Analysis of Apollo Command Module Afterbody Heating Part I: AS-202

Wright, Michael J.; Prabhu, Dinesh K.; Martinez, Edward

doi:10.2514/1.15873

Cited by 44 publications

(26 citation statements)

References 22 publications

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“…The freestream total enthalpy at the current condition (12:73 MJ=kg) is similar to that of Apollo AS-202 computations in [4], in which ionization is found to be negligible. Hence, air is modeled as a neutral mixture of five perfect species (N 2 , O 2 , NO, N, and O) with three dissociation and two exchange reactions.…”

Section: Simulation Methodologysupporting

confidence: 69%

“…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]. However, the majority of the trajectory points considered in [3,4] have Reynolds numbers low enough for the flow to be laminar.…”

Section: Introductionmentioning

confidence: 79%

“…The Reynolds number at the conditions considered here is more than twice of the highest value considered for AS-202 [4]. Transition to turbulence is therefore expected to occur earlier on the vehicle than in the Apollo case, possibly even on the forebody.…”

Section: Introductionmentioning

confidence: 94%

“…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]. However, the majority of the trajectory points considered in [3,4] have Reynolds numbers low enough for the flow to be laminar. The only exception is the AS-202 data during the second heat pulse, in which laminar calculations underpredict the heating measurements.…”

Section: Introductionmentioning

confidence: 98%

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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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Section: Simulation Methodologysupporting

confidence: 69%

Section: Introductionmentioning

confidence: 79%

Section: Introductionmentioning

confidence: 94%

Section: Introductionmentioning

confidence: 98%

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Hypersonic Turbulent Flow Simulation of Fire II Reentry Vehicle Afterbody

Reddy

Sinha

2009

Journal of Spacecraft and Rockets

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“…The maximum run-time of the facility is 300 s. Two 30 cm diameter schlieren windows in the side of the test section are available for optical acces. In the current experiments, the wind tunnel was operated in the range of M = 2 − 4 with a total pressure ranging from 2.7 − 12 bar and a total temperature of 288 K. The capsule geometry used for the definition of the wind tunnel model is a scaled version of the AS-202 outer mold line as defined in Wright et al 1 The model has a diameter of 50 mm and is fabricated out of Makrolon, it is side-mounted on a stainless steel sting. Two models are used for 0 and 25 degrees angle of attack, see Figure 2.…”

Section: Iia Flow Facility and Wind Tunnel Modelmentioning

confidence: 99%

Experimental Investigation of the Supersonic Wake of a Reentry Capsule

Schrijer

Walpot

2010

48th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

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The wake behind an Apollo shaped capsule is investigated in the framework of the 'afterbody heating' topic in the RTO WG043 working group. Measurements are performed by means of schlieren, shadowgraphy and stereo particle image velocimetry (SPIV) and are used for CFD validation purposes. The model geometry is a scaled version of the AS-202. It was found from the shadowgraph results that the angle of attack where separation occurs decreases with increasing Mach number and is not influenced by the Reynolds number. Furthermore, it was possible to correlate the shear layer transition location using Re2,x t , which is the Reynolds number based on post normal shock conditions where the length scale is the flow path from stagnation point to transition point. SPIV measurements were obtained at Mach 2 and 0• and 25• angle of attack. For the 0 • model, the wake was completely separated while for the 25• model the wake was partially separated and reattaches half way the model. Overall the PIV data return a quantitative three dimensional description of the velocity field around the capsule.

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Multifidelity domain-aware learning for the design of re-entry vehicles

Fiore

Maggiore

Mainini

2021

Struct Multidisc Optim

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The multidisciplinary design optimization (MDO) of re-entry vehicles presents many challenges associated with the plurality of the domains that characterize the design problem and the multi-physics interactions. Aerodynamic and thermodynamic phenomena are strongly coupled and relate to the heat loads that affect the vehicle along the re-entry trajectory, which drive the design of the thermal protection system (TPS). The preliminary design and optimization of re-entry vehicles would benefit from accurate high-fidelity aerothermodynamic analysis, which are usually expensive computational fluid dynamic simulations. We propose an original formulation for multifidelity active learning that considers both the information extracted from data and domain-specific knowledge. Our scheme is developed for the design of re-entry vehicles and is demonstrated for the case of an Orion-like capsule entering the Earth atmosphere. The design process aims to minimize the mass of propellant burned during the entry maneuver, the mass of the TPS, and the temperature experienced by the TPS along the re-entry. The results demonstrate that our multifidelity strategy allows to achieve a sensitive improvement of the design solution with respect to the baseline. In particular, the outcomes of our method are superior to the design obtained through a single-fidelity framework, as a result of the principled selection of a limited number of high-fidelity evaluations.

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Analysis of Apollo Command Module Afterbody Heating Part I: AS-202

Cited by 44 publications

References 22 publications

Hypersonic Turbulent Flow Simulation of Fire II Reentry Vehicle Afterbody

Hypersonic Turbulent Flow Simulation of Fire II Reentry Vehicle Afterbody

Experimental Investigation of the Supersonic Wake of a Reentry Capsule

Multifidelity domain-aware learning for the design of re-entry vehicles

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