LOX-Hydrocarbon Preparatory Thrust Chamber Technology Activities in Germany

Preclik, D.; Hagemann, G.; Knab, O.; Mäding, Chris; Haeseler, D.; Haidn, Oskar; Woschnak, A.; DeRosa, Marc L.

doi:10.2514/6.2005-4555

Cited by 17 publications

(9 citation statements)

References 1 publication

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“…[1][2][3][4][5] Methane is known to be a low-density hydrocarbon having advantages that are similar to those of both kerosene and hydrogen. It is also a cryogenic fuel having a density about six times higher than that of hydrogen, and it requires less insulation and fewer handling concerns than hydrogen.…”

Section: Introductionmentioning

confidence: 99%

Flowfield and Heat Transfer Characteristics of Cooling Channel Flows in a Methane-Cooled Thrust Chamber

Negishi¹,

Daimon²,

Kawashima³

et al. 2012

48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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In recent years, methane has attracted attention as a propellant for liquid rocket engines because of its various advantages compared to typical propellants such as hydrogen. When methane is used as a coolant for a regenerative cooling system, its near-critical thermodynamic and transport properties experience large variations because its critical pressure is higher than that of typical propellants; this significantly influences the flowfield and heat transfer characteristics. Therefore, adequate understanding of the flowfield and heat transfer characteristics of methane in regenerative cooling channels is a prerequisite for future engine development. In this study, conjugated coolant and heat transfer simulations were performed to investigate the flowfield and heat transfer characteristics of transcritical methane flows in a sub-scale methane-cooled thrust chamber. The computed results were validated against experimental data measured in hot firing tests. They compared well with the measured pressures and temperatures in cooling channels, and wall temperatures were within the permitted levels. Detailed flow analysis revealed peculiar flow structures in the cooling channel: a strong secondary flow induced in the concave-heated part in the channel throat section and the coexistence of two different gas phases-ideal and real-in a single cross-section in the cylindrical region. A high wall temperature appeared in the cylindrical region of the thrust chamber under the considered conditions; this was due to the heat transfer deterioration induced by an M-shaped velocity profile and a turbulentheat flux reduction. Nomenclature a, b,  = coefficients in cubic-type equation of state C p = constant pressure heat capacity G = mass flow rate H = channel height h = heat transfer coefficient Pr = Prandtl number p = pressure q = heat flux R = gas constant Re = Reynolds number R u = universal gas constant s = normal distance T = temperature u = streamwise velocity W = molecular weight x, y, z = coordinate system Z = compressibility factor y + = normalized wall distance 2  = thermal conductivity  = viscosity  = density Subscripts b = bulk c = coolant cr = critical cw = coolant-side wall gw = hot-gas-side wall in = inlet out = outlet t = turbulent throat = throat pcr = pseudo-critical w = wall

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Section: Introductionmentioning

confidence: 99%

Flowfield and Heat Transfer Characteristics of Cooling Channel Flows in a Methane-Cooled Thrust Chamber

Negishi¹,

Daimon²,

Kawashima³

et al. 2012

48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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“…The demanding issues in terms of high operational and handling costs of such propellants increased the attention for hydrocarbons in the predevelopment of future launch vehicles [1]. Liquid oxygen / hydrocarbons rocket engines have the advantage in fact of being relatively low cost, low pollution, and high performance.…”

Section: Introductionmentioning

confidence: 99%

Injector characterization for a gaseous oxygen-methane single element combustion chamber

Celano

Silvestri

Schlieben

et al. 2016

Progress in Propulsion Physics

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The results from an experimental investigation on an oxygenmethane single-injector combustion chamber are presented. They provide detailed information about the thermal loads at the hot inner walls of the combustion chamber at representative rocket engine conditions and pressures up to 20 bar. The present study aims to contribute to the understanding of the thermal transfer processes and to validate the in-house design tool Thermtest and a base for an attempt to simulate the §ame behavior with large-eddy simulation (LES). Due to the complex §ow phenomena linked to the use of cryogenic propellants, like extreme variation of §ow properties and steep temperature gradients, in combination with intensive chemical reactions, the problem has been partially simpli¦ed by injecting gaseous oxygen (GOx) and gaseous methane (GCH 4 ). external combustion chamber width, m c speci¦c heat capacity, J/(kg·K)' E in total energy entering the control volume, W ' E out total energy leaving the control volume, W

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“…ithin the last ten recent years, the propellant combination LOX/CH 4 has received considerable attraction worldwide as a propellant combination for space propulsion applications [1][2][3][4][5]. The advantages are numerous: a specific impulse better than that of oxygen/kerosene, reduced cost and complexity in in handling compared to hydrogen and reduced requirements for turbomachinery.…”

Section: Introductionmentioning

confidence: 99%

“…3 Senior research fellow, Chemical Kinetics Department, JanHendrik.Starcke@dlr.de. 4 Prof. of Technical University Munich, haidn@lfa.mw.tum.de , AIAA Senior Member W simulations of rocket combustion chamber under pressure 20 bar was developed through simulations and reactions/species local sensitivity analysis performed for 65 experimental targets for ignition delay times and 15 targets for laminar flame speed. These experimental measurements cover the next operating conditions: p 5 = 1-50 bar, T 5 = 940K -2100K, f = 0.5−2 ( for ignition delay time); p = 1-60 bar, T 0 = 300K, f = 0.6−1.4 (for laminar flame speed).…”

Section: Introductionmentioning

confidence: 99%

Skeletal Mechanism of the Methane Oxidation for Space Propulsion Applications

Slavinskaya¹,

Abbasi²,

Weinschenk³

et al. 2016

52nd AIAA/SAE/ASEE Joint Propulsion Conference

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A skeletal chemical kinetic model for СН 4 /air combustion with 100 reactions and 24 chemical species was developed from the detailed mechanism, with 42 species and 298 reactions. The mechanism reduction was performed with the multi target reduction strategy realized in the in-house developed DLR RedMaster code. RedMaster is able to analyze the different chemical processes (ignition delay time and laminar flame speed) in the given time and height points. The obtained reduced model describes satisfactory experimental data for ignition delay and flame speed under conditions: p 5 = 1-50 bar, T 5 = 940K-210K, f = 0.5−2; p = 1-60 bar, T 0 = 300K, f = 0.6−1.4. Some problems related to reaction mechanism reduction are analyzed.

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LOX-Hydrocarbon Preparatory Thrust Chamber Technology Activities in Germany

Cited by 17 publications

References 1 publication

Flowfield and Heat Transfer Characteristics of Cooling Channel Flows in a Methane-Cooled Thrust Chamber

Flowfield and Heat Transfer Characteristics of Cooling Channel Flows in a Methane-Cooled Thrust Chamber

Injector characterization for a gaseous oxygen-methane single element combustion chamber

Skeletal Mechanism of the Methane Oxidation for Space Propulsion Applications

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