During their operational life-time, actively cooled liners of cryogenic combustion chambers are known to exhibit a characteristic so-called doghouse deformation, pursued by formation of axial cracks. The present work aims at developing a model that quantitatively accounts for this failure mechanism. High-temperature material behaviour is characterised in a test programme and it is shown that stress relaxation, strain rate dependence, isotropic and kinematic hardening as well as material ageing have to be taken into account in the model formulation. From fracture surface analyses of a thrust chamber it is concluded that the failure mode of the hot wall ligament at the tip of the doghouse is related to ductile rupture. A material model is proposed that captures all stated effects. Basing on the concept of continuum damage mechanics, the model is further extended to incorporate softening effects due to material degradation. The model is assessed on experimental data and quantitative agreement is established for all tests available. A 3D finite element thermo-mechanical analysis is performed on a representative thrust chamber applying the developed material-damage model. The simulation successfully captures the observed accrued thinning of the hot wall and quantitatively reproduces the doghouse deformation.
Accurate information on the heat transfer evolution along the hot gas wall of a rocket chamber is a crucial prerequisite for optimizing an engine's regenerative cooling circuit design. To improve Dasa's available data basis in this field, a calorimeter model combustor was developed and tested. The main objective of these experiments was to characterize the heat transfer behaviour of different co-axial injection element configurations under various operational conditions (P c : 100-=-120 bar, O/F: 6.0-^7.6) first without and then with gaseous hydrogen wall film cooling.The experimental results disclosed that the co-axial injection designs under investigation produce local wall heat loads that differ up to 25 percent. These differences in heat transfer become more apparent at axial distances of about 100 mm from the injector and beyond. An important finding was also that hydrogen wall film cooling employing the present slot injection principle is in fact much more effective when compared to wall element mixture ratio trimming under similar operational conditions.A major interest of the current program was finally, to apply Dasa's axisymmetric, Euler/Lagrange CFD-code CryoROC to the calorimeter chamber tests and assess its capability in this field. In doing so, it turned out that basically two parameters are sufficient to adapt the simulation to the experiment of interest, i.e. a characteristic droplet size distribution parameter, which mainly affects the axial build-up of the wall heat flux, and a global scaling factor for fine tuning of the observed wall heat flux level. The CryoROC results also demonstrated that a realistic simulation the wall heat transfer problem in a cryogenic rocket engine requires to * Senior Member AIAA
The demand for a more comprehensive engineering tool for design and parametric investigations of thrust-chamber relevant heat transfer is pushing the improvement of coolant and hot gas side prediction tools. Regenerative Coolant Flow Simulation (RCFS) [1], Astrium in-house developed one-dimensional (1D) tool to compute hot gas and coolant side heat transfer in a coupled approach, is based on the hot gas side Cinjarew approach which has its origin in the late 1960s. This tool was used as a starting basis for the development and validation of a further improved method. Over the past years, Astrium Space Transportation (ST) has continuously expanded the knowledge in this ¦eld. In addition, subscale hot ¦rings, using di¨erent propellant combinations and injection conditions, relevant to open and closed cycle applications, were used for the second RCFS generation ¡ the RCFS-II.
The creation of validation data for CFD, thermal, structural and life time analyses of actively cooled thermally loaded wall structures of rocket engines by means of Thermo-Mechanical Fatigue (TMF)-tests is discussed in this paper. During a TMF test, only a small section of the hot gas wall of the real engine (the so called TMF panel) is tested. For such a TMF panel, realistic cooling conditions similar to a full scale rocket engine are chosen. The 2d measurement of the thermal field of the heat loaded structure provides (together with the measurement of the temperature, pressure and mass flow rate of the coolant of the TMF panel) data for the combined validation of the CFD analysis of the coolant flow and the thermal analysis of the wall structure. The measurement of the deformation of the thermally loaded structure provides (together with the already determined temperature distribution and the above mentioned pressure measurements of the cooling channels) data for the validation of the structural analysis of the thermally loaded structure. Counting the number of laser loading cycles (laser on-off) until the TMF panel fails (by cracks appearing on the laser loaded side of the cooling channels) provides data for the validation of (either post processing or damage parameter based) life time analyses of thermally loaded structures.
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