Three-dimensional flow analysis in a rocket engine coolant channel of high depth/width ratio

Froelich, A.; Immich, H.; Lebail, F.; Popp, M.; Scheuerer, G.

doi:10.2514/6.1991-2183

Cited by 16 publications

(9 citation statements)

References 6 publications

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“…3. The equations were, however, rewritten such that a general transformed coordinate system is now used, making the formulation more general and making the program easier to use (input, output, etc.…”

Section: Fi-mentioning

confidence: 99%

Numerical analysis of high aspect ratio cooling passage flow and heat transfer

Lebail¹,

Popp²

1993

29th Joint Propulsion Conference and Exhibit

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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

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Section: Fi-mentioning

confidence: 99%

Numerical analysis of high aspect ratio cooling passage flow and heat transfer

Lebail¹,

Popp²

1993

29th Joint Propulsion Conference and Exhibit

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“…The 3-D CFD program VADUCT ('variable area ducts') was developed by Dasa to simulate the three dimensional flow and heat transfer in typical coolant channel geometries [8], [9]. Model experiments with air in large scale coolant channels (scale 25:1) heated on the hot gas side performed at the Technical University Dresden [6] showed good agreement between calculated and measured flow behaviour.…”

Section: Heat Transfer Analysismentioning

confidence: 97%

Cryogenic liquid rocket engine technology developments within the German National Technology Programme

Immich

Mayer²

1997

33rd Joint Propulsion Conference and Exhibit

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In a joint technology programme Daimler-Benz Aerospace (Dasa) and the German Aerospace Research Establishment (DLR) are presently investigating technologies for advanced thrust chambers of reusable, high performance cryogenic rocket engines. This Technology Programme on Cryogenic Rocket Propulsion (TEKAN) has been established in 1996 under sponsorship of the German Space Agency DARA. In this programme the extensive experience of Dasa with regeneratively cooled liquid rocket engine combustion chambers is combined with the research oriented activities of DLR in liquid rocket propulsion in order to reach the targets.The present paper describes and discusses the technology developments within TEKAN which have been initiated.Rocket engine cycle analyses for different reusable launch vehicle concepts are performed in order to define the requirements for the thrust chamber components of future launch vehicle propulsion systems. In addition, a simulation tool is developed for transient engine analysis. Basic theoretical and experimental work on combustion chamber heat transfer, cooling, lifetime, and combustion modelling is continued in this programme. Technology developments of the main thrust chamber components as injectors and combustion chamber technologies for high lifetime (thermal barrier coatings, elastic liner technologies) are investigated with experimental test programmes on subscale level and theoretical analyses. Technologies for expander cycle engines are developed and the applicability of carbon fibre reinforced composite materials in cryogenic thrust chambers is investigated.

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“…Most investigations implemented coupled heat transfer to simulate asymmetrically heated cooling channels because the temperature fields of the fluid region obtained with and without wall heat conduction are significantly different [2]. Previous studies showed that the selected geometries, fluids (hydrogen, nitrogen, methane), and boundary conditions were close to those expected in the regenerative cooling system of rocket engines (Froelich et al [3], Neuner et al [4], Jung et al [5], DiValentin and Naraghi [6], Kang and Sun [7], Pizzarelli et al [2,[8][9][10][11][12][13][14], Negishi et al [15], Ruan and Meng [16], and Wang et al [17]). A high aspect ratio cooling channel (HARCC) has been widely used in the regenerative cooling system of rocket engines.…”

Section: Introductionmentioning

confidence: 94%

Coupled Heat Transfer of Supercritical n-Decane in a Curved Cooling Channel

Liang

Liu

Pan

2016

Journal of Thermophysics and Heat Transfer

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The coupled heat transfer of supercritical-pressure n-decane in a concave heated, 90 deg bend, square channel was investigated numerically. Both the geometry and boundary conditions were close to those expected in active-cooled scramjet engines. The effect of secondary motions and wall thermal conductivity on heat transfer was studied. The flowfields obtained by the Spalart-Allmaras model and Reynolds stress model were compared. The secondary flow of Prandtl's first kind caused by the crosswise pressure gradient in the curved section was 1.0% of the maximum axial velocity. The secondary flow of Prandtl's second kind caused by Reynolds stress was one magnitude less than that of Prandtl's first kind. The Spalart-Allmaras turbulence model produced reasonable results with less computational effort than the Reynolds stress model. Secondary flow improves heat transfer by enhancing the mixing of the fluid. Furthermore, higher wall thermal conductivity produces more uniformly distributed temperature and heat flux. Nomenclature c p = specific heat at constant pressure, J · kg −1 K −1 D h = hydraulic diameter of cooling channel, m G = mass flux, kg · m −2 s −1 k w = wall thermal conductivity, W·m −1 K −1 P = pressure, Pa R = radii of curved channel, m q w = heat flux, W·m −2 T = temperature, K u x , u y = x, y velocity component of secondary flow, ms −1 u z = axial velocity, m·s −1 u z0 = axial velocity at centerline of channel, m·s −1 ρ = density, kg · m −3

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Three-dimensional flow analysis in a rocket engine coolant channel of high depth/width ratio

Cited by 16 publications

References 6 publications

Numerical analysis of high aspect ratio cooling passage flow and heat transfer

Numerical analysis of high aspect ratio cooling passage flow and heat transfer

Cryogenic liquid rocket engine technology developments within the German National Technology Programme

Coupled Heat Transfer of Supercritical n-Decane in a Curved Cooling Channel

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