Dirk Greuel

Effusion Cooling of Throat Region in Rocket Engines Applying Fibre Reinforced Ceramics

This report summarizes the experimental work in the field of effusion cooling done with C/C (carbon fibres in carbon matrix) for a porous cylinder in combination with a porous nozle. The experiments presented here have been realized using gaseous hydrogen at ambient temperature as coolant. These recent DLR experiments have been performed for combustion chamber pressures ranging between o.8 MPa and 1.2 MPa at a constant oxidizer/fuel mixture ratio 6.5. For this basic experiments there is a micro combustor at DLR. The porosity for the cylinder and nozzle was e = 26%. The coolant mass flow for the porous cylinder and nozzle has been varied in a wide range to get informations about the influence of the cooling film from the cylinder to the nozzle.

Fabrication of TBC-armored rocket combustion chambers by EB-PVD methods and TLP assembling

Abstract A thermal barrier coating (TBC) system for rocket chambers made of Cu-based high strength alloys has been developed in a pilot project in line with EB-PVD (electron-beam physical vapor deposition) technology aiming at TBC application on Cu-based walls of real rocket combustion chambers. The TBC system consists of a metallic bond coating compatible with Cu-based material and an yttria partially stabilized zirconia TBC. The TBC overlayer is a distinctive ceramic structure designed for an exceptionally low Young’s modulus to withstand the extreme mismatch stresses between the internally LN-cooled high thermal expansion Cu metal base and the low thermal expansion hot ceramic shell. The TBC system has been qualified under close-to-service conditions on cylindrical LH2-cooled combustion chamber segments, where they have performed superior. As EB-PVD technology is a line-of-sight process that is rather able to coat internal cavities, a transient liquid phase (TLP) joining technique for fully coated parts has been developed, that allows to assemble complete components out of vapor-accessible fully coated parts. It is capable, e.g. to incorporate sinuous cooling passages in the throat areas of combustion chambers, and/or to assemble oversized parts out of smaller components by maintaining parent metal properties. A manufacturing process is outlined for making internal TBC armored combustion chambers.

Empirical Verification of Effusion Cooled CMC Rocket Thrust Chambers

Thrust chambers are one of the most sensitive components in rocket propulsion systems due to safety and efficiency related to reasonable costs. Competitive space transportation systems ask for such low cost and high sophisticated solutions. A very promising approach in this field is the development of an effusion cooled CMC combustion chamber design, which offers a new perspective on the way to reliable future cryogenic rocket engines. The undoubted advantages of transpiration cooling under the critical view of efficiency, damage tolerance and low cost aspects can be accomplished with a relatively simple and low weight concept using carbon fiber composite materials. In the recent years DLR works on the application of effusion technology. The paper illustrates the empirical development steps accompanied by numerical simulations until a break through in form of the latest high performance tests early in 2005.