A program for three-dimensional flow and heat transfer analysis in a rocket engine cooling channel of high aspect ratio is described. The program, which is based on the parabolized Navier-Stokes equations, was validated by comparing flow and heat transfer calculations with measurement data as found in the literature. Calculations were performed for cooling channel flow of an actual 100 bar engine. Calculated hulk temperature rise and pressure drop data agree well with data obtained from test results. A parameter study shows the effects of turbulence intensity, wall surface roughness, heat transfer boundary caridition, and near-wall flow model.It is shown that the program can be successfully used for three-dimensional analysis of flow and heat transfer in rocket engine cooling channels. It is sufficiently CPU-time efficient and easy to handle to be employed in support of the design process of present and future rocket engine combustion chdmbers.
NomenclatureA = cross-section area AR = aspect ratio = WW c = specific heat at constant pressure Dh= hydraulic diameter h =channel's height (radial direction) m =mass flow rate Nu = Nusselt number O/F= mixture ratio p = static pressure q, = wall heat flux Rc =channel's mean radius of curvature Re =Reynolds number based on hydraulic diameter T = static temperature Tbun = bulk temperature Tu =turbulence's intensity u,v,w = velocity components in cross-stream and streamwise uy = shear velocity w = channel's width (spanwise direction) X =coordinate along the combustion's chamber axis P directions, respectively * DASA Spa-Communications and Propulsion Systems Division. SOW MUnchm 80 e
The prediction of heat transfer characteristics in a regeneratively cooled rocket combustion chamber is one of the most important and most challenging tasks in the design work of a high performance rocket engine. This task becomes even more ambitious, when multipleuseand life capabilities ofseveral hours are requested for the engine.DASA, reyp. the design engineers of the former Boelkow rocket engine team were confronted with the problem from the very beginning of their staged combustion cycle demonstration program ( 1964-68 ), because this type of engine cycle, today well known to every rocket engineer, offers the potential of realizing chamher pressures up to 300 bar.Thisreport summarizesthe effortspentat DASAfrom thattimeup to now tocontinouslyimproveand validatethe heatflux prediction code by the information available from their combustion chamber development programs.The DASA heat transfer prediction code is now well suited for a straight forward design of hydrogedoxygen combustion chambers in thepressurerangefrom30to200harandth~stlevels from 50 to loo0 kN.However, new projects in the field of reuseable launchers and the general trend to most economic designs of space-transportation systems will require even more reliability in life prediction and therefore also further improvement of heat transfer prediction.
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