Analysis of the Heat-Shield Experiment on the Pioneer-Venus Entry Probes

Wakefield, R. M.; Pitts, William C.

doi:10.2514/6.1980-1494

Cited by 46 publications

(16 citation statements)

References 3 publications

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“…When compared with theoretical analysis, experimentally measured heating and ablating rates for the heat shields of Pioneer-Venus [1] and Galileo probes [2][3][4] showed disagreements. The rates were underpredicted at the stagnation point and overpredicted in the downstream region.…”

Section: Introductionmentioning

confidence: 74%

Numerical Study of Spallation Phenomenon in an Arc-Jet Environment

Davuluri

Zhang

Martin

2016

Journal of Thermophysics and Heat Transfer

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It is hypothesized that spallation could affect the aerodynamic heating rates of reentry vehicles using ablative heat shields. To investigate spallation effects, a code is developed to compute the dynamics of spalled particles. The code uses a finite-rate chemistry model to study the chemical interactions of the particles with the flow field. The spallation code is one-way coupled to a computational fluid dynamics solver that models the hypersonic flow field around an ablative sample. Spalled particle behavior is numerically studied in both argon and air flow fields. Nomenclature A p = area of cross section of the particle, m 2 A s = surface area of the particle, m 2 C D = drag coefficient c i = local concentration of species i at particle surface, mol∕m 3 c v = specific heat at constant volume, J∕mol · K d p = diameter of the particle, m E p = internal energy of the particle, J∕kg E R = energy barrier for reaction, J∕mol F D = drag force, N J C i = vapor flux of C i species, kg∕m 2 · s k f = forward reaction rate, m∕s M = Mach number M w = molar weight, kg∕mol m p = mass of the particle, kg _ m C∕i = mass rate of the particle due to reactions, kg∕s Nu = Nusselt number P = vapor pressure, N∕m 2 p drag = drag power, J∕s _ q = heat rate, J∕s R = universal gas constant Re = Reynolds number based on relative velocity S = molecular speed ratio T = temperature, K U = state vector U ∞ = freestream velocity, m∕s u; v; w = velocity components in axial, radial, and z directions, m∕s V r = relative velocity between particle and fluid, m∕s W = source vector x; y; z = position components in axial, radial, and z directions, m Y i = mass fraction, kg∕kg α v = vaporization coefficient γ 0 = reaction efficiency ΔG f = Gibbs free energy of formation, J∕mol Δh = enthalpy of formation, J∕mol ϵ = emissivity of the particle κ = thermal conductivity, W∕m · K μ = dynamic viscosity, kg∕m · s ν = mean thermal speed, m∕s ρ = density, kg∕m 3 σ = Stefan-Boltzmann constantSubscripts conv = convection f = flow field g = gas phase p = particle rad = radiation rxn = chemical reactions s = solid phase (graphite) sub = sublimation tr = translational-rotational energy mode ve = vibrational-electronic energy mode

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Section: Introductionmentioning

confidence: 74%

Numerical Study of Spallation Phenomenon in an Arc-Jet Environment

Davuluri

Zhang

Martin

2016

Journal of Thermophysics and Heat Transfer

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“…For Microprobe, a nose radius equal to half of the vehicle's overall base radius is an acceptable value. This is the same ratio used in the Pioneer-Venus and Galileo entry probes [7,10,11]. Similarly, rounding the vehicle's shoulders is performed to decrease local heating.…”

Section: Discussionmentioning

confidence: 99%

Mars Microprobe entry analysis

Braun

Mitcheltree²,

Cheatwood

1997

1997 IEEE Aerospace Conference

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Abstract{The Mars Microprobe mission will provide the rst opportunity for subsurface measurements, including water detection, near the south pole of Mars. In this paper, performance of the Microprobe aeroshell design is evaluated through development of a six-degreeof-freedom (6-DOF) aerodynamic database and ight dynamics simulation. Numerous mission uncertainties are quantied and a Monte-Carlo analysis is performed to statistically assess mission performance. Results from this 6-DOF Monte-Carlo simulation demonstrate that, in a majority of the cases (approximately 2{), the penetrator impact conditions are within current design tolerances. Several trajectories are identied in which the current set of impact requirements are not satised. From these cases, critical design parameters are highlighted and additional system requirements are suggested. In particular, a relatively large angle-of-attack range near peak heating is identied.

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“…This qualitatively explained the ight data. 14 In the previous study 11 the injection-induced turbulence model was incorporated in the ow solver through the use of the one-equation turbulence model proposed by Goldberg and Ramakrishnan. 15 In the model the damping function for low-Reynolds-numbereffect was de ned by comparing local eddy size with Kolmogorov scale.…”

Section: Nomenclaturementioning

confidence: 99%

Calculation of Surface Heat Transfer for a Sphere with Wall Injection

Izawa

Sawada

2000

Journal of Thermophysics and Heat Transfer

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A computational uid dynamics method is developed to calculate heat-transfer rate for a sphere with wall injection in hypersonic ow. Attention is paid to exploring the enhanced heat-transfer rate by wall injection. The concept of injection-induced turbulence effect is employed in which the injection ow is assumed to be inherently turbulent. A detailed comparison of the calculated heat-transfer distributions with existing experimental data is carried out. Although the quantitative agreement is not yet accomplished for the entire wall surface, the enhanced heat-transfer rate at the stagnation point is well reproduced, and the higher heating rate in the downstream region is indicated in the calculated results. NomenclatureVan Driest damping function d = mixing length of turbulence, cm d m = limiting mixing length of turbulence at the wall, cm d inj = mixing length of turbulence caused by the injection ow from the wall surface, cm M 1 = freestream Mach number Çm= local mass addition rate, g/cm 2 -s Çm = area-averaged mass addition rate to a given location, g/cm 2 -s m ¤ = mass-ow rate ratio ( Çm / Çm 1 ) Çm 1 = freestream mass-ow rate, g/cm 2 -s p = pressure, psi p t = total pressure of the wind tunnel, psi q = local heating rate, W/cm 2 q 0 = theoretical value of the stagnation heating rate for a sphere with 17.78 cm diameter, W/cm 2 R e = Reynolds number based on the nose diameter R N = nose radius, 8.89 cm s = distance along the wall surface from the stagnation point, cm T = temperature, K T t = total temperature of the wind tunnel, K u = velocity component tangential to the wall, cm/s v = velocity component normal to the wall, cm/s y = distance from the wall, cm y + = dimensionless wall coordinate g = dimensionless normal coordinate; Eq. (5) l = molecular viscosity, g/cm-s l t = eddy viscosity, g/cm-s (l t ) BL = eddy viscosity evaluated by the Baldwin-Lomax turbulence model, g/cm-s (l t ) inj = eddy viscosity caused by the injection ow from the wall surface, g/cm-s (l t ) Park = eddy viscosity of the injection ow at the wall according to Park's model, g/cm-s Presented as Paper 99-0734 at 01. Student Member AIAA. † Professor, Department of Aeronautics and Space Engineering, Aobayama 01; sawada@cdf.mech.tohoku.ac.jp. Senior Member AIAA. q = density, g/cm 3 s = time constant, s v = Karman constant, 0.4 W = stagnation point heating rate ratio (q / q 0 ) x = vorticity, 1/s Subscripts B = 0 = value for no mass addition e = edge of boundary layer s = local value at s w = wall value

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Analysis of the Heat-Shield Experiment on the Pioneer-Venus Entry Probes

Cited by 46 publications

References 3 publications

Numerical Study of Spallation Phenomenon in an Arc-Jet Environment

Numerical Study of Spallation Phenomenon in an Arc-Jet Environment

Mars Microprobe entry analysis

Calculation of Surface Heat Transfer for a Sphere with Wall Injection

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