Prediction and Analysis of Heat Transfer in Small Rocket Chambers

Kirchberger, Christoph; Wagner, Róbert; Kau, Hans-Peter; Soller, Sebastian; Martin, Philip; Bouchez, Marc; Bonzom, Christophe

doi:10.2514/6.2008-1260

Cited by 29 publications

(11 citation statements)

References 5 publications

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“…The advantage of this approach is a satisfying accuracy maintaining a reasonably fast simulation of the conjugate heat transfer from the hot gas into the cooling fluid. Information on THERMTEST as well as a comparison with experimental data and calculations from commonly available CFD code has been published [10][11][12] .…”

Section: A Simulation Toolmentioning

confidence: 99%

Assessment of Analytical Models for Film Cooling in a Hydrocarbon/GOX Rocket Combustion Chamber

Kirchberger¹,

Schlieben²,

Haidn³

2012

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

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The understanding of the mechanisms of heat transfer and advanced cooling techniques is a key for mastering the designing of reliable and efficient rocket combustion chambers. However, the effects determining film cooling efficiency as well as influences on the combustion process itself under typical conditions for rocket engines are still insufficiently understood. Therefore, the Institute for Flight Propulsion (LFA) of the Technische Universität München operates a rocket combustor test facility which allows testing at application-relevant conditions. The experimental work is complemented by numerical investigations, where in day-to-day operations also analytical and semi-empirical correlations are used. This paper presents an assessment of different analytical and semiempirical film cooling models known from open literature and their application to kerosene film cooling in a GOX/ kerosene combustion chamber. Nomenclature AcronymsCC = Combustion chamber CFD = Computational Fluid Dynamics GOX = Gaseous oxygen MR = Mixture ratio (here oxygen/kerosene mass flow ratio) NHFR = Net Heat Flux Reduction (cooling effectiveness) THERMTEST = Tool for thermal simulation of rocket combustion chamber by TUM Greek Letters  = heat transfer coefficient  = adiabatic film cooling effectiveness,     c cool c aw T T T x T x      = film part, relative film mass flow,  Injector , Kerosene Film Film m m / m       Latin Letters Indices m  = mass flow rate aw = adiabatic wall p = pressure c = chamber, core stream T = temperature cool = coolant (esp. film) O/F = mixture ratio e = entrainment r = radius inj = injector s = slot height t = total x = distance from film applicator throat = regarding nozzle throat

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Section: A Simulation Toolmentioning

confidence: 99%

Assessment of Analytical Models for Film Cooling in a Hydrocarbon/GOX Rocket Combustion Chamber

Kirchberger¹,

Schlieben²,

Haidn³

2012

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

Self Cite

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“…In the past, Thermtest has been validated and adapted for the use of kerosene/oxygen as propellants in small single-element combustion chambers featuring swirl or double swirl coaxial injector elements. Information on Thermtest as well as a comparison with experimental data and calculations from commonly available computational §uid dynamics (CFD) code has been published [19,20].…”

Section: Thermtest Simulationmentioning

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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“…State of the art and existing models have been used to define the configurations to be tested [11] below.…”

Section: Increased Engine Thermal Efficiency: Novel Cooling Conceptsmentioning

confidence: 99%

Achievements obtained on Aero-Thermal Loaded Materials for High-Speed Atmospheric Vehicles within ATLLAS

Steelant¹

2009

16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference

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The project ATLLAS (Aerodynamic and Thermal Load Interactions with Lightweight Advanced Materials for High Speed Flight) kicked off in October 2006. Its main objective is to evaluate and assess potential high-temperature resistant materials for sustained superand hypersonic flight. This covers both materials for the external geometry as well as for the internal combustor geometries. The project, led by ESA-ESTEC, is co-funded by the European Commission under the 6 th Framework Program and a consortium of 13 partners from industry, research institutions and universities.The objective is to identify and assess lightweight advanced materials which can withstand ultra high temperatures and heat fluxes enabling high-speed flight above Mach 3. At these high speeds, classical materials used for airframes and propulsion units are not longer feasible and need to be replaced preferably by high-temperature, lightweight materials and if required, some parts with active cooling. Both metallic and non-metallic materials are tested extensively including Ni-based Hollow Sphere Packings, Ultra High Temperature Composites and several (non)-oxide Ceramic Matrix Composites. Envisaged cooling techniques for CMC based combustion liners are film, effusion, transpiration and regenerative cooling.

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Prediction and Analysis of Heat Transfer in Small Rocket Chambers

Cited by 29 publications

References 5 publications

Assessment of Analytical Models for Film Cooling in a Hydrocarbon/GOX Rocket Combustion Chamber

Assessment of Analytical Models for Film Cooling in a Hydrocarbon/GOX Rocket Combustion Chamber

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

Achievements obtained on Aero-Thermal Loaded Materials for High-Speed Atmospheric Vehicles within ATLLAS

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