Engineering topics

Prabhu, Dinesh K.

Publications and source records attributed to Prabhu, Dinesh K..

At least 55 records · Page 3

NASA NTRS2012

Transition Experiments on Large Bluntness Cones with Distributed Roughness in Hypersonic Flight

Large bluntness cones with smooth nosetips and roughened frusta were flown in the NASA Ames hypersonic ballistic range at a Mach number of 10 through quiescent air environments. Global surface intensity (temperature) distributions were optically measured and analyzed to determine transition onset and progression over the roughened surface. Real-gas Navier-Stokes calculations of model flowfields, including laminar boundary layer development in these flowfields, were conducted to predict values of key dimensionless parameters used to correlate transition on such configurations in hypersonic flow. For these large bluntness cases, predicted axial distributions of the roughness Reynolds number showed (for each specified freestream pressure) that this parameter was a maximum at the physical beginning of the roughened zone and decreased with increasing run length along the roughened surface. Roughness-induced transition occurred downstream of this maximum roughness Reynolds number location, and progressed upstream towards the beginning of the roughened zone as freestream pressure was systematically increased. Roughness elements encountered at the upstream edge of the roughened frusta thus acted like a finite-extent trip array, consistent with published results concerning the tripping effectiveness of roughness bands placed on otherwise smooth surfaces.

Reda, Daniel. C.↗

NASA NTRS2012

Flowfield Analysis of a Small Entry Probe (SPRITE) Tested in an Arc Jet

A novel concept of small size (diameter less than 15 inches) entry probes named SPRITE (Small Probe Re-entry Investigation for TPS Engineering) has been developed at NASA Ames Research Center (ARC). These flight probes have on-board data acquisition systems that have also been developed in parallel at NASA ARC by Greg Swanson1. Flight probes of this size facilitate testing over a wide range of conditions in arc jets available at NASA ARC, thereby fulfilling a 'test what you fly' paradigm. As indicated by the acronym, these probes, with suitably tailored trajectories, are primarily meant to be robotic flight test beds for TPS materials, although the design is flexible enough to accommodate additional objectives of flight-testing other vehicle subsystems. A first step towards establishing the feasibility of the SPRITE concept is to arc-jet test fully instrumented models at flight scale. In a follow-on to the Large-Scale Article Tests (LSAT2) performed in the 60 MW Interaction Heating Facility (IHF) in late 2008/early 2009, a full-scale model of Deep Space-2 (DS23) made of red oak was tested in the 20 MW Aerodynamic Heating Facility (AHF). There were no issues with mass capture by the diffuser for blunt bodies of roughly 15 inches diameter tested in the 18-inch nozzle of the AHF. Building on this initial success, two identical test articles - SPRITE-T1-1 and SPRITE-T1-2 (T1 indicating the choice of back shell geometry) - were fabricated, and one of them, SPRITE-T1-1, was tested in the AHF recently. Both these test articles, 14 inches in diameter, have a 45deg sphere-cone (like DS2) made of PICA bonded on to a 1/8th inch thick aluminum shell using RTV. The aft portion of the test article is a conical frustum (15deg cone angle) with LI-2200 bonded on to the aluminum shell. Each model is fully instrumented with: (a) thermocouples imbedded in plugs in the heat shield, (b) thermocouples bonded to the aluminum substructure; the thermocouples are distributed over the entire shell, and (c) a few strain gages. Data from some of the thermocouples and gages are acquired by the on-board data acquisition system (DAS), while data from the others are routed to the facility-provided DAS, thereby enabling a cross check on the in situ measurement capability. as inputs to v2.6.1 of the in-house materials thermal response code, FIAT

atmospheric entry↗

NASA NTRS2012

Thermal Analysis of Small Re-Entry Probe

The Small Probe Reentry Investigation for TPS Engineering (SPRITE) concept was developed at NASA Ames Research Center to facilitate arc-jet testing of a fully instrumented prototype probe at flight scale. Besides demonstrating the feasibility of testing a flight-scale model and the capability of an on-board data acquisition system, another objective for this project was to investigate the capability of simulation tools to predict thermal environments of the probe/test article and its interior. This paper focuses on finite-element thermal analyses of the SPRITE probe during the arcjet tests. Several iterations were performed during the early design phase to provide critical design parameters and guidelines for testing. The thermal effects of ablation and pyrolysis were incorporated into the final higher-fidelity modeling approach by coupling the finite-element analyses with a two-dimensional thermal protection materials response code. Model predictions show good agreement with thermocouple data obtained during the arcjet test.

Agrawal, Parul↗

NASA NTRS2012

Stagnation Point Radiative Heating Relations for Venus Entry

Improved analytic expressions for calculating the stagnation point radiative heating during entry into the atmosphere of Venus have been developed. These analytic expressions can be incorporated into entry trajectory simulation codes. Together with analytical expressions for convective heating at the stagnation point, the time-integrated total heat load at the stagnation point is used in determining the thickness of protective material required, and hence the mass of the fore body heatshield of uniform thickness.

Tauber, Michael E.↗

NASA NTRS2011

Emission Spectroscopic Measurements with an Optical Probe in the NASA Ames IHF Arc Jet Facility

An optical probe was designed to measure radiation (from inside the arc heater) incident on a test sample immersed in the arc-heated stream. Currently, only crude estimates are available for this incident radiation. Unlike efforts of the past, where the probe line of sight was inclined to the nozzle centerline, the present development focuses on having the probe line of sight coincide with the nozzle centerline. A fiber-coupled spectrometer was used to measure the spectral distribution of incident radiation in the wavelength range of 225 to 900 nm. The radiation heat flux in this wavelength range was determined by integration of measured emission spectral intensity calibrated to incident irradiance from an integrating sphere. Two arc-heater conditions, corresponding to stream bulk enthalpy levels of 12 and 22 MJ/kg, were investigated in the 13-inch diameter nozzle of the Interaction Heating Facility at NASA Ames Research Center. With the probe placed at a distance of 10 inches from the nozzle exit plane, total radiative heat fluxes were measured to be 3.3 and 8.4 W/sq cm for the 12 and 22 MJ/kg conditions, respectively. About 17% of these radiative fluxes were due to bound-bound radiation from atoms and molecules, while the remaining 83% could be attributed to continua (bound-free and/or free-free). A comparison with spectral simulation based on CFD solutions for the arc-heater flow field and with spectroscopic measurements in the plenum region indicates that more than 95% of the measured radiation is generated in the arc region. The total radiative heat flux from the line radiation could increase by a factor of two through contributions from wavelengths outside the measured range, i.e., from the vacuum ultraviolet (wavelengths less than 225 nm) and the infrared (wavelengths greater than 900 nm). An extrapolation of the continuum radiation to these two wavelength regions was not attempted. In the tested configuration, the measured radiative heat flux accounts for only about 1.4% of the nominal heat flux on a flat face model and therefore is considered negligible. In the 6-inch diameter nozzle, on account of shorter path lengths, the radiation heat flux could be significant. Therefore, future tests in the 6-inch nozzle will have radiometers in addition to the optical probe.

Winter, Michael↗

NASA NTRS2011

Flowfield Analysis of a Small Entry Probe (SPRITE) Tested in an Arc Jet

Results of simulations of flow of an arc-heated stream around a 14-inch diameter 45 sphere-cone configuration are presented. Computations are first benchmarked against pressure and heat flux measurements made using copper slug calorimeters of different shapes and sizes. The influence of catalycity of copper on computed results is investigated. Good agreements between predictions and measurements are obtained by assuming the copper slug to be partially catalytic to atomic recombination. With total enthalpy estimates obtained from these preliminary computations, calculations are then performed for the test article, with the nozzle and test article considered as an integrated whole the same procedure adopted for calorimeter simulations. The resulting heat fluxes at select points on the test article (points at which fully instrumented plugs were placed) are used in material thermal response code calculations. Predicted time histories of temperature are compared against thermocouple data from the instrumented plugs, and recession determined. Good agreement is obtained for in-depth thermocouples.

Prabhu, Dinesh K.↗

NASA NTRS2011

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.

Johnston, Christopher O.↗

NASA NTRS2011

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

An assessment of computational uncertainties is presented for numerical methods used by NASA to predict laminar, convective aeroheating environments for Mars entry vehicles. A survey was conducted of existing experimental heat-transfer and shock-shape data for high enthalpy, reacting-gas CO2 flows and five relevant test series were selected for comparison to predictions. Solutions were generated at the experimental test conditions using NASA state-of-the-art computational tools and compared to these data. The comparisons were evaluated to establish predictive uncertainties as a function of total enthalpy and to provide guidance for future experimental testing requirements to help lower these uncertainties.

Hollis, Brian R.↗

NASA NTRS2011

Excited State Chemistry in the Free Stream of the NASA IHF Arc Jet Facility Observed by Emission Spectroscopy

Spectroscopic measurements of non-equilibrium emission were made in the free stream of the 60 megawatts Interaction Heating Facility at NASA Ames Research Center. In the visible near infrared wavelength region, the most prominent emission was from molecular N2, and in the ultra violet region, the spectra were dominated by emission from molecular NO. The only atomic lines observed were those of copper (an erosion product of the electrodes). The bands of the 1st Positive system of N2 (if B is true then A is true) differed significantly from spectra computed spectra assuming only thermal excitation, suggesting overpopulation of the high vibrational states of the B state of N2. Populations of these high vibrational levels (peaking at v (sub upper) equals 13) of the N2 B state were determined by scaling simulated spectra; calculations were performed for each upper vibrational state separately. The experimental-theoretical procedure was repeated for several radial positions away from the nozzle axis to obtain spatial distributions of the upper state populations; rotational symmetry of the flow was assumed in simulations. The overpopulation of the high vibrational levels has been interpreted as the effect of inverse pre-dissociation of neutral atoms in the N2 A state, which populates the N2 B state through a level crossing process at v (sub upper) is greater than 10.

arc-jet testing↗

NASA NTRS2011

Uncertainty Assessment of Hypersonic Aerothermodynamics Prediction Capability

The present paper provides the background of a focused effort to assess uncertainties in predictions of heat flux and pressure in hypersonic flight (airbreathing or atmospheric entry) using state-of-the-art aerothermodynamics codes. The assessment is performed for four mission relevant problems: (1) shock turbulent boundary layer interaction on a compression corner, (2) shock turbulent boundary layer interaction due a impinging shock, (3) high-mass Mars entry and aerocapture, and (4) high speed return to Earth. A validation based uncertainty assessment approach with reliance on subject matter expertise is used. A code verification exercise with code-to-code comparisons and comparisons against well established correlations is also included in this effort. A thorough review of the literature in search of validation experiments is performed, which identified a scarcity of ground based validation experiments at hypersonic conditions. In particular, a shortage of useable experimental data at flight like enthalpies and Reynolds numbers is found. The uncertainty was quantified using metrics that measured discrepancy between model predictions and experimental data. The discrepancy data is statistically analyzed and investigated for physics based trends in order to define a meaningful quantified uncertainty. The detailed uncertainty assessment of each mission relevant problem is found in the four companion papers.

Bose, Deepak↗

NASA NTRS2011

Flexible Ablators: Applications and Arcjet Testing

Flexible ablators were conceived in 2009 to meet the technology pull for large, human Mars Exploration Class, 23 m diameter hypersonic inflatable aerodynamic decelerators. As described elsewhere, they have been recently undergoing initial technical readiness (TRL) advancement by NASA. The performance limits of flexible ablators in terms of maximum heat rates, pressure and shear remain to be defined. Further, it is hoped that this emerging technology will vastly expand the capability of future NASA missions involving atmospheric entry systems. This paper considers four topics of relevance to flexible ablators: (1) Their potential applications to near/far term human and robotic missions (2) Brief consideration of the balance between heat shield diameter, flexible ablator performance limits, entry vehicle controllability and aft-body shear layer impingement of interest to designers of very large entry vehicles, (3) The approach for developing bonding processes of flexible ablators for use on rigid entry bodies and (4) Design of large arcjet test articles that will enable the testing of flexible ablators in flight-like, combined environments (heat flux, pressure, shear and structural tensile loading). Based on a review of thermal protection system performance requirements for future entry vehicles, it is concluded that flexible ablators have broad applications to conventional, rigid entry body systems and are enabling to large deployable (both inflatable and mechanical) heat shields. Because of the game-changing nature of flexible ablators, it appears that NASA's Office of the Chief Technologist (OCT) will fund a focused, 3-year TRL advancement of the new materials capable of performance in heat fluxes in the range of 200-600 W/sq. cm. This support will enable the manufacture and use of the large-scale arcjet test designs that will be a key element of this OCT funded activity.

Thermal Protection↗

NASA NTRS2010

Overset Techniques for Hypersonic Multibody Configurations with the DPLR Solver

Three unit problems in shock-shock/shock-boundary layer interactions are considered in the evaluation overset techniques with the Data Parallel Line Relaxation (DPLR) computational fluid dynamics solver, a three dimensional Navier-Stokes solver . The unit problems considered are those of two stacked hemispherical cylinders (of different diameters and lengths, and at various orientations relative to each other or relative to the nozzle axis) tested in a hypersonic wind tunnel. These problems are taken as representative of a Two-Stage-To-Orbit design. The objective of the present presentation would be to discuss the techniques used to develop suitable overset grid systems and then evaluate their respective solutions by comparing to corresponding point matched grid solutions and experimental data. Both successful and unsuccessful techniques would be discussed. All solutions would be calculated using the DPLR solver and SUGGAR will be used to develop the domain connectivity information.

Hyatt, Andrew James↗

NASA NTRS2010

Emission Spectroscopic Measurements in the Plenum Region of the NASA IHF Arc Jet Facility

A newly designed segment with optical access was installed in the plenum chamber of the 60 MW Interaction Heating arcjet Facility at NASA Ames Research Center. This special segment has ports located off axis, and the optical fibers can be inserted into these ports. The special segment allows for optical examination of the arc-heated flow as it enters the plenum, and thus assists in determining estimates of the thermodynamic state of the inflow to the convergent section of the nozzle. In the present work, optical emission measurements have been made in VIS-NIR region (wavelengths between 500 nm to 900 nm) for two settings of the arc heater - a 6000 A condition (high condition) with the minimum amount of radial injection of cold air in the plenum, and a 3300 A condition (low condition) with significant amount of cold air injection to reduce the enthalpy of the arc-heated stream. The results presented here were obtained using an Acton SP300i spectrometer coupled to a Princeton Instruments PI-max intensified camera. In addition to the optical emission measurements, computations were performed for the flow in the plenum and radiation along lines of sight corresponding to the optical ports. Along the centerline, i.e., the longest line of sight across the plenum cross-section, there is good agreement between computations and measurements for the high enthalpy condition, although the off-axis radial profiles show some differences. For the low enthalpy condition, there are significant differences between computations and measurements. The current working hypothesis is that the computational model does not capture details of the mixing process in the plenum.

Winter, Michael W.↗

NASA NTRS2010

Transition Experiments on Blunt Bodies with Isolated Roughness Elements in Hypersonic Free Flight

Smooth titanium hemispheres with isolated three-dimensional (3D) surface roughness elements were flown in the NASA Ames hypersonic ballistic range through quiescent CO2 and air environments. Global surface intensity (temperature) distributions were optically measured and thermal wakes behind individual roughness elements were analyzed to define tripping effectiveness. Real-gas Navier-Stokes calculations of model flowfields, including laminar boundary layer development in these flowfields, were conducted predict key dimensionless parameters used to correlate transition on blunt bodies in hypersonic flow. For isolated roughness elements totally immersed within the laminar boundary layer, critical roughness Reynolds numbers for flights in air were found to be higher than those measured for flights in CO2, i.e., it was easier to trip the CO2 boundary layer to turbulence. Tripping effectiveness was found to be dependent on trip location within the subsonic region of the blunt body flowfield, with effective tripping being most difficult to achieve for elements positioned closest to the stagnation point. Direct comparisons of critical roughness Reynolds numbers for 3D isolated versus 3D distributed roughness elements for flights in air showed that distributed roughness patterns were significantly more effective at tripping the blunt body laminar boundary layer to turbulence.

Reda, Daniel C.↗

NASA NTRS2008

Re-entry vehicle shape for enhanced performance

A convex shell structure for enhanced aerodynamic performance and/or reduced heat transfer requirements for a space vehicle that re-enters an atmosphere. The structure has a fore-body, an aft-body, a longitudinal axis and a transverse cross sectional shape, projected on a plane containing the longitudinal axis, that includes: first and second linear segments, smoothly joined at a first end of each the first and second linear segments to an end of a third linear segment by respective first and second curvilinear segments; and a fourth linear segment, joined to a second end of each of the first and second segments by curvilinear segments, including first and second ellipses having unequal ellipse parameters. The cross sectional shape is non-symmetric about the longitudinal axis. The fourth linear segment can be replaced by a sum of one or more polynomials, trigonometric functions or other functions satisfying certain constraints.

Brown, James L.↗

NASA NTRS2005

Shuttle Orbiter Contingency Abort Aerodynamics: Real-Gas Effects and High Angles of Attack

An important element of the Space Shuttle Orbiter safety improvement plan is the improved understanding of its aerodynamic performance so as to minimize the "black zones" in the contingency abort trajectories [1]. These zones are regions in the launch trajectory where it is predicted that, due to vehicle limitations, the Orbiter will be unable to return to the launch site in a two or three engine-out scenario. Reduction of these zones requires accurate knowledge of the aerodynamic forces and moments to better assess the structural capability of the vehicle. An interesting aspect of the contingency abort trajectories is that the Orbiter would need to achieve angles of attack as high as 60deg. Such steep attitudes are much higher than those for a nominal flight trajectory. The Orbiter is currently flight certified only up to an angle of attack of 44deg at high Mach numbers and has never flown at angles of attack larger than this limit. Contingency abort trajectories are generated using the data in the Space Shuttle Operational Aerodynamic Data Book (OADB) [2]. The OADB, a detailed document of the aerodynamic environment of the current Orbiter, is primarily based on wind-tunnel measurements (over a wide Mach number and angle-of-attack range) extrapolated to flight conditions using available theories and correlations, and updated with flight data where available. For nominal flight conditions, i.e., angles of attack of less than 45deg, the fidelity of the OADB is excellent due to the availability of flight data. However, at the off-nominal conditions, such as would be encountered on contingency abort trajectories, the fidelity of the OADB is less certain. The primary aims of a recent collaborative effort (completed in the year 2001) between NASA and Boeing were to determine: 1) accurate distributions of pressure and shear loads on the Orbiter at select points in the contingency abort trajectory space; and 2) integrated aerodynamic forces and moments for the entire vehicle and the control surfaces (body flap, speed brake, and elevons). The latter served the useful purpose of verification of the aerodynamic characteristics that went into the generation of the abort trajectories.

Prabhu, Dinesh K.↗

NASA NTRS1992

Numerical simulations of non-equilibrium shock layers with efficient implicit schemes

Current and future calculations of nonequilibrium shock layers require the use of a very large number of equations, due to a multiplicity of chemical species, excited states, and internal energy modes. The computational cost associated with the use of standard implicit methods becomes prohibitive; it is therefore desirable to examine the potential of several methods and determine if any can be projected to be more efficient and accurate for large systems of equations. Here, the performance of several implicit schemes on several simple practical examples of reacting flows is examined. The Euler equations are solved by three different implicit methods, and two methods of coupling between the fluid dynamics and the chemistry are studied. Several cases of stiffness are considered and both 1D and 2D examples are computed.

Cambier, Jean-Luc↗

NASA NTRS1992

Numerical simulation of unsteady flow in a hypersonic shock tunnel facility

This paper describes the computational work performed on the simulation of a 16-in shock-tunnel facility. The numerical problems encountered during the computation of these flows are discussed along with the validity of some approximations used, notably concerning the reduction of the problem into problems of smaller dimensionality. Quasi-1D simulations can be used to help design experiments, or to better understanding the characteristics of the facility. An application to the design of a nonintrusive diagnostic is shown. The multidimensional flow transients computed include the shock reflection at the end of the driven tube, the shock propagation down the nozzle, and the breaking of the main diaphragm.

Cambier, Jean-Luc↗