ALBERT

Description: A thermal protection system (TPS) comprising a mixture of silicon carbide and SiOx that has been converted from Si that is present in a collection of diatom frustules and at least one diatom has quasi-periodic pore-to-pore separation distance d(p-p) in a selected range. Where a heat shield comprising the converted SiC/SiOx frustules receives radiation, associated with atmospheric (re)entry, a portion of this radiation is reflected so that radiation loading of the heat shield is reduced.

Description: Computational fluid dynamics (CFD) computations are performed at time increments using structural properties of the nozzle and flow properties of combustion products flowing through the nozzle. Each CFD computation accounts for movement of the wall geometry of the rocket nozzle due to the flowfield. Structural dynamics computations are performed at each time increment using the CFD computations in order to describe the movement of the wall geometry. Mesh dynamics computations at each time increment redefine the flowfield to account for the movement of the wall geometry. The mesh dynamics computations are based on a spring analogy process. The computations are iterated to solution convergence at each time increment with results being output to an output device.

Description: We present a high-order finite-element method for moving body and fluid/structure interaction problems. Our solution strategy is based on a space-time discontinuous Galerkin (DG) spectral-element discretization which extends to arbitrary order of accuracy. The space-time DG discretization is a natural choice for moving body and fluid-structure interaction problems as moving surfaces are incorporated simply by considering curved space-time elements whose space-time faces align with the moving body. We present a discontinuous-Galerkin in time discretization for six-degree of motion modeling of rigid bodies, and a continuous-Galerkin discretization for equations of linear elasticity to generate curved space-time meshes. Numerical results for several simple 2D test cases are presented in order to verify the implementation of the different models. Finally we present a preliminary dynamic simulation of a parachute.

Description: Lattice Boltzmann (LB) and hybrid Reynolds-averaged Navier-Stokes/large eddy simulation (RANS/LES) methods within the Launch Ascent and Vehicle Aerodynamics (LAVA) solver framework are applied to NASA's Revolutionary Computational Aerosciences (RCA) standard test cases for separated flows. A detailed comparison between the performance and accuracy of the two emerging numerical methodologies for turbulence resolving simulations, i.e. the LB and hybrid RANS/LES methods will be presented. This contribution addresses the RCA technical challenge to identify and down-select critical turbulence, transition, and numerical method technologies for 40% reduction in predictive error for standard turbulence separated flow test cases. Results for the 2D NASA wall-mounted hump and the axisymmetric transonic bump including time-averaged pressure coefficient, skin friction, and velocity pro les, as well as resolved and modeled Reynolds stresses for both numerical approaches will be presented and differences between LB and hybrid RANS/LES will be discussed.

Description: The Hypersonic Materials Environmental Test System arc-jet facility located at the NASA Langley Research Center in Hampton, Virginia, is primarily used for the research, development, and evaluation of high-temperature thermal protection systems for hypersonic vehicles and reentry systems. In order to improve testing capabilities and knowledge of the test article environment, a detailed three-dimensional model of the arc-jet nozzle and free-jet portion of the flow field has been developed. The computational fluid dynamics model takes into account non-uniform inflow state profiles at the nozzle inlet as well as catalytic recombination efficiency effects at the probe surface. Results of the numerical simulations are compared to calibrated Pitot pressure and stagnation-point heat flux for three test conditions at low, medium, and high enthalpy. Comparing the results and test data indicates an effectively fully-catalytic copper surface on the heat flux probe of about 10% recombination efficiency and a 2-3 kPa pressure drop from the total pressure measured at the plenum section, prior to the nozzle. With these assumptions, the predictions are within the uncertainty of the stagnation pressure and heat flux measurements. The predicted velocity conditions at the nozzle exit were also compared and showed good agreement with radial and axial velocimetry data.

Description: After the Columbus Moderate Temperature Loop (MTL) InterFace Heat eXchanger (IFHX) low temperature event of GMT 345-2013, NASA investigated relevant transient scenarios involving IFHX rupture after water freezing and subsequent thawing. NASA recommended development of a Fault Detection Isolation and Recovery (FDIR) plan that would, in the event of a heat exchanger freeze event, close the Water On/Off Valves (WOOVs) to isolate the heat exchanger and prevent ammonia from the external flow loops from spreading into the cabin. NASA performed a preliminary simplified analysis for the reference case of IFHX rupture, but for a deeper understanding TAS developed detailed SINDA-FLUINT models of the Columbus ITCS that were built and run through the SINAPS GUI. This allowed simulation of the ammonia leakage physics including the variation of environmental parameters, thus providing more accurate and specific input to the FDIR under development. The result was finalization of the IFHX WOOVs closure sequence and wait times to contain the ammonia propagation to Columbus and allow identification of the leaking IFHX. In addition, the analysis results provided reference pressure profiles to be used on console and by the Engineering as support for the telemetry data assessment in case of failure.This paper gives an overview on the issue and focuses on the analytical aspects of the multiphase fluid dynamics involved.

Description: Disclosed herein is a cryogenic heat transfer system capable of transferring 50 W or more at cryogenic temperatures of 100.degree. K or less for use with cryocooler systems. In an embodiment, a cryogenic heat transfer system comprises a refrigerant contained within an inner chamber bound by a condenser in fluid communication with an evaporator through at least one flexible conduit, the condenser in thermal communication with the cold station of a cryocooler, and the evaporator positionable in thermal communication with a heat source, typically a radiation shield of a cryogenic chamber. A process to remove heat from a cryogenic chamber is also disclosed.

Description: Several quantitative measurements extracted from nitric oxide (NO) planar laser-induced fluorescence (PLIF) data obtained in a hypersonic boundary layer are reported: (a) off-body NO mole fraction; (b) surface heat flux; and (c) near-wall static temperature. The experimental data was obtained at NASA Langley Research Centers 31 in. Mach 10 air tunnel. NO was seeded into the flow through a spanwise slot on the surface of the 10 degree half-angle wedge model. An ultraviolet planar laser sheet was positioned perpendicular to the wedge surface, downstream of the seeding slot, to excite six fluorescence transitions. A method for extracting the relative NO mole fraction, based on spatial variations of the J= 0.5 PLIF signal, is presented. Combined with the principle of mass conservation, the absolute NO mole fraction is determined. These measurements were used to assess CFD diffusion modelling, correct previously reported PLIF thermometry results, and develop methods for NO-PLIF heat transfer measurements.

Description: The development and implementation of kL-based Reynolds Average Navier-Stokes (RANS) two-equation turbulence models are reported herein. The kL is based on Abdol-Hamid's closure and Menter's modification to Rotta's two-equation model. Rotta shows that a reliable transport equation can be formed from the turbulent length scale L, and the turbulent kinetic energy k. Rotta's kL equation is well suited for term-by-term modeling and displays useful features compared to other scale formulation. One of the important differences is the inclusion of higher order velocity derivatives in the source terms of the scale equation. This can enhance the ability of RANS solvers to simulate unsteady flows in URANS mode. The present report documents the formulation of two model levels of turbulence models as implemented in the computational fluid dynamics FUN3D code. The levels are the two-equation linear k-kL and the two-equation algebraic Reynolds stress model (ARSM). Free shear, separated and corner flow cases are documented and compared with experimental, and other turbulence model data. The results show generally very good comparisons with experimental data. The results from this formulation are similar or better than results using the SST two-equation turbulence model. ARSM shows great promise with similar level of computational resources as basic two-equation turbulence models.

Description: The near and very near wake of thin flat plates with both sharp and circular trailing edges (TEs) are investigated with direct numerical simulations (DNSs). The TE is circular in two of the cases (IN & NS) and sharp in one of them (ST). The separating boundary layers are turbulent in all cases. The objectives of this study are twofold. The first is to explore the effect of significantly reducing Re(sub D) (Reynolds number based on circular TE diameter, D) on the flow in the TE region, and the shedding process (Cases IN and NS). The second is to better understand the reasons underlying the findings of an earlier experimental wake investigation (sharp TE) where (1) the center-line values in normal intensity, and the peak in shear stress profiles in the cross-stream direction, were found to first increase in the streamwise direction (x), from that obtained at the TE, before diminishing further downstream, and (2) a broadband peak was observed in centerline cross-stream velocity (v) spectra (indicating quasi-periodicity, possibly due to vortices or wave-like motions). Case ST from the present study showed a near wake instability resulting in spanwise vortices (with a streamwise component). The instability is intermittent and contributes to both the broadband peak in the v spectrum and the initial increases in normal intensity and shear stress (as in the experiment). Case NS, with the lower value of ReD is an "essentially" non-shedding case where the flow in the TE region continually changes direction (upward/downward) because of turbulence. Case IN, with twice the value of ReD as Case NS, also exhibits a swaying motion in the TE region. In addition, vortex shedding is initiated during periods when the flow direction changes rapidly. Shedding in this case is intermittent. It results in a peak in the v spectrum obtained at the centerline (x/D = 1.0).

Description: In the context of Large- Eddy Simulations (LES), Boundary-layer inflow turbulence is simulated using both the Synthetic Eddy Model (SEM) and Digital Filtering (DF). The effects of the projection error are investigated. The effect of the prescribed length scales on the adjustment region was found to be negligible for length scales less than one-tenth of the boundary-layer thickness. While it was conjectured that one method of the two might be more robust than the other, our results show that both the Digital Filtering Method and the Synthetic Eddy Method accurately replicate the boundary layer while successfully accounting for inflow turbulence.

Description: Experiments were performed in NASAs SW-2 cascade facility to compare three trailing edge actuation concepts to a 3D airfoil section with no trailing edge treatment. At a Reynolds number of 105, trailing edge pulsed ejection using fluidic oscillator devices was shown to fill the momentum deficit in the wake more uniformly than other actuators tested, with expected benefits for tonal noise in engine fans. Furthermore, pulsed ejection was found to alter the acoustic signature of the wake to reduce broadband noise. In some locations in the wake, spectral components of velocity were found to be reduced by 2 to 5 dB across nearly all frequencies. Trailing edge pulsed ejection is established as a feasible concept to reduce both tonal and broadband noise emissions from aircraft engines.

Description: The development and implementation of kL-based Reynolds Average Navier-Stokes (RANS) turbulence models are reported herein. The kL is based on Abdol-Hamid's closure and Menter's modi cation to Rotta's two-equation model. Rotta shows that a reliable transport equation can be formed from the turbulent length scale L, and the turbulent kinetic energy k. Rotta's kL equation is well suited for term-by-term modeling and displays useful features compared to other scale formulation. One of the important di erences is the inclusion of higher order velocity derivatives in the source terms of the scale equation. This can enhance the ability of RANS solvers to simulate unsteady ows in URANS mode. The present report documents the formulation of three model levels of turbulence models as implemented in the CFD code FUN3D. Methodology and calibration examples are shown in detail. The levels are the linear k-kL and the two-equation algebraic Reynolds stress model (ARSM) as well as the full Reynolds Stress Model (RSM). Attached, separated and corner ow cases are documented and compared with experimental, theoretical and other turbulence model data. The results show generally very good comparisons with canonical and experimental data. The results from this formulation are similar or better than results using the SST two- equation turbulence model. ARSM shows great promise with similar level of computational resources as general two equation turbulence models.

Description: No abstract available

Description: Lattice Boltzmann (LB) based Large Eddy Simulation (LES), Reynolds-averaged Navier-Stokes (RANS) as well as hybrid RANS/LES methods within the Launch Ascent and Vehicle Aerodynamics (LAVA) solver framework are applied to NASA's wall-mounted hump. Computational results are compared with experiments performed by Greenblatt et al. A detailed comparison between the accuracy and resolution requirements of the two approaches for turbulence resolving simulations, as well as the suitability of different grid paradigms (body-fitted curvilinear and block structured Cartesian) are presented. This test case is part of NASA's Revolutionary Computational Aerosciences (RCA) sub-project which addresses the technical challenge of predicting flow separation and reattachment accurately. Improvements in predictive accuracy by as much as 90% are demonstrated using LB as well as hybrid RANS/LES approaches compared to state-of-the-art steady state RANS simulations.

Description: Electrowetting heat pipes (EHPs) are a newly conceptualized class of heat pipes, wherein the adiabatic wick section is replaced by electrowetting-based pumping of the condensate (as droplets) to the evaporator. Specific advantages include the ability to transport high heat loads over long distances, low thermal resistance and power consumption, and the absence of moving mechanical parts. In this work, we describe characterization of key microfluidic operations (droplet motion and splitting) underlying the EHP on the International Space Station (ISS). The testing was performed under the Advanced Passive Thermal eXperiment (APTx) project, a project to test a suite of passive thermal control devices funded by the ISS Technology Demonstration Office at NASA JSC (Johnson Space Center). A rapid manufacturing method was used to fabricate the electrowetting device on a printed circuit board. Key device-related considerations were to ensure reliability and package the experimental hardware within a confined space. Onboard the ISS, experiments were conducted to study electrowetting-based droplet motion and droplet splitting, by imaging droplet manipulation operations via pre-programmed electrical actuation sequences. An applied electric field of 36 Volts per micron resulted in droplet speeds approaching 10 millimeters per second. Droplet splitting dynamics were observed and the time required to split droplets was quantified. Droplet motion data was analyzed to estimate the contact line friction coefficient. Overall, this demonstration is the first-ever electrowetting experiment in space. The obtained results are useful for future design of the EHP and other electrowetting-based systems for microgravity applications.

Description: Introduction: NASAs next mission to Mars, the Mars 2020, will use the same heatshield of the Mars Science Laboratory (MSL) for thermal protection during entry, descent and landing. The heatshield is a tiled system made of Phenolic Impregnated Carbon Ablators (PICA) blocks [1]. PICA is a lightweight carbon fiber/polymeric resin material that offers excellent performances for protecting probes during planetary entry. The Mars Entry Descent and Landing Instrument (MEDLI) suite on MSL offers unique in-flight validation data for models of atmospheric entry and material response. MEDLI recorded, among others, time-resolved in-depth temperature data of PICA using thermocouple sensors assembled in the MEDLI Integrated Sensor Plugs (MISP). The objective of this work is to compare the thermal response of the MSL heatshield to the MISP flight data. In preparation to Mars 2020 post-flight analysis, the predictive material response capability is benchmarked against MEDLI flight data.

Description: Protecting a spacecraft during atmospheric entry is one of highest risk factors that needs to be mitigated during design of a space exploration mission. At entry speeds from space, air turns into high-temperature plasma, and spacecraft Thermal Protection Systems (TPS) are needed to protect the vehicle payload. Modern successful material architectures of spacecraft shields use a porous carbon fiber substrate impregnated with phenolic as an ablator material. In the lecture, efforts to build a Predictive Material Modeling framework for porous ablators from micro-scale to macro-scale will be presented. Several numerical methods and techniques will be summarized that use voxelized images to compute geometrical properties of the porous substrate. These computed properties include porosity, specific surface area and tortuosity that are otherwise indirectly measured through experimental techniques. Direct simulation Monte Carlo (DSMC), a particle-based method for approximating the Boltzmann equation, is used to compute the permeability coefficient of the porous substrate based on its digitized representation. The method computes the flow within the microstructure, where the size of the pores may approach the mean-free-path of the flow. Finally, a high-fidelity model implemented in PATO (Porous-material Analysis Toolbox) is discussed, and some examples of ablative material response are presented, including for the first time 3D simulations of the full tiled heat shield for the Mars Science Laboratory (MSL) capsule.

Description: No abstract available

Description: Recent introduction of Coaxial Thermocouple type calorimeters into the NASA Ames arc jet facilities has inspired an analysis of 2D conduction effects internal to this type of calorimeter. The 1D finite slab inverse analysis (which is typically used to deduce the heat transfer to the calorimeter) relies on the assumption that lateral conduction (i.e., 2D effects) is negligible. Most calorimeter bodies have a spherical nose, which in itself is a violation of the 1D finite slab analysis assumption. Secondly most calorimeters experience a variation in heating across the face of the body which is also a violation of the 1D finite slab analysis assumption. It turns out that these two effects tend to cancel each other to some extent. This paper shows the extent to which error exists in the analysis of the Coaxial Thermocouple type calorimeters, and also offers analysis strategies for reducing the errors.

Description: Due to the unique thermal vacuum testing requirements for a Mars Rover instrument, NASA Goddard developed a low cost, high fidelity thermal control system utilizing Thermal Electric Coolers (TECs) combined with a heat rejection fluid loop to actively control 8 independent payload thermal boundary zones in a simulated Mars pressure vacuum chamber with a Carbon Dioxide atmosphere. These zones could control instrument components to a specific temperature as a function of time to simulate exact temporal flight boundary predictions.The Mars Organic Molecule Analyzer (MOMA) instrument is a dual source (pyrolysis gas chromatograph and laser desorption) mass spectrometer (MS) based package that detects and characterizes organic molecules, as part of ESA's 2020 ExoMars Rover mission to seek the signs of life on Mars.Due to the unique thermal vacuum testing requirements for a Mars Rover instrument, NASA Goddard developed a low cost, high fidelity thermal control system utilizing Thermal Electric Coolers (TECs) combined with a heat rejection fluid loop to actively control 8 independent payload thermal boundary zones in a simulated Mars pressure vacuum chamber with a Carbon Dioxide atmosphere. These zones could control instrument components to a specific temperature as a function of time to simulate exact temporal flight boundary predictions.The Mars Organic Molecule Analyzer (MOMA) instrument is a dual source (pyrolysis gas chromatograph and laser desorption) mass spectrometer (MS) based package that detects and characterizes organic molecules, as part of ESA's 2020 ExoMars Rover mission to seek the signs of life on Mars.

Description: We present key features of the CGNS standard, focusing on its two main elements, the data model (CGNS/SIDS) and its implementations (CGNS/HDF5 and CGNS/Python). The data model is detailed to emphasize how the topological user oriented information, such as families, are separated from the actual meshing that could be split or modified during the CFD work flow, and how this topological information is traced during the meshing process. We also explain why the same information can be described in multiple ways and how to handle such alternatives in an application. Two implementations, using HDF5 and Python, are illustrated in several use examples, both for archival and interoperability purposes. The CPEX extension formalized process is explained to show how to add new features to the standard in a consensual way; we present some of the next extensions to come. Finally we conclude by showing how powerful a consensual public approach like CGNS can be, as opposed to a stand-alone private one. All throughout the paper, we demonstrate how the use of CGNS could be of great benefit for both the meshing and CFD solver communities.

Description: Performance tests of non-contacting finger seal designs were conducted at 300, 700, 922 K (70, 800 and 1200 F) at pressure differentials up to 517 kPa (75 psid) and surface speeds up to 366 m/s (1200 ft/s). Room temperature, static analysis of the seal was performed. A simplified CFD model was developed to examine pressure loads within the seal. Results from the CFD model were used as input to a finite element analysis model of a six-finger segment of the non-contacting finger seal. Examination of predicted deflections of individual components of the seal gives insight into the seal behavior. Wear patterns from testing verify the pattern of radial deflection. The models are used to predict maximum pressure differential capability of the seal and compared to experimental results. The CFD model slightly under-predicts the measured leakage flow factor, but has the same trend as the measured flow factor versus pressure differential.

Description: Electrically Driven Thermal Management is an active research and technology development initiative incorporating ISS technology flight demonstrations (STP-H5), development of Microgravity Science Glovebox (MSG) flight experiment, and laboratory-based investigations of electrically based thermal management techniques. The program targets integrated thermal management for future generations of RF electronics and power electronic devices. This presentation reviews four program elements: i.) results from the Electrohydrodynamic (EHD) Long Term Flight Demonstration launched in February 2017 ii.) development of the Electrically Driven Liquid Film Boiling Experiment iii.) two University based research efforts iv.) development of Oscillating Heat Pipe evaluation at Goddard Space Flight Center.

Description: It is known that ice nucleating particles (INP) immersed within supercooled droplets promote the formation of ice. Common theoretical models used to represent this process assume that the immersed particle lowers the work of ice nucleation without significantly affecting the dynamics of water in the vicinity of the particle. This is contrary to evidence showing that immersed surfaces significantly affect the viscosity and diffusivity of vicinal water. To study how this may affect ice formation this work introduces a model linking the ice nucleation rate to the modification of the dynamics and thermodynamics of vicinal water by immersed particles. It is shown that INP that significantly reduce the work of ice nucleation also pose strong limitations to the growth of the nascent ice germs. This leads to the onset of a new ice nucleation regime, called spinodal ice nucleation, where the dynamics of ice germ growth instead of the ice germ size determines the nucleation rate. Nucleation in this regime is characterized by an enhanced sensitivity to particle area and cooling rate. Comparison of the predicted ice nucleation rate against experimental measurements for a diverse set of species relevant to cloud formation suggests that spinodal ice nucleation may be common in nature.

Description: High-fidelity Computational Fluid Dynamics (CFD) simulations have been carried out for several multi-rotor Unmanned Aerial Vehicles (UAVs). Three vehicles have been studied: the classic quadcopter DJI Phantom 3, an unconventional quadcopter specialized for forward flight, the SUI Endurance, and an innovative concept for Urban Air Mobility (UAM), the Elytron 4S UAV. The three-dimensional unsteady Navier-Stokes equations are solved on overset grids using high-order accurate schemes, dual-time stepping, and a hybrid turbulence model. The DJI Phantom 3 is simulated with different rotors and with both a simplified airframe and the real airframe including landing gear and a camera. The effects of weather are studied for the DJI Phantom 3 quadcopter in hover. The SUI En- durance original design is compared in forward flight to a new configuration conceived by the authors, the hybrid configuration, which gives a large improvement in forward thrust. The Elytron 4S UAV is simulated in helicopter mode and in airplane mode. Understanding the complex flows in multi-rotor vehicles will help design quieter, safer, and more efficient future drones and UAM vehicles.

Description: Experimental measurements were performed on a swept flat-plate model with an airfoil leading edge and imposed chordwise pressure gradient to determine the effects of a backward-facing step on transition in a low-speed stationary crossflow-dominated boundary layer. Detailed hot-wire measurements were performed for three step heights ranging from 36 to 49% of the boundary-layer thickness at the step and corresponding to subcritical, nearly critical, and critical cases. In general, the step had a small localized effect on the growth of the stationary crossflow vortex, whereas the unsteady disturbance amplitudes increased with increasing step height. Intermittent spikes in instantaneous velocity began to appear for the two larger step heights. A physical explanation was provided for the mechanism leading to transition and the sudden movement in the transition front due to the critical steps. The large localized velocity spikes, which ultimately led to an intermittent breakdown of the boundary layer, were the result of nonlinear interactions of the different types of unsteady instabilities with each other and with the stationary crossflow vortices. Thus, the unsteady disturbances played the most important role in transition, but the stationary crossflow vortices also had a significant role via the modulation and the increased amplitude of the unsteady disturbances.

Description: A parametric experimental study was performed with sweeping jet actuators (fluidic oscillators) to determine their effectiveness in controlling flow separation on an adverse pressure gradient ramp. Actuator parameters that were investigated include blowing coefficients, operation mode, pitch and spreading angles, streamwise location, and size. Surface pressure measurements and surface oilflow visualization were used to characterize the effects of these parameters on the actuator performance. 2D Particle Image Velocimetry measurements of the flow field over the ramp and hot-wire measurements of the actuators jet flow were also obtained for selective cases. In addition, the sweeping jet actuators were compared to other well-known flow control techniques such as micro-vortex generators, steady blowing, and steady vortex-generating jets. The results confirm that the sweeping jet actuators are more effective than steady blowing and steady vortex-generating jets for this ramp configuration. The results also suggest that an actuator with a wider jet spreading (110 vs. 70 degrees) placed closer (2.3 vs. 7 boundary layer thickness upstream) to the flow separation location provides better performance. Different actuator sizes obtained by scaling down the actuator geometry produced different jet spreading. Scaling down the actuator (based on the throat dimensions) from 6.35 3.18 mm to 3.81 1.9 mm resulted in similar flow control performance; however, scaling down the actuator further to 1.9 0.95 mm reduced the actuator efficiency by reducing the jet spreading considerably. The results of this study provide insight that can be used to design and select the optimal sweeping jet actuator configuration for flow control applications.

Description: Supersonic boundary-layer receptivity to freestream acoustic disturbances is investigated by solving the NavierStokes equations for Mach 3.5 flow over a 7 deg half-angle cone. The freestream disturbances are generated from a wavy wall placed at the nozzle wall. The freestream acoustic disturbances radiated by the wavy wall are obtained by solving the linearized Euler equations. The results show that no noticeable instability modes are generated when the acoustic disturbances impinge the cone obliquely. The results show that the perturbations generated inside the boundary layer by the acoustic disturbances are the response of the boundary layer to the external forcing. The amplitude of the forced disturbances inside the boundary layer are about 2.5 times larger than the incoming field for zero azimuthal wave number, and they are about 1.5 times for large azimuthal wave numbers.

Description: A swept flat plate model with an imposed pressure gradient was experimentally investigated in a low-speed flow to determine the effect of a backward-facing step on transition in a stationary crossflowdominated flow. Detailed hotwire measurements of boundary-layer flow were performed to investigate the upstream shift in transition due to a step height of 49% of the local unperturbed boundary-layer thickness. Increasing the initial stationary crossflow amplitude caused an upstream movement of the transition front for the backward-facing step case. The step caused a local increase in the growth of the stationary crossflow instabilities, but the stationary crossflow amplitude at transition was sufficiently low (〈0.04U(sub e)) so that stationary crossflow was not solely responsible for transition. The unsteady velocity spectra downstream of the step were rich with unsteady disturbances in the 80- to 1500-Hz range. Three distinct families of disturbances were identified based on phase speed and wave angle, namely, a highly oblique disturbance (possibly traveling-crossflow-like), a TollmienSchlichting-wave-like disturbance, and a shear-layer instability. The stationary crossflow disturbances caused a modulation of the unsteady disturbances, resulting in spatially concentrated peaks in unsteady disturbance amplitude. This modulation of the unsteady disturbances is believed to be the reason for the upstream movement of the transition front with increasing stationary crossflow amplitude.

Description: The abilities of two different Reynolds-Averaged Navier-Stokes codes to predict the effects of an active flow control device are evaluated. The flow control device consists of a blowing slot located on the upper surface of an NACA 0018 airfoil, near the leading edge. A second blowing slot present on the airfoil near mid-chord is not evaluated here. Experimental results from a wind tunnel test show that a slot blowing with high momentum coefficient will increase the lift of the airfoil (compared to no blowing) and delay flow separation. A slot with low momentum coefficient will decrease the lift and induce separation even at low angles of attack. Two codes, CFL3D and FUN3D, are used in two-dimensional computations along with several different turbulence models. Two of these produced reasonable results for this flow, when run fully turbulent. A more advanced transition model failed to predict reasonable results, but warrants further study using different inputs. Including inviscid upper and lower tunnel walls in the simulations was found to be important in obtaining pressure distributions and lift coefficients that best matched experimental data. A limited number of three-dimensional computations were also performed.

Description: A wind tunnel test program has been conducted to define convective heating environments on the back-face of a Hypersonic Inflatable Aerodynamic Decelerator aeroshell. Wind tunnel testing was conducted at Mach 6 and Mach 10 at unit Reynolds numbers from 0.510(exp 6)/ft to 3.910(exp 6)/ft on a 6.3088 in diameter aeroshell model. Global heating data were obtained through phosphor thermography on the aeroshell back face, as well as on the payload and the aeroshell front face. For all test conditions, laminar flow was produced on the aeroshell front face, while the separated wake shear layer and aeroshell back-face boundary layer were transitional or turbulent. Along the leeward centerline of the aeroshell back face and payload centerbody, heating levels increased with both free stream Reynolds number and angle of attack. The Reynolds number dependency was due to increasing strength of wake turbulence with Reynolds number. The angle-of-attack dependency was due to movement of the wake-vortex reattachment point on the aeroshell back face. The maximum heating levels on the aeroshell back face and payload were approximately 5% to 6%, respectively, of the aeroshell front-face stagnation point. To allow for extrapolation of the ground test data to flight conditions, the back face and payload heating levels were correlated as a function of aeroshell front-face peak momentum thickness Reynolds numbers.

Description: In support of Topic A.2.8 of NASA NRA NNH10ZEA001N, the University of Florida (UF) has investigated the use of flow field optical diagnostic and micromachined sensor-based techniques for assessing the wall shear stress on an acoustic liner. Stereoscopic particle image velocimetry (sPIV) was used to study the velocity field over a liner in the Grazing Flow Impedance Duct (GFID). The results indicate that the use of a control volume based method to determine the wall shear stress is prone to significant error. The skin friction over the liner as measured using velocity curve fitting techniques was shown to be locally reduced behind an orifice, relative to the hard wall case in a streamwise plane centered on the orifice. The capacitive wall shear stress sensor exhibited a linear response for a range of shear stresses over a hard wall. PIV over the liner is consistent with lifting of the near wall turbulent structure as it passes over an orifice, followed by a region of low wall shear stress.

Description: Spontaneous rotational Raman scattering spectroscopy is used to acquire the first ever high quality, spatially-resolved measurements of the mean and root mean square (rms) temperature fluctuations in turbulent, high-velocity heated jets. Raman spectra in air were obtained across a matrix of radial and axial locations downstream from a 50 mm diameter nozzle operating from subsonic to supersonic conditions over a wide range of temperatures and Mach numbers, in accordance with the Tanna matrix frequently used in jet noise studies. These data were acquired in the hostile, high noise (115 dB) environment of a large scale open air test facility at NASA Glenn Research Center (GRC). Temperature estimates were determined by performing nonlinear least squares fitting of the single shot spectra to the theoretical rotational Stokes spectra of N2 and O2, using a custom in-house code developed specifically for this investigation. The laser employed in this study was a high energy, long-pulsed, frequency doubled Nd:YAG laser. One thousand single-shot spectra were acquired at each spatial coordinate. Mean temperature and rms temperature variations were calculated at each measurement location. Excellent agreement between the averaged and single-shot temperatures was observed with an accuracy better than 2.5 percent for temperature, and rms variations in temperature between +/-2.2 percent at 296 K and +/-4.5 percent at 850 K. The results of this and planned follow-on studies will support NASA GRC's development of physics-based jet noise prediction, turbulence modeling and aeroacoustic source modeling codes.

Description: Grid convergence studies are performed to establish reference solutions for benchmark three dimensional turbulent flows in support of the ongoing turbulence model verification and validation e ort at the Turbulence Modeling Resource website curated by NASA. The bench- mark cases are a subsonic flow around a hemisphere cylinder and a transonic flow around the ONERA M6 wing with a sharp trailing edge. The study applies widely-used computational fluid dynamics codes developed and supported at the NASA Langley Research Center: FUN3D, USM3D, and CFL3D. Reference steady-state solutions are computed for the Reynolds-Averaged Navier-Stokes equations with the Spalart-Allmaras turbulence model on families of consistently-refined grids composed of different types of cells. Coarse-to- ne and code-to-code solution variation is described in detail.

Description: Launch Ascent and Vehicle Aerodynamics (LAVA) (Computational Fluid Dynamics (CFD) solver) contribution to HiLiftPW-3 (3rd AIAA CFD High Lift Prediction Workshop). Including HL-CRM (NASA High Lift Common Research Model) and JSM (Joint Statistical Modeling) case studies.

Description: To investigate air radiation in expanding flows and provide experimental data for validating associated computational models, experiments were conducted in the X2 expansion tunnel facility at the Centre for Hypersonics of the University of Queensland. A 54deg turning angle wedge model was employed to generate steady expanding flows with in flow total enthalpies of 50.7, 63.4 and 75.4 MJ/kg. VUV spectra from 118 to 180 nm were acquired across the wedge at three equispaced distances away from the top of the model, as well as through its top surface. High speed filtered images were also obtained by coupling a Shimadzu 1 MHz high speed camera to a bandpass filter to obtain calibrated images of the 777 nm oxygen triplet. Both the across-wedge VUV spectra and filtered images of the 777 nm atomic oxygen were compared with NEQAIR simulations, which were performed using flow field data from two-dimensional CFD simulations with two-temperature 11-species air chemistry utilising the in-house Navier-Stokes flow solver Eilmer3. Data extracted from consecutive frames of the filtered high speed images confirmed up to 8 s of available test time for the flow conditions tested. For the strongly radiating 149 and 174 nm atomic nitrogen lines, large disagreement between experimental data and NEQAIR predictions can be observed from the start of the expansion fan where the electron-ion recombination process commences. The spatial extent, or spans of the radiance profiles of the 149 and 174 nm N lines are significantly underpredicted by NEQAIR, and are very close to those of N, N+ and electron number density profiles, which follow that of flow density. The electron-ion recombination process is proposed as the main reason for these discrepancies. The comparisons between NEQAIR simulations and filtered images of the 777 nm oxygen triplet show good agreement in the post-shock compression region and the start of the expansion fan for the 63.4 MJ/kg condition, but with up to a factor of three overprediction by NEQAIR further downstream, which is attributed to electron-impact excitation. Similar trends are found with the 75.4 MJ/kg condition, with reduced level of agreement in the compression region, which can be due to uncertainties in inflow condition.

Description: Velocity profiles are measured using molecular tagging velocimetry (MTV) in the high temperature test facility (HTTF) at Oregon State University during a depressurized conduction cooldown (DCC) event. The HTTF is a quarter scale electrically heated nuclear reactor simulator designed to replicate various accident scenarios. During a DCC, a double ended guillotine break results in the reactor pressure vessel (RPV) depressurizing into the reactor cavity and ultimately leading to air ingress in the reactor core (lock-exchange and gas diffusion). It is critical to understand the resulting buoyancy-driven flow to characterize the reactor self-cooling capacity through natural circulation. During tests conducted at ambient pressure and temperature, the RPV containing helium is opened (via the hot and cold legs) to a large vessel filled with nitrogen to simulate the atmosphere. The velocity profile on the hot leg pipe centerline is recorded at 10 Hz with MTV based on NO tracers. The precision of the velocimetry was measured to be 0.02 m/s in quiescent flow prior to the tests. A helium flow from the RPV is initially observed in the top quarter of the pipe. During the first 20 seconds of the event, helium flows out of the RPV with a maximum velocity below 2 m/s. The velocity profile transitions from parabolic to linear in character and decays slowly over the rest of the recording; peak velocities of 0.2 m/s are observed after 30 min. A counter-flow of nitrogen is also observed intermittently, which occurs at lower velocities (〉0.1 m/s).

Description: Water hammer analysis in pipe lines, in particularly during priming into evacuated lines is important for the design of spacecraft and other in-space application. In the current study, a finite volume network flow analysis code is used for modeling three different geometrical configurations: the first two being straight pipe, one with atmospheric air and other with evacuated line, and the third case is a representation of a complex flow network system. The numerical results show very good agreement qualitatively and quantitatively with measured data available in the literature. The peak pressure and impact time in case of straight pipe priming in evacuated line shows excellent agreement.

Description: A 4BMS-X (Four Bed Molecular Sieve - Exploration) design and heater optimization study for CO2 sorbent beds in proposed exploration system architectures is presented. The primary objectives of the study are to reduce heater power and thermal gradients within the CO2 sorbent beds while minimizing channeling effects. Some of the notable changes from the ISS (International Space Station) CDRA (Carbon Dioxide Removal Assembly) to the proposed exploration system architecture include cylindrical beds, alternate sorbents and an improved heater core. Results from both 2D and 3D sorbent bed thermal models with integrated heaters are presented. The 2D sorbent bed models are used to optimize heater power and fin geometry while the 3D models address end effects in the beds for more realistic thermal gradient and heater power predictions.

Description: Multilayer insulation (MLI) is considered the state of the art insulation for cryogenic propellant tanks in the space environment. MLI traditionally consists of multiple layers of metalized films separated by low conductivity spacers. In order to better understand some of the details within MLI design and construction, GRC has been investigating the heat loads caused by multiple types of seams. To date testing has been completed with 20 layer and 50 layer blankets. Although a truly seamless blanket is not practical, a blanket lay-up where each individual layer was overlapped and tapped together was used as a baseline for the other seams tests. Other seams concepts tested included: an overlap where the complete blanket was overlapped on top of itself; a butt joint were the blankets were just trimmed and butted up against each other, and a staggered butt joint where the seam in the out layers is offset from the seam in the inner layers. Measured performance is based on a preliminary analysis of rod calibration tests conducted prior to the start of seams testing. Baseline performance for the 50 layer blanket showed a measured heat load of 0.46 Watts with a degradation to about 0.47 Watts in the seamed blankets. Baseline performance for the 20 layer blanket showed a measured heat load of 0.57 Watts. Heat loads for the seamed tests are still begin analyzed. So far analysis work has suggested the need for corrections due to heat loads from both the heater leads and the instrumentation wires. A careful re-examination of the calibration test results with these factors accounted for is also underway. This presentation will discuss the theory of seams in MLI, our test results to date, and the uncertainties in our measurements.

Description: Due to the variety of requirements across aerospace platforms, and one off projects, the repeatability of cryogenic multilayer insulation has never been fully established. The objective of this test program is to provide a more basic understanding of the thermal performance repeatability of MLI systems that are applicable to large scale tanks. There are several different types of repeatability that can be accounted for: these include repeatability between multiple identical blankets, repeatability of installation of the same blanket, and repeatability of a test apparatus. The focus of the work in this report is on the first two types of repeatability. Statistically, repeatability can mean many different things. In simplest form, it refers to the range of performance that a population exhibits and the average of the population. However, as more and more identical components are made (i.e. the population of concern grows), the simple range morphs into a standard deviation from an average performance. Initial repeatability testing on MLI blankets has been completed at Florida State University. Repeatability of five GRC provided coupons with 25 layers was shown to be +/- 8.4 whereas repeatability of repeatedly installing a single coupon was shown to be +/- 8.0. A second group of 10 coupons have been fabricated by Yetispace and tested by Florida State University, through the first 4 tests, the repeatability has been shown to be +/- 16. Based on detailed statistical analysis, the data has been shown to be statistically significant.

Description: For large in-space cryogenic upper stages, substantial axial heat removal from a forward skirt by vapor-based heat interception may not be achieved by simple attachment methods unless sufficient thermal conductance from the skirt to the cooling fluid can be achieved. Preferable methods would allow for the addition of the cooling system to existing structure with minimal impact on the structure. Otherwise, significant modification to the basic structural design andor novel and complex attachment mechanisms with high effective thermal conductance are likely to be required. The approach being pursued by evolvable Cryogenics (eCryo) is to increase the thermal performance of a relatively simple attachment system by applying metallic or other thermally conductive material coatings to the mating surface area of the fluid channel where it is attached the skirt wall. The expectation of candidate materials is that the dramatic increase in conductivity of pure metals at temperatures close to liquid hydrogen vapor temperature will compensate for the reduced actual contact area typical of mechanical joints. Basic contact conductance data at low temperatures for candidate interface materials is required to enable the test approach. A test rig was designed at NASA Glenn Research Center to provide thermal contact resistance testing between small sample coupons coated with conductive material via electron beam evaporation, a low-temperature option that will not affect physical properties of base materials. Average coating thicknesses were 10 k. The test fixture was designed to mount directly to a cryocooler cold head within a vacuum test chamber. The purpose of this test was to determine qualitative contact conductance between various test samples. Results from this effort will be implemented in a sub-scale vapor-based heat interception test, where the applicability for increased heat removal on large structural skirts will be considered.

Description: The main penetrations (supports and piping) through multilayer insulation systems for cryogenic tanks have been previously addressed by heat flow measurements. Smaller penetrations due to fasteners and attachments are now experimentally investigated. The use of small pins or plastic garment tag fasteners to each the handling and construction of multilayer insulation (MLI) blankets goes back many years. While it has long been understood that penetrations and other discontinuities degrade the performance of the MLI blanket, quantification of this degradation has generally been lumped into gross performance multipliers (often called degradation factors or scale factors). Small penetrations contribute both solid conduction and radiation heat transfer paths through the blanket. The conduction is down the stem of the structural element itself while the radiation is through the hole formed during installation of the pin or fastener. Analytical models were developed in conjunction with MLI perforation theory and Fouriers Law. Results of the analytical models are compared to experimental testing performed on a 10 layer MLI blanket with approximately 50 small plastic pins penetrating the test specimen. The pins were installed at 76-mm spacing inches in both directions to minimize the compounding of thermal effects due to localized compression or lateral heat transfer. The testing was performed using a liquid nitrogen boil-off calorimeter (Cryostat-100) with the standard boundary temperatures of 293 K and 78 K. Results show that the added radiation through the holes is much more significant than the conduction down the fastener. The results are shown to be in agreement with radiation theory for perforated films.

Description: The need to demonstrate and evaluate the effectiveness of heat interception methods for use on a relevant cryogenic propulsion stage at a system level has been identified. Evolvable Cryogenics (eCryo) Structural Heat Intercept, Insulation and Vibration Evaluation Rig (SHIIVER) will be designed with vehicle specific geometries (SLS Exploration Upper Stage (EUS) as guidance) and will be subjected to simulated space environments. One method of reducing structure-born heat leak being investigated utilizes vapor-based heat interception. Vapor-based heat interception could potentially reduce heat leak into liquid hydrogen propulsion tanks, increasing potential mission length or payload capability. Due to the high number of unknowns associated with the heat transfer mechanism and integration of vapor-based heat interception on a realistic large-scale skirt design, a sub-scale investigation was developed. The sub-project effort is known as the Small-scale Laboratory Investigation of Cooling Enhancements (SLICE). The SLICE aims to study, design, and test sub-scale multiple attachments and flow configuration concepts for vapor-based heat interception of structural skirts. SLICE will focus on understanding the efficiency of the heat transfer mechanism to the boil-off hydrogen vapor by varying the fluid network designs and configurations. Various analyses were completed in MATLAB, Excel VBA, and COMSOL Multiphysics to understand the optimum flow pattern for heat transfer and fluid dynamics. Results from these analyses were used to design and fabricate test article subsections of a large forward skirt with vapor cooling applied. The SLICE testing is currently being performed to collect thermal mechanical performance data on multiple skirt heat removal designs while varying inlet vapor conditions necessary to intercept a specified amount of heat for a given system. Initial results suggest that applying vapor-cooling provides a 50 heat reduction in conductive heat transmission along the skirt to the tank. The information obtained by SLICE will be used by the SHIIVER engineering team to design and implement vapor-based heat removal technology into the SHIIVER forward skirt hardware design.

Description: Previously we identified the receding flow, where two fluid streams recede from each other, as an open numerical problem, because all well-known numerical fluxes give an anomalous temperature rise, thus called the overheating problem. This phenomenon, although presented in several textbooks, and many previous publications, has scarcely been satisfactorily addressed and the root cause of the overheating problem not well understood. We found that this temperature rise was solely connected to entropy rise and proposed to use the method of characteristics to eradicate the problem. However, the root cause of the entropy production was still unclear. In the present study, we identify the cause of this problem: the entropy rise is rooted in the pressure flux in a finite volume formulation and is implanted at the first time step. It is found theoretically inevitable for all existing numerical flux schemes used in the finite volume setting, as confirmed by numerical tests. This difficulty cannot be eliminated by manipulating time step, grid size, spatial accuracy, etc, although the rate of overheating depends on the flux scheme used. Finally, we incorporate the entropy transport equation, in place of the energy equation, to ensure preservation of entropy, thus correcting this temperature anomaly. Its applicability is demonstrated for some relevant 1D and 2D problems. Thus, the present study validates that the entropy generated ab initio is the genesis of the overheating problem.

Description: A computational study of the Adaptive Deployable Entry and Placement Technology (ADEPT) Sounding Rocket (SR-1) Test is presented using the US3D flow solver. ADEPT SR-1 is intended, in part, to assess the dynamic stability of this entry vehicle architecture. Given that no dynamic stability data exists for the ADEPT geometry, a limited ballistic range campaign has been performed to characterize the vehicle's stability characteristics pre-flight for Mach numbers between 1.21 and 2.5. Here, this data is used to assess the accuracy of US3D's free-flight CFD capability. Computed trajectories from US3D and experimental data show that the flow solver compares well in vehicle oscillation frequency, downrange distance, and oscillatory amplitude during high Mach number flight (Mavg = 2.36). For Mach numbers below 1.5, the solver under predicts total angle-of-attack by an average of 16%, but compares well in oscillatory frequency and downrange distance. Additionally, a capability for simulating the trajectory of the flight article through the atmosphere using CFD is presented. This capability couples US3D's free-flight capability to an atmosphere model that accounts for changes in free-stream density and temperature as the vehicle descends. Two simulations for the purpose of demonstrating the capability and viability of this approach are applied to SR-1 flight article, and some unique challenges are discussed.

Description: Abstract: NASA is currently investigating methods to reduce the boil-off rate on large cryogenic upper stages. Two such methods to reduce the total heat load on existing upper stages are vapor cooling of the cryogenic tank support structure and integration of thick multilayer insulation systems to the upper stage of a launch vehicle. Previous efforts have flown a 2-layer MLI blanket and shown an improved thermal performance, and other efforts have ground-tested blankets up to 70 layers thick on tanks with diameters between 2 3 meters. However, thick multilayer insulation installation and testing in both thermal and structural modes has not been completed on a large scale tank. Similarly, multiple vapor cooled shields are common place on science payload helium dewars; however, minimal effort has gone into using boil-off to intercept heat on large structural surfaces associated with rocket stages. A majority of the vapor cooling effort focuses on metallic cylinders called skirts, which are the most common structural components for launch vehicles. In order to provide test data for comparison with analytical models, a 4 meter diameter test tank and assembly is currently being designed to include skirt structural systems with integral vapor cooling. In order to give a more representative result for actual applications, the technologies will be designed for an 8.4 m and scaled down to a 4 meter tank. In order to prove that the insulation system and vapor cooling attachment methods are structurally sound, acoustic testing will also be performed on the system. The test tank with insulation and vapor cooled shield installed will be tested thermally in the B2 test facility at NASAs Plum Brook Station both before and after being vibration tested at Plum Brooks Space Power Facility.

Description: Abstract: This paper covers cryogenic, tensile testing and research completed on a number of epoxies used in cryogenic applications. Epoxies are used in many different applications; however, this research focused on the use of epoxy used to bond MLI standoffs to cryogenic storage tanks and the loads imparted to the tank through the MLI. To conduct testing, samples were made from bare stainless steel, aluminum and primed aluminum. Testing involved slowly cooling test samples with liquid nitrogen then applying gradually increasing tensile loads to the epoxy. The testing evaluated the strength and durability of epoxies at cryogenic temperatures and serves as a base for future testing. The results of the tests showed that some epoxies withstood the harsh conditions while others failed. The two epoxies yielding the best results were Masterbond EP29LPSP and Scotch Weld 2216. For all metal surfaces tested, both epoxies had zero failures for up to 11.81 kg of mass.

Description: High-powered lasers were used to induce ablation and to form fusion crusts in the lab on Tamdakht H5 chondrites and basalt. These ground tests were undertaken to improve our understanding, and ultimately improve our abilty to model and predict, meteoroid ablation during atmospheric entry. The infrared fiber laser at the LHMEL facilty, operated in the continuous wave (i.e. non-pulsed) mode, provided radiation surface heat flux at levels similar to meteor entry for these tests. Results are presented from the first round of testing on samples of Tamdakht H5 ordinary chondrite which were ex-posed to entry-relevant heating rates between 2 and 10 kWcm2.

Description: This paper covers cryogenic, tensile testing and research completed on a number of epoxies used in cryogenic applications. Epoxies are used in many different applications; however, this research focused on the use of epoxy used to bond MLI standoffs to cryogenic storage tanks and the loads imparted to the tank through the MLI. To conduct testing, samples were made from bare stainless steel, aluminum and primed aluminum. Testing involved slowly cooling test samples with liquid nitrogen then applying gradually increasing tensile loads to the epoxy. The testing evaluated the strength and durability of epoxies at cryogenic temperatures and serves as a base for future testing. The results of the tests showed that some epoxies withstood the harsh conditions while others failed. The two epoxies yielding the best results were Masterbond EP29LPSP and Scotch Weld 2216. For all metal surfaces tested, both epoxies had zero failures for up to 11.81 kg of mass..

Description: NASAs current Mars architectures are assuming the production and storage of 23 tons of liquid oxygen on the surface of Mars over a duration of 500+ days. In order to do this in a mass efficient manner, an energy efficient refrigeration system will be required. Based on previous analysis NASA has decided to do all liquefaction in the propulsion vehicle storage tanks. In order to allow for transient Martian environmental effects, a propellant liquefaction and storage system for a Mars Ascent Vehicle (MAV) was modeled using Thermal Desktop. The model consisted of a propellant tank containing a broad area cooling loop heat exchanger integrated with a reverse turbo Brayton cryocooler. Cryocooler sizing and performance modeling was conducted using MAV diurnal heat loads and radiator rejection temperatures predicted from a previous thermal model of the MAV. A system was also sized and modeled using an alternative heat rejection system that relies on a forced convection heat exchanger. Cryocooler mass, input power, and heat rejection for both systems were estimated and compared against sizing based on non-transient sizing estimates.

Description: Detailed spectrally and spatially resolved radiance has been measured in the Electric Arc Shock Tube for conditions relevant to Titan entry, varying atmospheric composition, free-stream density (equivalent to altitude) and shock velocity. Permutations in atmospheric composition include 1.1, 2, 5 and 8.6 CH4 by mole with a balance of N2 and 1.5 CH4 0.5 Ar 98 N2 by mole, which is consistent with the current understanding of Titan's atmosphere. The effect of gas impurities identified in previous shock tube studies were also examined by testing in pure N2 and deliberate addition of air to the CH4N2 mixtures. The test campaign measured radiation at velocities from 4.7 kms to 8 kms and free-stream pressures from 0.1 to 0.47 Torr. These conditions cover a range of potential trajectories for flight missions, including a direct ballistic trajectory, a fly by or an extremely high speed entry. Radiances measured in this work are substantially larger compared to that reported both in past EAST test campaigns and other shock tube facilities. Depending on the metric used for comparison, the discrepancy can be as high as an order of magnitude. Potential causes for the discrepancy, such as the effect of oxygen due to Air leakage, gas composition and purity are discussed. The present work provides a new benchmark set of data to replace those published in previous studies.

Description: To investigate air radiation in expanding flows and provide experimental data for validating associated computational models, experiments were conducted in the X2 expansion tunnel facility at the Centre for Hypersonics of the University of Queensland. A 54 turning angle wedge model was employed to generate steady expanding flows with in flow total enthalpies of 50.7, 63.4 and 75.4 MJkg. VUV spectra from 118 to 180 nm were acquired across the wedge at three equispaced distances away from the top of the model, as well as through its top surface. High speed filtered images were also obtained by coupling a Shimadzu 1 MHz high speed camera to a bandpass filter to obtain calibrated images of the 777 nm oxygen triplet. Both the across-wedge VUV spectra and filtered images of the 777 nm atomic oxygen were compared with NEQAIR simulations, which were performed using flow field data from two-dimensional CFD simulations with two-temperature 11-species air chemistry utilizing the in-house Navier-Stokes flow solver Eilmer3. Data extracted from consecutive frames of the filtered high speed images confirmed up to 8 s of available test time for the flow conditions tested. For the strongly radiating 149 and 174 nm atomic nitrogen lines, large disagreement between experimental data and NEQAIR predictions can be observed from the start of the expansion fan where the electron-ion recombination process commences. The spatial extent, or spans of the radiance profiles of the 149 and 174 nm N lines are significantly under predicted by NEQAIR, and are very close to those of N, N+ and electron number density profiles, which follow that of flow density. The electron-ion recombination process is proposed as the main reason for these discrepancies. The comparisons between NEQAIR simulations and filtered images of the 777 nm oxygen triplet show good agreement in the post-shock compression region and the start of the expansion fan for the 63.4 MJkg condition, but with up to a factor of three over prediction by NEQAIR further downstream, which is attributed to electron-impact excitation. Similar trends are found with the 75.4 MJkg condition, with reduced level of agreement in the compression region, which can be due to uncertainties in inflow condition.

Description: NASAs current Mars architectures are assuming the production and storage of 23 tons of liquid oxygen on the surface of Mars over a duration of 500+ days. In order to do this in a mass efficient manner, an energy efficient refrigeration system will be required. Based on previous analysis NASA has decided to do all liquefaction in the propulsion vehicle storage tanks. In order to allow for transient Martian environmental effects, a propellant liquefaction and storage system for a Mars Ascent Vehicle (MAV) was modeled using Thermal Desktop. The model consisted of a propellant tank containing a broad area cooling loop heat exchanger integrated with a reverse turbo Brayton cryocooler. Cryocooler sizing and performance modeling was conducted using MAV diurnal heat loads and radiator rejection temperatures predicted from a previous thermal model of the MAV. A system was also sized and modeled using an alternative heat rejection system that relies on a forced convection heat exchanger. Cryocooler mass, input power, and heat rejection for both systems were estimated and compared against sizing based on non-transient sizing estimates.

Description: Large-eddy simulations were used to investigate turbulent temperature fluctuations and turbulent heat flux in hot jets. A high-resolution finite-difference Navier-Stokes solver, WRLES, was used to compute the flow from a 2-inch round nozzle. Several different flow conditions, consisting of different jet Mach numbers and temperature ratios, were examined. Predictions of mean and fluctuating velocities were compared to previously obtained particle image velocimetry data. Predictions of mean and fluctuating temperature were compared to new data obtained using Raman spectroscopy. Based on the good agreement with experimental data for the individual quantities, the combined quantity turbulent heat flux was examined.

Description: The Reynolds-stress and triple product Lag models were created with a normal stress distribution which was defined by the accepted 4:3:2 distribution of streamwise, spanwise and wall normal stresses, and a ratio of (sub w) = 0.3k in the log layer region of high Reynolds number flat plate flow, which implies R11(+) = 4/(9/2).3 2.96. More recent measurements show a more complex picture of the log layer region at high Reynolds numbers. The first cut at improving these models along with the direction for future refinements is described. Comparison with recent high Reynolds number data shows areas where further work is needed, but also shows inclusion of the modeled turbulent transport terms improve the prediction where they influence the solution. Additional work is needed to develop a model that better matches experiments, but there is significant improvement in many of the details of the log layer behavior.

Description: The Reynolds-stress and triple product Lag models were created with a normal stress distribution which was defined by the accepted 4:3:2 distribution of streamwise, spanwise and wall normal stresses, and a ratio of (sub w) = 0.3k in the log layer region of high Reynolds number flat plate flow, which implies R11(+) = 4/(9/2).3 2.96. More recent measurements show a more complex picture of the log layer region at high Reynolds numbers. The first cut at improving these models along with the direction for future refinements is described. Comparison with recent high Reynolds number data shows areas where further work is needed, but also shows inclusion of the modeled turbulent transport terms improve the prediction where they influence the solution. Additional work is needed to develop a model that better matches experiments, but there is significant improvement in many of the details of the log layer behavior.

Description: A flow conditioning device for incrementally stepping down pressure within a piping system is presented. The invention includes an outer annular housing, a center element, and at least one intermediate annular element. The outer annular housing includes an inlet end attachable to an inlet pipe and an outlet end attachable to an outlet pipe. The outer annular housing and the intermediate annular element(s) are concentrically disposed about the center element. The intermediate annular element(s) separates an axial flow within the outer annular housing into at least two axial flow paths. Each axial flow path includes at least two annular extensions that alternately and locally direct the axial flow radially outward and inward or radially inward and outward thereby inducing a pressure loss or a pressure gradient within the axial flow. The pressure within the axial flow paths is lower than the pressure at the inlet end and greater than the vapor pressure for the axial flow. The invention minimizes fluidic instabilities, pressure pulses, vortex formation and shedding, and/or cavitation during pressure step down to yield a stabilized flow within a piping system.

Description: One of the challenges of designing and flying a scramjet-powered vehicle is the difficulty of preflight testing. Ground tests at realistic flight conditions introduce several sources of uncertainty to the flow that must be addressed. For example, the scales of the available facilities limit the size of vehicles that can be tested and so performance metrics for larger flight vehicles must be extrapolated from ground tests at smaller scales. To create the correct flow enthalpy for higher Mach number flows, most tunnels use a heater that introduces vitiates into the flow. At these conditions, the effects of the vitiates on the combustion process is of particular interest to the engine designer, where the ground test results must be extrapolated to flight conditions. In this paper, the uncertainty of the cracked JP-7 chemical kinetics used in the modeling of a hydrocarbon-fueled scramjet was investigated. The factors that were identified as contributing to uncertainty in the combustion process were the level of flow vitiation, the uncertainty of the kinetic model coefficients and the variation of flow properties between ground testing and flight. The method employed was to run simulations of small, unit problems and identify which variables were the principal sources of uncertainty for the mixture temperature. Then using this resulting subset of all the variables, the effects of the uncertainty caused by the chemical kinetics on a representative scramjet flow-path for both vitiated (ground) and nonvitiated (flight) flows were investigated. The simulations showed that only a few of the kinetic rate equations contribute to the uncertainty in the unit problem results, and when applied to the representative scramjet flowpath, the resulting temperature variability was on the order of 100 K. Both the vitiated and clean air results showed very similar levels of uncertainty, and the difference between the mean properties were generally within the range of uncertainty predicted.

Description: The current work compares experimentally and computationally obtained nitric oxide (NO) planar laser induced fluorescence (PLIF) images of the mixing flowfields for three types of high-speed fuel injectors: a strut, a ramp, and a rectangular flushwall. These injection devices, which exhibited promising mixing performance at lower flight Mach numbers, are currently being studied as a part of the Enhanced Injection and Mixing Project (EIMP) at the NASA Langley Research Center. The EIMP aims to investigate scramjet fuel injection and mixing physics, and improve the understanding of underlying physical processes relevant to flight Mach numbers greater than eight. In the experiments, conducted in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF), the injectors are placed downstream of a Mach 6 facility nozzle, which simulates the high Mach number air flow at the entrance of a scramjet combustor. Helium is used as an inert substitute for hydrogen fuel. Both schlieren and PLIF techniques are applied to obtain mixing flowfield flow visualizations. The experimental PLIF is obtained by using a UV laser sheet to interrogate a plane of the flow by exciting fluorescence from the NO molecules, which are present in the AHSTF air. Consequently, the absence of signal in the resulting PLIF images is an indication of pure helium (fuel). The computational PLIF is obtained by applying a fluorescence model for NO to the results of the Reynolds-averaged simulations (RAS) of the mixing flow field carried out using the VULCAN-CFD solver. This approach is required because the PLIF signal is a nonlinear function of not only NO concentration, but also pressure, temperature, and the flow velocity. This complexity allows additional flow features to be identified and compared with those obtained from the computational fluid dynamics (CFD) simulations, however, such comparisons are only semiquantitative. Three-dimensional image reconstruction, similar to that used in magnetic resonance imaging, is also used to obtain images in the streamwise and spanwise planes from select cross-stream PLIF plane data. Synthetic schlieren is also computed from the RAS data. Good agreement between the experimental and computational results provides increased confidence in the CFD simulations for investigations of injector performance.

Description: NASA Mars Science Laboratory (MSL), which landed the Curiosity rover on the surface of Mars on August 5th, 2012, was the largest and heaviest Mars entry vehicle representing a significant advancement in planetary entry, descent and landing capability. Hypersonic flight performance data was collected using MSLs on-board sensors called Mars Entry, Descent and Landing Instrumentation (MEDLI). This talk will give an overview of MSL entry and a description of MEDLI sensors. Observations from flight data will be examined followed by a discussion of analysis efforts to reconstruct surface heating from heatshields in-depth temperature measurements. Finally, a brief overview of MEDLI2 instrumentation, which will fly on NASAs Mars2020 mission, will be presented with a discussion on how lessons learned from MEDLI data affected the design of MEDLI2 instrumentation.

Description: This paper covers cryogenic, tensile testing and research completed on a number of epoxies used in cryogenic applications. Epoxies are used in many different applications; however, this research focused on the use of epoxy used to bond MLI standoffs to cryogenic storage tanks and the loads imparted to the tank through the MLI. To conduct testing, samples were made from bare stainless steel, aluminum and primed aluminum. Testing involved slowly cooling test samples with liquid nitrogen then applying gradually increasing tensile loads to the epoxy. The testing evaluated the strength and durability of epoxies at cryogenic temperatures and serves as a base for future testing. The results of the tests showed that some epoxies withstood the harsh conditions while others failed. The two epoxies yielding the best results were Masterbond EP29LPSP and Scotch Weld 2216. For all metal surfaces tested, both epoxies had zero failures for up to 11.81 kg of mass.

Description: No abstract available

Description: We report upon the development and testing of a 4-stage adiabatic demagnetization refrigerator (ADR) capable of continuous cooling at 0.100 Kelvin. This cooler is being built to cool the detector array aboard NASA's Primordial Inflation Polarization Explorer (PIPER) observatory. The goal of this balloon mission is to measure the primordial gravitational waves that should exist if the theory of cosmological inflation is correct. At altitude, the ADR will hold the array of transition-edge sensors at 100 mK continuously while periodically rejecting heat to a 1.2 K pumped helium bath. During testing on ground, the array is held at the same temperature but heat is rejected to a 4.2 K helium bath indicating the flexibility in this coolers design.

Description: A fluidic oscillator having independent frequency and amplitude control includes a fluidic-oscillator main flow channel having a main flow inlet, a main flow outlet, and first and second control ports disposed at opposing sides thereof. A fluidic-oscillator controller has an inlet and outlet. A volume defined by the main flow channel is greater than the volume defined by the controller. A flow diverter coupled to the outlet of the controller defines a first fluid flow path from the controller's outlet to the first control port and defines a second fluid flow path from the controller's outlet to the second control port.

Description: One or more embodiments of techniques or systems for shaped recess flow control are provided herein. A shaped recess or cavity can be formed on a surface associated with fluid flow. The shaped recess can be configured to create or induce fluid effects, temperature effects, or shedding effects that interact with a free stream or other structures. The shaped recess can be formed at an angle to a free stream flow and may be substantially "V" shaped. The shaped recess can be coupled with a cooling channel, for example. The shaped recess can be upstream or downstream from a cooling channel and aligned in a variety of manners. Due to the fluid effects, shedding effects, and temperature effects created by a shaped recess, lift-off or separation of cooling jets of cooling channels can be mitigated, thereby enhancing film cooling effectiveness.

Description: The present invention is a system and method of visualizing fluid flow around an object, such as an aircraft or wind turbine, by aligning the object between an imaging system and a celestial object having a speckled background, taking images, and comparing those images to obtain fluid flow visualization.

Description: A system and method is provided for traversing inside one or more pipes. In an embodiment, a fluid is injected into the one or more pipes thereby promoting a fluid flow. An inspection device is deployed into the one or more pipes at least partially filled with a flowing fluid. The inspection device comprises a housing wherein the housing is designed to exploit the hydrokinetic effects associated with a fluid flow in one or more pipes as well as maneuver past a variety of pipe configurations. The inspection device may contain one or more sensors capable of performing a variety of inspection tasks.

Description: Unstructured grid solvers have well-known issues predicting surface heat fluxes when strong shocks are present. Various efforts have been made to address the underlying numerical issues that cause the erroneous predictions. The present work addresses some of the shortcomings of unstructured grid solvers, not by addressing the numerics, but by applying structured grid best practices to unstructured grids. A methodology for robust shock detection and shock fitting is outlined and applied to production relevant cases. Results achieved by using the Loci-CHEM Computational Fluid Dynamics solver are provided.

Description: Stereo particle image velocimetry measurements were performed downstream of a backward-facing step in a stationary-cross flow dominated flow. The PIV measurements exhibit excellent quantitative and qualitative agreement with the previously acquired hotwire data. Instantaneous PIV snapshots reveal new information about the nature and cause of the \spikes" that occurred prior to breakdown in both the hotwire and PIV data. The PIV snapshots show that the events occur simultaneously across multiple stationary cross flow wavelengths, indicating that this is not simply a local event, but is likely caused by the 2D Tollmien-Schlichting instability that is introduced by the step. While the TS instability is a 2D instability, it is also modulated in the spanwise direction due to interactions with the stationary cross flow, as are the other unsteady disturbances present. Because of this modulation, the "spike" events cause an instantaneous increase of the spanwise modulation of the streamwise and spanwise velocity initially caused by the stationary cross flow. Breakdown appears to be caused by this instantaneous modulation, possibly due to a high-frequency secondary instability similar to a traveling-cross flow breakdown scenario. These results further illuminate the respective roles of the stationary cross flow and unsteady disturbances in transition downstream of a backward-facing step.

Description: Detailed spectrally and spatially resolved radiance has been measured in the Electric Arc Shock Tube (EAST) facility for conditions relevant to high speed entry into a variety of atmospheres, including Earth, Venus, Titan, Mars and the Outer Planets. The tests that measured radiation relevant for Earth re-entry are the focus of this work and are taken from campaigns 47, 50, 52 and 57. These tests covered conditions from 8 km/s to 15.5 km/s at initial pressures ranging from 0.05 Torr to 1 Torr, of which shots at 0.1 and 0.2 Torr are analyzed in this paper. These conditions cover a range of points of interest for potential fight missions, including return from Low Earth Orbit, the Moon and Mars. The large volume of testing available from EAST is useful for statistical analysis of radiation data, but is problematic for identifying representative experiments for performing detailed analysis. Therefore, the intent of this paper is to select a subset of benchmark test data that can be considered for further detailed study. These benchmark shots are intended to provide more accessible data sets for future code validation studies and facility-to-facility comparisons. The shots that have been selected as benchmark data are the ones in closest agreement to a line of best fit through all of the EAST results, whilst also showing the best experimental characteristics, such as test time and convergence to equilibrium. The EAST data are presented in different formats for analysis. These data include the spectral radiance at equilibrium, the spatial dependence of radiance over defined wavelength ranges and the mean non-equilibrium spectral radiance (so-called 'spectral non-equilibrium metric'). All the information needed to simulate each experimental trace, including free-stream conditions, shock time of arrival (i.e. x-t) relation, and the spectral and spatial resolution functions, are provided.

Description: Unstructured grid solvers have well-known issues predicting surface heat fluxes when strong shocks are present. Various efforts have been made to address the underlying numerical issues that cause the erroneous predictions. The present work addresses some of the shortcomings of unstructured grid solvers, not by addressing the numerics, but by applying structured grid best practices to unstructured grids. A methodology for robust shock detection and shock-fitting is outlined and applied to production-relevant cases. Results

Description: A preliminary verification and validation of a new material response model is presented. This model, Icarus, is intended to serve as a design tool for the thermal protection systems of re-entry vehicles. Currently, the capability of the model is limited to simulating the pyrolysis of a material as a result of the radiative and convective surface heating imposed on the material from the surrounding high enthalpy gas. Since the major focus behind the development of Icarus has been model extensibility, the hope is that additional physics can be quickly added. The extensibility is critical since thermal protection systems are becoming increasing complex, e.g. woven carbon polymers. Additionally, as a three-dimensional, unstructured, finite-volume model, Icarus is capable of modeling complex geometries as well as multi-dimensional physics, which have been shown to be important in some scenarios and are not captured by one-dimensional models. In this paper, the mathematical and numerical formulation is presented followed by a discussion of the software architecture and some preliminary verification and validation studies.

Description: A numerical model is being developed using Python which characterizes the conversion and temperature profiles of a packed bed reactor (PBR) that utilizes the Sabatier process; the reaction produces methane and water from carbon dioxide and hydrogen. While the specific kinetics of the Sabatier reaction on the RuAl2O3 catalyst pellets are unknown, an empirical reaction rate equation1 is used for the overall reaction. As this reaction is highly exothermic, proper thermal control is of the utmost importance to ensure maximum conversion and to avoid reactor runaway. It is therefore necessary to determine what wall temperature profile will ensure safe and efficient operation of the reactor. This wall temperature will be maintained by active thermal controls on the outer surface of the reactor. Two cylindrical PBRs are currently being tested experimentally and will be used for validation of the Python model. They are similar in design except one of them is larger and incorporates a preheat loop by feeding the reactant gas through a pipe along the center of the catalyst bed. The further complexity of adding a preheat pipe to the model to mimic the larger reactor is yet to be implemented and validated; preliminary validation is done using the smaller PBR with no reactant preheating. When mapping experimental values of the wall temperature from the smaller PBR into the Python model, a good approximation of the total conversion and temperature profile has been achieved. A separate CFD model incorporates more complex three-dimensional effects by including the solid catalyst pellets within the domain. The goal is to improve the Python model to the point where the results of other reactor geometry can be reasonably predicted relatively quickly when compared to the much more computationally expensive CFD approach. Once a reactor size is narrowed down using the Python approach, CFD will be used to generate a more thorough prediction of the reactors performance.

Description: An effort was undertaken to perform CFD analysis of fluid flow in Lean-Direct Injection (LDI) combustors with axial swirl-venturi elements for next-generation LDI-3 combustor design. The National Combustion Code (NCC) was used to perform non-reacting and two-phase reacting flow computations for a nineteen-element injector array arranged in a three-module, 7-5-7 element configuration. All computations were performed with a consistent approach of mesh-optimization, spray-modeling, ignition and kinetics-modeling with the NCC. Computational predictions of the aerodynamics of the injector were used to arrive at an optimal injector design that meets effective area and fuel-air mixing criteria. LDI-3 emissions (EINOx, EICO and UHC) were compared with the previous generation LDI-2 combustor experimental data at representative engine cycle conditions.

Description: The current work compares experimentally and computationally obtained nitric oxide (NO) planar laser-induced fluorescence (PLIF) images of the mixing flowfields for three types of high-speed fuel injectors: a strut, a ramp, and a rectangular flush-wall. These injection devices, which exhibited promising mixing performance at lower flight Mach numbers, are currently being studied as a part of the Enhanced Injection and Mixing Project (EIMP) at the NASA Langley Research Center. The EIMP aims to investigate scramjet fuel injection and mixing physics, and improve the understanding of underlying physical processes relevant to flight Mach numbers greater than eight. In the experiments, conducted in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF), the injectors are placed downstream of a Mach 6 facility nozzle, which simulates the high Mach number air flow at the entrance of a scramjet combustor. Helium is used as an inert substitute for hydrogen fuel. The PLIF is obtained by using a tunable laser to excite the NO, which is present in the AHSTF air as a direct result of arc-heating. Consequently, the absence of signal is an indication of pure helium (fuel). The PLIF images computed from the computational fluid dynamics (CFD) simulations are obtained by combining a fluorescence model for NO with the Reynolds-Averaged Simulation results carried out using the VULCAN-CFD solver to obtain a computational equivalent of the experimentally measured PLIF signal. The measured NO PLIF signal is mainly a function of NO concentration allowing for semi-quantitative comparisons between the CFD and the experiments. The PLIF signal intensity is also sensitive to pressure and temperature variations in the flow, allowing additional flow features to be identified and compared with the CFD. Good agreement between the PLIF and the CFD results provides increased confidence in the CFD simulations for investigations of injector performance.

Description: A loop heat pipe (LHP) is a very versatile heat transfer device that can transport a large heat load over a long distance with a small temperature difference. All LHPs currently servicing orbiting spacecraft are designed to operate in the room temperature range. Future space telescopes and space-based Earth resource imaging satellites require passive cryogenic heat transport devices that can thermally couple remote cryocoolers to sensor or instrument of interest while providing the capability of payload vibration jitter isolation, implementation of redundant coolers, and coupling of multiple sensors to a common heat sink. All of these requirements can be satisfied by using a cryogenic LHP (CLHP). Although the development of CLHPs faces several technical challenges, NASA Goddard Space Flight Center has devoted extensive efforts in developing CLHP technology over the past decade and has made significant progress. In particular, the combination of the innovative ideas of using a secondary capillary pump to manage the parasitic heat gain and using a hot reservoir to reduce the system pressure under the ambient condition has led to the successful development of the CLHP. Several CLHPs charged with nitrogen and hydrogen were built and tested in thermal vacuum chambers. These CLHPs demonstrated reliable start-up and robust operation during power cycle and sink temperature cycle tests.

Description: AZ-2000-IECW white paint and StaMet black Kapton have been evaluated for the Kevlar/polyimide shells that enclose the Solar Probe Plus Magnetometer (MAG) sensors and multilayer insulation. Flight qualification testing on AZ-2000-IECW painted Kevlar/polyimide laminate was completed at Goddard Space Flight Center. This paint potentially meets all the requirements. However, it has no flight heritage. StaMet is hotter in the sun, and is specular. The results of the MAG thermal balance test show StaMet meets the thermal requirement and heater power budget. The mission prefers to fly StaMet after evaluating the risks of AZ-2000-IECW flaking and glint from StaMet to the Star Trackers.

Description: Normal gravity flame blowoff limits in an axisymmetric pmma rod geometry in upward axial stagnation flow are compared with microgravity Burning and Suppression of Solids II (BASS-II) results recently obtained aboard the International Space Station. This testing utilized the same BASS-II concurrent rod geometry, but with the addition of normal gravity buoyant flow. Cast polymethylmethacrylate (pmma) rods of diameters ranging from 0.635 cm to 3.81 cm were burned at oxygen concentrations ranging from 14 to 18 by volume. The forced flow velocity where blowoff occurred was determined for each rod size and oxygen concentration. These blowoff limits compare favorably with the BASS-II results when the buoyant stretch is included and the flow is corrected by considering the blockage factor of the fuel. From these results, the normal gravity blowoff boundary for this axisymmetric rod geometry is determined to be linear, with oxygen concentration directly proportional to flow speed. We describe a new normal gravity upward flame spread test method which extrapolates the linear blowoff boundary to the zero stretch limit to resolve microgravity flammability limits, something current methods cannot do. This new test method can improve spacecraft fire safety for future exploration missions by providing a tractable way to obtain good estimates of material flammability in low gravity.

Description: The four 1,280 bolometer detector arrays that will fly on the balloon borne PIPER mission will be cooled by a 4-stage adiabatic demagnetization refrigerator (ADR). Two of the three mechanically independent ADR assemblies provide thermal isolation to their salt pills through Kevlar suspensions while the other provides thermal isolation to its salt pill through the use of bellows and Vespel material. The ADR integrates with the detector arrays and it sits in a large bucket Dewar containing superfluid liquid helium. This paper will describe the complex mechanical design of the PIPER ADR, and summarize the mechanical analysis done to validate the design.The four 1,280 bolometer detector arrays that will fly on the balloon borne PIPER mission will be cooled by a 4-stage adiabatic demagnetization refrigerator (ADR). Two of the three mechanically independent ADR assemblies provide thermal isolation to their salt pills through Kevlar suspensions while the other provides thermal isolation to its salt pill through the use of bellows and Vespel material. The ADR integrates with the detector arrays and it sits in a large bucket Dewar containing superfluid liquid helium. This paper will describe the complex mechanical design of the PIPER ADR, and summarize the mechanical analysis done to validate the design.

Description: Understanding, predicting, and controlling fluid slosh dynamics is critical to safety and improving performance of space missions when a significant percentage of the spacecrafts mass is a liquid. Computational fluid dynamics simulations can be used to predict the dynamics of slosh, but these programs require extensive validation. Many CFD programs have been validated by slosh experiments using various fluids in earth gravity, but prior to the ISS SPHERES-Slosh experiment, little experimental data for long-duration, zero-gravity slosh existed. This paper presents the current status of an ongoing CFD validation study using the ISS SPHERES-Slosh experimental data.

Description: Understanding, predicting, and controlling fluid slosh dynamics is critical to safety and improving performance of space missions when a significant percentage of the spacecrafts mass is a liquid. Computational fluid dynamics simulations can be used to predict the dynamics of slosh, but these programs require extensive validation. Many CFD programs have been validated by slosh experiments using various fluids in earth gravity, but prior to the ISS SPHERES-Slosh experiment1, little experimental data for long-duration, zero-gravity slosh existed. This paper presents the current status of an ongoing CFD validation study using the ISS SPHERES-Slosh experimental data.

Description: A high-fidelity approach for simulating the aerothermodynamic environments of meteor entries is developed. Two primary components of this model are coupled radiation and coupled ablation. Coupled radiation accounts for the impact of radiation on the flow field energy equations, while coupled ablation explicitly models the injection of ablation products within the flow field and radiation simulations. For a meteoroid with a velocity of 20 km/s, coupled radiation reduces the stagnation point radiative heating by over 60%. For altitudes below 40 km, the impact of coupled radiation on the flow field structure is shown to be fundamentally different, as a result of the large optical thicknesses, than that seen for reentry vehicles, which do not reach such altitudes at velocities greater than 10 km/s. The impact of coupled ablation (with coupled radiation) is shown to provide at least a 70% reduction in the radiative heating relative to the coupled-radiation-only cases. This large reduction is partially the result of the low ionization energies, relative to air species, of ablation products. The low ionization energies of ablation products, such as Mg and Ca, provide strong photoionization and atomic line absorption in regions of the spectrum that air species do not. MgO and CaO are also shown to provide significant absorption. Turbulence is shown to impact the distribution of ablation products through the shock- layer, which results in up to a 100% increase in the radiative heating downstream of the stagnation point. To create a database of heat transfer coefficients the developed model was applied to a range of cases. This database considered velocities ranging from 14 to 20 km/s, altitudes ranging from 20 to 50 km, and nose radii ranging from 1 to 100 m. The heat transfer coefficients from these simulations are below 0.045 for the range of cases (with turbulence), which is significantly lower than the canonical value of 0.1.

Description: Spacecraft and instruments on space missions are built using a wide variety of carefully-chosen materials. In addition to having mechanical properties appropriate for surviving the launch environment, these materials generally must have thermal conductivity values which meet specific requirements in their operating temperature ranges. Space missions commonly propose to include materials for which the thermal conductivity is not well known at cryogenic temperatures. We developed a test facility in 2004 at NASAs Goddard Space Flight Center to measure material thermal conductivity at temperatures between 4 and 300 Kelvin, and we have characterized many candidate materials since then. The measurement technique is not extremely complex, but proper care to details of the setup, data acquisition and data reduction is necessary for high precision and accuracy. We describe the thermal conductivity measurement process and present results for several materials.

Description: Roughness is unavoidable on practical high-speed vehicles, so it is critical to determine its impact on boundary layer transition. The flow field downstream of a streamwise array of cylindrical roughness elements is probed with hot-wire anemometry in this experiment. Mean flow distortion is examined in several measurement planes in the wake of the cylindrical roughness using the streak strength profiles and contour plots of the mass flux and total temperature. The roughness element heights and spacings were varied and their instability modes were examined. Cylindrical roughness elements approximately 140 micron tall produce an odd instability mode that grows weakly with downstream distance in the measurement range of this experiment. Cylindrical roughness elements approximately 280 micron tall produce an even instability mode that grows, becomes nonlinear, and then breaks down. Transition onset remains constant relative to the most downstream roughness in the streamwise array when the 280 micron roughness elements are spaced 2 diameters apart. Transition onset occurs at an earlier upstream location relative to the most downstream roughness in the streamwise array when the roughness elements are spaced 4 diameters appear to recover before the next downstream roughness element, so the location of transition shifts with the location of the most downstream roughness element in the array. When the rough- apart. The wake behind roughness elements spaced 2 diameters apart do not ness elements are spaced 4 diameters apart, the flow behind the first roughness element has enough space to recover before feeding into the second roughness element, and thus, moves transition forward.

Description: Flow separation control over a wall-mounted hump model was studied experimentally to assess the performance of sweeping jet actuators. Results were compared to that of the 2004 CFD validation experiment (CFDVAL2004), which examined flow separation control with steady suction and unsteady zero-net-mass-flow actuators. Comparisons were carried out at low and high amplitude excitations. In addition to the active flow control methods, a passive flow control method (i.e., vortex generator) was used to complement the dataset. Steady/unsteady surface pressure measurements and surface oilflow visualization were used in the performance assessment of the actuators. The results indicated that the sweeping jet actuators are more effective than the steady suction and unsteady zero-net-mass-flow actuators. For the same momentum coefficient, the sweeping jet actuators produced more flow acceleration upstream of separation, more pressure recovery downstream, and consistently a smaller separation bubble.

Description: Low-pressure blowoff experiments were conducted with a stagnation flame stabilized on the forward tip of cast PMMA rods in a vertical wind tunnel. Pressure, forced flow velocity, gravity, and ambient oxygen concentration were varied. Stagnation flame blowoff is determined from a time-stamped video recording of the test. The blowoff pressure is determined from test section pressure transducer data that is synchronized with the time stamp. The forced flow velocity is also determined from the choked flow orifice pressure. Most of the tests were performed in normal gravity, but a handful of microgravity tests were also conducted to determine the influence of buoyant flow velocity on the blowoff limits. The blowoff limits are found to have a linear dependence between the partial pressure of oxygen and the total pressure, regardless of forced flow velocity and gravity level. The flow velocity (forced and/or buoyant) affects the blowoff pressure through the critical Damkohler number residence time, which dictates the partial pressure of oxygen at blowoff. This is because the critical stretch rate increases linearly with increasing pressure at low pressure (sub-atmospheric pressures) since a second-order overall reaction rate with two-body reactions dominates in this pressure range.

Description: An anti-vortex baffle is a liquid propellant management device placed adjacent to an outlet of the propellant tank. Its purpose is to substantially reduce or eliminate the formation of free surface dip and vortex, as well as prevent vapor ingestion into the outlet, as the liquid drains out through the flight. To design an effective anti-vortex baffle, Computational Fluid Dynamic (CFD) simulations were undertaken for the NASA Ares I vehicle LOX tank subjected to the simulated flight loads with and without the anti-vortex baffle. The Six Degree-Of-Freedom (6-DOF) dynamics experienced by the Crew Launch Vehicle (CLV) during ascent were modeled by modifying the momentum equations in a CFD code to accommodate the extra body forces from the maneuvering in a non-inertial frame. The present analysis found that due to large moments, the CLV maneuvering has a significant impact on the vortical flow generation inside the tank. Roll maneuvering and side loading due to pitch and yaw are shown to induce swirling flow. The vortical flow due to roll is symmetrical with respect to the tank centerline, while those induced by pitch and yaw maneuverings showed two vortices side by side. The study found that without the anti-vortex baffle, the swirling flow caused surface dip during the late stage of drainage and hence early vapor ingestion. The flow can also be non-uniform in the drainage pipe as the secondary swirling flow velocity component can be as high as 10% of the draining velocity. An analysis of the vortex dynamics shows that the swirling flow in the drainage pipe during the Upper Stage burn is mainly the result of residual vortices inside the tank due to the conservation of angular momentum. The study demonstrated that the swirling flow in the drainage pipe can be effectively suppressed by employing the anti-vortex baffle.

Description: Vortex generators within a streamline-traced, external-compression supersonic inlet for Mach 1.66 were investigated to determine their ability to increase total pressure recovery and reduce total pressure distortion. The vortex generators studied were rectangular vanes arranged in counter-rotating and co-rotating arrays. The vane geometric factors of interest included height, length, spacing, angle-of-incidence, and positions upstream and downstream of the inlet terminal shock. The flow through the inlet was simulated numerically through the solution of the steady-state, Reynolds-averaged Navier-Stokes equations on multi-block, structured grids using the Wind-US flow solver. The vanes were simulated using a vortex generator model. The inlet performance was characterized by the inlet total pressure recovery and the radial and circumferential total pressure distortion indices at the engine face. Design of experiments and statistical analysis methods were applied to quantify the effect of the geometric factors of the vanes and search for optimal vane arrays. Co-rotating vane arrays with negative angles-of-incidence positioned on the supersonic diffuser were effective in sweeping low-momentum flow from the top toward the sides of the subsonic diffuser. This distributed the low-momentum flow more evenly about the circumference of the subsonic diffuser and reduced distortion. Co-rotating vane arrays with negative angles-of-incidence or counter-rotating vane arrays positioned downstream of the terminal shock were effective in mixing higher-momentum flow with lower-momentum flow to increase recovery and decrease distortion. A strategy of combining a co-rotating vane array on the supersonic diffuser with a counter-rotating vane array on the subsonic diffuser was effective in increasing recovery and reducing distortion.

Description: Recent demonstration of advanced liquid hydrogen storage techniques using Integrated Refrigeration and Storage (IRAS) technology at NASA Kennedy Space Center led to the production of large quantities of solid densified liquid and slush hydrogen in a 125,000 L tank. Production of densified hydrogen was performed at three different liquid levels and LH2 temperatures were measured by twenty silicon diode temperature sensors. System energy balances and solid mass fractions are calculated. Experimental data reveal hydrogen temperatures dropped well below the triple point during testing (up to 1 K), and were continuing to trend downward prior to system shutdown. Sub-triple point temperatures were seen to evolve in a time dependent manner along the length of the horizontal, cylindrical vessel. Twenty silicon diode temperature sensors were recorded over approximately one month for testing at two different fill levels (33 67). The phenomenon, observed at both two fill levels, is described and presented detailed and explained herein., and The implications of using IRAS for energy storage, propellant densification, and future cryofuel systems are discussed.

Description: The reliance on non-permanent coatings in Condensing Heat Exchanger (CHX) designs is a significant technical issue to be solved before long-duration spaceflight can occur. Therefore, high reliability CHXs have been identified by the Evolvable Mars Campaign (EMC) as critical technologies needed to move beyond low earth orbit. The Laser Processed Condensing Heat Exchanger project aims to solve these problems through the use of femtosecond laser processed surfaces, which have unique wetting properties and potentially exhibit anti-microbial growth properties. These surfaces were investigated to identify if they would be suitable candidates for a replacement CHX surface. Among the areas researched in this project include microbial growth testing, siloxane flow testing in which laser processed surfaces were exposed to siloxanes in an air stream, and manufacturability.

Description: Time-resolved particle image velocimetry (TRPIV) measurements are performed down-stream of a swept backward-facing step, with a height of 49% of the boundary-layer thickness. The results agree well qualitatively with previously reported hotwire measurements, though the amplitudes of the fluctuating components measured using TRPIV are higher. Nonetheless, the low-amplitude instabilities in the flow are fairly well resolved using TR- PIV. Proper orthogonal decomposition is used to study the development of the traveling cross flow and Tollmien-Schlichting (TS) instabilities downstream of the step and to study how they interact to form the large velocity spikes that ultimately lead to transition. A secondary mode within the traveling cross flow frequency band develops with a wavelength close to that of the stationary cross flow instability, so that at a certain point in the phase, it causes an increase in the spanwise modulation initially caused by the stationary cross flow mode. This increased modulation leads to an increase in the amplitude of the TS mode, which, itself, is highly modulated through interactions with the stationary cross flow. When the traveling cross flow and TS modes align in time and space, the large velocity spikes occur. Thus, these three instabilities, which are individually of low amplitude when the spikes start to occur (U'rms/Ue 〈0.03), interact and combine to cause a large flow disturbance that eventually leads to transition.

Description: Obtaining measurements of flight environments on ablative heat shields is both critical for spacecraft development and extremely challenging due to the harsh heating environment and surface recession. Thermocouples installed several millimeters below the surface are commonly used to measure the heat shield temperature response, but an ill-posed inverse heat conduction problem must be solved to reconstruct the surface heating environment from these measurements. Ablation can contribute substantially to the measurement response making solutions to the inverse problem strongly dependent on the recession model, which is often poorly characterized. To enable efficient surface reconstruction for recession model sensitivity analysis, a method for decoupling the surface recession evaluation from the inverse heat conduction problem is presented. The decoupled method is shown to provide reconstructions of equivalent accuracy to the traditional coupled method but with substantially reduced computational effort. These methods are applied to reconstruct the environments on the Mars Science Laboratory heat shield using diffusion limit and kinetically limited recession models.

Description: Increasing integration density of electronic components has exacerbated the thermal management challenges facing electronic system developers. The high power, heat flux, and volumetric heat generation of emerging devices are driving the transition from remote cooling, which relies on conduction and spreading, to embedded cooling, which facilitates direct contact between the heat-generating device and coolant flow. Microgap coolers employ the forced flow of dielectric fluids undergoing phase change in a heated channel between devices. While two phase microcoolers are used routinely in ground-based systems, the lack of acceptable models and correlations for microgravity operation has limited their use for spacecraft thermal management. Previous research has revealed that gravitational acceleration plays a diminishing role as the channel diameter shrinks, but there is considerable variation among the proposed gravity-insensitive channel dimensions and minimal research on rectangular ducts. Reliable criteria for achieving gravity-insensitive flow boiling performance would enable spaceflight systems to exploit this powerful thermal management technique and reduce development time and costs through reliance on ground-based testing. In the present effort, the authors have studied the effect of evaporator orientation on flow boiling performance of HFE7100 in a 218 m tall by 13.0 mm wide microgap cooler. Similar heat transfer coefficients and critical heat flux were achieved across five evaporator orientations, indicating that the effect of gravity was negligible.

Description: OBJECTIVES: To evaluate proposed anti-vortex design in suppressing swirling flow during US burn. APPROACH: Include two major body forces in the analysis a)Vehicle acceleration (all three components); b)Vehicle maneuvers (roll, pitch, and yaw). Perform two drainage analyses of Ares I LOX tank using 6 DOF body forces predicted by GN&C analysis (Guidance Navigation and Control) during vehicle ascent: one with baffle, one without baffle. MODEL: Use Ares I defined geometry. O-Grid for easy fitting of baffle. In this preliminary analysis the holes are sealed. Use whole 360 deg. model with no assumption of symmetry or cyclic boundary conditions. Read in 6DOF data vs time from a file.

Description: Full seven-equation Reynolds stress turbulence models are promising tools for todays aerospace technology challenges. This paper examines two such models for computing challenging turbulent flows including shock-wave boundary layer interactions, separation and mixing layers. The Wilcox and the SSG/LRR full second-moment Reynolds stress models have been implemented into the FUN3D (Fully Unstructured Navier-Stokes Three Dimensional) unstructured Navier-Stokes code and were evaluated for four problems: a transonic two-dimensional diffuser, a supersonic axisymmetric compression corner, a compressible planar shear layer, and a subsonic axisymmetric jet. Simulation results are compared with experimental data and results computed using the more commonly used Spalart-Allmaras (SA) one-equation and the Menter Shear Stress Transport (SST-V) two-equation turbulence models.

Description: A typical entry vehicle design depends on the synthesis of many essential subsystems, including thermal protection system (TPS), structures, payload, avionics, and propulsion, among others. The ability to incorporate aerothermodynamic considerations and TPS design into the early design phase is crucial, as both are closely coupled to the vehicle's aerodynamics, shape and mass. In the preliminary design stage, reasonably accurate results with rapid turn-representative entry envelope was explored. Initial results suggest that for Mach numbers ranging from 9-20, a few inviscid solutions could reasonably sup- port surface heating predictions at Mach numbers variation of +/-2, altitudes variation of +/-10 to 20 kft, and angle-of-attack variation of +/- 5. Agreement with Navier-Stokes solutions was generally found to be within 10-15% for Mach number and altitude, and 20% for angle of attack. A smaller angle-of-attack increment than the 5 deg around times for parametric studies and quickly evolving configurations are necessary to steer design decisions. This investigation considers the use of an unstructured 3D inviscid code in conjunction with an integral boundary-layer method; the former providing the flow field solution and the latter the surface heating. Sensitivity studies for Mach number, angle of attack, and altitude, examine the feasibility of using this approach to populate a representative entry flight envelope based on a limited set of inviscid solutions. Each inviscid solution is used to generate surface heating over the nearby trajectory space. A subset of a considered in this study is recommended. Results of the angle-of-attack sensitivity studies show that smaller increments may be needed for better heating predictions. The approach is well suited for application to conceptual multidisciplinary design and analysis studies where transient aeroheating environments are critical for vehicle TPS and thermal design. Concurrent prediction of aeroheating environments, coupled with the use of unstructured methods, is considered enabling for TPS material selection and design in conceptual studies where vehicle mission, shape, and entry strategies evolve rapidly.

Description: A thermal protection system that reduces a mismatch of thermal expansion coefficients CTE between a first material layer (CTE1) and a second material layer (CTE2) at a first layer-second layer interface. A portion of aluminum borosilicate (abs) or another suitable additive (add), whose CTE value, CTE(add), satisfies (CTE(add)-CTE1)(CTE(add)-CTE2)〈0, is distributed with variable additive density,.rho.(z;add), in the first material layer and/or in the second material layer, with.rho.(z;add) near the materials interface being relatively high (alternatively, relatively low) and.rho.(z;add) in a region spaced apart from the interface being relatively low (alternatively, relatively high).

Description: An experimental investigation of the effects of distributed surface roughness on boundary-layer transition and turbulent heating has been conducted. Hypersonic wind tunnel testing was performed using hemispherical models with surface roughness patterns simulating those produced by heat shield ablation. Global aeroheating and transition onset data were obtained using phosphor thermography at Mach 6 and Mach 10 over a range of roughness heights and free stream Reynolds numbers sufficient to produce laminar, transitional and turbulent flow. Upstream movement of the transition onset location and increasing heating augmentation over predicted smooth-wall levels were observed with both increasing roughness heights and increasing free stream Reynolds numbers. The experimental heating data are presented herein, as are comparisons to smooth-wall heat transfer distributions from computational flow-field simulations. The transition onset data are also tabulated, and correlations of these data are presented.