William L. Kleb

Exploring XP for scientific research

Can we successfully apply XP (Extreme Programming) in a scientific research context? A pilot project at the NASA Langley Research Center tested XPs applicability in this context. Since the cultural environment at a government research center differs from the customer-centric business view, eight of XPs 12 practices seemed incompatible with the existing research culture. Despite initial awkwardness, the authors determined that XP can function in situations for which it appears to be ill suited.

Efficient construction of discrete adjoint operators on unstructured grids using complex variables

A methodology is developed and implemented to mitigate the lengthy software development cycle typically associated with constructing a discrete adjoint solver for aerodynamic simulations. The approach is based on a complex-variable formulation that enables straightforward differentiation of complicated real-valued functions. An automated scripting process is used to create the complex-variable form of the set of discrete equations. An efficient method for assembling the residual and cost function linearizations is developed. The accuracy of the implementation is verified through comparisons with a discrete direct method as well as a previously developed handcoded discrete adjoint approach. Comparisons are also shown for a large-scale configuration to establish the computational efficiency of the present scheme. To ultimately demonstrate the power of the approach, the implementation is extended to high temperature gas flows in chemical nonequilibrium. Finally, several fruitful research and development avenues enabled by the current work are suggested.

Aeroheating Predictions for X-34 Using an Inviscid Boundary-Layer Method

Radiative equilibrium surface temperatures and surface heating rates from a combined inviscid-boundary layer method are presented for the X-34 Reusable Launch V ehicle for several points along the hypersonic descent portion of its trajectory. I n viscid, perfect-gas solutions are generated with the Langley Aerothermodynamic Upwind Relaxation Algorithm LAURA and the Data-Parallel Lower-Upper Relaxation DPLUR code. Surface temperatures and heating rates are then computed using the Langley Approximate Three-Dimensional Convective Heating LATCH engineering code employing both laminar and turbulent ow models. The combined inviscid-boundary layer method provides accurate predictions of surface temperatures over most of the vehicle and requires much less computational eeort than a Navier-Stokes code. This enables the generation of a more thorough aerothermal database which is necessary to design the thermal protection system and specify the vehicles ight limits.

Assessment of Radiative Heating Uncertainty for Hyperbolic Earth Entry

This paper investigates the shock-layer radiative heating uncertainty for hyperbolic Earth entry, with the main focus being a Mars return. In Part I of this work, a baseline simulation approach involving the LAURA Navier-Stokes code with coupled ablation and radiation is presented, with the HARA radiation code being used for the radiation predictions. Flight cases representative of peak-heating Mars or asteroid return are de ned and the strong influence of coupled ablation and radiation on their aerothermodynamic environments are shown. Structural uncertainties inherent in the baseline simulations are identified, with turbulence modeling, precursor absorption, grid convergence, and radiation transport uncertainties combining for a +34% and ..24% structural uncertainty on the radiative heating. A parametric uncertainty analysis, which assumes interval uncertainties, is presented. This analysis accounts for uncertainties in the radiation models as well as heat of formation uncertainties in the flow field model. Discussions and references are provided to support the uncertainty range chosen for each parameter. A parametric uncertainty of +47.3% and -28.3% is computed for the stagnation-point radiative heating for the 15 km/s Mars-return case. A breakdown of the largest individual uncertainty contributors is presented, which includes C3 Swings cross-section, photoionization edge shift, and Opacity Project atomic lines. Combining the structural and parametric uncertainty components results in a total uncertainty of +81.3% and ..52.3% for the Mars-return case. In Part II, the computational technique and uncertainty analysis presented in Part I are applied to 1960s era shock-tube and constricted-arc experimental cases. It is shown that experiments contain shock layer temperatures and radiative ux values relevant to the Mars-return cases of present interest. Comparisons between the predictions and measurements, accounting for the uncertainty in both, are made for a range of experiments. A measure of comparison quality is de ned, which consists of the percent overlap of the predicted uncertainty bar with the corresponding measurement uncertainty bar. For nearly all cases, this percent overlap is greater than zero, and for most of the higher temperature cases (T >13,000 K) it is greater than 50%. These favorable comparisons provide evidence that the baseline computational technique and uncertainty analysis presented in Part I are adequate for Mars-return simulations. In Part III, the computational technique and uncertainty analysis presented in Part I are applied to EAST shock-tube cases. These experimental cases contain wavelength dependent intensity measurements in a wavelength range that covers 60% of the radiative intensity for the 11 km/s, 5 m radius flight case studied in Part I. Comparisons between the predictions and EAST measurements are made for a range of experiments. The uncertainty analysis presented in Part I is applied to each prediction, and comparisons are made using the metrics defined in Part II. The agreement between predictions and measurements is excellent for velocities greater than 10.5 km/s. Both the wavelength dependent and wavelength integrated intensities agree within 30% for nearly all cases considered. This agreement provides confidence in the computational technique and uncertainty analysis presented in Part I, and provides further evidence that this approach is adequate for Mars-return simulations. Part IV of this paper reviews existing experimental data that include the influence of massive ablation on radiative heating. It is concluded that this existing data is not sufficient for the present uncertainty analysis. Experiments to capture the influence of massive ablation on radiation are suggested as future work, along with further studies of the radiative precursor and improvements in the radiation properties of ablation products.

Collaborative Software Development in Support of Fast Adaptive AeroSpace Tools (FAAST)

A collaborative software development approach is described. The software product is an adaptation of proven computational capabilities combined with new capabilities to form the Agencys next generation aerothermodynamic and aerodynamic analysis and design tools. To efficiently produce a cohesive, robust, and extensible software suite, the approach uses agile software development techniques; specifically, project retrospectives, the Scrum status meeting format, and a subset of Extreme Programmings coding practices are employed. Examples are provided which demonstrate the substantial benefits derived from employing these practices. Also included is a discussion of issues encountered when porting legacy Fortran 77 code to Fortran 95 and a Fortran 95 coding standard.

Comparison of CFD Predictions with Shuttle Global Flight Thermal Imagery and Discrete Surface Measurements

Surface temperature measurements from the STS-119 boundary-layer transition experiment on the space shuttle orbiter Discovery provide a rare opportunity to assess turbulent CFD models at hypersonic flight conditions. This flight data was acquired by on-board thermocouples and by infrared images taken off-board by the Hypersonic Thermodynamic Infrared Measurements (HYTHIRM) team, and is suitable for hypersonic CFD turbulence assessment between Mach 6 and 14. The primary assessment is for the Baldwin-Lomax and Cebeci-Smith algebraic turbulence models in the DPLR and LAURA CFD codes, respectively. A secondary assessment is made of the Shear-Stress Transport (SST) two-equation turbulence model in the DPLR code. Based upon surface temperature comparisons at eleven thermocouple locations, the algebraic-model turbulent CFD results average 4% lower than the measurements for Mach numbers less than 11. For Mach numbers greater than 11, the algebraic-model turbulent CFD results average 5% higher than the three available thermocouple measurements. Surface temperature predictions from the two SST cases were consistently 3 4% higher than the algebraic-model results. The thermocouple temperatures exhibit a change in trend with Mach number at about Mach 11; this trend is not reflected in the CFD results. Because the temperature trends from the turbulent CFD simulations and the flight data diverge above Mach 11, extrapolation of the turbulent CFD accuracy to higher Mach numbers is not recommended.

COMPUTATIONAL AEROHEATING PREDICTIONS FOR X-34

Radiative equilibrium surface temperatures, heating rates, streamlines, surface pressures, and flow-field features as predicted by the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) are presented for the X-34 Technology Demonstrator. Results for two trajectory points corresponding to entry peak heating and two control surface deflections are discussed. This data is also discussed in the context of Thermal Protection System (TPS) design issues. The work presented in this report is part of a larger effort to define the X-34 aerothermal environment, including the application of engineering codes and wind-tunnel studies.

Supersonic retro-propulsion experimental design for computational fluid dynamics model validation

The development of supersonic retro-propulsion, an enabling technology for heavy payload exploration missions to Mars, is the primary focus for the present paper. A new experimental model, intended to provide computational fluid dynamics model validation data, was recently designed for the Langley Research Center Unitary Plan Wind Tunnel Test Section 2. Pre-test computations were instrumental for sizing and refining the model, over the Mach number range of 2.4 to 4.6, such that tunnel blockage and internal flow separation issues would be minimized. A 5-in diameter 70-deg sphere-cone forebody, which accommodates up to four 4∶1 area ratio nozzles, followed by a 10-in long cylindrical aftbody was developed for this study based on the computational results. The model was designed to allow for a large number of surface pressure measurements on the forebody and aftbody. Supplemental data included high-speed Schlieren video and internal pressures and temperatures. The run matrix was developed to allow for the quantification of various sources of experimental uncertainty, such as random errors due to run-to-run variations and bias errors due to flow field or model misalignments. Some preliminary results and observations from the test are presented, although detailed analyses of the data and uncertainties are still on going.

Analysis of Navier-Stokes codes applied to Supersonic Retro-Propulsion wind tunnel test

This paper describes the pre-test analysis of three Navier-Stokes codes applied to a Supersonic Retro- Propulsion (SRP) wind tunnel test.1 2 Advancement of SRP as a technology hinges partially on the ability of computational methods to accurately predict vehicle aerodynamics during the SRP phase of atmospheric descent. A wind tunnel test at the Langley Unitary Plan Wind Tunnel was specifically designed to validate Navier-Stokes codes for SRP applications. The test consisted of a 5-inch diameter, 70-degree sphere-cone forebody with cylindrical afterbody, with four configurations spanning 0 to 4 jets. Test data include surface pressure (including high-frequency response), flowfield imagery, and internal pressure and temperature measurements. Three computational fluid dynamics (CFD) codes (DPLR, FUN3D, and OVERFLOW) are exercised for both single and multiple-nozzle configurations for a range of Mach (M) numbers and thrust coefficients. Comparisons to test data will be used to evaluate accuracy, identify modeling shortcomings, and gain insight into the computational requirements necessary for computing these complex flows.

Computational and Experimental Unsteady Pressures for Alternate SLS Booster Nose Shapes

Delayed Detached Eddy Simulation (DDES) predictions of the unsteady transonic flow about a Space Launch System (SLS) configuration were made with the Fully UNstructured Three-Dimensional (FUN3D) flow solver. The computational predictions were validated against results from a 2.5% model tested in the NASA Ames 11-Foot Transonic Unitary Plan Facility. The peak Cp,rms value was under-predicted for the baseline, Mach 0.9 case, but the general trends of high Cp,rms levels behind the forward attach hardware, reducing as one moves away both streamwise and circumferentially, were captured. Frequency of the peak power in power spectral density estimates was consistently under-predicted. Five alternate booster nose shapes were assessed, and several were shown to reduce the surface pressure fluctuations, both as predicted by the computations and verified by the wind tunnel results.