Kerry Trumble

Post-flight Aerothermal Analysis of the Stardust Sample Return Capsule

The reentry of the Stardust sample return capsule was captured by several optical instruments through an observation campaign aboard the NASA DC-8 airborne observatory. Flow environments obtained from computational fluid dynamics solutions are loosely coupled with material response modeling to predict the surface temperature and the observed continuum emission of Stardust throughout the reentry. The calculated surface temperatures are compared with the data from several spectral instruments onboard the airborne observatory, including the ECHELLE (echelle-based spectrograph for the crisp and high efficient detection of low light emission) camera and conventional spectrometer in Czerny–Turner configuration. The ECHELLE camera recorded spectral intensity at a period in the trajectory before peak heating. The graybody curves corresponding to the average and area-averaged surface temperatures predicted by the computational fluid dynamics and material response coupled simulation have excellent agreement with the recorded data at altitudes lower than 74 km. At these altitudes, the computational fluid dynamics and material response coupling agrees with the surface temperature to within 50 K. The computational fluid dynamics calculation without the material response modeling overestimates surface temperatures because it does not take into account such things as ablation. The overprediction of the computational fluid dynamics and material response simulated surface temperature early in the trajectory coincides with highemission intensity lines corresponding to thermal paint products. The presence of paint on the heat shield could have contributed to the lower observed surface temperatures and could explain the overprediction by the simulated data, which does not account for the paint. The average surface temperatures resulting from the spectrometer in Czerny– Turner configuration telescope analysis agree to within less than 5% with the average surface temperatures predicted by the material response. This observation period included the point of peak heating. The calculated flux based on the surface temperature agrees well with the observed flux. Surface temperature is one of the critical parameters used in the design of thermal protection systems, because it is an indicator of material performance. The coupled computational fluid dynamics and material response approach employed in the present analysis increases confidence for future missions such as the crew exploration vehicle Orion.

Toward Supersonic Retropropulsion CFD Validation

This paper begins the process of verifying and validating computational fluid dynamics (CFD) codes for supersonic retropropulsive flows. Four CFD codes (DPLR, FUN3D, OVERFLOW, and US3D) are used to perform various numerical and physical modeling studies toward the goal of comparing predictions with a wind tunnel experiment specifically designed to support CFD validation. Numerical studies run the gamut in rigor from code-to-code comparisons to observed order-of-accuracy tests. Results indicate that this complex flowfield, involving time-dependent shocks and vortex shedding, design order of accuracy is not clearly evident. Also explored is the extent of physical modeling necessary to predict the salient flowfield features found in high-speed Schlieren images and surface pressure measurements taken during the validation experiment. Physical modeling studies include geometric items such as wind tunnel wall and sting mount interference, as well as turbulence modeling that ranges from a RANS (Reynolds-Averaged Navier-Stokes) 2-equation model to DES (Detached Eddy Simulation) models. These studies indicate that tunnel wall interference is minimal for the cases investigated; model mounting hardware effects are confined to the aft end of the model; and sparse grid resolution and turbulence modeling can damp or entirely dissipate the unsteadiness of this self-excited flow.

Nonequilibrium Particle and Continuum Analyses of Stardust Entry for Near{Continuum Conditions

The Stardust Sample Return Capsule (SRC) entered the Earth’s atmosphere at a velocity of 12.6 km/s. At high altitude, the o w eld is expected to be in a strong state of thermochemical nonequilibrium. In the present study, both continuum (CFD) and particle (DSMC) methods are used to analyze the forebody o w of the SRC at an altitude of 81 km where the freestream Knudsen number is about 0.005. The very large entry velocity represents a highly energetic condition for which the thermochemistry models are not well calibrated. Direct comparisons between baseline CFD and DSMC models give enormous dierences in basic o w eld properties. To study the discrepancy between the solutions, dieren t methods for determining the temperature used by CFD to control the dissociation and ionization reactions are investigated. Also, a new model is introduced for the DSMC technique that makes it possible to simulate reverse direction chemical reactions in a manner more consistent with that used in CFD. While the revised CFD and DSMC results are in better agreement with each other, under these highly-energetic, near-continuum o w conditions, signican t dierences remain between continuum and particle solutions. Additional CFD computations performed at lower altitude indicate, as expected, that o w eld results become less sensitive to details of the chemistry modeling further into the continuum regime.

Modeling of Stardust Entry at High Altitude, Part 1: Flowfield Analysis

DOI: 10.2514/1.37360 The Stardust sample return capsule entered the Earth’s atmosphere at a very energetic velocity of 12:6 km=s .I n the present study, both continuum (computational fluid dynamics) and particle (direct simulation Monte Carlo) methods are used to analyze the forebody flow of the Stardust sample return capsule at altitudes of 81 and 71 km, where the flow is in the near-continuum regime. At the higher altitude, direct comparisons between baseline computational fluiddynamicsanddirectsimulationMonteCarlomodelsgiveenormousdifferencesinbasic flowfield properties. To study the discrepancy between the solutions, a modified approach for determining the temperature used by computational fluid dynamics to control the dissociation and ionization reactions is investigated. The modified computational fluid dynamics and direct simulation Monte Carlo results are in significantly better agreement with each other, illustrating the strong sensitivity to chemistry modeling under these highly energetic conditions. Significant differences persist in temperatures near the capsule surface and in surface heat flux. Evaluation of local Knudsen numbers indicates that the flow experiences noncontinuum behavior in the shock front andatthecapsulesurfacethatexplainsthesmallerheat fluxpredictedbydirectsimulationMonteCarlo.Atthelower altitude, the flowfield results become less sensitive to details of the chemistry modeling, although noncontinuum effects are again predicted at the stagnation point.

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.

Near-Ultraviolet Emission Spectroscopy During an Airborne Observation of the Stardust Reentry

Thermal radiation of the heatshield and the emission of the postshock layer around the Stardust capsule, during its reentry, were detected by a NASA-led observation campaign aboard NASA’s DC-8 airborne observatory involving teams from several nations. The German SLIT experiment used a conventional spectrometer, in a Czerny–Turner configuration (300 mm focal length and a 600 lines=mm grating), fed by fiber optics, to cover a wavelength range from 324 to 456 nm with a pixel resolution of 0.08 nm. The reentering spacecraft was tracked manually using a camerawith a viewangle of 20 deg, and light from the capsulewas collected using a smallmirror telescopewith a view angle of only 0.45 deg. Data were gathered with a measurement frequency of 5 Hz in a 30-s time interval around the point of maximum heating until the capsule left the field of view. The emission of carbon nitride (as a major ablation product), N 2 and different atoms were monitored successfully during that time. Because of the nature of the experimental setup, spatial resolution of the radiation field was not possible. Therefore, all measured values represent an integration of radiation from the visible part of the glowing heatshield, and from the plasma in the postshock region. Further, due to challenges in tracking, not every spectrum gathered contained data. Themeasured spectra can be split up into two parts: 1) continuum spectra, which represent a superposition of the heatshield radiation and the continuum radiation of particles due to microspallation in the plasma, and 2) line spectra from the plasma in the shock layer. Planck temperatures (interpreted as the surface temperatures of the Stardust heatshield) were determined assuming either a constant surface temperature, or a temperature distribution deduced from numerical simulation. The constant surface temperatures are in good agreement with numerical simulations, but the peak values at the stagnation point are significantly lower than those in the numerical simulation if a temperature distribution over the surface is assumed. Emission bands of carbon nitride and N 2 were tracked along the visible trajectory and compared with a spectral simulation with satisfying agreement. Values for the integrated radiation of the transitions of interest for these species were extracted from this comparison.

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.

Continuing Validation of Computational Fluid Dynamics for Supersonic Retropropulsion

A large step in the validation of Computational Fluid Dynamics (CFD) for Supersonic Retropropulsion (SRP) is shown through the comparison of three Navier-Stokes solvers (DPLR, FUN3D, and OVERFLOW) and wind tunnel test results. The test was designed specifically for CFD validation and was conducted in the Langley supersonic 4 x4 Unitary Plan Wind Tunnel and includes variations in the number of nozzles, Mach and Reynolds numbers, thrust coefficient, and angles of orientation. Code-to-code and code-to-test comparisons are encouraging and possible error sources are discussed.

Development of Supersonic Retro-Propulsion for Future Mars Entry, Descent, and Landing Systems

1 ____________________________________________ * Aerospace Engineer, Atmospheric Flight & Entry Systems Branch, MS 489, Karl.T.Edquist@nasa.gov, Senior Member. † Aerospace Engineer, Atmospheric Flight & Entry Systems Branch, MS 489, Member. ‡ Propulsion Systems Engineer, Propulsion Systems Branch, MS EP4. § Senior Engineer, Aeroscience & Flight Mechanics Division, MS EG5, Member. ¶ Systems Engineer, Entry, Descent, and Landing Systems and Advanced Technologies Group, MS 321-220. # Research Scientist, Aerothermodynamics Branch, MS 230-2, Member. ** Aerospace Engineer, Systems Analysis Branch, MS 258-1, Member. †† Graduate Research Assistant, Daniel Guggenheim School of Aerospace Engineering, Student Member. Development of Supersonic Retro-Propulsion for Future Mars Entry, Descent, and Landing Systems