A Galilean-invariant field equation is proposed and tested on standard turbulence model test cases. The field equation provides an additional non-dimensional outer scale which allows the turbulence models to reproduce the axial normal stress increase with Re seen in high Reynolds numbers experiments. The field equation provides a Reynolds number for every point based on the length of turbulent flow upstream of that point in the domain. This outer scale equation can be considered an odometer that gives a length scale conjectured to be related to the large stream wise structures that are seen in turbulent flow and that require run length to develop. A new RANS model using this additional scale is able to match the Reynolds number variation of the normal stresses seen at high Reynolds number. Furthermore, the good attached flow prediction capabilities of current RANS models appears to be attained. Using this scale equation, the entire Reynolds-stress state appears to be predicted correctly, over a large run length Reynolds number range such as experienced in aircraft design.

Adaptive Mesh Refinement (AMR) promises a much more computationally efficient means to obtain a discrete approximation to a continuous boundary value problem of a specified accuracy than classic isotropic refinement. The AMR capability of OVERFLOW (a computational fluid dynamics (CFD) code) is utilized to provide estimates of the exact analytical solutions to problems of interest to turbulence modeling. Predictions of surface pressure and skin friction, essentially the state of stress at the surface, shows little difference with grids believed to be grid resolved. Velocity profiles, on the other hand, show marked differences in flows with shocks. The AMR method, as implemented in OVERFLOW 2.2k, appears to provide the ability to produce arbitrarily accurate solutions at a predictable cost much smaller than classic uniform mesh refinement.

Revised versions of Lag methodology Reynolds-stress and triple product models are applied to accepted test cases to assess the improvement, or lack thereof, in the prediction capability of the models. The Bachalo-Johnson bump flow is shown as an example for this abstract submission.

This paper presents the results of a NASA initiated Agency-wide assessment to better characterize the risks and potential mitigation approaches associated with landing human class payloads on Mars. Due to the criticality and long-lead nature of advancing Entry, Descent, and Landing (EDL) techniques, it is necessary to determine an appropriate strategy to improve the capability to land large payloads. A key focus of this study was to understand the key EDL risks with a focus on determining what “must” be tested at Mars. This process identified the various risks and potential risk mitigation strategies, along with the required key near-term technology development efforts and in what environment those technology demonstrations were best suited. The study identified key risks along with advantages to each entry technology. In addition, it was determined that with the EDL concept of operations (con ops) which minimized large scale transition events during entry, there was no technology requirement for a Mars pre-cursor demonstration as a necessary risk-mitigation test. Instead, NASA should take a direct path to a human-scale lander.

Computational assessments were performed to size boundary layer trips for a scaled Apollo capsule model in the High Enthalpy Shock Tunnel (HIEST) facility at the JAXA Kakuda Space Center in Japan. For stagnation conditions between 2 MJ/kg and 20 MJ/kg and between 10 MPa and 60 MPa, the appropriate trips were determined to be between 0.2 mm and 1.3 mm high, which provided kappa/delta values on the heatshield from 0.15 to 2.25. The tripped configuration consisted of an insert with a series of diamond shaped trips along the heatshield downstream of the stagnation point. Surface heat flux measurements were obtained on a capsule with a 250 mm diameter, 6.4% scale model, and pressure measurements were taken at axial stations along the nozzle walls. At low enthalpy conditions, the computational predictions agree favorably to the test data along the heatshield centerline. However, agreement becomes less favorable as the enthalpy increases conditions. The measured surface heat flux on the heatshield from the HIEST facility was under-predicted by the computations in these cases. Both smooth and tripped configurations were tested for comparison, and a post-test computational analysis showed that kappa/delta values based on the as-measured stagnation conditions ranged between 0.5 and 1.2. Tripped configurations for both 0.6 mm and 0.8 mm trip heights were able to effectively trip the flow to fully turbulent for a range of freestream conditions.

No AccessTechnical NoteExperimental Study of High-Enthalpy Heat Flux Augmentation in Shock TunnelsHideyuki Tanno, Tomoyuki Komuro, Randolph P. Lillard and Joseph OlejniczakHideyuki TannoJapan Aerospace Exploration Agency, Kakuda, Miyagi 9811525, Japan*Senior Researcher, Research Unit IV, Research and Development Directorate, Kakuda Space Center; . Member AIAA.Search for more papers by this author, Tomoyuki KomuroJapan Aerospace Exploration Agency, Kakuda, Miyagi 9811525, Japan†Senior Adviser, Research Unit IV, Research and Development Directorate, Kakuda Space Center.Search for more papers by this author, Randolph P. LillardNASA Johnson Space Center, Houston, Texas 77058‡Technology Demonstrations Mission Program Executive, Space Technology Mission Directorate.Search for more papers by this author and Joseph OlejniczakNASA Ames Research Center, Moffett Field, California 94035§Manager, MPCV Orion Aerosciences.Search for more papers by this authorPublished Online:8 Sep 2015https://doi.org/10.2514/1.T4478SectionsRead Now ToolsAdd to favoritesDownload citationTrack citations ShareShare onFacebookTwitterLinked InRedditEmail About References [1] Olenjniczak J., Wright M. J., Laurence S. J. and Hornung H. G., “Computational Modeling of T5 Laminar and Turbulent Heating Data on Blunt Cones, Part 1: Titan Applications,” 43rd AIAA Aerospace Sciences Meeting and Exhibit, AIAA Paper 2005-0176, Jan. 2005. LinkGoogle Scholar[2] Marineau E. C., Laurence S. J. and Hornung H. G., “Apollo-Shaped Capsule Boundary Layer Transition at High-Enthalpy in T5,” 48th AIAA Aerospace Sciences Meeting, AIAA Paper 2010-0446, Jan. 2010. LinkGoogle Scholar[3] Hollis B. R. and Prabhu D. K., “Assessment of Laminar, Convective Aeroheating Prediction Uncertainties for Mars Entry Vehicles,” 42nd AIAA Thermophysics Conference, AIAA Paper 2011-3144, June 2011. LinkGoogle Scholar[4] Itoh K., Ueda S., Tanno H., Komuro T. and Sato K., “Hypersonic Aerothermodynamic and Scramjet Research Using High Enthalpy Shock Tunnel,” Shock Waves, Vol. 12, No. 2, 2002, pp. 93–98. doi:https://doi.org/10.1007/s00193-002-0147-0 SHWAEN 0938-1287 CrossrefGoogle Scholar[5] Tanno H., Sato K., Komuro T., Takahashi M., Itoh K., Ishihara T., Ogino Y. and Sawada K., “Heat Flux Measurement of Re-Entry Capsule Models Under High-Enthalpy High-Pressure Condition,” Proceedings of the 7th European Symposium on Aerothermodynamics, ESA Communications ESA-SP-692, ESTEC, Noordwijk, The Netherlands, Aug. 2011. Google Scholar[6] Detra R. W., Kemp N. H. and Riddell F. R., “Addendum to Heat Transfer to Satellite Vehicles Reentering the Atmosphere,” Jet Propulsion, Vol. 27, No. 12, 1957, pp. 1256–1257. doi:https://doi.org/10.2514/8.12510 JETPAV 0095-8751 Google Scholar[7] Imada T., “Concept Study of HTV-R (HTV-Return),” Journal of Space Technology and Science, Vol. 27, No. 2, 2013, pp. 2_1–2_8. doi:https://doi.org/10.11230/jsts.27.2_1 Google Scholar[8] Sanderson S. and Sturtevant B., “Transient Heat Flux Measurement Using a Surface Junction Thermocouple,” Review of Scientific Instruments, Vol. 73, No. 7, 2002, pp. 2781–2787. doi:https://doi.org/10.1063/1.1484255 RSINAK 0034-6748 CrossrefGoogle Scholar[9] Prabhu R. K. and Erickson W. D., “A Rapid Method for the Computation of Equilibrium Chemical Composition of Air to 15000 K,” NASA TP-2792, 1988. Google Scholar[10] Takahashi M., Kodera M., Itoh K., Komuro T., Sato K. and Tanno H., “Influence of Thermal Non-Equilibrium on Nozzle Flow Condition of High Enthalpy Shock Tunnel HIEST,” 16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference, AIAA Paper 2009-7267, 2009. LinkGoogle Scholar Previous article Next article

Radiance measurements in air at enthalpies from 8-20 MJkg have been made over a 250mm diameter flat-faced test article in Japan Aerospace Exploration Agency's HIgh-Enthalpy Shock Tunnel (HIEST). Measurements were made in the ultraviolet region (200-400 nm wavelength) in an attempt to resolve the long-standing discrepancy between theoryand measurements of heat flux over a blunt body; this discrepancy is often attributed toradiation. The spectra obtained indicate the presence of atomic iron vapor in the flowfield.At the highest enthalpies, the radiance is at the blackbody limit. An attempt to model theradiance is made by taking a nominal CFD flowfield without any contamination productsand processing it through a line-by-line radiation simulation tool. Iron vapor is introducedinto the shocked gas ahead of the model and radiation computations are repeated; the molefraction of iron vapor is adjusted to match the data. For the higher enthalpy conditions, theradiance was strongly absorbed and it was necessary to adjust the temperature and NOdensity in the freestream to match the signal below 300 nm. Once the observed spectrawere satisfactorily matched, the radiance to the stagnation point was then computed. It isshown that the impurity radiation is sufficiently large to explain the discrepancy.

The performance of the new LagRST models is compared with curvature corrected eddyviscosity RANS models. The Reynolds-stress models provide very good predictions of thisow without any additional curvature corrections. Along with comparison of predictionsof the vortex itself, some of the other measurements included in the oweld are comparedwith the predictions in the wake and over the wing. Dierences in predictions between thevarious RANS models in the wake region are small - not substantial, and all predictionsare in line with experimental measurements from 0.005c to 0.67c downstream of the trailingedge, at odds with what self-similar wake analysis would predict.

An aeroheating measurement test campaign of an Apollo capsule model with laminar and turbulent boundary layer was performed in the free-piston shock tunnel HIEST at JAXA Kakuda Space Center. A 250mm-diameter 6.4%-scaled Apollo CM capsule model made of SUS-304 stainless steel was applied in this study. To measure heat flux distribution, the model was equipped with 88 miniature co-axial Chromel-Constantan thermocouples on the heat shield surface of the model. In order to promote boundary layer transition, a boundary layer trip insert with 13 "pizza-box" isolated roughness elements, which have 1.27mm square, were placed at 17mm below of the model geometric center. Three boundary layer trip inserts with roughness height of k=0.3mm, 0.6mm and 0.8mm were used to identify the appropriate height to induce transition. Heat flux records with or without roughness elements were obtained for model angles of attack 28 under stagnation enthalpy between H(sub 0)=3.5MJ/kg to 21MJ/kg and stagnation pressure between P(sub 0)=14MPa to 60MPa. Under the condition above, Reynolds number based on the model diameter was varied from 0.2 to 1.3 million. With roughness elements, boundary layer became fully turbulent less than H(sub 0)=9MJ/kg condition. However, boundary layer was still laminar over H(sub 0)=13MJ/kg condition even with the highest roughness elements. An additional experiment was also performed to correct unexpected heat flux augmentation observed over H(sub 0)=9MJ/kg condition.

Development of boundary layer trip for capsule shaped reentry vehicles under high-enthalpy hypersonic condition was developed in the free-piston shock tunnel HIEST. Wind tunnel test campaign with an 6% scaled Apollo CM model was conducted with stagnation pressure of 30MPa to 60Mpa and with stagnation enthalpy from 3.5MJ/kg to 22MJ/kg. Reynolds number based on model diameter was varied from 0.2 to 1.3 million. To measure heat flux around the model, eighty-four miniature co-axial thermocouples were instrumented on the heat shield surface of the model. Angle-of-attack of the model was fixed to 28degree. A pizza-box configuration boundary layer trip was mounted on the heat shield surface of the model, which trip height k were 0.3mm, 0.6mm and 0.8mm. With k=0.6mm height trip, boundary layer became fully turbulent less than H0=10MJ/kg. However, boundary layer was still laminar over H0=15MJ/kg even with k=0.8mm trip.

This paper describes the addition of an overset grid capability to the DPLR flow solver for hypersonic flow in thermochemical nonequilibrium. Modifications to the preexisting flow solver were simplified through the use of DiRTlib, a “solver neutral” library of overset utilities. The new capability is demonstrated on a series of examples, including the Orion Crew Module and other reentry vehicles. For the overset grids used in these examples, the hole cutting and interpolation stencils were determined using SUGGAR, a generalized grid assembly code that can naturally accommodate both the three-dimensional and true two-dimensional cell-centered discretization schemes in DPLR. First a series of building-block examples are presented which highlight aspects of the new capability and assess the technique with comparisons to baseline, block-structured discretizations. The new capability is then exercised for the specific case of a tension tie geometry protruding from the Orion heatshield at both wind tunnel and flight conditions. The addition of overset capability to the DPLR flow solver is seen to be an essential feature for analyzing increasingly complex geometries in thermochemical nonequilibrium.

The Project Orion Crew Exploration Vehicle aerothermodynamic experimentation strategy, as it relates to flight database development, is reviewed. Experimental data has been obtained to both validate the computational predictions utilized as part of the database and support the development of engineering models for issues not adequately addressed with computations. An outline is provided of the working groups formed to address the key deficiencies in data and knowledge for blunt reentry vehicles. The facilities utilized to address these deficiencies are reviewed, along with some of the important results obtained thus far. For smooth wall comparisons of computational convective heating predictions against experimental data from several facilities, confidence was gained with the use of algebraic turbulence model solutions as part of the database. For cavities and protuberances, experimental data is being used for screening various designs, plus providing support to the development of engineering models. With the reaction-control system testing, experimental data were acquired on the surface in combination with off-body flow visualization of the jet plumes and interactions. These results are being compared against predictions for improved understanding of aftbody thermal environments and uncertainties.

An investigation of the aeroheating environment of the Project Orion Crew Exploration Vehicle was performed in the Arnold Engineering Development Center Hypervelocity Wind Tunnel 9 Mach 8 and Mach 10 nozzles and in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel. Heating data were obtained using a thermocouple-instrumented similar to 0.035-scale model [0.1778 m (7 in.) diameter] of the flight vehicle. Runs were performed in the Tunnel 9 Mach 10 nozzle at freestream unit Reynolds numbers of 1 x 10(6) to 20 x 10(6)/ft, in the Tunnel 9 Mach 8 nozzle at freestream unit Reynolds numbers of 8 x 10(6) to 48 x 10(6)/ft, and in the 20-Inch Mach 6 Air Tunnel at freestream unit Reynolds numbers of 1 x 10(6) to 7 x 10(6)/ft. In both facilities, enthalpy levels were low and the test gas (N(2) in Tunnel 9 and air in the 20-Inch Mach 6 Air Tunnel) behaved as a perfect gas. These test conditions produced laminar, transitional, and turbulent data in the Tunnel 9 Mach 10 nozzle; transitional and turbulent data in the Tunnel 9 Mach 8 nozzle; and laminar and transitional data in the 20-Inch Mach 6 Air Tunnel. Laminar and turbulent predictions were generated for all wind-tunnel test conditions, and comparisons were performed with the experimental data to help define the accuracy of the computational method. In general, it was found that both laminar data and predictions and turbulent data and predictions agreed to within less than the estimated +/- 12% experimental uncertainty estimate. Laminar heating distributions from all three data sets were shown to correlate well and demonstrated Reynolds numbers independence when expressed in terms of the Stanton number based on adiabatic-wall-recovery enthalpy. Transition-onset locations on the lee-side centerline were determined from the data and correlated in terms of boundary-layer parameters. Finally, turbulent heating augmentation ratios were determined for several body-point locations and correlated in terms of the boundary-layer momentum Reynolds number.

An investigation of the aeroheating environment of the Project Orion Crew Entry Vehicle has been performed in the Arnold Engineering Development Center Tunnel 9. Data were measured on a approx. 3.5% scale model (0.1778m/7-inch diam.) of the vehicle using coaxial thermocouples in the Mach 8 and Mach 10 nozzles of Tunnel 9. Runs were performed at free stream Reynolds numbers of 1 106/ft to 20 10(exp 6)/ft in the Mach 10 nozzle and 8 10(exp 6)/ft to 48 10(exp 6)/ft in the Mach 8 nozzle. The test gas in Tunnel 9 is pure N2, which at these operating conditions remains un-dissociated and may be treated as a perfect gas. At these conditions, laminar, transitional, and turbulent flow was produced on the model at Mach 10, and transitional and turbulent conditions were produced on the model at Mach 8. The majority of runs were made on a clean, smooth-surface model configuration and a limited number of runs were made in which inserts with varying boundary-layer trips configurations were used to force the occurrence of transition. Laminar and turbulent predictions were generated for all wind tunnel test conditions and comparisons were performed with the data for the purpose of helping to define uncertainty margins for the computational method. Data from both the wind tunnel test and the computations are presented herein. Figure 1 shows a schematic of the thermocouple locations on the model and figures 2 and 3 show a photo and schematic of the AEDC Hypervelocity Tunnel 9. Figure 4 shows a typical grid used in the computations. From the comparisons shown in figures 5 through 8 it was concluded that for perfect-gas conditions, the computations could predict either fully-laminar or full-turbulent flow to within +/-10% of the experimental data. The experimental data showed that transition began on the leeside of the heatshield at a free stream Reynolds number of 9 10(exp 6)/ft in the Mach 10 nozzle and fully-developed turbulent flow was produced at 20 10(exp 6)/ft. In the Mach 8 nozzle, transition on the leeside of the heat-shield was observed for all test conditions, and full-developed turbulent flow occurred at a free stream Reynolds number of 18 10(exp 6)/ft. On the aftbody of the vehicle no evidence of turbulence was detected at Mach 10 conditions, and at Mach 8 conditions, transition appeared to begin on the windside of the aftbody at free stream Reynolds number of 18x10(exp 6)/ft with fully-developed turbulent flow occurring only at the highest test condition of 48x10(exp 6)/ft.

An investigation of the aeroheating environment of the Project Orion Crew Exploration Vehicle was performed in the Arnold Engineering Development Center Hypervelocity Wind Tunnel No. 9 Mach 8 and Mach 10 nozzles and in the NASA Langley Research Center 20 - Inch Mach 6 Air Tunnel. Heating data were obtained using a thermocouple-instrumented approx.0.035-scale model (0.1778-m/7-inch diameter) of the flight vehicle. Runs were performed in the Tunnel 9 Mach 10 nozzle at free stream unit Reynolds numbers of 1x10(exp 6)/ft to 20x10(exp 6)/ft, in the Tunnel 9 Mach 8 nozzle at free stream unit Reynolds numbers of 8 x 10(exp 6)/ft to 48x10(exp 6)/ft, and in the 20-Inch Mach 6 Air Tunnel at free stream unit Reynolds numbers of 1x10(exp 6)/ft to 7x10(exp 6)/ft. In both facilities, enthalpy levels were low and the test gas (N2 in Tunnel 9 and air in the 20-Inch Mach 6) behaved as a perfect-gas. These test conditions produced laminar, transitional and turbulent data in the Tunnel 9 Mach 10 nozzle, transitional and turbulent data in the Tunnel 9 Mach 8 nozzle, and laminar and transitional data in the 20- Inch Mach 6 Air Tunnel. Laminar and turbulent predictions were generated for all wind tunnel test conditions and comparisons were performed with the experimental data to help define the accuracy of computational method. In general, it was found that both laminar data and predictions, and turbulent data and predictions, agreed to within less than the estimated 12% experimental uncertainty estimate. Laminar heating distributions from all three data sets were shown to correlate well and demonstrated Reynolds numbers independence when expressed in terms of the Stanton number based on adiabatic wall-recovery enthalpy. Transition onset locations on the leeside centerline were determined from the data and correlated in terms of boundary-layer parameters. Finally turbulent heating augmentation ratios were determined for several body-point locations and correlated in terms of the boundary-layer momentum Reynolds number.