Thermochemical Erosion Analysis for Carbon-Carbon Rocket Nozzles

Bianchi, Daniele; Nasuti, Francesco; Onofri, Marcello; Martelli, Emanuele

doi:10.2514/6.2009-4977

Cited by 14 publications

(34 citation statements)

References 10 publications

Supporting

Mentioning

Contrasting

Order By: Relevance

“…To obtain a better evaluation of the ablating flowfield, the mass and energy exchange mechanisms must be carefully modeled using the information from the Navier-Stokes solver; this has been performed integrating the numerical solver with a computational surface thermochemistry technique. 11,14,15 …”

Section: Theoretical Modelingmentioning

confidence: 98%

Aerothermodynamic Analysis of Reentry Flows with Coupled Ablation

Bianchi¹,

Nasuti

Onofri

2011

17th AIAA International Space Planes and Hypersonic Systems and Technologies Conference

Self Cite

View full text Add to dashboard Cite

A study of graphite ablation in reentry flows is carried out by a fully-coupled surface ablation approach integrated with a two-dimensional axisymmetric Navier-Stokes solver. The gas-surface interaction model is based on both chemical equilibrium and finite-rate ablation modeling and can account for both surface oxidation and sublimation. The surface temperature is obtained from the steady-state ablation approximation. This numerical procedure can predict aerothermal heating, chemical species concentrations, and carbon material ablation rate over the heat-shield surface of reentry vehicles. Two dimensional axi-symmetric simulations have been performed to numerically reproduce the ablation of a graphite sphere-cone for which literature data are available from testing performed in the Interaction Heating Facility at the NASA Ames Research Center. The freestream conditions of the selected test case are typical for Earth reentry from a planetary mission. The predicted ablation rate and surface temperature assuming frozen chemistry in the flow show a good agreement with the available experimental data. The agreement is further improved by the use of finite-rate ablation modeling. Nomenclature A = cross section area, m 2 D im = effective diffusion coefficient, m 2 /s h = enthalpy, J/kg j = diffusional mass flux, kg/m 2 · s k = thermal conductivity, W/m · K m = mass blowing rate per unit area, kg/m 2 · s N c = number of species p = pressure, N/m 2 q = heat flux per unit area, W/m 2 r = distance from surface, ṁ s = recession rate, m/s T = temperature, K t = time, s u = streamwise velocity, m/s v = mass motion of the mixture, m/s v = velocity component normal to surface, m/s W = molecular weight, kg/kmole y = species mass fractionSubscripts i = species s = solid (graphite) properties at gas/solid interface s i = solid (graphite) properties at initial conditions w = gas properties at gas/solid interface Symbols ∆H abl = heat of ablation, J/kg ǫ = surface emissivity η = outward coordinate normal to surface µ = viscosity, kg/m · s ρ = density, kg/m 3 σ = Stefan-Boltzmann constanṫ ω = species source term, kg/m 2 · s

show abstract

Section: Theoretical Modelingmentioning

confidence: 98%

Aerothermodynamic Analysis of Reentry Flows with Coupled Ablation

Bianchi¹,

Nasuti

Onofri

2011

17th AIAA International Space Planes and Hypersonic Systems and Technologies Conference

Self Cite

View full text Add to dashboard Cite

show abstract

“…For the analysis of the flight, however, no direct measure of the final nozzle throat can be performed, and therefore, we assumed the semi-empirical correlations gained from the SFTs analysis as a functional dependence of the nozzle throat erosion rate (or mass flow rate per unit area) as a function of the SRM operating pressure, since in the typical regime of diffusion limited erosion of the throat insert, its is known [13][14][15] that this is the main functional dependence of the nozzle throat erosion. This means that each SFT provides a nozzle throat erosion correlation for the analysis of the SRM flight data.…”

Section: Srm Static Firing Test Reconstruction Modelmentioning

confidence: 99%

Analysis and Performance Reconstruction of VEGA Solid Rocket Motors Qualification Flights

Cavallini¹,

Favini²,

Neri³

2014

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

View full text Add to dashboard Cite

VEGA is the launch vehicle developed by the European Space Agency, qualified with its first two maiden flights on February, 13rd 2012 and on May, 7th 2013 and recently delivered to the commercial market with the third flight held on April, 30th 2014.\ud During the launcher development, a total of eight static firing tests have been performed for the three solid stages, which compose the launcher: the first stage P80, the second stage Zefiro 23 and the third stage Z9.\ud In this work, the analysis and performance reconstruction of the solid rocket motors of the VEGA launch vehicle for the two qualification flights is carried out with a post-firing reconstruction model, developed for the purpose.\ud The aim is to use the measures acquired during the flights and the experience gained from the static firing tests analysis, in order to evaluate the actual behavior of the VEGA solid stages, through the non-ideal parameters: combustion efficiency, thrust efficiency, hump, scale factor and nozzle throat erosion law, which define the actual performance parameters of the solid rocket motors.\ud The purpose of the work is to assess the SRMs performance parameters from the flight data, comparing the outcomes of the flight data analysis with the ones provided by the static firing test data analysis.\ud The final aim is to consolidate the methodology for the analysis and reconstruction of the solid stage flight data, in order to define/characterize the scattering of the motor performance, reducing the uncertainties for the prediction methodologies in the upcoming and following VEGA flights

show abstract

“…[24][25][26] The solver is based on a second-order finite-difference approach, validated and verified for non-pyrolyzing and pyrolyzing ablative materials. [8][9][10][11][12][13] The Spalart-Allmaras one-equation model 27 is used to compute the turbulent viscosity µ t . Turbulent conductivity and turbulent mass diffusivity are computed from µ t , specific heat at constant pressure, turbulent Prandtl number, and turbulent Schmidt number.…”

Section: Numerical Modelmentioning

confidence: 99%

“…The purpose of the present study is to extend a CFD approach validated for the study of carbonbased pyrolyzing and non-pyrolyzing materials [8][9][10][11][12][13] to the case of silica-based materials and to perform a comprehensive analysis of TPS materials behavior in oxygen/methane thruster environment. After the description of the theoretical and numerical models, the first part of the analysis is dedicated to the validation, by means of comparison with experimental data, of the newly developed boundary condition and gassurface interaction model for silica-based materials.…”

Section: Introductionmentioning

confidence: 99%

Ablative Material Behavior in Oxygen/Methane Thruster Environment

Turchi

Bianchi

Nasuti

et al. 2012

43rd AIAA Thermophysics Conference

Self Cite

View full text Add to dashboard Cite

Ablative materials represent a low cost and reliable means to insulate rocket engine components from high-temperature, corrosive combustion product environments. Besides their diffuse application in solid rocket nozzles, their use also emerges as a valid alternative in liquid rocket engines. Together with the growing interest in oxygen/methane liquid rocket engines, these materials have gained attention as possible insulator for small upperstage engines or in-space thrusters. In this framework, a validated approach for the study of carbon-based pyrolyzing and non-pyrolyzing materials, together with a novel boundary condition developed to analyze the silica-based material behavior, has been used to numerically reproduce the material response in the highly oxidizing environment generated by the combustion of oxygen and methane. At first, the validation against experimental data of the silica-based material erosion model is presented. Subsequently, the behavior and the response of different ablators in a oxygen/methane environment is numerically investigated for a wide range of operating conditions. Commonly made assumptions in the simulation of the material response are thoroughly analyzed and a critical overview of the results is presented. Nomenclaturebinary diffusion coefficient, m 2 /s h = enthalpy, J/kg k = thermal conductivity, W/m · K N c = number of specieṡ m = mass blowing rate per unit area, kg/m 2 · s p = pressure, N/m 2 q radout = wall radiative heat flux, W/m 2 s = erosion rate, m/s T = temperature, K t = time, s v = flow velocity, m/s v = velocity component normal to surface, m/s * Ph.D. Student, Dipartimento di Ingegneria Meccanica e Aerospaziale, Via Eudossiana 18. AIAA Student Member. alessandro.turchi@uniroma1.it. ẇ i = species source term in control volume, kg/m 3 · s x = mole fraction y = mass fraction y + = dimensionless wall distance Subscripts c = charred state f = failed material g = pyrolysis gas i = species s = solid state v = virgin state w = gas properties at gas/solid interface Symbols η = outward coordinate normal to surface Γ m = mass resin fraction Γ V = volume resin fraction µ = viscosity, kg/m · s µ t = turbulent viscosity, kg/m · s ρ = density, kg/m 3 ρ A = primary resin component density, kg/m 3 ρ B = secondary resin component density, kg/m 3 ρ R = reinforcing material density, kg/m 3

show abstract

Thermochemical Erosion Analysis for Carbon-Carbon Rocket Nozzles

Cited by 14 publications

References 10 publications

Aerothermodynamic Analysis of Reentry Flows with Coupled Ablation

Aerothermodynamic Analysis of Reentry Flows with Coupled Ablation

Analysis and Performance Reconstruction of VEGA Solid Rocket Motors Qualification Flights

Ablative Material Behavior in Oxygen/Methane Thruster Environment

Contact Info

Product

Resources

About