growth rates are evaluated whereas damping effects are neglected as they are not subject to this study. Characteristics of the model are assessed and compared to experimental observations to check the validity and the applicability of the theory.

No abstract

The increasing importance of decentralized energy production based on renewable resources requires gas turbine systems due to their low emissions and flexible energy conversion. Therefore, a suitable hybrid power plant demonstrator consisting of an SOFC (solid oxide fuel cell) coupled to an MGT (micro gas turbine) is being set up at the German Aerospace Center (DLR). This facility requires a burner concept for low calorific gases capable of combusting the exhaust products of the fuel cell system anode side, here referred to as SOFC off-gas. The combustor behavior for the demonstrator system is investigated using an atmospheric combustor test rig at DLR. The main aspect of this work is the combustor operation inside the power plant system with varying power demands and also varying methane contents, representing biogas operation. This is leading to operating points with very low heating values (LHV) which require a flame stabilization strategy via direct addition of natural gas / biogas into the SOFC off-gas before entering the combustor. This is tested in view of expected impacts on electrical system efficiency and other critical system parameters. The combustion system is furthermore investigated in view of CO emissions in various significant operating points.

No abstract

Recognizing the attention currently devoted to the environmental impact of aviation, this three-part publication series introduces two new aircraft propulsion concepts for the timeframe beyond 2030. The first part focuses on the novel steam injecting and recovering aero engine concept. In the second part, the free-piston composite cycle engine concept is presented. Complementary to the two technical publications, this third part describes the cooperative project, which was initiated by an interdisciplinary consortium, aiming at the demonstration and the proof of concept of both aforementioned aero engine concepts. At the beginning of the project, simulations on propulsion, aircraft system, and test bench level will be conducted. On this basis, preliminary tests and fundamental experiments are planned in order to establish a solid basis for the demonstration. Finally, a system demonstration will be carried out at the laboratory level. Thus, the project allows for a final judgement on both the feasibility of the new concepts and the attainability of the requirements for future aircraft propulsion systems.

A lean, swirl-stabilized gas turbine model combustor is simulated with a stochastic approach for combustion noise prediction. The employed hybrid and particle based method, FRPM-CN (Fast Random Particle Method for Combustion Noise Prediction) reconstructs temperature variance based direct combustion noise sources from local CFD-RANS turbulence and flow field statistics. Those monopole sound sources are used as right hand side forcing of the Linearized Euler Equations. First, findings from steady state CFD simulations are validated with experimental results. It is shown that the employed RANS models accurately reproduce the experimental flow field and combustion. Turbulence is treated with a two equation model and a global reaction mechanism is utilized for combustion. Subsequently, the specifications of the CCA (Computational Combustion Acoustics) setup is introduced and selected pressure spectra of the acoustics simulations are compared to experimental results, showing that FRPM-CN is able to deliver absolute combustion noise levels for the investigated burner at low computational costs.

Combustion noise in the laboratory scale PRECCINSTA burner is simulated with a new, robust and highly efficient approach for combustion noise prediction. The applied hybrid method FRPM-CN (Fast Random Particle Method for Combustion Noise prediction) relies on a stochastic, particle based sound source reconstruction approach. Turbulence statistics from reacting CFD-RANS simulations are used as input for the stochastic method, where turbulence is synthesized based on a first order Langevin ansatz. Sound propagation is modeled in the time domain with a modified set of linearized Euler equations and monopole sound sources are incorporated as right hand side forcing of the pressure equation at every timestep of the acoustics simulations. First, reacting steady state CFD simulations are compared to experimental data, showing very good agreement. Subsequently, the computational combustion acoustics setup is introduced, followed by comparisons of numerical with experimental pressure spectra. It is shown that FRPM-CN accurately captures absolute combustion noise levels without any artificial correction. Benchmark runs show that the computational costs of FRPM-CN are much lower than that of direct simulation approaches. The robustness and reliability of the method is demonstrated with parametric studies regarding source grid refinement, the choice of either RANS or URANS statistics and the employment of different global reaction mechanisms.

A new highly efficient, hybrid CFD/CAA approach for broadband combustion noise modeling is introduced. The inherent sound source generation mechanism is based on turbulent flow field statistics, which are determined from reacting RANS calculations. The generated sources form the right-hand side of the linearized Euler equations for the calculation of sound fields. The stochastic time-domain source reconstruction algorithm is briefly described with emphasis on two different ways of spatial discretization, RPM (Random Particle Method) and the newly developed FRPM (Fast RPM). The application of mainly the latter technique to combustion noise (CN) prediction and several methodical progressions are presented in the paper. (F)RPM-CN is verified in terms of its ability to accurately reproduce prescribed turbulence-induced one- and two-point statistics for a generic test and the DLR-A jet flame validation case. Former works on RPM-CN have been revised and as a consequence methodical improvements are introduced along with the progression to FRPM-CN: A canonical CAA setup for the applications DLR-A, -B and H3 flame is used. Furthermore, a second order Langevin decorrelation model is introduced for FRPM-CN, to avoid spurious high frequency noise. A new calibration parameter set for reacting jet noise prediction with (F)RPM-CN is proposed. The analysis shows the universality of the data set for 2D jet flame applications and furthermore the method’s accountance for Reynolds scalability. In this context, a Mach number scaling law is used to conserve Strouhal similarity of the jet flame spectra. Finally, the numerical results are compared to suitable similarity spectra.