New experimental results on transpiration cooling for hydrogen/oxygen rocket combustion chambers

Abstract

Introduction While previous experimental activities of DLR focused on low pressure applications of combustion chamber segments cooled with either water or gaseous hydrogen wich have been performed at the DLR test facilities P6.1 and M3, this paper presents the most recent experimental results obtained at the high pressure GH^/LO^ test facility P8. For this purpose, DLR model combustor B, a 10 MPa regeneratively cooled experimental combustion chamber having modular design was applied to implement transpiration cooled segments at different downstream positions of the combustion chamber. These recent DLR experiments have been performed for combustion chamber pressures ranging between 3.5 and 10 MPa and a constant oxidizer/fuel mixture ratio of 6.5. Screening tests at lower pressures showed nickel alloys to be favorable for transpiration cooling purposes. Hence, for the high pressure tests reported here only the material porosity was changed with a variation of the typical pore size between 1 /^m and 5 /im. The mass flow rate of the coolant, ambient gaseous hydrogen, has been varied in a wide range to increase the existing data base for modeling and verification purposes as well as to gain experience and to check the limits of operability of this cooling technique especially at the low end of coolant mass flow rates. After a brief description of the governing equations and some available models as well as a short presentation of the experimental setup and the operating conditions, the paper focuses on the experimental results and their comparison with the models derived for transpiration cooling applications. Based on the results of this comparison a modification of the model of Kays and Crawford is suggested. •Member AIAA Copyright ©1998 Erhan Serbest, O.Haidn, K.Frohlke. Published by the American Institute of Aeronautics and Astronautics,Inc. with permission For future space transportation systems, advanced cryogenic combustion chambers are necessary for both Reusable Launch Vehicles (RLV) and Expendable Launch Vehicles (ELV) to meet the requirements for higher payload capabilities. Independent of either ELV or RLV applications, the major requirement for these high pressure combustion chambers is reliability. In future liquid rocket engines with combustion chamber pressures up to 25 MPa thermal loads of the chamber will lay beyond 140 MW/m? in the throat region [1]. Conventional regeneratively cooled combustion chambers made from milled slotted liners of high conductivity copper alloys, will hardly fulfill the requirements for a further increase of both chamber reliability and life. Presently, various candidate techniques such as thermal barrier coatings, improved liner materials, microchannel cooling structures, film cooling via injector trimming and transpiration cooling are discussed and tested by industry and research laboratories. All these methods seem to be means to overcome the problem of the severe thermal loads to the combustion chamber walls. Transpiration cooling, although a rather old idea [2], has not yet been applied for chamber liner cooling purposes in flight engines due to insufficient knowledge about local heat flux distribution, coolant film stability, and coolant mass flow rate controlability. Furthermore, technologies for the production of porous materials having reliable and reproducible parameters are still under development. The benefit of this cooling technique is that plastic deformation of the chamber wall is reduced and hence life time problems due to fatique may be negliable. A problematic issue might be the manufacturing process of the porous material according to the needs in a combustion chamber with varying combustion chamber wall curvature. For investigations in this field there are two facilities available at DLR Lampoldshausen. For basic reasearch there is the M3 Micro Combustor with pressures up to 2 MPa, see [3], [4], [6]. For pressures and mass flow rates the P8, a facility which allows for combustion chamber pressures above 30 MPa and mass flow rates which exceed 11 kg/s, is in use. Governing Equations and Models Reference Heat Flux Without Blowing The Bartz equation (1) in its modified form is a quite standard way for heat flux determination in rocket combustion chambers. Based on this equation a reference heat flux for the non-blowing case is defined which will be later used to evaluate the performance of the transpiration cooling. This reference heat flux is presented in form of a Stanton number StQ