Design and Development of a LOX/LCH4 Technology Demonstrator

Salvatore, Vito; Battista, Francesco; Votta, Raffaele; Clemente, Marco Di; Ferraiuolo, Michele; Roncioni, Pietro; Ricci, D.; Natale, Pasquale; Panelli, M; Cardillo, Daniela; Matteis, Paolo

doi:10.2514/6.2012-3935

Cited by 8 publications

(8 citation statements)

References 1 publication

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“…The calculations are performed to second order accuracy with the criteria for convergence being a stabilised integral of total surface heat flux along the chamber wall. (1) as 1031 kgm -3 . The corresponding temperature is calculated at the injector upstream pressure of 10.93 MPa for this LOX density as 118K using NIST 11 tables.…”

Section: Figure 2 Grid 1 44792 Cellsmentioning

confidence: 99%

Comparison of Reactive Flow Simulations for a LOX/CH4 Uni-element Rocket

French

Roncioni

2013

49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

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This study details the reactive flow simulation of a single injector LOX/CH 4 rocket similar to a comparable device, the CIRA Sub Scale Bread Board (SSBB) planned within the context of the HYPROB program funded by the Italian Ministry of University and Research (MIUR). In order to identify a suitable numerical method for use in the design of the SSBB calculations have been performed on the NASA sponsored Penn State LOX/CH 4 uni-element rocket engine for which heat flux data are available. The main features of this rocket are discussed in a former article relating to the experimental setup, methodology and results. The current reactive flow model is based on a twostage injector-chamber and chamber-throat-nozzle approach. In the first stage the injector-chamber region is modelled using a pressure based solver approach. In the second stage the converged results for the chemical distribution at the outlet calculated in the first stage are used as upstream boundary conditions to model the chamber-throat and nozzle region using a density based solver. In the first stage reactive flow is modelled using the Peng-Robinson equation of state whereas the second stage assumes an ideal gas. The reason for this two-stage approach is to reduce CPU time and circumvent problems associated with numerical stability which occur when modelling the complete domain in a single stage and which is validated by direct comparison of predicted wall heat fluxes with experiment. Comparison is also made between the results of this two-stage method and calculations for the complete configuration using different chemistry models. These simulations are performed with a commercial CFD solver to solve the steady state axi-symmetric Navier-Stokes equations. The Laminar Finite Rate and the Eddy Dissipation Concept models are used for the turbulent chemical kinetic coupling and the Jones-Lindstedt model is used for the chemical reactions for liquid oxygenmethane combustion. The results presented indicate the qualitative effects of different turbulence and chemical kinetic models on flow structures as well as a quantitative comparison of wall heat fluxes for different model combinations which are compared with experimental data.

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Section: Figure 2 Grid 1 44792 Cellsmentioning

confidence: 99%

Comparison of Reactive Flow Simulations for a LOX/CH4 Uni-element Rocket

French

Roncioni

2013

49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

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“…3 The main objective is to improve the National system and technology capabilities on liquid rocket engines for future space applications, with specific regard to LOX/LCH 4 technology. The final aim of the program is building a rocket engine demonstrator whose high-level requirements can be summarized as follows: thrust class of 30 kN; relevant to future applications of space propulsion; pressurefed testing; regenerative cooling with liquid methane; ability to reproduce an expander cycle engine.…”

Section: Introductionmentioning

confidence: 99%

Cooling Channel Analysis of a LOX/LCH4 Rocket Engine Demonstrator

Pizzarelli¹,

Betti

Nasuti

et al. 2014

50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

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A computational procedure able to describe the coupled hot-gas/wall/coolant environment that occurs in most liquid rocket engines is presented and demonstrated. The coupled analysis is performed by loose coupling of the two-dimensional axisymmetric Reynolds-Averaged Navier-Stokes equations for the hot-gas flow and the conjugate three-dimensional model for the coolant flow and solid material heat transfer in the regenerative cooling circuit. The latter model is in turn based on the coupled Reynolds-Averaged Navier-Stokes equations for the coolant flow and Fourier equation for the thermal conduction in the solid material. In this study, the thermal behavior of a regeneratively cooled oxygen/methane engine demonstrator is analyzed in detail. Starting from a nominal operative condition of the engine, different levels of channel surface roughness and coolant mass flow rate are considered in order to understand their influence on the heat transfer capability of the cooling system. Results show that the heat transfer can be markedly impaired if the operating parameters undergo rather minor changes with respect to the nominal condition

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“…In this paper performance and stability analyses are conducted on the Penn State LOX/CH4 uni-element rocket, the details of which are described by Locke, Pal and Woodward 2 with performance and stability assessments by Hulka and Jones 3 as well as on the sub scale breadboard (SSBB) experimental rocket engine under development at CIRA 4 Of particular interest are the chamber pressures around 800 psia (5.5MPa), see Table 1 since these correspond to the nominal chamber pressures associated with the CIRA engine, and which are not fully documented by Hulka and Jones. These cases will be addressed in this paper and will be presented as an alternative step towards evaluating the performance of the CIRA SSBB.…”

Section: Introductionmentioning

confidence: 99%

Comparison of Performance and Stability Analyses for LOX/CH4 Rocket Engines

French¹,

Battista²,

Salvatore³

et al. 2014

50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

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Liquid rocket engine development using oxygen and methane propellants is currently underway at CIRA within the context of the Hyprob program funded by the Italian Ministry of University and Research (MIUR). Different methods for evaluating the performance and combustion stability of oxygen and methane propellants are being analyzed. This study details performance and stability analyses of the NASA sponsored Penn State LOX/CH 4 Uni-element rocket as well as those for a similar rocket engine under development at CIRA. Simulations are conducted with two one-dimensional liquid rocket combustion analysis programs, The Rocket Combustor Interactive Design and Analysis code program and the Generalized Instability Model, the latter of which uses an analytical model to study combustion instabilities by solving an inhomogeneous pressure wave equation using a modified Galerkin method. Nomenclature GIM = Generalized Instability Model ROCCID = Rocket Combustor Interactive Design and Analysis Code LOX = Liquid Oxygen SSBB = Sub Scale Bread Board dV = differential volume E′ = energy sources m E = Rupe Mixing Efficiency 2 n E = integral of square of normal modes f = frequency n F = source terms in temporal ODE GR = growth rate M ′ = momentum sources MR = mixture ratio LEE = Linearised Euler Solver L 1 = first longitudinal mode L 2 = second longitudinal mode L 3 = third longitudinal mode C 1 = first coupled mode C 2 = second coupled mode C 3 = third coupled mode a = sonic velocity w f = inhomogenous term of wall boundary condition h = step height HRL , 0 x = heat release location Downloaded by UNIVERSITY OF TENNESSEE on August 13, 2015 | http://arc.aiaa.org | 2 k = wave number K = complex wave number L = length of combustor section l = mode number m = mass n = normal vector u = velocity p = pressure q = heat release t = time x = axial coordinate Y = admittance Z = impedance β = amplitude parameter in velocity lag heat response model w δ = frequency shift γ = ratio of specific heats ε = error η = temporal modes Ω = perturbed frequency ρ = density τ = time lag, sensitive time lag T τ = total time lag i τ = insensitive time lag ψ = spatial, acoustic or normal modes ω = angular frequency Superscripts ' = perturbation variable -= average . = first temporal derivative .. = second temporal derivative Subscripts c = chamber j = injector b = burning n = mode number bdry = boundary term _ nl = non linear term _ mf = mean flow term OX = oxidizerF = fuel Downloaded by UNIVERSITY OF TENNESSEE on August 13, 2015 | http://arc.aiaa.org |

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Design and Development of a LOX/LCH4 Technology Demonstrator

Cited by 8 publications

References 1 publication

Comparison of Reactive Flow Simulations for a LOX/CH4 Uni-element Rocket

Comparison of Reactive Flow Simulations for a LOX/CH4 Uni-element Rocket

Cooling Channel Analysis of a LOX/LCH4 Rocket Engine Demonstrator

Comparison of Performance and Stability Analyses for LOX/CH4 Rocket Engines

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