Brian R. Hollis

Spectrum Modeling for Air Shock-Layer Radiation at Lunar-Return Conditions

Anew air-radiationmodel is presented for the calculation of the radiative flux from lunar-return shock layers. For modeling atomic lines, the data from a variety of theoretical and experimental sources are compiled and reviewed. A line model is chosen that consists of oscillator strengths from the National Institute of Standards and Technology database and theOpacity Project (formany lines not listedby theNational Institute of Standards andTechnology), as well as Stark broadening widths obtained from the average of available values. Uncertainties for the oscillator strengths and Stark broadening widths are conservatively chosen from the reviewed data, and for the oscillator strengths, the chosen uncertainties are found to be larger than those listed in the National Institute of Standards and Technology database. This new atomic line model is compared with previous models for equilibrium constantproperty layers chosen to approximately represent a lunar-return shock layer. It is found that the new model increases the emission resulting from the 1–6-eV spectral range by up to 50%.This increase is due to both the increase in oscillator strengths for some important lines and to the addition of lines from the Opacity Project, which are not commonly treated in shock-layer radiation predictions. Detailed theoretical atomic bound–free cross sections obtained from the Opacity Project’s TOPbase are applied for nitrogen and oxygen. An efficient method of treating these detailed cross sections is presented.The emission fromnegative ions is considered and shown to contribute up to 10% to the total radiative flux. The modeling of the molecular-band systems using the smeared-rotational-band approach is reviewed. The validity of the smeared-rotational-band approach for both emitting and absorbing-band systems is shown through comparisons with the computationally intensive line-by-line approach. The absorbingband systems are shown to reduce the radiative flux by up to 10%, whereas the emitting-band systems are shown to contribute less than a 5% increase in the flux. The combined models chosen for the atomic line, atomic bound–free, negative-ion, and molecular-band components result in a computationally efficient model that is ideal for coupled solutions with a Navier–Stokes flowfield. It is recommended that the notable increases shown, relative to previous models, for the atomic line and negative-ion continuum should be included in future radiation predictions for lunarreturn vehicles.

X-33 Hypersonic Boundary Layer Transition

Boundary layer and aeroheating characteristics of several X-33 configurations have been experimentally examined in the Langley 20-Inch Mach 6 Air Tunnel. Global surface heat transfer distributions, surface streamline patterns, and shock shapes were measured on 0.013-scale models at Mach 6 in air. Parametric variations include angles-of-attack of 20-deg, 30-deg, and 40-deg; Reynolds numbers based on model length of 0.9 to 6.6 million; and body-flap deflections of 0, 10 and 20-deg. The effects of discrete and distributed roughness elements on boundary layer transition, which included trip height, size, location, and distribution, both on and off the windward centerline, were investigated. The discrete roughness results on centerline were used to provide a transition correlation for the X-33 flight vehicle that was applicable across the range of reentry angles of attack. The attachment line discrete roughness results were shown to be consistent with the centerline results, as no increased sensitivity to roughness along the attachment line was identified. The effect of bowed panels was qualitatively shown to be less effective than the discrete trips; however, the distributed nature of the bowed panels affected a larger percent of the aft-body windward surface than a single discrete trip.

A Review of Aerothermal Modeling for Mars Entry Missions

The current status of aerothermal analysis for Mars entry missions is reviewed. The aeroheating environment of all Mars missions to date has been dominated by convective heating. Two primary uncertainties in our ability to predict forebody convective heating are turbulence on a blunt lifting cone and surface catalysis in a predominantly CO2 environment. Future missions, particularly crewed vehicles, will encounter additional heating from shock-layer radiation due to a combination of larger size and faster entry velocity. Localized heating due to penetrations or other singularities on the aeroshell must also be taken into account. The physical models employed to predict these phenomena are reviewed, and key uncertainties or deficiencies inherent in these models are explored. Capabilities of existing ground test facilities to support aeroheating validation are also summarized. Engineering flight data from the Viking and Pathfinder missions, which may be useful for aerothermal model validation, are discussed, and an argument is presented for obtaining additional flight data. Examples are taken from past, present, and future Mars entry missions, including the twin Mars Exploration Rovers and the Mars Science Laboratory, scheduled for launch by NASA in 2011.

Non-Boltzmann Modeling for Air Shock-Layer Radiation at Lunar-Return Conditions

This paper investigates the non-Boltzmann modeling of the radiating atomic and molecular electronic states present in lunar-return shock-layers. The Master Equation is derived for a general atom or molecule while accounting for a variety of excitation and de-excitation mechanisms. A new set of electronic-impact excitation rates is compiled for N, O, and N2+, which are the main radiating species for most lunar-return shock-layers. Based on these new rates, a novel approach of curve-fitting the non-Boltzmann populations of the radiating atomic and molecular states is developed. This new approach provides a simple and accurate method for calculating the atomic and molecular non-Boltzmann populations while avoiding the matrix inversion procedure required for the detailed solution of the Master Equation. The radiative flux values predicted by the present detailed non-Boltzmann model and the approximate curve-fitting approach are shown to agree within 5% for the Fire 1634 s case.

Radiative Heating Methodology for the Huygens Probe

The radiative heating environment for the Huygens probe near peak heating conditions for Titan entry is investigated in this paper. The task of calculating the radiation-coupled flowfield, accounting for non-Boltzmann and non-optically thin radiation, is simplified to a rapid yet accurate calculation. This is achieved by using the viscous-shock layer (VSL) technique for the stagnation-line flowfield calculation and a modified smeared rotational band (SRB) model for the radiation calculation. These two methods provide a computationally efficient alternative to a Navier-Stokes flowfield and line-by-line radiation calculation. The results of the VSL technique are shown to provide an excellent comparison with the Navier-Stokes results of previous studies. It is shown that a conventional SRB approach is inadequate for the partially optically-thick conditions present in the Huygens shock-layer around the peak heating trajectory points. A simple modification is proposed to the SRB model that improves its accuracy in these partially optically-thick conditions. This modified approach, labeled herein as SRBC, is compared throughout this study with a detailed line-by-line (LBL) calculation and is shown to compare within 5% in all cases. The SRBC method requires many orders-of-magnitude less computational time than the LBL method, which makes it ideal for coupling to the flowfield. The application of a collisional-radiative (CR) model for determining the population of the CN electronic states, which govern the radiation for Huygens entry, is discussed and applied. The non-local absorption term in the CR model is formulated in terms of an escape factor, which is then curve-fit with temperature. Although the curve-fit is an approximation, it is shown to compare well with the exact escape factor calculation, which requires a computationally intensive iteration procedure.

Nonequilibrium Stagnation-Line Radiative Heating for Fire II

This paper presents a detailed analysis of the shock-layer radiative heating to the Fire II vehicle using a new air radiation model and a viscous shock-layer flowfield model. This new air radiation model contains the most up-to-date properties for modeling the atomic-line, atomic photoionization, molecular band, and non-Boltzmann processes. The applied viscous shock-layer flowfield analysis contains the same thermophysical properties and nonequilibrium models as the LAURA Navier-Stokes code. Radiation-flowfield coupling, or radiation cooling, is accounted for in detail in this study. It is shown to reduce the radiative heating by about 30% for the peak radiative heating points, while reducing the convective heating only slightly. A detailed review of past Fire II radiative heating studies is presented. It is observed that the scatter in the radiation predicted by these past studies is mostly a result of the different flowfield chemistry models and the treatment of the electronic state populations. The present predictions provide, on average throughout the trajectory, a better comparison with Fire II flight data than any previous study. The magnitude of the vacuum ultraviolet (VUV) contribution to the radiative flux is estimated from the calorimeter measurements. This is achieved using the radiometer measurements and the predicted convective heating. The VUV radiation predicted by the present model agrees well with the VUV contribution inferred from the Fire II calorimeter measurement, although only when radiation-flowfield coupling is accounted for. This agreement provides evidence that the present model accurately models the VUV radiation, which is shown to contribute significantly to the Fire II radiative heating.

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

Aerodynamic heating tests were conducted on a 70-deg sphere ‐cone Mars entry vehicle cone 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-point heating, and it was concluded that the wake e ow remained laminar. The wake e ow establishment process was quantie ed and was found to require approximately 40 ‐70 e 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.

Mars Science Laboratory Heat Shield Aerothermodynamics: Design and Reconstruction

The Mars Science Laboratory heat shield was designed to withstand a fully turbulent heat pulse using information from ground testing and computational analysis on a preflight design trajectory. Instrumentation on the flight heat shield measured in-depth temperatures to permit reconstruction of the surface heating. The data indicate that boundary-layer transition occurred at five of seven measurement locations before peak heating. Data oscillations at three pressure measurement locations may also indicate transition. This paper presents the heat shield temperature and pressure data, possible explanations for the timing of boundary-layer transition, and a comparison of reconstructed and computational heating on the actual trajectory. A smooth-wall boundary-layer Reynolds number that was used to predict transition is compared with observed transition at various heat shield locations. A single transition Reynolds number criterion does not uniformly explain the timing of boundary-layer transition observed duri...

Assessment of Laminar, Convective Aeroheating Prediction Uncertainties for Mars-Entry Vehicles

DOI: 10.2514/1.A32257An assessment of computational uncertainties is presented for numerical methods used by NASA to predictlaminar, convective aeroheating environments for Mars-entry vehicles. A survey was conducted of existingexperimental heat transfer and shock-shape data for high-enthalpy reacting-gas CO

Turbulent Aeroheating Testing of Mars Science Laboratory Entry Vehicle

An experimental investigation of turbulent aeroheating on the Mars Science Laboratory entry vehicle heat shield has been conducted in the Arnold Engineering Development Center Hypervelocity Wind Tunnel No. 9. Testing was performed on a 6-in. (0.1524 m) diameter MSL model in pure N2 gas in the tunnels Mach 8 and Mach 10 nozzles at free stream Reynolds numbers of 4.1 x 10(exp 6)/ft to 49 x 10(exp 6)/ft (1.3 x 10(exp 7)/m to 19 x 10(exp 6/ft) and 1.2 x 10(exp 6)/ft to 19 x 10(exp 6)/ft (0.39 x 10(exp 7)/m to 62 x 10(exp 7)/m), respectively. These conditions were sufficient to span the regime of boundary-layer flow from completely laminar to fully-developed turbulent flow over the entire forebody. A supporting aeroheating test was also conducted in the Langley Research Center 20-Inch Mach 6 Air Tunnel at free stream Reynolds number of 1 x 10(exp 6)/ft to 7 x 10(exp 6)/ft (0.36 x 10(exp 7)/m to 2.2 x 10(exp 7)/m) in order to help corroborate the Tunnel 9 results. A complementary computational fluid dynamics study was conducted in parallel to the wind tunnel testing. Laminar and turbulent predictions were generated for the wind tunnel test conditions and comparisons were performed with the data for the purpose of helping to define uncertainty margins on predictions for aeroheating environments during entry into the Martian atmosphere. Data from both wind tunnel tests and comparisons with the predictions are presented herein. It was concluded from these comparisons that for perfect-gas conditions, the computational tools could predict fully-laminar or fully-turbulent heating conditions to within 12% or better of the experimental data.