ALBERT

Description: Submitted in partial fulfillment of the requirements for the degree of Doctor of Philosophy at the Massachusetts Institute of Technology and the Woods Hole Oceanographic Institution May 1998

Description: Planktonic protozoan grazers have the potential to significantly affect the chemistry of particle-associated trace metals. This is due both to the importance of protists as consumers of bacterial-sized particles, and to the unique low-pH, enzyme-rich microenvironment of the grazer food vacuole. This thesis examines the role of protozoan grazers in the marine geochemistry of strongly hydrolyzed, particle-reactive trace metals, in particular Th and Fe. A series of tracer experiments was carried out in model systems in order to determine the effect of grazer-mediated transformations on the chemical speciation and partitioning of radioisotopes C9Fe, 234Th, 51Cr) associated with prey cells. Results indicate that protozoan grazers are equally able to mobilize intracellular and extracellular trace metals. In some cases, protozoan regeneration of trace metals appears to lead to the formation of metal-organic complexes. Protozoan grazing may generate colloidal material that can scavenge trace metals and, via aggregation, lead to an increase in the metal/organic carbon ratio of aggregated particles. Model system experiments were also conducted in order to determine the effect of grazers on mineral phases, specifically colloidal iron oxide (ferrihydrite). Several independent techniques were employed, including size fractionation ors9Fe-labeled colloids, competitive ligand exchange, and iron-limited diatoms as "probes" for bioavailable Fe. Experimental evidence strongly suggests that protozoan grazing can affect the surface chemistry and increase the dissolution rate of iron oxide phases through phagotrophic ingestion. In further work on protozoan-mediated dissolution of colloidal Fe oxides, a novel tracer technique was developed based on the synthesis of colloidal ferrihydrite impregnated with 133Ba as an inert tracer. This technique was shown to be a sensitive, quantitative indicator for the extent of ferrihydrite dissolution/alteration by a variety of mechanisms, including photochemical reduction and ligand-mediated dissolution. In field experiments using this technique, grazing by naturally occuring protistan assemblages was shown to significantly enhance the dissolution rate of colloidal ferrihydrite over that in non-grazing controls. Laboratory and field results indicate that, when integrated temporally over the entire euphotic zone, protozoan grazing may equal or exceed photoreduction as a pathway for the dissolution of iron oxides.

Description: This work was financially supported by a Department of Defense ONR-NDSEG Graduate Fellowship, Office ofNaval Research AASERT Award (N00014-94-1-0711), and the National Science Foundation EGB Program (OCE-9523910).

Description: No abstract available

Description: Marshall Space Flight Center (MSFC) is the National Aeronautics and Space Administration (NASA)-designated center for the development of space launch systems. MSFC is particularly known for propulsion system development. Many engineering skills and technical disciplines are needed to accomplish this mission. This presentation will focus on the work of the Fluid Dynamics Branch (ER42). ER42 resides in the Propulsion Systems Department at MSFC. The branch is responsible for all aspects of the discipline of fluid dynamics applied to propulsion or propulsion-induced loads and environments. This work begins with design trades and parametric studies, and continues through development, risk assessment, anomaly investigation and resolution, and failure investigations. Applications include the propellant delivery system including the main propulsion system (MPS) and turbomachinery; combustion devices for liquid engines and solid rocket motors; coupled systems; and launch environments. An advantage of the branch is that it is neither analysis nor test centric, but discipline centric. Fluid dynamics assessments are made by analysis, from lumped parameter modeling through unsteady computational fluid dynamics (CFD); testing, which can be cold flow or hot fire; or a combination of analysis and testing. Integration of all discipline methods into one branch enables efficient and accurate support to the projects. To accomplish this work, the branch currently employs approximately fifty engineers divided into four teams -- Propellant Delivery CFD, Combustion Driven Flows CFD, Unsteady and Experimental Flows, and Acoustics and Stability. This discussion will highlight some of the work performed in the branch and the direction in which the branch is headed.

Description: There was interest in understanding the impact of out-of-round nozzle extension on the nozzle side load during transient startup operations. The out-of-round nozzle extension could be the result of asymmetric internal stresses, deformation induced by previous tests, and asymmetric loads induced by hardware attached to the nozzle. The objective of this study was therefore to computationally investigate the effect of out-of-round nozzle extension on the nozzle side loads during an engine startup transient. The rocket engine studied encompasses a regeneratively cooled chamber and nozzle, along with a film cooled nozzle extension. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and transient inlet boundary flow properties derived from an engine system simulation. Six three-dimensional cases were performed with the out-of-roundness achieved by three different degrees of ovalization, elongated on lateral y and z axes: one slightly out-of-round, one more out-of-round, and one significantly out-of-round. The results show that the separation line jump was the primary source of the peak side loads. Comparing to the peak side load of the perfectly round nozzle, the peak side loads increased for the slightly and more ovalized nozzle extensions, and either increased or decreased for the two significantly ovalized nozzle extensions. A theory based on the counteraction of the flow destabilizing effect of an exacerbated asymmetrical flow caused by a lower degree of ovalization, and the flow stabilizing effect of a more symmetrical flow, created also by ovalization, is presented to explain the observations obtained in this effort.

Description: The traditional design and analysis practice for advanced propulsion systems relies heavily on expensive full-scale prototype development and testing. Over the past decade, use of high-fidelity analysis and design tools such as CFD early in the product development cycle has been identified as one way to alleviate testing costs and to develop these devices better, faster and cheaper. In the design of advanced propulsion systems, CFD plays a major role in defining the required performance over the entire flight regime, as well as in testing the sensitivity of the design to the different modes of operation. Increased emphasis is being placed on developing and applying CFD models to simulate the flow field environments and performance of advanced propulsion systems. This necessitates the development of next generation computational tools which can be used effectively and reliably in a design environment. The turbomachinery simulation capability presented here is being developed in a computational tool called Loci-STREAM [1]. It integrates proven numerical methods for generalized grids and state-of-the-art physical models in a novel rule-based programming framework called Loci [2] which allows: (a) seamless integration of multidisciplinary physics in a unified manner, and (b) automatic handling of massively parallel computing. The objective is to be able to routinely simulate problems involving complex geometries requiring large unstructured grids and complex multidisciplinary physics. An immediate application of interest is simulation of unsteady flows in rocket turbopumps, particularly in cryogenic liquid rocket engines. The key components of the overall methodology presented in this paper are the following: (a) high fidelity unsteady simulation capability based on Detached Eddy Simulation (DES) in conjunction with second-order temporal discretization, (b) compliance with Geometric Conservation Law (GCL) in order to maintain conservative property on moving meshes for second-order time-stepping scheme, (c) a novel cloud-of-points interpolation method (based on a fast parallel kd-tree search algorithm) for interfaces between turbomachinery components in relative motion which is demonstrated to be highly scalable, and (d) demonstrated accuracy and parallel scalability on large grids (approx 250 million cells) in full turbomachinery geometries.

Description: No abstract available

Description: Detached-eddy simulation (DES) based on the v(sup 2)-f Reynolds-averaged Navier-Stokes (RANS) model is developed and tested. The v(sup 2)-f model incorporates the anisotropy of near-wall turbulence which is absent in other RANS models commonly used in the DES community. The v(sup 2)-f RANS model is modified in order the proposed v(sup 2)-f-based DES formulation reduces to a transport equation for the subgrid-scale kinetic energy isotropic turbulence. First, three coefficients in the elliptic relaxation equation are modified, which is tested in channel flows with friction Reynolds number up to 2000. Then, the proposed v(sup 2)-f DES model formulation is derived. The constant, C(sub DES), required in the DES formulation was calibrated by simulating both decaying and statistically-steady isotropic turbulence. After C(sub DES) was calibrated, the v(sub 2)-f DES formulation is tested for flow around a circular cylinder at a Reynolds number of 3900, in which case turbulence develops after separation. Simulations indicate that this model represents the turbulent wake nearly as accurately as the dynamic Smagorinsky model. Spalart-Allmaras-based DES is also included in the cylinder flow simulation for comparison.

Description: No abstract available

Description: A Presentation is given on Modeling Contact Line Dynamics in Evaporating Menisci.

Description: This study demonstrates that coupling of a material thermal response code and a flow solver using finite-rate gas/surface interaction model provides time-accurate solutions for multidimensional ablation of carbon based charring ablators. The material thermal response code used in this study is the Two-dimensional Implicit Thermal Response and Ablation Program (TITAN), which predicts charring material thermal response and shape change on hypersonic space vehicles. Its governing equations include total energy balance, pyrolysis gas momentum conservation, and a three-component decomposition model. The flow code solves the reacting Navier-Stokes equations using Data Parallel Line Relaxation (DPLR) method. Loose coupling between material response and flow codes is performed by solving the surface mass balance in DPLR and the surface energy balance in TITAN. Thus, the material surface recession is predicted by finite-rate gas/surface interaction boundary conditions implemented in DPLR, and the surface temperature and pyrolysis gas injection rate are computed in TITAN. Two sets of gas/surface interaction chemistry between air and carbon surface developed by Park and Zhluktov, respectively, are studied. Coupled fluid-material response analyses of stagnation tests conducted in NASA Ames Research Center arc-jet facilities are considered. The ablating material used in these arc-jet tests was a Phenolic Impregnated Carbon Ablator (PICA). Computational predictions of in-depth material thermal response and surface recession are compared with the experimental measurements for stagnation cold wall heat flux ranging from 107 to 1100 Watts per square centimeter.

Description: The unsteady convective flow effects in a transonic compressor rotor with a circumferential-groove casing treatment are investigated in this paper. Experimental results show that the circumferential-groove casing treatment increases the compressor stall margin by almost 50% for the current transonic compressor rotor. Steady flow simulation of the current casing treatment, however, yields only a 15% gain in stall margin. The flow field at near-stall operation is highly unsteady due to several self-induced flow phenomena. These include shock oscillation, vortex shedding at the trailing edge, and interaction between the passage shock and the tip clearance vortex. The primary focus of the current investigation is to assess the effects of flow unsteadiness and unsteady flow convection on the circumferential-groove casing treatment. Unsteady Reynolds-averaged Navier-Stokes (URANS) and Large Eddy Simulation (LES) techniques were applied in addition to steady Reynolds-averaged Navier-Stokes (RANS) to simulate the flow field at near-stall operation and to determine changes in stall margin. The current investigation reveals that unsteady flow effects are as important as steady flow effects on the performance of the circumferential grooves casing treatment in extending the stall margin of the current transonic compressor rotor. The primary unsteady flow mechanism is unsteady flow injection from the grooves into the main flow near the casing. Flows moving into and out of the grooves are caused due to local pressure difference near the grooves. As the pressure field becomes transient due to self-induced flow oscillation, flow injection from the grooves also becomes unsteady. The unsteady flow simulation shows that this unsteady flow injection from the grooves is substantial and contributes significantly to extending the compressor stall margin. Unsteady flows into and out of the grooves have as large a role as steady flows in the circumferential grooves. While the circumferential-groove casing treatment seems to be a steady flow device, unsteady flow effects should be included to accurately assess its performance as the flow is transient at near-stall operation.

Description: Turbine blades in rocket and jet engine turbomachinery experience enormous harmonic loading conditions. These loads result from the integer number of upstream and downstream stator vanes as well as the other turbine stages. The standard technique for forced response analysis to assess structural integrity is to decompose a CFD generated flow field into its harmonic components, and to then perform a frequency response analysis at the problematic natural frequencies. Recent CFD analysis and water-flow testing at NASA/MSFC, though, indicates that this technique may miss substantial harmonic and non-harmonic excitation sources that become present in complex flows. These complications suggest the question of whether frequency domain analysis is capable of capturing the excitation content sufficiently. Two studies comparing frequency response analysis with transient response analysis, therefore, have been performed. The first is of a bladed disk with each blade modeled by simple beam elements. It was hypothesized that the randomness and other variation from the standard harmonic excitation would reduce the blade structural response, but the results showed little reduction. The second study was of a realistic model of a bladed-disk excited by the same CFD used in the J2X engine program. The results showed that the transient analysis results were up to 10% higher for "clean" nodal diameter excitations and six times larger for "messy" excitations, where substantial Fourier content around the main harmonic exists.

Description: Several experiments have been performed in the Boeing/AFOSR Mach-6 Quiet Tunnel at Purdue University. A 7 degree half angle cone at 6 degree angle of attack with temperature-sensitive paint (TSP) and PCB pressure transducers was tested under quiet flow. The stationary crossflow vortices appear to break down to turbulence near the lee ray for sufficiently high Reynolds numbers. Attempts to use roughness elements to control the spacing of hot streaks on a flared cone in quiet flow did not succeed. Roughness was observed to damp the second-mode waves in areas influenced by the roughness, and wide roughness spacing allowed hot streaks to form between the roughness elements. A forward-facing cavity was used for proof-of-concept studies for a laser perturber. The lowest density at which the freestream laser perturbations could be detected was 1.07 x 10(exp -2) kilograms per cubic meter. Experiments were conducted to determine the transition characteristics of a streamwise corner flow at hypersonic velocities. Quiet flow resulted in a delayed onset of hot streak spreading. Under low Reynolds number flow hot streak spreading did not occur along the model. A new shock tube has been built at Purdue. The shock tube is designed to create weak shocks suitable for calibrating sensors, particularly PCB-132 sensors. PCB-132 measurements in another shock tube show the shock response and a linear calibration over a moderate pressure range.

Description: Hydroxyl radical (OH) planar laser-induced fluorescence (PLIF) measurements were performed in the University of Virginia s dual-mode scramjet experiment. The test section was set up in configuration A, which includes a Mach 2 nozzle, combustor, and extender section. Hydrogen fuel was injected through an unswept compression ramp at two different equivalence ratios. Through the translation of the optical system and the use of two separate camera views, the entire optical range of the combustor was accessed. Single-shot, average, and standard deviation images of the OH PLIF signal are presented at several streamwise locations. The results show the development of a highly turbulent flame structure and provide an experimental database to be used for numerical model assessment.

Description: An approximate equation was derived to predict the mass of insulation required to limit the maximum temperature reached by an insulated structure subjected to a transient heating pulse. In the course of the derivation two figures of merit were identified. One figure of merit correlates to the effectiveness of the heat capacity of the underlying structural material in reducing the amount of required insulation. The second figure of merit provides an indicator of the mass efficiency of the insulator material. An iterative, one dimensional finite element analysis was used to size the external insulation required to protect the structure at a single location on the Space Shuttle Orbiter and a reusable launch vehicle. Required insulation masses were calculated for a range of different materials for both structure and insulator. The required insulation masses calculated using the approximate equation were shown to typically agree with finite element results within 10 to 20 percent over the range of parameters studied. Finite element results closely followed the trends indicated by both figures of merit.

Description: Recent experimental supersonic retropropulsion tests were conducted at the NASA Langley Research Center Unitary Plan Wind Tunnel Test Section 2 for a range of Mach numbers from 2.4 to 4.6. A 5-inch 70-degree sphere-cone forebody model with a 10-inch cylindrical aftbody experimental model was used which is capable of multiple retrorocket configurations. These configurations include a single central nozzle on the center point of the forebody, three nozzles at the forebody half-radius, and a combination of the first two configurations with no jets being plugged. A series of measurements were achieved through various instrumentation including forebody and aftbody pressure, internal pressures and temperatures, and high speed Schlieren visualization. Specifically, several high speed pressure transducers on the forebody and in the plenum were implemented to look at unsteady flow effects. The following work focuses on analyzing frequency traits due to the unsteady flow for a range of thrust coefficients for single, tri, and quad-nozzle test cases at freestream Mach 4.6 and angle of attack ranging from -8 degrees to +20 degrees. This analysis uses Matlab s fast Fourier transform, Welch's method (modified average of a periodogram), to create a power spectral density and analyze any high speed pressure transducer frequency traits due to the unsteady flow.

Description: Coupling computational fluid dynamics (CFD) with a controls analysis tool elegantly allows for high accuracy predictions of the interaction between sloshing liquid propellants and th e control system of a launch vehicle. Instead of relying on mechanical analogs which are not valid during aU stages of flight, this method allows for a direct link between the vehicle dynamic environments calculated by the solver in the controls analysis tool to the fluid flow equations solved by the CFD code. This paper describes such a coupling methodology, presents the results of a series of test cases, and compares said results against equivalent results from extensively validated tools. The coupling methodology, described herein, has proven to be highly accurate in a variety of different cases.

Description: The Semi-Span Supersonic Transport (S4T) is an aeroelastically scaled wind-tunnel model built to test active controls concepts for large flexible supersonic aircraft in the transonic flight regime. It is one of several models constructed in the 1990's as part of the High Speed Research (HSR) Program. Control laws were developed for the S4T by M4 Engineering, Inc. and by Zona Technologies, Inc. under NASA Research Announcement (NRA) contracts. The model was tested in the NASA-Langley Transonic Dynamics Tunnel (TDT) four times from 2007 to 2010. The first two tests were primarily for plant identification. The third entry was used for testing control laws for Ride Quality Enhancement, Gust Load Alleviation, and Flutter Suppression. Whereas the third entry only tested FS subcritically, the fourth test demonstrated closed-loop operation above the open-loop flutter boundary. The results of the third entry are reported elsewhere. This paper reports on flutter suppression results from the fourth wind-tunnel test. Flutter suppression is seen as a way to provide stability margins while flying at transonic flight conditions without penalizing the primary supersonic cruise design condition. An account is given for how Controller Performance Evaluation (CPE) singular value plots were interpreted with regard to progressing open- or closed-loop to higher dynamic pressures during testing.

Description: Rapid reduced-order numerical models are being investigated as candidates to simulate the dynamics of a flexible launch vehicle during atmospheric ascent. There has also been the extension of these new approaches to include gust response. These methods are used to perform aeroelastic and gust response analyses at isolated Mach numbers. Such models require a method to time march through a succession of ascent Mach numbers. An approach is presented for interpolating reduced-order models of the unsteady aerodynamics at successive Mach numbers. The transonic Mach number range is considered here since launch vehicles can suffer the highest dynamic loads through this range. Realistic simulations of the flexible vehicle behavior as it traverses this Mach number range are presented. The response of the vehicle due to gusts is computed. Uncertainties in root mean square and maximum bending moment and crew module accelerations are presented due to assumed probability distributions in design parameters, ascent flight conditions, gusts. The primary focus is on the uncertainty introduced by modeling fidelity. It is found that an unsteady reduced order model produces larger excursions in the root mean square loading and accelerations than does a quasi-steady reduced order model.

Description: In September 1995, a project was initiated to replace the existing drive line at NASA's most unique transonic wind tunnel, the National Transonic Facility (NTF), with a single 101 MW synchronous motor driven by a Load Commutated Inverter (LCI). This Adjustable Speed Drive (ASD) system also included a custom four-winding transformer, harmonic filter, exciter, switch gear, control system, and feeder cable. The complete system requirements and design details have previously been presented and published [1], as well as the commissioning and acceptance test results [2]. The NTF was returned to service in December 1997 with the new drive system powering the fan. Today, this installation still represents the world s largest horizontal single motor/drive combination. This paper describes some significant events that occurred with the drive system during the first 15 years of service. These noteworthy issues are analyzed and root causes presented. Improvements that have substantially increased the long term viability of the system are given.

Description: A capillary flow liquid acquisition device (LAD) for cryogenic propellants has been developed and tested in NASA Glenn Research Center to meet the requirements of transferring cryogenic liquid propellants from storage tanks to an engine in reduced gravity environments. The prototypical mesh screen channel LAD was fabricated with a mesh screen, covering a rectangular flow channel with a cylindrical outlet tube, and was tested with liquid oxygen (LOX). In order to better understand the performance in various gravity environments and orientations at different liquid submersion depths of the screen channel LAD, a series of computational fluid dynamics (CFD) simulations of LOX flow through the LAD screen channel was undertaken. The resulting velocity vector field visualization for the flow in the channel has been used to reveal the gravity effects on the flow in the screen channel.

Description: High resolution infrared observations made from a mobile ground based optical system captured the laminar-to-turbulent boundary layer transition process as it occurred during Space Shuttle Endeavour's return to earth following its final mission in 2011. The STS-134 imagery was part of a larger effort to demonstrate an emerging and reliable non-intrusive global thermal measurement capability and to complement a series of boundary layer transition flight experiments that were flown on the Shuttle. The STS-134 observations are believed to be the first time that the development and movement of a hypersonic boundary layer transition front has been witnessed in flight over the entire vehicle surface and in particular, at unprecedented spatial resolution. Additionally, benchmark surface temperature maps of the Orbiter lower surface collected over multiple flights and spanning a Mach range of 18 to 6 are now available and represent an opportunity for collaborative comparison with computational techniques focused on hypersonic transition and turbulence modeling. The synergy of the global temperature maps with the companion in-situ thermocouple measurements serve as an example of the effective leveraging of resources to achieve a common goal of advancing our understanding of the complex nature of high Mach number transition. It is shown that quantitative imaging can open the door to a multitude of national and international opportunities for partnership associated with flight-testing and subsequent validation of numerical simulation techniques. The quantitative imaging applications highlighted in this paper offer unique and complementary flight measurement alternatives and suggest collaborative instrumentation opportunities to advance the state of the art in transition prediction and maximize the return on investment in terms of developmental flight tests for future vehicle designs.

Description: The generation, propagation and radiation of sound from a perfectly expanded Mach 2.5 cold supersonic jet flowing through an enclosed rigid-walled duct with an upstream J-deflector have been numerically simulated with the aid of OVERFLOW Navier-Stokes CFD code. A one-equation turbulence model is considered. While the near-field sound sources are computed by the CFD code, the far-field sound is evaluated by Kirchhoff surface integral formulation. Predictions of the farfield directivity of the OASPL (Overall Sound Pressure Level) agree satisfactorily with the experimental data previously reported by the author. Calculations also suggest that there is significant entrainment of air into the duct, with the mass flow rate of entrained air being about three times the jet exit mass flow rate.

Description: While externally mixed, or separate flow, nozzle systems are most common in high bypass-ratio aircraft, they are not as attractive for use in lower bypass-ratio systems and on aircraft that will fly supersonically. The noise of such propulsion systems is also dominated by jet noise, making the study and noise reduction of these exhaust systems very important, both for military aircraft and future civilian supersonic aircraft. This paper presents particle image velocimetry of internally mixed nozzle with different area ratios between core and bypass, and nozzles that are ideally expanded and convergent. Such configurations independently control the geometry of the internal mixing layer and of the external shock structure. These allow exploration of the impact of shocks on the turbulent mixing layers, the impact of bypass ratio on broadband shock noise and mixing noise, and the impact of temperature on the turbulent flow field. At the 2009 AIAA/CEAS Aeroacoustics Conference the authors presented data and analysis from a series of tests that looked at the acoustics of supersonic jets from internally mixed nozzles. In that paper the broadband shock and mixing noise components of the jet noise were independently manipulated by holding Mach number constant while varying bypass ratio and jet temperature. Significant portions of that analysis was predicated on assumptions regarding the flow fields of these jets, both shock structure and turbulence. In this paper we add to that analysis by presenting particle image velocimetry measurements of the flow fields of many of those jets. In addition, the turbulent velocity data documented here will be very useful for validation of computational flow codes that are being developed to design advanced nozzles for future aircraft.

Description: Numerical simulations of gas-seeding strategies required for planar laser-induced fluorescence (PLIF) in a Mach 10 air flow were performed. The work was performed to understand and quantify adverse effects associated with gas seeding and to compare different flow rates and different types of seed gas. The gas was injected through a slot near the leading edge of a flat plate wedge model used in NASA Langley Research Center's 31- Inch Mach 10 Air Tunnel facility. Nitric oxide, krypton, and iodine gases were simulated at various injection rates. Simulation results showing the deflection of the velocity field for each of the cases are presented. Streamwise distributions of velocity and concentration boundary layer thicknesses as well as vertical distributions of velocity, temperature, and mass distributions are presented for each of the cases. Relative merits of the different seeding strategies are discussed.

Description: No abstract available

Description: JPL has captured its experience from over four decades of robotic space exploration into a set of design rules. These rules have gradually changed into explicit requirements and are now formally implemented and verified. Over an extended period of time, the initial understanding of intent and rationale for these rules has faded and rules are now frequently applied without further consideration. In the meantime, mission classes and their associated risk postures have evolved, coupled with resource constraints and growing design diversity, bringing into question the current "one size fits all" thermal margin approach. This paper offers a systematic review of the heat flow path from an electronic junction to the eventual heat rejection to space. This includes the identification of different regimes along this path and the associated requirements. The work resulted in a renewed understanding of the intent behind JPL requirements for hot thermal margins and a framework for relevant considerations, which in turn enables better decision making when a deviation to these requirements is considered.

Description: This paper details the results of an experimental investigation into the cavitation instabilities created by a circular orifice conducted at the University of Alabama in Huntsville Propulsion Research Center. This experiment was conducted in concert with a computational simulation to serve as a reference point for the simulation. Testing was conducted using liquid nitrogen as a cryogenic propellant simulant. A 1.06 cm diameter thin orifice with a rounded inlet was tested in an approximately 1.25 kg/s flow with inlet pressures ranging from 504.1 kPa to 829.3 kPa. Pressure fluctuations generated by the orifice were measured using a high frequency pressure sensor located 0.64 tube diameters downstream of the orifice. Fast Fourier Transforms were performed on the high frequency data to determine the instability frequency. Shedding resulted in a primary frequency with a cavitation related subharmonic frequency. For this experiment, the cavitation instability ranged from 153 Hz to 275 Hz. Additionally, the strength of the cavitation occur red as a function of cavitation number. At lower cavitation numbers, the strength of the cavitation instability ranged from 2.4 % to 7 % of the inlet pressure. However, at higher cavitation numbers, the strength of the cavitation instability ranged from 0.6 % to 1 % of the inlet pressure.

Description: Plasma diagnostic measurement campaigns in the NASA Ames Interaction Heating Facility (IHF) have been conducted over the last several years with a view towards characterizing the flow in the arc jet facility by providing data necessary for modeling and simulation. Optical emission spectroscopy has been used in the plenum and in the free jet of the nozzle. Radiation incident over a probe surface has also been measured using radiometry. Plenum measurements have shown distinct radial profiles of temperature over a range of operating conditions. For cases where large amounts of cold air are added radially to the main arc-heated stream, the temperature profiles are higher by as much as 1500 K than the profiles assumed in flow simulations. Optical measurements perpendicular to the flow direction in the free jet showed significant contributions to the molecule emission through inverse pre-dissociation, thus allowing determination of atom number densities from molecular emission. This has been preliminarily demonstrated with the N2 1st Positive System. Despite the use of older rate coefficients, the resulting atom densities are reasonable and surprisingly close to flow predictions.

Description: Forced periodic flows arise in a broad range of aerodynamic applications such as rotorcraft, turbomachinery, and flapping wing configurations. Standard practice involves solving the unsteady flow equations forward in time until the initial transient exits the domain and a statistically stationary flow is achieved. It is often required to simulate through several periods to remove the initial transient making unsteady design optimization prohibitively expensive for most realistic problems. An effort to reduce the computational cost of these calculations led to the development of the Harmonic Balance method [1, 2] which capitalizes on the periodic nature of the solution. The approach exploits the fact that forced temporally periodic flow, while varying in the time domain, is invariant in the frequency domain. Expanding the temporal variation at each spatial node into a Fourier series transforms the unsteady governing equations into a steady set of equations in integer harmonics that can be tackled with the acceleration techniques afforded to steady-state flow solvers. Other similar approaches, such as the Nonlinear Frequency Domain [3,4,5], Reduced Frequency [6] and Time-Spectral [7, 8, 9] methods, were developed shortly thereafter. Additionally, adjoint-based optimization techniques can be applied [10, 11] as well as frequency-adaptive methods [12, 13, 14] to provide even more flexibility to the method. The Fourier temporal basis functions imply spectral convergence as the number of harmonic modes, and correspondingly number of time samples, N, is increased. Some elect to solve the equations in the frequency domain directly, while others choose to transform the equations back into the time domain to simplify the process of adding this capability to existing solvers, but each harnesses the underlying steady solution in the frequency domain. These temporal projection methods will herein be collectively referred to as Time-Spectral methods. Time-Spectral methods have demonstrated marked success in reducing the computational costs associated with simulating periodic forced flows, but have yet to be fully applied to overset or Cartesian solvers for arbitrary motion with dynamic hole-cutting. Overset and Cartesian grid methodologies are versatile techniques capable of handling complex geometry configurations in practical engineering applications, and the combination of the Time-Spectral approach with this general capability potentially provides an enabling new design and analysis tool. In an arbitrary moving-body scenario for these approaches, a Lagrangian body moves through a fixed Eulerian mesh and mesh points in the Eulerian mesh interior to the solid body are removed (cut or blanked), leaving a hole in the Eulerian mesh. During the dynamic motion some gridpoints in the domain are blanked and do not have a complete set of time-samples preventing a direct implementation of the Time-Spectral method. Murman[6] demonstrated the Time-Spectral approach for a Cartesian solver with a rigid domain motion, wherein the hole cutting remains constant. Similarly, Custer et al. [15, 16] used the NASA overset OVERFLOW solver and limited the amount of relative motion to ensure static hole-cutting and interpolation. Recently, Mavriplis and Mundis[17] demonstrated a qualitative method for applying the Time-Spectral approach to an unstructured overset solver for arbitrary motion. The goal of the current work is to develop a robust and general method for handling arbitrary motion with the Time-Spectral approach within an overset or Cartesian mesh method, while still approaching the spectral convergence rate of the original Time-Spectral approach. The viscous OVERFLOW solver will be augmented with the new Time-Spectral algorithm and the capability of the method for benchmark problems in rotorcraft and turbomachinery will be demonstrated. This abstract begins with a brief synopsis of the Time-Spectral approach for overset grids and provides details of e current approach to allow for arbitrary motion. Model problem results in one and two dimensions are included to demonstrate the viability of the method and the convergence properties. Section IV briefly outlines the implementation into the OVERFLOW solver, and the abstract closes with a description of the benchmark test cases which will be included in the final paper.

Description: Advances in automation and robustness of the X-rays approach to domain connectivity for overset grids are presented. Given the surface definition for each component that makes up a complex configuration, the determination of hole points with appropriate hole boundaries is automatically and efficiently performed. Improvements made to the original X-rays approach for identifying the minimum hole include an automated closure scheme for hole-cutters with open boundaries, automatic determination of grid points to be considered for blanking by each hole-cutter, and an adaptive X-ray map to economically handle components in close proximity. Furthermore, an automated spatially varying offset of the hole boundary from the minimum hole is achieved using a dual wall-distance function and an orphan point removal iteration process. Results using the new scheme are presented for a number of static and relative motion test cases on a variety of aerospace applications.

Description: This paper describes a collaborative and cost-shared approach to reducing fuel burn under the NASA Environmentally Responsible Aviation project. NASA and General Electric (GE) Aviation are working together aa an integrated team to obtain compressor aerodynamic data that is mutually beneficial to both NASA and GE Aviation. The objective of the High OPR Compressor Task is to test a single stage then two stages of an advanced GE core compressor using state-of-the-art research instrumentation to investigate the loss mechanisms and interaction effects of embedded transonic highly-loaded compressor stages. This paper presents preliminary results from NASA's in-house multistage computational code, APNASA, in preparation for this advanced transonic compressor rig test.

Description: Three high order shock-capturing schemes are compared for large eddy simulations (LES) of temporally evolving mixing layers (TML) for different convective Mach numbers (Mc) ranging from the quasi-incompressible regime to highly compressible supersonic regime. The considered high order schemes are fifth-order WENO (WENO5), seventh-order WENO (WENO7) and the associated eighth-order central spatial base scheme with the dissipative portion of WENO7 as a nonlinear post-processing filter step (WENO7fi). This high order nonlinear filter method (H.C. Yee and B. Sjogreen, Proceedings of ICOSAHOM09, June 22-26, 2009, Trondheim, Norway) is designed for accurate and efficient simulations of shock-free compressible turbulence, turbulence with shocklets and turbulence with strong shocks with minimum tuning of scheme parameters. The LES results by WENO7fi using the same scheme parameter agree well with experimental results of Barone et al. (2006), and published direct numerical simulations (DNS) work of Rogers & Moser (1994) and Pantano & Sarkar (2002), whereas results by WENO5 and WENO7 compare poorly with experimental data and DNS computations.

Description: Engineers working to understand and reduce cryogenic boil-off must solve a variety of transport problems. An important class of nonlinear problems involves the thermal and mechanical design of cryogenic struts. These classic problems are scattered about the literature and typically require too many resources to obtain. So, to save time for practicing engineers, the author presents this essay. Herein, a variety of new, old, and revisited analytical and finite difference solutions of the thermal problem are covered in this essay, along with commentary on approach and assumptions. This includes a few thermal radiation and conduction combined mode solutions with a discussion on insulation, optimum emissivity, and geometrical phenomenon. Solutions to cooling and heat interception problems are also presented, including a discussion of the entropy generation. The literature on the combined mechanical and thermal design of cryogenic support struts is reviewed with an introduction to the associated numerical methods.

Description: No abstract available

Description: In compressible turbulent combustion/nonequilibrium flows, the constructions of numerical schemes for (a) stable and accurate simulation of turbulence with strong shocks, and (b) obtaining correct propagation speed of discontinuities for stiff reacting terms on coarse grids share one important ingredient - minimization of numerical dissipation while maintaining numerical stability. Here coarse grids means standard mesh density requirement for accurate simulation of typical non-reacting flows. This dual requirement to achieve both numerical stability and accuracy with zero or minimal use of numerical dissipation is most often conflicting for existing schemes that were designed for non-reacting flows. The goal of this paper is to relate numerical dissipations that are inherited in a selected set of high order shock-capturing schemes with the onset of wrong propagation speed of discontinuities as a function of stiffness of the source term and the grid spacing.

Description: The present research deals with thermal degradation of polyurethane foam (PUF) during flight test. Model of thermal decomposition was developed that accounts for polyurethane kinetics parameters extracted from thermogravimetric analyses and radial heat losses to the surrounding environment. The model predicts mass loss of foam, the temperature and kinetic of release of the exhaust gases and char as function of heat and radiation loads. When PUF is heated, urethane bond break into polyol and isocyanate. In the first stage, isocyanate pyrolyses and oxidizes. As a result, the thermo-char and oil droplets (yellow smoke) are released. In the second decomposition stage, pyrolysis and oxidization of liquid polyol occur. Next, the kinetics of chemical compound release and the information about the reactions occurring in the base area are coupled to the CFD simulations of the base flow in a single first stage motor vertically stacked vehicle configuration. The CFD simulations are performed to estimate the contribution of the hot out-gassing, chemical reactions, and char oxidation to the temperature rise of the base flow. The results of simulations are compared with the flight test data.

Description: Mass Estimating Relationships (MERs) have been developed for use in the Program to Optimize Simulated Trajectories II (POST2) as part of NASA's multi-mission Earth Entry Vehicle (MMEEV) concept. MERs have been developed for the thermal protection systems of PICA and of Carbon Phenolic atop Advanced Carbon-Carbon on the forebody and for SIRCA and Acusil II on the backshell. How these MERs were developed, the resulting equations, model limitations, and model accuracy are discussed herein.

Description: A study on the effect of inflow turbulence boundary conditions on the local flow on and around a body in flight has been carried out. The study has been carried out using OVERFLOW2 flow solver using the default Spalart-Allmaras turbulence model in OVERFLOW. Many OVERFLOW turbulent flow simulations have been reported using the SA-fv3 model1 over the years. The present study demonstrates that the turbulence levels imposed as a boundary condition (b.c.) in the far-field as implemented in OVERFLOW are not correct. In fact, very low level of turbulence at the far-field boundary as implemented in OVERFLOW results in the SA-fv3 model predicting transition-like profiles on a given body. By choosing sufficiently high levels of the Reynolds number of turbulence, Rt, as a boundary condition, this anomalous behavior of SA-fv3 model is eliminated. Since numerous papers using OVERFLOW have been presented in the literature including that by the author2 using low level of inflow turbulence, it will be beneficial to the CFD community at large and in particular to the OVERFLOW community to understand this effect of the inflow b.c. in the SA-fv3 model. Various results reported over the years in domains such as high-lift applications, drag prediction applications and rotorcraft flow applications can be revisited using the new turbulence inflow boundary condition prescription as suggested in this study. It has been demonstrated in the literature that the effect of inflow turbulence level on the downstream development of flow is too important to be ignored. Demonstrative results are shown and compared with experiment for the case of hovering XV-15 rotor flow at a tip Reynolds number of 4.9 x 10(exp 6) and Mach number of 0.69 in Figs. 1 and 2 below. Fig. 1(a,b) shows skin friction predictions at radial station of r/R = 0.28 corresponding to 5 different inflow boundary conditions obtained with the SA-fv3 model. As the inflow turbulence is increased monotonically, from a value of Rt = 0.1 to 100.0 as shown in Fig. 1, a fully developed turbulent skin friction profile is realized progressively as shown in Fig. 1(b). Similarly skin friction results shown in Fig. 2(a,b) corresponding to radial stations of r/R = 0.72 reinforce the postulate made here that as the inflow turbulence level is increased gradually, fully turbulent on-the-blade simulation is realized uniformly. In the full paper extensive results on skin friction, pressure and velocity profiles as well as the Figure of Merit will be presented and compared with experiment that will illustrate the substantial effect of the inflow boundary conditions on the development of hovering rotor flow downstream.

Description: Propellant slosh is a potential source of disturbance critical to the stability of space vehicles. The slosh dynamics are typically represented by a mechanical model of a spring mass damper. This mechanical model is then included in the equation of motion of the entire vehicle for Guidance, Navigation and Control analysis. Our previous effort has demonstrated the soundness of a CFD approach in modeling the detailed fluid dynamics of tank slosh and the excellent accuracy in extracting mechanical properties (slosh natural frequency, slosh mass, and slosh mass center coordinates). For a practical partially-filled smooth wall propellant tank with a diameter of 1 meter, the damping ratio is as low as 0.0005 (or 0.05%). To accurately predict this very low damping value is a challenge for any CFD tool, as one must resolve a thin boundary layer near the wall and must minimize numerical damping. This work extends our previous effort to extract this challenging parameter from first principles: slosh damping for smooth wall and for ring baffle. First the experimental data correlated into the industry standard for smooth wall were used as the baseline validation. It is demonstrated that with proper grid resolution, CFD can indeed accurately predict low damping values from smooth walls for different tank sizes. The damping due to ring baffles at different depths from the free surface and for different sizes of tank was then simulated, and fairly good agreement with experimental correlation was observed. The study demonstrates that CFD technology can be applied to the design of future propellant tanks with complex configurations and with smooth walls or multiple baffles, where previous experimental data is not available.

Description: The Multi-Mission Earth Entry Vehicle project is developing an integrated tool called Multi Mission System Analysis for Planetary Entry Descent and Landing that will provide key technology solutions including mass sizing, aerodynamics, aerothermodynamics, and thermal and structural analysis for any given sample return mission. Thermal soak analysis and temperature predictions of various components including the payload container of the entry vehicle are part of the solution that this tool will offer to mission designers. The present paper focuses on the thermal soak analysis of an entry vehicle design based on the Mars Sample Return entry vehicle geometry and discusses a technical approach to develop parametric models for thermal soak analysis that will be integrated into the tool.

Description: No abstract available

Description: The presentation describes supersonic flight testing accomplished on a novel mixed-compression axisymmetric inlet utilizing channels for off-design flow matching rather than a translating centerbody concept.

Description: A high-quality model validation experiment was performed in the NASA Langley Research Center Unitary Plan Wind Tunnel to assess the predictive accuracy of computational fluid dynamics (CFD) models for a blunt-body supersonic retro-propulsion configuration at Mach numbers from 2.4 to 4.6. Static and fluctuating surface pressure data were acquired on a 5-inch-diameter test article with a forebody composed of a spherically-blunted, 70-degree half-angle cone and a cylindrical aft body. One non-powered configuration with a smooth outer mold line was tested as well as three different powered, forward-firing nozzle configurations: a centerline nozzle, three nozzles equally spaced around the forebody, and a combination with all four nozzles. A key objective of the experiment was the determination of experimental uncertainties from a range of sources such as random measurement error, flowfield non-uniformity, and model/instrumentation asymmetries. This paper discusses the design of the experiment towards capturing these uncertainties for the baseline non-powered configuration, the methodology utilized in quantifying the various sources of uncertainty, and examples of the uncertainties applied to non-powered and powered experimental results. The analysis showed that flowfield nonuniformity was the dominant contributor to the overall uncertainty a finding in agreement with other experiments that have quantified various sources of uncertainty.

Description: Active flow control technology is finding increasing use in aerospace applications to control flow separation and improve aerodynamic performance. In this paper we examine the characteristics of a class of fluidic actuators that are being considered for active flow control applications for a variety of practical problems. Based on recent experimental work, such actuators have been found to be more efficient for controlling flow separation in terms of mass flow requirements compared to constant blowing and suction or even synthetic jet actuators. The fluidic actuators produce spanwise oscillating jets, and therefore are also known as sweeping jets. The frequency and spanwise sweeping extent depend on the geometric parameters and mass flow rate entering the actuators through the inlet section. The flow physics associated with these actuators is quite complex and not fully understood at this time. The unsteady flow generated by such actuators is simulated using the lattice Boltzmann based solver PowerFLOW R . Computed mean and standard deviation of velocity profiles generated by a family of fluidic actuators in quiescent air are compared with experimental data. Simulated results replicate the experimentally observed trends with parametric variation of geometry and inflow conditions.

Description: The HIFiRE-1 flight experiment provided a valuable database pertaining to boundary layer transition over a 7-degree half-angle, circular cone model from supersonic to hypersonic Mach numbers, and a range of Reynolds numbers and angles of attack. This paper reports selected findings from the ongoing computational analysis of the measured in-flight transition behavior. Transition during the ascent phase at nearly zero degree angle of attack is dominated by second mode instabilities except in the vicinity of the cone meridian where a roughness element was placed midway along the length of the cone. The growth of first mode instabilities is found to be weak at all trajectory points analyzed from the ascent phase. For times less than approximately 18.5 seconds into the flight, the peak amplification ratio for second mode disturbances is sufficiently small because of the lower Mach numbers at earlier times, so that the transition behavior inferred from the measurements is attributed to an unknown physical mechanism, potentially related to step discontinuities in surface height near the locations of a change in the surface material. Based on the time histories of temperature and/or heat flux at transducer locations within the aft portion of the cone, the onset of transition correlated with a linear N-factor, based on parabolized stability equations, of approximately 13.5. Due to the large angles of attack during the re-entry phase, crossflow instability may play a significant role in transition. Computations also indicate the presence of pronounced crossflow separation over a significant portion of the trajectory segment that is relevant to transition analysis. The transition behavior during this re-entry segment of HIFiRE-1 flight shares some common features with the predicted transition front along the elliptic cone shaped HIFiRE-5 flight article, which was designed to provide hypersonic transition data for a fully 3D geometric configuration. To compare and contrast the crossflow dominated transition over the HIFiRE-1 and HIFiRE-5 configurations, this paper also analyzes boundary layer instabilities over a subscale model of the HIFiRE-5 flight configuration that was tested in the Mach 6 quiet tunnel facility at Purdue University.

Description: Laminar stagnation region heating augmentation is investigated in the AEDC Tunnel 9 at Mach 10 by performing high frequency surface pressure and heat transfer measurements on the Orion CEV capsule at zero degree angle-of-attack for unit Reynolds numbers between 0.5 and 15 million per foot. Heating augmentation increases with Reynolds number, but is also model size dependent as it is absent on a 1.25-inch diameter model at Reynolds numbers where it reaches up to 15% on a 7-inch model. Heat transfer space-time correlations on the 7-inch model show that disturbances convect at the boundary layer edge velocity and that the streamwise integral scale increases with distance. Therefore, vorticity amplification due to stretching and piling-up in the stagnation region appears to be responsible for the stagnation point heating augmentation on the larger model. This assumption is reinforced by the f(exp -11/3) dependence of the surface pressure spectrum compared to the f(exp -1) dependence in the free stream. Vorticity amplification does not occur on the 1.25- inch model because the disturbances are too large. Improved free stream fluctuation measurements will be required to determine if significant vorticity is present upstream or mostly generated behind the bow shock.

Description: This paper describes a finite volume based numerical algorithm that allows multi-dimensional computation of fluid flow within a system level network flow analysis. There are several thermo-fluid engineering problems where higher fidelity solutions are needed that are not within the capacity of system level codes. The proposed algorithm will allow NASA's Generalized Fluid System Simulation Program (GFSSP) to perform multi-dimensional flow calculation within the framework of GFSSP s typical system level flow network consisting of fluid nodes and branches. The paper presents several classical two-dimensional fluid dynamics problems that have been solved by GFSSP's multi-dimensional flow solver. The numerical solutions are compared with the analytical and benchmark solution of Poiseulle, Couette and flow in a driven cavity.

Description: Background oriented schlieren (BOS) is a recent development of the schlieren and shadowgraph methods. The BOS technique has the ability to provide visualizations of the density gradient in both the axial and radial directions. The resultant magnitude of the density gradients allows for comparison with shadowgraph images. This paper first compares data obtained by the BOS and shadowgraph techniques at identical conditions in a free jet. The patterns and spacing of the shock trains obtained by the two techniques are found to be consistent with one another. This provides confidence in the shock spacing measurement by the BOS technique. Due to its simpler setup, BOS is then applied to investigate the shock spacing associated with the screech phenomenon, especially during stage jumps. Screech frequencies from a 37.6 mm convergent nozzle, as a function of jet Mach number (M(sub j)), are shown to exhibit various stages. As many as eight stages are identified with the present nozzle over the range 1.0 〈 M(sub j) 〈1.7. BOS images are acquired at various screech conditions and the shock spacing is examined as a function of M(sub j).

Description: During developmental testing of the Ascent Abort 1 (AA-1) cold gas thruster system, unexpected behavior was detected. Upon further review the design as it existed may not have met the requirements. To determine the best approach for modifying the design, the system was modeled with a dynamic fluid analysis tool (EASY5). The system model consisted of the nitrogen storage tank, pressure regulator, thruster valve, nozzle, and the associated interconnecting line lengths. The regulator and thruster valves were modeled using a combination of the fluid and mechanical modules available in EASY5. The simulation results were then compared against actual system test data. The simulation results exhibited behaviors similar to the test results, such as the pressure regulators response to thruster firings. Potential design solutions were investigated using the analytical model parameters, including increasing the volume downstream of the regulator and increasing the orifice area. Both were shown to improve the regulator response.

Description: A series of experimental tests, using both qualitative and quantitative techniques, were conducted to characterize both surface and off-surface flow characteristics of an axisymmetric, modified-cosine-shaped, wall-mounted hill named "FAITH" (Fundamental Aero Investigates The Hill). Two separate models were employed: a 6" high, 18" base diameter machined aluminum model that was used for wind tunnel tests and a smaller scale (2" high, 6" base diameter) sintered nylon version that was used in the water channel facility. Wind tunnel and water channel tests were conducted at mean test section speeds of 165 fps (Reynolds Number based on height = 500,000) and 0.1 fps (Reynolds Number of 1000), respectively. The ratio of model height to boundary later height was approximately 3 for both tests. Qualitative techniques that were employed to characterize the complex flow included surface oil flow visualization for the wind tunnel tests, and dye injection for the water channel tests. Quantitative techniques that were employed to characterize the flow included Cobra Probe to determine point-wise steady and unsteady 3D velocities, Particle Image Velocimetry (PIV) to determine 3D velocities and turbulence statistics along specified planes, Pressure Sensitive Paint (PSP) to determine mean surface pressures, and Fringe Imaging Skin Friction (FISF) to determine surface skin friction (magnitude and direction). This initial report summarizes the experimental set-up, techniques used, data acquired and describes some details of the dataset that is being constructed for use by other researchers, especially the CFD community. Subsequent reports will discuss the data and their interpretation in more detail

Description: The primary objective of Observation and Analysis of Smectic Islands in Space (OASIS) experiment is to exploit the unique characteristics of freely suspended liquid crystals in a microgravity environment to advance the understanding of fluid state physics

Description: This DVD collection includes the complete proceedings of Minnowbrook Workshops I through VI. Titles include Minnowbrook I - 1993 Workshop on End-Stage Boundary Layer Transition (NASA/CP-2007-214667, CASI ID 20070038942), Minnowbrook II - 1997 Workshop on Boundary Layer Transition in Turbomachines (NASA/CP-1998-206958, CASI ID 19980206205), Minnowbrook III - 2000 Workshop on Boundary Layer Transition and Unsteady Aspects of Turbomachinery Flow (NASA/CP-2001-210888, CASI ID 20020067662), Minnowbrook IV - 2003 Workshop on Transition and Unsteady Aspects of Turbomachinery Flows (NASA TM-2004-212913, CASI ID 20040121174), Minnowbrook V - 2006 Workshop on Unsteady Flows in Turbomachinery (NASA/CP-2006-214484, CASI ID 20070024781), and Minnowbrook VI - 2009 Workshop on Flow Physics and Control for Internal and External Aerodynamics (NACA/CP-2010-216112, CASI ID 20100018557).

Description: During the five years of this project, the AMTOC team developed an adjoint-based methodology for design and optimization of complex time-dependent flows, implemented AMTOC in a testbed environment, directly assisted in implementation of this methodology in the state-of-the-art NASA's unstructured CFD code FUN3D, and successfully demonstrated applications of this methodology to large-scale optimization of several supersonic and other aerodynamic systems, such as fighter jet, subsonic aircraft, rotorcraft, high-lift, wind-turbine, and flapping-wing configurations. In the course of this project, the AMTOC team has published 13 refereed journal articles, 21 refereed conference papers, and 2 NIA reports. The AMTOC team presented the results of this research at 36 international and national conferences, meeting and seminars, including International Conference on CFD, and numerous AIAA conferences and meetings. Selected publications that include the major results of the AMTOC project are enclosed in this report.

Description: The Fluid Dynamics Branch's (ER42) at MSFC mission is to support NASA and other customers with discipline expertise to enable successful accomplishment of program/project goals. The branch is responsible for all aspects of the discipline of fluid dynamics, analysis and testing, applied to propulsion or propulsion-induced loads and environments, which includes the propellant delivery system, combustion devices, coupled systems, and launch and separation events. ER42 supports projects from design through development, and into anomaly and failure investigations. ER42 is committed to continually improving the state-of-its-practice to provide accurate, effective, and timely fluid dynamics assessments and in extending the state-of-the-art of the discipline.

Description: The Tropospheric Emission Spectrometer (TES) instrument heat rejection system has been operating in space for nearly 8 years since launched on NASA's EOS Aura Spacecraft. The instrument is an infrared imaging fourier transform spectrometer with spectral coverage of 3.2 to 15.4 microns. The loop heat pipe (LHP) based heat rejection system manages all of the instrument components waste heat including the two mechanical cryocoolers and their drive electronics. Five propylene LHPs collect and transport the instrument waste heat to the near room temperature nadir viewing radiators. During the early months of the mission, ice contamination of the cryogenic surfaces including the focal planes led to increased cryocooler loads and the need for periodic decontamination cycles. Focal plane decontamination cycles require power cycling both cryocoolers which also requires the two cryocooler LHPs to turn off and on during each cycle. To date, the cryocooler LHPs have undergone 24 start-ups in orbit successfully. This paper reports on the TES cryocooler loop heat pipe based heat rejection system performance. After a brief overview of the instrument thermal design, the paper presents detailed data on the highly successful space operation of the loop heat pipes since instrument turn-on in 2004. The data shows that the steady-state and transient operation of the LHPs has not changed since 2004 and shows consistent and predictable performance. The LHP based heat rejection system has provided a nearly constant heat rejection heat sink for all of its equipment which has led to exceptional overall instrument performance with world class science.

Description: ven with great advances in computational techniques and computing power during recent decades, the modeling of unsteady separated flows, such as those encountered in the wake of a re-entry vehicle, continues to be one of the most challenging problems in CFD. Of most interest to the aerothermodynamics community is accurately predicting transient heating loads on the base of a blunt body, which would result in reduced uncertainties and safety margins when designing a re-entry vehicle. However, the prediction of heat transfer can vary widely depending on the turbulence model employed. Therefore, selecting a turbulence model which realistically captures as much of the flow physics as possible will result in improved results. Reynolds Averaged Navier Stokes (RANS) models have become increasingly popular due to their good performance with attached flows, and the relatively quick turnaround time to obtain results. However, RANS methods cannot accurately simulate unsteady separated wake flows, and running direct numerical simulation (DNS) on such complex flows is currently too computationally expensive. Large Eddy Simulation (LES) techniques allow for the computation of the large eddies, which contain most of the Reynolds stress, while modeling the smaller (subgrid) eddies. This results in models which are more computationally expensive than RANS methods, but not as prohibitive as DNS. By complimenting an LES approach with a RANS model, a hybrid LES/RANS method resolves the larger turbulent scales away from surfaces with LES, and switches to a RANS model inside boundary layers. As pointed out by Bertin et al., this type of hybrid approach has shown a lot of promise for predicting turbulent flows, but work is needed to verify that these models work well in hypersonic flows. The very limited amounts of flight and experimental data available presents an additional challenge for researchers. Recently, a joint study by NASA and CUBRC has focused on collecting heat transfer data on the backshell of a scaled model of the Orion Multi-Purpose Crew Vehicle (MPCV). Heat augmentation effects due to the presence of cavities and RCS jet firings were also investigated. The high quality data produced by this effort presents a new set of data which can be used to assess the performance of CFD methods. In this work, a hybrid LES/RANS model developed at North Carolina State University (NCSU) is used to simulate several runs from these experiments, and evaluate the performance of high fidelity methods as compared to more typical RANS models. .

Description: Airframe noise corresponds to the acoustic radiation due to turbulent flow in the vicinity of airframe components such as high-lift devices and landing gears. The combination of geometric complexity, high Reynolds number turbulence, multiple regions of separation, and a strong coupling with adjacent physical components makes the problem of airframe noise highly challenging. Since 2010, the American Institute of Aeronautics and Astronautics has organized an ongoing series of workshops devoted to Benchmark Problems for Airframe Noise Computations (BANC). The BANC workshops are aimed at enabling a systematic progress in the understanding and high-fidelity predictions of airframe noise via collaborative investigations that integrate state of the art computational fluid dynamics, computational aeroacoustics, and in depth, holistic, and multifacility measurements targeting a selected set of canonical yet realistic configurations. This paper provides a brief summary of the BANC effort, including its technical objectives, strategy, and selective outcomes thus far.

Description: The goal is to softly land high mass vehicles (10s of metric tons) on Mars. Supersonic Retropropulsion (SRP) is a potential method of deceleration. Current method of supersonic parachutes does not scale past ~1 metric ton. CFD is of increasing importance since flight and experimental data at these conditions is difficult to obtain. CFD must first be validated at these conditions.

Description: A novel technique to measure heat transfer and liquid film thickness distributions over relatively large areas for two-phase flow and heat transfer phenomena using infrared (IR)thermometry is described. IR thermometry is an established technology that can be used to measure temperatures when optical access to the surface is available in the wavelengths of interest. In this work, a midwave IR camera (3.6-5.1 microns) is used to determine the temperature distribution within a multilayer consisting of a silicon substrate coated with a thin insulator. Since silicon is largely transparent to IR radiation, the temperature of the inner and outer walls of the multilayer can be measured by coating selected areas with a thin, IR opaque film. If the fluid used is also partially transparent to IR, the flow can be visualized and the liquid film thickness can be measured. The theoretical basis for the technique is given along with a description of the test apparatus and data reduction procedure. The technique is demonstrated by determining the heat transfer coefficient distributions produced by droplet evaporation and flow boiling heat transfer.

Description: Conference paper on supersonic retropropulsion CFD post-test analysis.

Description: CR Tech's Thermal Desktop-SINDA/FLUINT software was used in the thermal analysis of a flame deflector design for Launch Complex 39B at Kennedy Space Center, Florida. The analysis of the flame deflector takes into account heat transfer due to plume impingement from expected vehicles to be launched at KSC. The heat flux from the plume was computed using computational fluid dynamics provided by Ames Research Center in Moffet Field, California. The results from the CFD solutions were mapped onto a 3-D Thermal Desktop model of the flame deflector using the boundary condition mapping capabilities in Thermal Desktop. The ablation subroutine in SINDA/FLUINT was then used to model the ablation of the refractory material.

Description: An exploratory study designed to examine the viability of making air density measurements in a Mach 10 flow using laser-induced fluorescence of the iodine Cordes bands is presented. Experiments are performed in the NASA Langley Research Center 31 in. Mach 10 air wind tunnel in the hypersonic near wake of a multipurpose crew vehicle model. To introduce iodine into the wake, a 0.5% iodine/nitrogen mixture is seeded using a pressure tap at the rear of the model. Air density was measured at 56 points along a 7 mm line and three stagnation pressures of 6.21, 8.62, and 10.0 MPa (900, 1250, and 1450 psi). Average results over time and space show rho(sub wake)/rho(sub freestream) of 0.145 plus or minus 0.010, independent of freestream air density. Average off-body results over time and space agree to better than 7.5% with computed densities from onbody pressure measurements. Densities measured during a single 60 s run at 10.0 MPa are time-dependent and steadily decrease by 15%. This decrease is attributed to model forebody heating by the flow.

Description: An analytical solution was derived for the transient response of an insulated aerospace vehicle structure subjected to a simplified heat pulse. This simplified problem approximates the thermal response of a thermal protection system of an atmospheric entry vehicle. The exact analytical solution is solely a function of two non-dimensional parameters. A simpler function of these two parameters was developed to approximate the maximum structural temperature over a wide range of parameter values. Techniques were developed to choose constant, effective properties to represent the relevant temperature and pressure-dependent properties for the insulator and structure. A technique was also developed to map a time-varying surface temperature history to an equivalent square heat pulse. Using these techniques, the maximum structural temperature rise was calculated using the analytical solutions and shown to typically agree with finite element simulations within 10 to 20 percent over the relevant range of parameters studied.

Description: There are inherent uncertainties and errors associated with using Computational Fluid Dynamics (CFD) to predict the flow field and there is no standard method for evaluating uncertainty in the CFD community. This paper describes an approach to -validate the . uncertainty in using CFD. The method will use the state of the art uncertainty analysis applying different turbulence niodels and draw conclusions on which models provide the least uncertainty and which models most accurately predict the flow of a backward facing step.

Description: Release of ammonia into the International Space Station (ISS) cabin atmosphere can occur if the water/ammonia barrier breach of the active thermal control system (ATCS) interface heat exchanger (IFHX) happens. After IFHX breach liquid ammonia is introduced into the water-filled internal thermal control system (ITCS) and then to the cabin environment through a ruptured gas trap. Once the liquid water/ammonia mixture exits ITCS, it instantly vaporizes and mixes with the U.S. Laboratory cabin air that results in rapid deterioration of the cabin conditions. The goal of the study is to assess ammonia propagation in the Station after IFHX breach to plan the operation procedure. A Computational Fluid Dynamics (CFD) model for accurate prediction of airflow and ammonia transport within each of the modules in the ISS cabin was developed. CFD data on ammonia content in the cabin aisle way of the ISS and, in particular, in the Russian On- Orbit Segment during the period of 15 minutes after gas trap rupture are presented for four scenarios of rupture response. Localized effects of ammonia dispersion and risk mitigation are discussed.

Description: Results of aerothermal material response calculations for the 5th Ablation Workshop Ablation Test-Case Series 2.

Description: In an exploratory investigation, quantitative unclustered laser Rayleigh scattering measurements of density were performed in the air in the NASA Langley Research Center's 31 in. Mach 10 wind tunnel. A review of 20 previous years of data in supersonic and Mach 6 hypersonic flows is presented where clustered signals typically overwhelmed molecular signals. A review of nucleation theory and accompanying nucleation calculations are also provided to interpret the current observed lack of clustering. Data were acquired at a fixed stagnation temperature near 990Kat five stagnation pressures spanning 2.41 to 10.0 MPa (350 to 1454 psi) using a pulsed argon fluoride excimer laser and double-intensified charge-coupled device camera. Data averaged over 371 images and 210 pixels along a 36.7mmline measured freestream densities that agree with computed isentropic-expansion densities to less than 2% and less than 6% at the highest and lowest densities, respectively. Cluster-free Mach 10 results are compared with previous clustered Mach 6 and condensation-free Mach 14 results. Evidence is presented indicating vibrationally excited oxygen and nitrogen molecules are absorbed as the clusters form, release their excess energy, and inhibit or possibly reverse the clustering process. Implications for delaying clustering and condensation onset in hypersonic and hypervelocity facilities are discussed.

Description: Foil bearings are self-acting hydrodynamics devices used to support high speed rotating machinery. The advantages that they offer to process fluid lubricated machines include: high rotational speed capability, no auxiliary lubrication system, non-contacting high speed operation, and improved damping as compared to rigid hydrodynamic bearings. NASA has had a sporadic research program in this technology for almost 6 decades. Advances in the technology and understanding of foil journal bearings have enabled several new commercial products in recent years. These products include oil-free turbochargers for both heavy trucks and automobiles, high speed electric motors, microturbines for distributed power generation, and turbojet engines. However, the foil thrust bearing has not received a complimentary level of research and therefore has become the weak link of oil-free turbomachinery. In an effort to both provide machine designers with basic performance parameters and to elucidate the underlying physics of foil thrust bearings, NASA Glenn Research Center has completed an effort to experimentally measure the performance of simple gas foil thrust bearing in air. The database includes simple bump foil supported thrust bearings with full geometry and manufacturing techniques available to the user. Test conditions consist of air at ambient pressure and temperatures up to 500 C and rotational speeds to 55,000 rpm. A complete set of axial load, frictional torque, and rotational speed is presented for two different compliant sub-structures and inter-pad gaps. Data obtained from commercially available foil thrust bearings both with and without active cooling is presented for comparison. A significant observation made possible by this data set is the speed-load capacity characteristic of foil thrust bearings. Whereas for the foil journal bearing the load capacity increases linearly with rotational speed, the foil thrust bearing operates in the hydrodynamic high speed limit. In this case, the load capacity is constant and in fact often decreases with speed if other factors such as thermal conditions and runner distortions are permitted to dominate the bearing performance.

Description: A new capability to test active flow control concepts and propulsion simulations at high Reynolds numbers in the National Transonic Facility at the NASA Langley Research Center is being developed. The first active flow control experiment was completed using the new FAST-MAC semi-span model to study Reynolds number scaling effects for several circulation control concepts. Testing was conducted over a wide range of Mach numbers, up to chord Reynolds numbers of 30 million. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. Preliminary analysis of the uncorrected lift data showed that the circulation control increased the low-speed maximum lift coefficient by 33%. At transonic speeds, the circulation control was capable of positively altering the shockwave pattern on the upper wing surface and reducing flow separation. Furthermore, application of the technique to only the outboard portion of the wing demonstrated the feasibility of a pneumatic based roll control capability.

Description: Numerical simulations of reacting and non-reacting flows within a scramjet combustor configuration experimentally mapped at the University of Virginia s Scramjet Combustion Facility (operating with Configuration A ) are described in this paper. Reynolds-Averaged Navier-Stokes (RANS) and hybrid Large Eddy Simulation / Reynolds-Averaged Navier-Stokes (LES / RANS) methods are utilized, with the intent of comparing essentially blind predictions with results from non-intrusive flow-field measurement methods including coherent anti-Stokes Raman spectroscopy (CARS), hydroxyl radical planar laser-induced fluorescence (OH-PLIF), stereoscopic particle image velocimetry (SPIV), wavelength modulation spectroscopy (WMS), and focusing Schlieren. NC State's REACTMB solver was used both for RANS and LES / RANS, along with a 9-species, 19- reaction H2-air kinetics mechanism by Jachimowski. Inviscid fluxes were evaluated using Edwards LDFSS flux-splitting scheme, and the Menter BSL turbulence model was utilized in both full-domain RANS simulations and as the unsteady RANS portion of the LES / RANS closure. Simulations were executed and compared with experiment at two equivalence ratios, PHI = 0.17 and PHI = 0.34. Results show that the PHI = 0.17 flame is hotter near the injector while the PHI = 0.34 flame is displaced further downstream in the combustor, though it is still anchored to the injector. Reactant mixing was predicted to be much better at the lower equivalence ratio. The LES / RANS model appears to predict lower overall heat release compared to RANS (at least for PHI = 0.17), and its capability to capture the direct effects of larger turbulent eddies leads to much better predictions of reactant mixing and combustion in the flame stabilization region downstream of the fuel injector. Numerical results from the LES/RANS model also show very good agreement with OH-PLIF and SPIV measurements. An un-damped long-wave oscillation of the pre-combustion shock train, which caused convergence problems in some RANS simulations, was also captured in LES / RANS simulations, which were able to accommodate its effects accurately.

Description: Several projects have been completed or are nearing completion at the NASA Langley Research Center (LaRC) National Transonic Facility (NTF). The addition of a Model Flow-Control/Propulsion Simulation test capability to the NTF provides a unique, transonic, high-Reynolds number test capability that is well suited for research in propulsion airframe integration studies, circulation control high-lift concepts, powered lift, and cruise separation flow control. A 1992 vintage Facility Automation System (FAS) that performs the control functions for tunnel pressure, temperature, Mach number, model position, safety interlock and supervisory controls was replaced using current, commercially available components. This FAS upgrade also involved a design study for the replacement of the facility Mach measurement system and the development of a software-based simulation model of NTF processes and control systems. The FAS upgrades were validated by a post upgrade verification wind tunnel test. The data acquisition system (DAS) upgrade project involves the design, purchase, build, integration, installation and verification of a new DAS by replacing several early 1990's vintage computer systems with state of the art hardware/software. This paper provides an update on the progress made in these efforts. See reference 1.

Description: A discrete adjoint-based design methodology for unsteady turbulent flows on three-dimensional dynamic overset unstructured grids is formulated, implemented, and verified. The methodology supports both compressible and incompressible flows and is amenable to massively parallel computing environments. The approach provides a general framework for performing highly efficient and discretely consistent sensitivity analysis for problems involving arbitrary combinations of overset unstructured grids which may be static, undergoing rigid or deforming motions, or any combination thereof. General parent-child motions are also accommodated, and the accuracy of the implementation is established using an independent verification based on a complex-variable approach. The methodology is used to demonstrate aerodynamic optimizations of a wind turbine geometry, a biologically-inspired flapping wing, and a complex helicopter configuration subject to trimming constraints. The objective function for each problem is successfully reduced and all specified constraints are satisfied.

Description: The results of a series of Perfectly Stirred Reactor (PSR) and Partially Stirred Reactor (PaSR) simulations are compared to each other over a wide range of operating conditions. It is found that the PaSR results can be simulated by a PSR solution with just an adjusted chemical reaction rate. A simple expression has been developed that gives the required change in reaction rate for a PSR solution to simulate the PaSR results. This expression is the basis of a simple turbulence-chemistry interaction model. The interaction model that has been developed is intended for use with simple one-step global reaction mechanisms and for steady-state flow simulations. Due to the simplicity of the model there is very little additional computational cost in adding it to existing CFD codes.

Description: A thermal protection system and a method of manufacturing are disclosed. The thermal protection system may be configured to protect a movable joint, for example, a flexible bearing of a rocket motor nozzle. The thermal protection system includes a series of annular shims separated by a plurality of discrete spacers. Each shim of the series of annular shims may have a larger diameter than the previous shim, and the shims may nest. The shims may comprise a thermally stable material, and the discrete spacers may comprise an elastomer. Optionally, an annular bearing protector may separate the annular shims from the flexible bearing.

Description: Experiments are performed concerning frost growth and densification in laminar flow over a flat surface under conditions of constant and variable humidity. The flat plate test specimen is made of aluminum-6031, and has dimensions of 0.3 mx0.3 mx6.35 mm. Results for the first variable humidity case are obtained for a plate temperature of 255.4 K, air velocity of 1.77 m/s, air temperature of 295.1 K, and a relative humidity continuously ranging from 81 to 54%. The second variable humidity test case corresponds to plate temperature of 255.4 K, air velocity of 2.44 m/s, air temperature of 291.8 K, and a relative humidity ranging from 66 to 59%. Results for the constant humidity case are obtained for a plate temperature of 263.7 K, air velocity of 1.7 m/s, air temperature of 295 K, and a relative humidity of 71.6 %. Comparisons of the data with the author's frost model extended to accommodate variable humidity suggest satisfactory agreement between the theory and the data for both constant and variable humidity.

Description: A Large Eddy Simulation (LES) is performed of a high blowing ratio (M = 1.7) film cooling flow with density ratio of unity. Mean results are compared with experimental data to show the degree of fidelity achieved in the simulation. While the trends in the LES prediction are a noticeable improvement over Reynolds-Averaged Navier-Stokes (RANS) predictions, there is still a lack a spreading on the underside of the lifted jet. This is likely due to the inability of the LES to capture the full range of influential eddies on the underside of the jet due to their smaller structure. The unsteady structures in the turbulent coolant jet are also explored and related to turbulent mixing characteristics

Description: A gas-charging plug can be easily analyzed for random vibration. The design features two steeped O-rings in a radial configuration at two different diameters, with a 0.050-in. (.1.3-mm) diameter through-hole between the two O-rings. In the charging state, the top O-ring is engaged and sealing. The bottom O-ring outer diameter is not squeezed, and allows air to flow by it into the tank. The inner diameter is stretched to plug the gland diameter, and is restrained by the O-ring groove. The charging port bushing provides mechanical stop to restrain the plug during gas charge removal. It also prevents the plug from becoming a projectile when removing gas charge from the accumulator. The plug can easily be verified after installation to ensure leakage requirements are met.

Description: The formation of scour patterns at bridge piers is driven by the forces at the boundary of the water flow. In most experimental scour studies, indirect processes have been applied to estimate the shear stress using measured velocity profiles. The estimations are based on theoretical models and associated assumptions. However, the turbulence flow fields and boundary layer in the pier-scour region are very complex and lead to low-fidelity results. In addition, available turbulence models cannot account accurately for the bed roughness effect. Direct measurement of the boundary shear stress, normal stress, and their fluctuations are attractive alternatives. However, most direct-measurement shear sensors are bulky in size or not compatible to fluid flow. A sensor has been developed that consists of a floating plate with folded beam support and an optical grid on the back, combined with a high-resolution optical position probe. The folded beam support makes the floating plate more flexible in the sensing direction within a small footprint, while maintaining high stiffness in the other directions. The floating plate converts the shear force to displacement, and the optical probe detects the plate s position with nanometer resolution by sensing the pattern of the diffraction field of the grid through a glass window. This configuration makes the sensor compatible with liquid flow applications.

Description: A non-toxic, non-flammable, low-freezing heat transfer fluid is being developed for drop-in replacement within current and future heat transfer loops currently using water or alcohol-based coolants. Numerous water-soluble compounds were down-selected and screened for toxicological, physical, chemical, compatibility, thermodynamic, and heat transfer properties. Two fluids were developed, one with a freezing point near 0 C, and one with a suppressed freezing point. Both fluids contain an additive package to improve material compatibility and microbial resistance. The optimized sub-zero solution had a freezing point of 30 C, and a freezing volume expansion of 10-percent of water. The toxicity of the solutions was experimentally determined as LD(50) greater than 5g/kg. The solutions were found to produce minimal corrosion with materials identified by NASA as potentially existing in secondary cooling loops. Thermal/hydrodynamic performance exceeded that of glycol-based fluids with comparable freezing points for temperatures Tf greater than 20 C. The additive package was demonstrated as a buffering agent to compensate for CO2 absorption, and to prevent microbial growth. The optimized solutions were determined to have physically/chemically stable shelf lives for freeze/thaw cycles and longterm test loop tests.

Description: No abstract available

Description: 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.

Description: Methods of utilizing magnetic particles or beads (MBs) in droplet-based (or digital) microfluidics are disclosed. The methods may be used in enrichment or separation processes. A first method employs the droplet meniscus to assist in the magnetic collection and positioning of MBs during droplet microfluidic operations. The sweeping movement of the meniscus lifts the MBs off the solid surface and frees them from various surface forces acting on the MBs. A second method uses chemical additives to reduce the adhesion of MBs to surfaces. Both methods allow the MBs on a solid surface to be effectively moved by magnetic force. Droplets may be driven by various methods or techniques including, for example, electrowetting, electrostatic, electromechanical, electrophoretic, dielectrophoretic, electroosmotic, thermocapillary, surface acoustic, and pressure.

Description: A high-fidelity computational fluid dynamics simulation of a next generation heavy lift space vehicle during launch is presented. The purpose of the simulation is to evaluate the acoustic overpressures during ignition to permit re-design of the launch site to safely handle heavy lift vehicles. The simulation is performed using the Launch, Ascent, and Vehicle Aerodynamics (LAVA) code, an immersed boundary block-structured Cartesian adaptive mesh refinement based solver. A verification and validation study of LAVA in the launch environment context is also performed, comparing to flight data and previous simulations of a Space Shuttle launch

Description: A dual-pump Coherent Anti-Stokes Raman Spectroscopy (CARS) instrument has been developed to obtain simultaneous measurements of temperature and absolute mole fractions of N2, O2 and H2 in supersonic combustion and generate databases for validation and development of CFD codes. Issues that compromised previous attempts, such as beam steering and high irradiance perturbation effects, have been alleviated or avoided. Improvements in instrument precision and accuracy have been achieved. An axis-symmetric supersonic combusting coaxial jet facility has been developed to provide a simple, yet suitable flow to CFD modelers. Approximately one million dual-pump CARS single shots have been collected in the supersonic jet for varying values of flight and exit Mach numbers at several locations. Data have been acquired with a H2 co-flow (combustion case) or a N2 co-flow (mixing case). Results are presented and the effects of the compressibility and of the heat release are discussed.

Description: The objective of this task was to quantify thermal losses involving integrating MLI into real life situations. Testing specifically focused on the effects of penetrations (including structural attachments, electrical conduit/feedthroughs, and fluid lines) through MLI. While there have been attempts at quantifying these losses both analytically and experimentally, none have included a thorough investigation of the methods and materials that could be used in such applications. To attempt to quantify the excess heat load coming into the system due to the integration losses, a calorimeter was designed to study two dimensional heat transfer through penetrated MLI. The test matrix was designed to take as many variables into account as was possible with the limited test duration and system size. The parameters varied were the attachment mechanism, the buffer material (for buffer attachment mechanisms only), the thickness of the buffer, and the penetration material. The work done under this task is an attempt to measure the parasitic heat loads and affected insulation areas produced by system integration, to model the parasitic loads, and from the model produce engineering equations to allow for the determination of parasitic heat loads in future applications. The methods of integration investigated were no integration, using a buffer to thermally isolate the strut from the MLI, and temperature matching the MLI on the strut. Several materials were investigated as a buffer material including aerogel blankets, aerogel bead packages, cryolite, and even an evacuated vacuum space (in essence a no buffer condition).

Description: No abstract available

Description: The Resource Prospector Mission (RPM) is an in-situ resource utilization (ISRU) technology demonstration mission planned to launch in 2018. The mission will use the Regolith and Environment Science & Oxygen and Lunar Volatile Extraction (RESOLVE) Payload to prospect for lunar volatiles such as water, oxygen, and carbon dioxide. These compounds will validate ISRU capability. The payload, particularly the Lunar Advanced Volatile Analysis (LAVA) subsystem, requires numerous temperature measurements to accurately control on-board heaters that keep the volatiles in the vapor phase to allow quantification and prevent the clogging of delivery lines. Previous spaceflight missions have proven that Resistive Temperature Detector (RTD) failure impedes mission success. The research resulted in a recommendation for a flight-forward RTD. The recommendation was based on accuracy, consistency, and ease of installation of RTDs procured from IST, QTI, and Honeywell.

Description: It is widely known that graphene and many of its derivative nanostructures have exceedingly high reported thermal conductivities (up to 4000 W/mK at 300 K). Such attractive thermal properties beg the use of these structures in practical devices; however, to implement these materials while preserving transport quality, the influence of structure on thermal conductivity should be thoroughly understood. For graphene nanostructures, having average phonon mean free paths on the order of one micron, a primary concern is how size influences the potential for heat conduction. To investigate this, we employ a novel technique to evaluate the lattice thermal conductivity from the Green-Kubo relations and equilibrium molecular dynamics in systems where phonon-boundary scattering dominates heat flow. Specifically, the thermal conductivities of graphene nanoribbons and carbon nanotubes are calculated in sizes up to 3 microns, and the relative influence of boundary scattering on thermal transport is determined to be dominant at sizes less than 1 micron, after which the thermal transport largely depends on the quality of the nanostructure interface. The method is also extended to carbon nanostructures (fullerenes) where phonon confinement, as opposed to boundary scattering, dominates, and general trends related to the influence of curvature on thermal transport in these materials are discussed.

Description: This report documents the work performed from March 2010 to March 2012. The Integrated Design and Engineering Analysis (IDEA) environment is a collaborative environment based on an object-oriented, multidisciplinary, distributed framework using the Adaptive Modeling Language (AML) as a framework and supporting the configuration design and parametric CFD grid generation. This report will focus on describing the work in the area of parametric CFD grid generation using novel concepts for defining the interaction between the mesh topology and the geometry in such a way as to separate the mesh topology from the geometric topology while maintaining the link between the mesh topology and the actual geometry.

Description: The present disclosure relates in part to a flow field structure comprising a hydrophilic part and a hydrophobic part communicably attached to each other via a connecting interface. The present disclosure further relates to electrochemical cells comprising the aforementioned flow fields.

Description: Polystyrene latex microspheres (PSLs) have been used for particle image velocimetry (PIV) and laser Doppler velocimetry (LDV) measurements for several decades. With advances in laser technologies, instrumentation, and data processing, the capability to collect more information about fluid flow beyond velocity is possible using new seed materials. To provide additional measurement capability, PSLs were synthesized with temperature-sensitive fluorescent dyes incorporated within the particle. These multifunctional PSLs would have the greatest impact if they could be used in large scale facilities with minimal modification to the facilities or the existing instrumentation. Consequently, several potential dyes were identified that were amenable to existing laser systems currently utilized in wind tunnels at NASA Langley Research Center as well as other wind and fluid (water) tunnels. PSLs incorporated with Rhodamine B, dichlorofluorescein (DCF, also known as fluorescein 548 or fluorescein 27) and other dyes were synthesized and characterized for morphology and spectral properties. The resulting particles were demonstrated to exhibit fluorescent emission, which would enable determination of both fluid velocity and temperature. They also would allow near-wall velocity measurements whereas laser scatter from surfaces currently prevents near-wall measurements using undoped seed materials. Preliminary results in a wind tunnel facility located at Virginia Polytechnic Institute and State University (Virginia Tech) have verified fluorescent signal detection and temperature sensitivity of fluorophore-doped PSLs.

Description: Thermocapillary convection in a liquid bridge, which is suspended between two coaxial disks under zero gravity, has been investigated numerically. The Navier-Stokes equations coupled with the energy conservation equation are solved on a staggered grid, and the level set approach is used to capture the free surface deformation of the liquid bridge. The velocity and temperature distributions inside the liquid bridge are analyzed. It is shown from this work that as the development of the thermocapillary convection, the center of the vortex inside the liquid bridge moves down and reaches an equilibrium position gradually. The temperature gradients in the regions near the upper center axis and the bottom cold corner are higher than those in the other regions.

Description: At modest temperatures, the thermal energy of atmospheric diatomic gases such as nitrogen is primarily distributed between only translational and rotational energy modes. Furthermore, these energy modes are fully excited such that the specific heat at constant volume is well approximated by the simple expression C(sub v) = 5/2 R. As a result, classical mechanics provides a suitable approximation at such temperatures of the true quantum mechanical behavior of the inter-molecular collisions of such molecules. Using classical mechanics, the transfer of energy between rotational and translation energy modes is studied. The approach of Lordi and Mates is adopted to compute the trajectories and time dependent rotational orientations and energies during the collision of two non-polar diatomic molecules. A Monte-Carlo analysis is performed collecting data from the results of many such simulations in order to estimate the rotational relaxation time. A Graphical Processing Unit (GPU) is employed to improve the performance of the Monte-Carlo analysis. A comparison of the performance of the GPU implementation to an implementation on traditional computer architecture is made. Effects of the assumed inter-molecular potential on the relaxation time are studied. The seminar will also present highlights of computational analyses performed at NASA Johnson Space Center of heat transfer in rarefied gases.

Description: Projects: (1) Boeing Launch Abort Analysis My first project for the summer was analyzing the Boeing CCDev Vehicle's abort aerodynamics using an inviscid solver (CART3D). The goal of the project was to develop the grid and CFD inputs necessary to use CART3D as a quick tool for investigating loading trends at various points along abort trajectories. As a supplementary task, I analyzed a few cases and compared them to the aerodatabase from the last generation geometry. (2) My second project for the summer dealt with investigating how heating changes as the height of a protuberance on top of a flat plate changes. The goal of this investigation is to better understand how to properly model heating on and around a protuberance. This is one of the biggest challenges when designing a re ]entry vehicle because very small changes in the shape and conditions leading up to a protuberance, not to mention the protuberance geometry, will greatly impact the local heating.

Description: An analytical solution was derived for the transient response of an insulated structure subjected to a simplified heat pulse. The solution is solely a function of two nondimensional parameters. Simpler functions of these two parameters were developed to approximate the maximum structural temperature over a wide range of parameter values. Techniques were developed to choose constant, effective thermal properties to represent the relevant temperature and pressure-dependent properties for the insulator and structure. A technique was also developed to map a time-varying surface temperature history to an equivalent square heat pulse. Equations were also developed for the minimum mass required to maintain the inner, unheated surface below a specified temperature. In the course of the derivation, two figures of merit were identified. Required insulation masses calculated using the approximate equation were shown to typically agree with finite element results within 10%-20% over the relevant range of parameters studied.

Description: As a follow-on activity to the HyBoLT flight experiment, a six degree half angle wedge-cone model at zero angle of attack has been employed to experimentally and computationally study the boundary layer crossflow instability at Mach 3.5 under low disturbance freestream conditions. Computed meanflow and linear stability analysis results are presented along with corresponding experimental Pitot probe data. Using a model-mounted probe survey apparatus, data acquired to date show a well defined stationary crossflow vortex pattern on the flat wedge surface. This effort paves the way for additional detailed, calibrated flow field measurements of the crossflow instability, both stationary and traveling modes, and transition-to-turbulence under quiet flow conditions as a means of validating existing stability theory and providing a foundation for dynamic flight instrumentation development.

Description: The purpose of this report is to summarize and document the work done to enable a NASA CFD code to model laminar-turbulent transition process on an isolated turbine blade. The ultimate purpose of the present work is to down-select a transition model that would allow the flow simulation of a variable speed power turbine to be accurately performed. The flow modeling in its final form will account for the blade row interactions and their effects on transition which would lead to accurate accounting for losses. The present work only concerns itself with steady flows of variable inlet turbulence. The low Reynolds number k- model of Wilcox and a modified version of the same model will be used for modeling of transition on experimentally measured blade pressure and heat transfer. It will be shown that the k- model and its modified variant fail to simulate the transition with any degree of accuracy. A case is thus made for the adoption of more accurate transition models. Three-equation models based on the work of Mayle on Laminar Kinetic Energy were explored. The three-equation model of Walters and Leylek was thought to be in a relatively mature state of development and was implemented in the Glenn-HT code. Two-dimensional heat transfer predictions of flat plate flow and two-dimensional and three-dimensional heat transfer predictions on a turbine blade were performed and reported herein. Surface heat transfer rate serves as sensitive indicator of transition. With the newly implemented model, it was shown that the simulation of transition process is much improved over the baseline k- model for the single Reynolds number and pressure ratio attempted; while agreement with heat transfer data became more satisfactory. Armed with the new transition model, total-pressure losses of computed three-dimensional flow of E3 tip section cascade were compared to the experimental data for a range of incidence angles. The results obtained, form a partial loss bucket for the chosen blade. In time the loss bucket will be populated with losses at additional incidences. Results obtained thus far will be discussed herein.

Description: An exhaust nozzle includes a conical duct terminating in an annular outlet. A row of vortex generating duplex tabs are mounted in the outlet. The tabs have compound radial and circumferential aft inclination inside the outlet for generating streamwise vortices for attenuating exhaust noise while reducing performance loss.

Description: A thermal expansion compensator is provided and includes a first electrode structure having a first surface, a second electrode structure having a second surface facing the first surface and an elastic element bonded to the first and second surfaces and including a conductive element by which the first and second electrode structures electrically and/or thermally communicate, the conductive element having a length that is not substantially longer than a distance between the first and second surfaces.