Combustion Instabilities in Rocket Engines

Supervisor	Subject
Prof. Dr.-Ing. Thomas Sattelmayer	combustion instabilities, modelling
Editor	Cooperation/Funding
Chemnitz Alexander, M.Sc.; Kings Nancy Dr.-Ing.	This project is funded by the DFG in the framework of the Sonderforschungsbereich Transregio 40

Motivation

Fig.1: Pressure amplitude distribution of the T1L2 mode in a rocket combustion chamber.

Combustion instabilities are a crucial issue in rocket engine development. The coupling between pressure oscillations in the combustion chamber and heat release fluctuations can lead to high mechanical and thermal loads, with possible consequences ranging from performance loss to the destruction of the engine. Since instabilities are usually perceived in late stages of the design process, remedies for this issue are costly, potentially even requiring extensive full scale testing.

Project Goals

This project aims at developing methods to reliably predict the combustion stability behavior of an engine already in early design stages. Based on a linear stability analysis procedure developed previously in this project, instability mechanisms are studied numerically and the modeling approach is enhanced further.

Modelling Approach

The approach followed for the characterization of a rocket combustor's thermoacoustic stability behavior relies on the description of fluctuating fields via perturbation equations. Several sub-models are employed to account for different aspects of thermoacoustics. The flame response is modeled via a Flame Transfer Function (FTF), acting as source term in the energy equation. Additional components like dampers or injection elements are efficiently included via impedance boundary conditions and scattering matrices.

An eigenvalue analysis in frequency space is performed using the Linearized Euler Equations (LEE) to describe the chamber acoustics. The approach is based on finite element computations conducted with the software COMSOL. This allows for an efficient characterization of a chamber's stability behavior.

Experimental Studies

In addition to the numerical works, experimental investigations are carried out on a cold-flow subscale rocket combustion chamber (Fig. 3). Several dynamic pressure sensors allow for the characterization of the internal sound fields.

Fig. 3:Experimental cold flow test configuration [1].

Results and Ongoing Work

Fig. 4: Frequency of the first transversal eigenmode via absorber ring reflection coefficient [1].

After the binary stability behavior of a rocket combustion chamber run by the DLR was correctly reproduced, current efforts focus on two aspects: increasing the understanding of the underlying mechanisms of rocket engine combustion instabilities by application as well as enhancing the stability analysis approach.

For the cold-flow setup the influence of a generalized absorber ring has been studied numerically by a systematic variation of the ring's reflection coefficient (Fig. 4). An enhanced model for the frequency dependent impedance of a single absorber has been developed and validation with experimental data showed good agreement [1]

The influence of radial mean flow gradients on the thermoacoustic stability of a combustion chamber is under investigation [2]. Moreover, stability analyses are carried out for Virtual Thrust Chamber Demonstrators designed by ArianeGroup to feature key technology trends for next generation rocket engines.

Referencs

[1] Chemnitz, Alexander; Kings, Nancy; Sattelmayer, Thomas: Modification of Eigenmodes in a Cold-Flow Rocket Combustion Chamber by Acoustic Resonators, Journal of Propulsion and Power, 35, 4, 765–779, 2019

[2] Chemnitz, Alexander and Sattelmayer, Thomas: Influence of Radial Stratification on Eigenfrequency Computations in Rocket Combustion Chambers, 8th EUCASS, 2019