Calibration, validation, and evaluation studies in the LENS facility

Aerodynamic heating tests were conducted on a 70-deg sphere-cone Mars entry vehicle con guration in a high-enthalpy impulse facility in both carbon dioxide and air test gases. The purpose of these tests was to obtain heat transfer data for comparison with results of Navier-Stokes computations. Surface heat transfer rates were determined for both the forebody and afterbody of the test models and for the stings that supported the models in the facility test section. Little difference was observed between normalized heating distributions for the air and carbon dioxide test conditions. For both cases, peak sting heating was on the order of 4-5% of the forebody stagnation-pointheating, and it was concluded that the wake ow remained laminar. The wake ow establishment process was quanti ed and was found to require approximately 40-70 ow path lengths, which corresponded to approximately 75% of the available facility test time. The repeatability of facility test conditions was estimated to vary between § 3% and § 10%. The overall experimental uncertainty of the data was estimated to be § 10-11% for forebody heating and § 17-22% for wake heating. NomenclatureB = measurement bias error h = enthalpy, J/kg k = thermal conductivity, W/m ¢ K M = Mach number P = measurement precision error p = pressure, N/m 2 q = heat transfer rate, W/m 2 R = radius, m Re = Reynolds number S = distance along model surface, m T = temperature, K t = time, s U 1 = freestream velocity, m/s U tot = total combined uncertainty y ref = reference length, m ® = thermal diffusivity, m 2 /s = thermal product, ®/ p k, W ¢ s 1=2 /m 2 ¢ Ķ = correction factor, K ¡1 ½ = density, kg/m 3 ¾ = heat transfer residual ¿ = nondimensional time, U 1 1t est =y ref Subscripts est = establishment w = wall 0 = stagnation 1 = freestream 2 = post-normal shock

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“…[11][12][13][14]. In air, the peak sting heating was 5% of that at the forebody stagnationpoint,and was 4% of that in CO 2 .…”

Section: Resultsmentioning

confidence: 99%

High-Enthalpy Aerothermodynamics of a Mars Entry Vehicle Part 1: Experimental Results

Hollis

Perkins

1997

Journal of Spacecraft and Rockets

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“…The peak heating on the model sting varied from 8% of the forebody stagnation point heating at the lowest Reynolds number to 15% at the highest Reynolds number. Relative peak sting heating rates of this level are consistent with turbulent wakes 6,7,14 , whereas relative laminar peak sting heating levels are in the 4-5% range [12][13][14][15] . Several surface oil flow tests were also carried out using the uninstrumented aluminum model.…”

Section: Resultsmentioning

confidence: 55%

Transition effects on heating in the wake of a blunt body

Hollis

Perkins

1997

32nd Thermophysics Conference

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A series of aerodynamic heating tests was conducted on a 70-deg sphere-cone planetary entry vehicle model in a Mach 10 perfect-gas wind tunnel at freestream Reynolds numbers based on diameter of 8.23x10 4 to 3.15x10 5 . Surface heating distributions were determined from temperature time-histories measured on the model and on its support sting using thin-film resistance gages. The experimental heating data were compared to computations made using an axisymmetric/2D, laminar, perfect-gas Navier-Stokes solver. Agreement between computational and experimental heating distributions to within, or slightly greater than, the experimental uncertainty was obtained on the forebody and afterbody of the entry vehicle as well as on the sting upstream of the free-shear-layer reattachment point. However, the distributions began to diverge near the reattachment point, with the experimental heating becoming increasingly greater than the computed heating with distance downstream from the reattachment point. It was concluded that this divergence was due to transition of the wake free shear layer just upstream of the reattachment point on the sting. Nomenclature B C,Hheating bias error

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“…16 shows the excellent agreement obtained between DSMC and the measured data. Another experiment was conducted using the same geometry in the Large Energy National Shocktunnel (LENS) facility again in nitrogen at a Mach number of 15.6 and a Knudsen number of 0.002 [62]. This higher-enthalpy experiment had a total temperature of 4350 K so that vibration was activated but there was still no chemistry.…”

Section: A Analyses Of Hypersonic Laboratory Experimentsmentioning

confidence: 99%

Computation of Hypersonic Flows Using the Direct Simulation Monte Carlo Method

Boyd

2015

Journal of Spacecraft and Rockets

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The direct simulation Monte Carlo method has evolved over 50 years into a powerful numerical technique for the computation of complex, nonequilibrium gas flows. In this context, "nonequilibrium" means that the velocity distribution function is not in an equilibrium form due to a low number of intermolecular collisions within a fluid element. In hypersonic flow, nonequilibrium conditions occur at high altitude and in regions of flowfields with small length scales. In this paper, the theoretical basis of the direct simulation Monte Carlo technique is discussed. In addition, the methods used in direct simulation Monte Carlo are described for simulation of high-temperature, real gas effects and gas-surface interactions. Several examples of the application of direct simulation Monte Carlo to flows around blunt hypersonic vehicles are presented to illustrate current capabilities. Areas are highlighted where further research on the direct simulation Monte Carlo technique is required. Nomenclature A= surface element of area, m 2 C = particle velocity vector, m∕s e r = specific rotational energy, J∕kg f = probability density function g = relative velocity, m∕s Kn = Knudsen number k = Boltzmann's constant, J∕kg L = characteristic length scale, m m = mass of a particle, kg N = average number of particles in a cell N p = number of particles in a cell N t = number of iterations n = number density, m −3 P c = collision probability r = particle position vector, m T = temperature, K t = time, s V = cell volume, m 3 Δf = rate of change due to collision processes ε act = activation energy, J ε r = rotational energy, J ε tot = total collision energy, J λ = mean free path, m ζ = number of rotational degrees of freedom σ = collision cross section, m 2 τ r = rotational relaxation time, s ω = viscosity temperature exponent

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Calibration, validation, and evaluation studies in the LENS facility

Cited by 19 publications

References 1 publication

High-Enthalpy Aerothermodynamics of a Mars Entry Vehicle Part 1: Experimental Results

High-Enthalpy Aerothermodynamics of a Mars Entry Vehicle Part 1: Experimental Results

Transition effects on heating in the wake of a blunt body

Computation of Hypersonic Flows Using the Direct Simulation Monte Carlo Method

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