Chemical Erosion of Carbon-Carbon/Graphite Nozzles in Solid-Propellant Rocket Motors

Thakre, Piyush; Yang, Vigor

doi:10.2514/1.34946

Cited by 122 publications

(77 citation statements)

References 34 publications

Supporting

Mentioning

Contrasting

Unclassified

Order By: Relevance

“…22,28,29,36 However, for the present analyzed conditions, radiation emitted from the wall can be relevant due to the high wall temperatures, the high emissivity of graphite, and the moderate level of convective heat transfer due to the reduced pressure level (10 bar). Therefore, radiation is not considered a significant factor at the nozzle throat and is usually neglected.…”

Section: Ivd Effect Of Wall Radiationmentioning

confidence: 93%

“…The wall temperature must satisfy the surface energy balance and is therefore the result of different factors among which the most important are the flame temperature, the level of erosion rate, and the energy absorption (or release) by the different oxidizing reactions. Finally, oxidizing species such as atomic and molecular oxygen, whose erosion contribution is traditionally neglected for solid fuels 11,12,22,28,29 (due to their negligible presence in the combustion gases), are shown to be important oxidizing species in hybrid rocket engines, contributing on average between 10 and 20% to total erosion. 5(c), it can be seen that the choice of the oxidizer can significantly alter the heat released by the oxidizing reactions: hydrogen peroxide, where the erosion is dominated by chemical attack by water vapor, shows the highest heat of ablation while oxygen, where exothermic oxidizing reactions with atomic and molecular oxygen are significant, shows the lowest heat of ablation.…”

Section: Iva Effect Of Propellant Combinationmentioning

confidence: 99%

See 1 more Smart Citation

Numerical Analysis of Nozzle Material Thermochemical Erosion in Hybrid Rocket Engines

Bianchi¹,

Nasuti²

2012

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

View full text Add to dashboard Cite

Ablative materials are commonly used to protect the nozzle metallic housing and to provide the internal contour to expand the exhaust gases in both solid rocket motors (SRM) and hybrid rocket engines (HRE). Due to the extremely harsh environment in which these materials operate, they are chemically eroded during motor firing with a resulting nominal performance reduction. The objective of the present work is to study the erosion behavior of graphite nozzles in HRE at different operating conditions and comparing results with those obtained for SRM. A mean distinctive feature of HRE operating conditions is, in fact, a greater concentration of oxygen-containing combustion products than SRM. The adopted approach relies on a validated full Navier-Stokes flow solver coupled with a thermochemical ablation model which takes into account heterogeneous chemical reactions at the nozzle surface, rate of diffusion of the species through the boundary-layer, ablation species injection in the boundary layer, heat conduction inside the nozzle material, and variable multispecies thermophysical properties. The parametric analysis performed in this study allows to assess the impact of various parameters that affect the nozzle erosion rate, taking into account various combinations of fuels and oxidizers operating at different conditions.

show abstract

Section: Ivd Effect Of Wall Radiationmentioning

confidence: 93%

Section: Iva Effect Of Propellant Combinationmentioning

confidence: 99%

Numerical Analysis of Nozzle Material Thermochemical Erosion in Hybrid Rocket Engines

Bianchi¹,

Nasuti²

2012

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

View full text Add to dashboard Cite

show abstract

“…The high temperature fluid flow in the inner contour of the nozzle system creates a soaring thermomechanical stresses through the liner cross section. This degrades the nozzle throat and widens the throat area of cross section [6]. This phenomenon causes a reduction in thrust and nozzle operational efficiency.…”

Section: Structure Description Of Hot Zone Assemblymentioning

confidence: 99%

“…Interface region of the fiber/matrix and the defective regions of the composite are more prone to ablation. This is due to the lower activation energy and high reactivity of the defective region in the composite system [64,65]. The ablation always developed along the direction of interface-to-fiber and interface-tomatrix due to the high oxidization tendency of the interface region.…”

Section: Ablative Mechanismsmentioning

confidence: 99%

Hybrid Carbon-Carbon Ablative Composites for Thermal Protection in Aerospace

Sanoj

Kandasubramanian

2014

Journal of Composites

View full text Add to dashboard Cite

Composite materials have been steadily substituting metals and alloys due to their better thermomechanical properties. The successful application of composite materials for high temperature zones in aerospace applications has resulted in extensive exploration of cost effective ablative materials. High temperature heat shielding to body, be it external or internal, has become essential in the space vehicles. The heat shielding primarily protects the substrate material from external kinetic heating and the internal insulation protects the subsystems and helps to keep coefficient of thermal expansion low. The external temperature due to kinetic heating may increase to about maximum of 500 ∘ C for hypersonic reentry space vehicles while the combustion chamber temperatures in case of rocket and missile engines range between 2000 ∘ C and 3000 ∘ C. Composite materials of which carbon-carbon composites or the carbon allotropes are the most preferred material for heat shielding applications due to their exceptional chemical and thermal resistance.

show abstract

“…(17) for the three reactions are taken from Ref. [20,23] and are listed in Table 1. The total erosion rate of carbon due to the surface heterogeneous reactions is:ṁ…”

Section: Iie Thermochemical Ablation Modelmentioning

confidence: 99%

Ablative Material Behavior in Oxygen/Methane Thruster Environment

Turchi

Bianchi

Nasuti

et al. 2012

43rd AIAA Thermophysics Conference

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

Chemical Erosion of Carbon-Carbon/Graphite Nozzles in Solid-Propellant Rocket Motors

Cited by 122 publications

References 34 publications

Numerical Analysis of Nozzle Material Thermochemical Erosion in Hybrid Rocket Engines

Numerical Analysis of Nozzle Material Thermochemical Erosion in Hybrid Rocket Engines

Hybrid Carbon-Carbon Ablative Composites for Thermal Protection in Aerospace

Ablative Material Behavior in Oxygen/Methane Thruster Environment

Contact Info

Product

Resources

About