Aerothermal Analysis of the Project Fire II Afterbody Flow

Wright, Michael J.; Loomis, Mark; Papadopoulos, Periklis

doi:10.2514/2.6757

Cited by 68 publications

(19 citation statements)

References 28 publications

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“…A super-catalytic wall boundary condition was employed with a surface emissivity of 0.89 for compatibility with the LAURA solutions. DPLR afterbody heating simulations have compared favorably to Earth entry flight data 32,33 .…”

Section: Dplr Cfd Codementioning

confidence: 99%

Aerothermodynamic Environments Definition for the Mars Science Laboratory Entry Capsule

Edquist

Dyakonov

Wright

et al. 2007

45th AIAA Aerospace Sciences Meeting and Exhibit

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Section: Dplr Cfd Codementioning

confidence: 99%

Aerothermodynamic Environments Definition for the Mars Science Laboratory Entry Capsule

Edquist

Dyakonov

Wright

et al. 2007

45th AIAA Aerospace Sciences Meeting and Exhibit

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“…Most notably, the FIRE II and Apollo AS-202 flowfields are computed at a series of time points spanning their respective heat pulses. The axisymmetric simulation of the zeroangle-of-attack FIRE II configuration matches the experimental data well [3]. Similarly, computation of the three-dimensional flowfield around the AS-202 vehicle yields heat transfer rates that are within the scatter of in-flight measurements [4].…”

Section: Introductionmentioning

confidence: 83%

Hypersonic Turbulent Flow Simulation of Fire II Reentry Vehicle Afterbody

Reddy

Sinha

2009

Journal of Spacecraft and Rockets

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This paper presents a numerical investigation of the hypersonic reacting flow around the FIRE II reentry capsule. At the chosen freestream conditions, the forebody boundary layer and the separated flow on the afterbody are turbulent. The Reynolds-averaged Navier-Stokes method along with two commonly used turbulence models are used to compute the flowfield. Accurate prediction of turbulent separated flow at hypersonic conditions is challenging due to the limitations of the underlying turbulence models. The presence of turbulent eddy viscosity in the flow simulation results in a smaller separation bubble than the laminar solution at identical conditions. Also, the two turbulence models predict different levels of eddy viscosity in the neck region. This has a dominant effect on the separation bubble size and the surface pressure. On the other hand, the eddy viscosity values in the near-wall region determine the heat transfer rate to the body. The two models predict comparable heating rates on the conical frustum, and the results match in-flight measurement well. By comparison, surface pressure predictions are appreciably higher than the data. Nomenclature D = diameter of the vehicle, m k = turbulent kinetic energy, J=kg M = Mach number p = pressure, Pa q = heat transfer rate, W=cm 2 Re D = Reynolds number based on freestream conditions and body diameter s = arc length from the nose stagnation point, m T = translational-rotational temperature, K T v = vibrational temperature, K U = velocity, m=s = molecular viscosity, Pa s T = turbulent eddy viscosity, Pa s T = turbulent kinematic viscosity, m 2 =s = density, kg=m 3 ! = specific turbulent dissipation rate, 1=s Subscripts T = turbulent w = wall 1 = freestream

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“…Total and radiative heating measurements were taken for about 37 s during the vehicle's entry into the Earth's atmosphere. 7 The FIRE II calorimeter and radiometer data have been used in several recent computational fluid dynamics 3,8,9 and DSMC 10 studies to assess the general level of aerothermal predictive capability in the reentry community. The present study does not make direct use of the FIRE II heating data, but rather utilizes the vehicle characteristics and reentry conditions in the shock layer model.…”

Section: Physics Of Fluids 22 106101 ͑2010͒mentioning

confidence: 99%

Modeling of the plasma generated in a rarefied hypersonic shock layer

Farbar

Boyd

2010

Physics of Fluids

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In this study, a rigorous numerical model is developed to simulate the plasma generated in a rarefied, hypersonic shock layer. The model uses the direct simulation Monte Carlo ͑DSMC͒ method to treat the particle collisions and the particle-in-cell ͑PIC͒ method to simulate the plasma dynamics in a self-consistent manner. The model is applied to compute the flow along the stagnation streamline in front of a blunt body reentering the Earth's atmosphere at very high velocity. Results from the rigorous DSMC-PIC model are compared directly to the standard DSMC modeling approach that uses the ambipolar diffusion approximation to simulate the plasma dynamics. It is demonstrated that the self-consistent computation of the plasma dynamics using the rigorous DSMC-PIC model captures many physical phenomena not accurately predicted by the standard modeling approach. These computations represent the first assessment of the validity of the ambipolar diffusion approximation when predicting the rarefied plasma generated in a hypersonic shock layer.

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Aerothermal Analysis of the Project Fire II Afterbody Flow

Cited by 68 publications

References 28 publications

Aerothermodynamic Environments Definition for the Mars Science Laboratory Entry Capsule

Aerothermodynamic Environments Definition for the Mars Science Laboratory Entry Capsule

Hypersonic Turbulent Flow Simulation of Fire II Reentry Vehicle Afterbody

Modeling of the plasma generated in a rarefied hypersonic shock layer

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