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At least 595 records · Page 33

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NASA NTRS2018

Larch (Larix dahurica Turcz) growth response to climate change in the Siberian permafrost zone

Larch-dominant communities are the most extensive high-latitude forests in Eurasia and are experiencing the strongest impacts from warming temperatures. We analyzed larch (Larix dahurica Turcz) growth index (GI) response to climate change. The studied larch-dominant communities are located within the permafrost zone of Northern Siberia at the northern tree limit (ca. N 67° 38′, E 99° 07′). Methods included dendrochronology, analysis of climate variables, root zone moisture content, and satellitederived gross (GPP) and net (NPP) primary productivity. It was found that larch response to warming included a period of increased annual growth increment (GI) (from the 1970s to ca. 1995) with a follow on GI decline. Increase in GI correlated with summer air temperature, whereas an observed decrease in GI was caused by water stress (vapor pressure deficit and drought increase). Water stress impact on larch growth in permafrost was not observed before the onset of warming (ca. 1970). Water limitation was also indicated by GI dependence on soil moisture stored during the previous year. Water stress was especially pronounced for stands growing on rocky soils with low water-holding capacity. GPP of larch communities showed an increasing trend, whereas NPP stagnated. A similar pattern of GI response to climate warming has also been observed for Larix sibirica Ledeb, Pinus sibirica Du Tour, and Abies sibirica Ledeb in the forests of southern Siberia. Thus, warming in northern Siberia permafrost zone resulted in an initial increase in larch growth from the 1970s to the mid-1990s. After that time, larch growth increment has decreased. Since ca. 1990, water stress at the beginning of the vegetative period became, along with air temperature, a main factor affecting larch growth within the permafrost zone.

Larix dahurica↗

NASA NTRS1990

Optimal trajectories for an aerospace plane. Part 1: Formulation, results, and analysis

The optimization of the trajectories of an aerospace plane is discussed. This is a hypervelocity vehicle capable of achieving orbital speed, while taking off horizontally. The vehicle is propelled by four types of engines: turbojet engines for flight at subsonic speeds/low supersonic speeds; ramjet engines for flight at moderate supersonic speeds/low hypersonic speeds; scramjet engines for flight at hypersonic speeds; and rocket engines for flight at near-orbital speeds. A single-stage-to-orbit (SSTO) configuration is considered, and the transition from low supersonic speeds to orbital speeds is studied under the following assumptions: the turbojet portion of the trajectory has been completed; the aerospace plane is controlled via the angle of attack and the power setting; the aerodynamic model is the generic hypersonic aerodynamics model example (GHAME). Concerning the engine model, three options are considered: (EM1), a ramjet/scramjet combination in which the scramjet specific impulse tends to a nearly-constant value at large Mach numbers; (EM2), a ramjet/scramjet combination in which the scramjet specific impulse decreases monotonically at large Mach numbers; and (EM3), a ramjet/scramjet/rocket combination in which, owing to stagnation temperature limitations, the scramjet operates only at M approx. less than 15; at higher Mach numbers, the scramjet is shut off and the aerospace plane is driven only by the rocket engines. Under the above assumptions, four optimization problems are solved using the sequential gradient-restoration algorithm for optimal control problems: (P1) minimization of the weight of fuel consumed; (P2) minimization of the peak dynamic pressure; (P3) minimization of the peak heating rate; and (P4) minimization of the peak tangential acceleration.

Miele, Angelo↗

NASA NTRS1980

Conceptual model of turbulent flameholding for scramjet combustors

New concepts and approaches to scramjet combustor design are presented. Blowoff was from failure of the recirculation-zone (RZ) flame to reach the dividing streamline (DS) at the rear stagnation zone. Increased turbulent exchange across the DS helped flameholding due to forward movement of the flame anchor point inside the RZ. Modeling of the blowoff phenomenon was based on a mass conservation concept involving the traverse of a flame element across the RZ and a flow element along the DS. The scale required to achieve flameholding, predicted by the model, showed a strong adverse effect of low pressure and low fuel equivalence ratio, moderate effect of flight Mach number, and little effect of temperature recovery factor. Possible effects of finite rate chemistry on flameholding and flamespreading in scramjets are discussed and recommendations for approaches to engine combustor design as well as for needed research to reduce uncertainties in the concepts are made.

Huber, P. W.↗

DOE OSTI2023

A reduced-model (nSOLT) simulation of neutral recycling effects on plasma turbulence in the divertor region of MAST-U

The 2D scrape-off-layer turbulence code (nSOLT), which includes 1D kinetic neutral–plasma interactions, is applied to study effects of neutral recycling on plasma turbulence for parameters illustrative of the MAST-U divertor region. Neutral recycling is modeled by injecting a fraction of the parallel plasma flux to the divertor back into the simulation domain as a source of Franck–Condon-distributed neutrals. Stationary sources, concentrated at the magnetic separatrix, model plasma streaming into the divertor region from the upstream scrape-off-layer and sustain plasma turbulence absent neutral recycling. Starting from one such no-neutrals equilibrium, we initiate recycling in a numerical experiment designed to diagnose and identify the effects of various neutral–plasma interactions on the divertor plasma, divertor turbulence, and plasma exhaust. The onset of recycling triggers an initial burst of enhanced cross field plasma transport that is quelled by ionization cooling and charge–exchange (CX) friction, with growing neutral pressure, leading to a quiescent, turbulence-free state. Diagnosis of this transient burst reveals that (1) the sudden increase in plasma density due to ionization dominates the onset of the burst, (2) electron cooling due to ionization increases collisionality and disconnects blob filaments from the sheath, and (3) CX friction drives tripole polarization of a blob that can dominate the curvature-driven dipole polarization, leading to the stagnation of blob propagation and reduced radial turbulent transport. It is shown that CX friction is negligible compared to sheath physics in determining equilibrium mean flow shearing rates, for parameters considered herein (specifically a short connection length to the divertor target), while it can significantly reduce interchange-instability growth rates.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗

NASA NTRS1996

Blunt body near wake flow field at Mach 6

Tests were conducted in a Mach 6 flow to examine the reattachment process of an axisymmetric free shear layer associated with the near wake of a 70 deg. half angle, spherically blunted cone with a cylindrical after body. Model angle of incidence was fixed at 0 deg. and free-stream Reynolds numbers based on body diameter ranged from 0.5 x 10(exp 6) to 4 x 10(exp 6). The sensitivity of wake shear layer transition on reattachment heating was investigated. The present perfect gas study was designed to compliment results obtained previously in facilities capable of producing real gas effects. The instrumented blunted cone model was designed primarily for testing in high enthalpy hypervelocity shock tunnels in both this country and abroad but was amenable for testing in conventional hypersonic blowdown wind tunnels as well. Surface heating rates were inferred from temperature - time histories from coaxial surface thermocouples on the model forebody and thin film resistance gages along the model base and cylindrical after body. General flow feature (bow shock, wake shear layer, and recompression shock) locations were visually identified by schlieren photography. Mean shear layer position and growth were determined from intrusive pitot pressure surveys. In addition, wake surveys with a constant temperature hot-wire anemometer were utilized to qualitatively characterize the state of the shear layer prior to reattachment. Experimental results were compared to laminar perfect gas predictions provided by a 3-D Navier Stokes code (NSHYP). Shear layer impingement on the instrumented cylindrical after body resulted in a localized heating maximum that was 21 to 29 percent of the forebody stagnation point heating. Peak heating resulting from the reattaching shear layer was found to be a factor of 2 higher than laminar predictions, which suggested a transitional shear layer. Schlieren flow visualization and fluctuating voltage time histories and spectra from the hot wire surveys across the shear layer substantiate this observation. The sensitivity of surface heating to forebody roughness was characterized for a reattaching shear layer. For example, at R(sub infinity), d = 4 x 10(exp 6), when the shear layer was transitional, the magnitude of peak heating from shear layer impingement was reduced by approximately 24 percent when transition grit was applied to the forebody. The spatial location of the local peak, however, remained unchanged.

Horvath, Thomas J.↗

DOE OSTI2023

A direct numerical simulation study for confined non-isothermal jet impingement at moderate nozzle-to-plate distances: Capturing jet-to-ambient density effects

A direct numerical simulation (DNS) campaign is deployed for a series of confined downward oriented, non-isothermal turbulent impinging jet configurations. A baseline Reynolds number of 9960 is obtained through a precursor DNS pipe flow simulation (Re τ = 505). Three jet temperature configurations (confinement height to nozzle diameter of three) enter a cylindrical domain that share ambient and impingement plate temperatures (298.15 K). The range of jet temperatures are crafted such that the ratio of inlet to ambient density varies from unity to 0.52, showcasing the effect of density disparity on flow characteristics such as core collapse, radial mixing of momentum and energy, near-wall stagnation behavior, wall-jet profiles, and large-scale vortical structures. Surface quantities provided include mean radial heat flux and wall-shear stress profiles, and heat flux histograms at select radial stations. Results showcase increased radial normal stresses for higher temperature jets that support increased mixing, resulting in large-scale recirculation structures that are smaller, while retaining similar normalized radial wall profiles for shear stress, heat flux and pressure. Furthermore, radial plots for wall shear stress and Nusselt number showcase strong radial decay as compared to previous configurations that share similar jet and ambient temperatures. For the 373.15 K case, a Gaussian-like histogram for heat fluxes at the impingement plate transitions to a log-normal profile as radial distances increase. In contrast, the 573.15 K configuration displays a bi-modal heat flux characteristic at the impingement plate, and in similar manner to the moderate temperature counterpart, transitions to a log-normal profile at larger radial distances.

33 ADVANCED PROPULSION SYSTEMS↗

NASA NTRS1993

Time-dependent simulation of reflected-shock/boundary layer interaction

An initial numerical/experimental investigation has been done to better understand multi-dimensional flow phenomena inside pulse facilities. Time-dependent quasi-one-dimensional and axisymmetric numerical simulations of complete shock tube flow are compared with experimental pressure traces recorded at the NASA Ames electric-arc driven shock tube facility (from cold driver shots). Of particular interest is the interaction between the reflected shock wave and the boundary layer. Evidence of the shock bifurcation caused by this interaction is clearly seen in the present experimental data. The axisymmetric simulations reproduce this phenomenon and demonstrate how this interaction can provide a mechanism for driver gas to contaminate the stagnation region. The simulations incorporate finite-rate chemistry, a moving mesh and laminar viscosity.

Wilson, Gregory J.↗

NASA NTRS1987

Ascent aeromaneuvering capabilities of transatmospheric vehicles

The low-earth orbit rendezvous capability of a conceptual transatmospheric vehicle is analyzed for two endo-/exo-atmospheric ascent missions. Both cases involve coasting aerodynamic maneuvers starting from the burn-out conditions corresponding to air-breathing propulsion systems that achieve orbital velocity within the atmosphere. The powered phase of the ascent trajectories approximate constant dynamic pressure, fuel-efficient flightpaths typically flown by supersonic aircraft. The aeromaneuvering coast phases of the ascent include both coplanar (to determine altitude capability without plane-inclination changes) and aeroturning to LEO rendezvous at 400 km altitude (to assess plane-change capability). The coast-phase ascent maneuvers are correlated with G-load requirements and aerothermodynamic heating characteristics at two critical locations on the vehicle surface (i.e., the nose stagnation point and the body centerline). The results are correlated and recommendations are made concerning thermal protection and structural requirements.

Menees, Gene P.↗

NASA NTRS1991

Flow downstream of the heliospheric terminal shock - Magnetic field kinematics

A kinematic model of the interplanetary magnetic field in the heliosheath beyond the solar wind terminal shock is presented in order to evaluate the possible importance of MHD effects in that region of space. The need for this evaluation arises because the interplanetary magnetic field is compressed across the terminal shock and further amplified by the decreasing flow speed beyond the shock. Streamlines which approach the stagnation point before turning in the downstream direction lead to the strongest effects due to the extreme slowing of the solar wind and consequent compression of the embedded magnetic field. The magnetic volume force therefore cannot be neglected on streamlines that approach the heliopause in the upstream direction, where the volume containing them is a large fraction of the overall of the heliosheath in the upstream direction. The increase in the magnetic pressure may act to bring the upstream terminal shock significantly closer to the sun, potentially reconciling a conflict between models and observations.

Nerney, S.↗

NASA NTRS1974

Lee surface flow phenomena over space shuttle at large angles of attack at M sub infinity equal 6

Surface pressure and heat transfer, flow separation, flow field, and oil flow patterns on the leeward side of a space shuttle orbiter model are investigated at a free stream Mach number of 6. The free stream Reynolds numbers are between 1.64 times 10 to the 7th power and 1.31 times 10 to the 8th power per meter, and the angle of attack is varied between 0 deg and 40 deg for the present experiments. The stagnation temperatures for the tests are approximately 500 K and the wall temperature is maintained at 290 K. Existing numerical methods of three-dimensional inviscid supersonic flow theory and compressible boundary layer theory are used to predict the present experimental measurements. Results of the present tests indicate two distinct types of flow separation and surface peak heating depending on the angle of attack.

Zakkay, V.↗

NASA NTRS2015

The Role of Cooling in Pahohoe Emplacement on Planetary Surfaces.

Abundant evidence is emerging that many lavas on Mars were emplaced as slow-moving pahoehoe flows. Models for such scenarios contrast sharply with those for steep-sloped applications where gravity is the dominant force. The mode of flow emplacement on low slopes is characterized by toe formation and inflation. In the latter phase of pahoehoe flow emplacement, stagnation, inflation, and toe formation are most closely tied to the final topography, dimensions, and morphologic features. This mode of emplacement is particularly relevant to the low slopes of planetary surfaces such as the plains of Mars, Io and the Moon.

Pahoehoe Lava Flows; Pressure Dependent Simulation↗

NASA NTRS2009

Validation of PICA Ablation and Thermal-Response Model at Low Heat Flux

Phenolic Impregnated Carbon Ablator (PICA) was the forebody heatshield material on the Stardust sample-return capsule and is also a primary candidate material for the Mars Science Lander (MSL), the Orion Crew Module, and the SpaceX Dragon vehicle. As part of the heatshield qualification for Orion, physical and thermal properties of virgin and charred PICA were measured, and an ablation and thermal response model was developed. We validated the model by comparing it with recession and temperature data from stagnation arcjet tests conducted over a wide range of stagnation heat flux of 107 to 1102 W/sq cm. The effect of orthotropic thermal conductivity was evident in the thermal response of the arcjet models. In general, model predictions compared well with the data; however, the uncertainty of the recession prediction was greatest for heat fluxes below 200 W/sq cm. More recent MSL testing focused on the low heat flux regime of 45 to 250 W/sq cm. The new results confirm the recession uncertainty, especially for pressures below 6 kPa. In this work we focus on improving the model predictions for MSL and Orion tests below 250 W/sq cm.

Milos, Frank S.↗

NASA NTRS1975

Thermostructural testing of instrumented ATJ-S graphite nosetip models in the AFFDL 50 MW rent facility

Instrumented ATJ-S graphite shell nosetip models were tested in the AFFDL 50 MW RENT Facility. The inside surface of the models were instrumented with seven thermocouples and three clip-on extensometers. One thermocouple was located at the inside stagnation point and three thermocouples were located at 120 degree intervals on the aft portion of the skirt of the model. In addition, three thermocouples at 120 degree intervals plus three clip-on extensometers at 120 degree intervals offset 60 degrees from the thermocouples were located in the critical region. The outside surface temperature of the models was monitored in two locations by optical pyrometers. Finally, the tests were recorded by three high speed motion picture cameras. Pressure models and calorimeters were also tested in order to help define the flow field condition and heating distribution around the models.

Budde, C. L.↗

NASA NTRS2010

Analysis of Unsteady Tip and Endwall Heat Transfer in a Highly Loaded Transonic Turbine Stage

In a previous study, vane-rotor shock interactions and heat transfer on the rotor blade of a highly loaded transonic turbine stage were simulated. The geometry consists of a high pressure turbine vane and downstream rotor blade. This study focuses on the physics of flow and heat transfer in the rotor tip, casing and hub regions. The simulation was performed using the Unsteady Reynolds-Averaged Navier-Stokes (URANS) code MSU-TURBO. A low Reynolds number k-epsilon model was utilized to model turbulence. The rotor blade in question has a tip gap height of 2.1 percent of the blade height. The Reynolds number of the flow is approximately 3x10(exp 6) per meter. Unsteadiness was observed at the tip surface that results in intermittent "hot spots". It is demonstrated that unsteadiness in the tip gap is governed by inviscid effects due to high speed flow and is not strongly dependent on pressure ratio across the tip gap contrary to published observations that have primarily dealt with subsonic tip flows. The high relative Mach numbers in the tip gap lead to a choking of the leakage flow that translates to a relative attenuation of losses at higher loading. The efficacy of new tip geometry is discussed to minimize heat flux at the tip while maintaining choked conditions. In addition, an explanation is provided that shows the mechanism behind the rise in stagnation temperature on the casing to values above the absolute total temperature at the inlet. It is concluded that even in steady mode, work transfer to the near tip fluid occurs due to relative shearing by the casing. This is believed to be the first such explanation of the work transfer phenomenon in the open literature. The difference in pattern between steady and time-averaged heat flux at the hub is also explained.

Shyam, Vikram↗

NASA NTRS1976

Flow and heat transfer measurements in a pseudo-shock region with surface cooling

An experimental investigation was conducted to acquire information on the flow structure, mean flowfield, and temperature distributions in a pseudo-shock region in a supersonic diffuser with surface cooling. The Mach number upstream was 2.9, and the wall to stagnation temperature ratio was 0.44. A Mach-disk-like shock wave emanated from the thin separated flow region near the beginning of the compression region, and reattachment occurred one diameter downstream so that the flow was not separated over most of the pseudo-shock region. The flow compression was a shock-free, predominantly viscous process. Along the pseudo-shock region the measured heat-transfer coefficient increased approximately as the 0.8 power of the measured wall static pressure. The estimated wall shear stress increased downstream of flow attachment, but was still considerably less than the upstream value.

Cuffel, R. F.↗

NASA NTRS2001

Dual-Mode Combustion of Hydrogen in a Mach 5, Continuous-Flow Facility

Results of an experimental and numerical study of a dual-mode scramjet combustor are reported. The experiment consisted of a direct-connect test of a Mach 2 hydrogen-air combustor with a single unswept-ramp fuel injector. The flow stagnation enthalpy simulated a flight Mach number of 5. Measurements were obtained using conventional wall instrumentation and a particle-imaging laser diagnostic technique. The particle imaging was enabled through the development of a new apparatus for seeding fine silicon dioxide particles into the combustor fuel stream. Numerical simulations of the combustor were performed using the GASP code. The modeling, and much of the experimental work, focused on the supersonic combustion mode. Reasonable agreement was observed between experimental and numerical wall pressure distributions. However, the numerical model was unable to predict accurately the effects of combustion on the fuel plume size, penetration, shape, and axial growth.

Goyne, C. P.↗

DOE OSTI2022

First graded metal pushered single shell capsule implosions on the National Ignition Facility

Graded metal pushered single shell (PSS) capsules are predicted to be a viable alternative to low-Z capsule indirect drive inertial confinement fusion (ICF) implosions for achieving high fusion yields [MacLaren et al., Phys. Plasmas 28, 122710 (2021)]. The first experiments with Be/Cr-graded metal PSS capsules indicate that the implementation of the principle design feature, the graded density inner metal layer, has succeeded in producing a stable implosion with performance in agreement with predictions. With 50% Cr concentration in the pusher, PSS capsules have greater than ~2× higher shell densities during stagnation for enhanced core confinement and radiation trapping at ~35% lower shell implosion velocities than low-Z capsules. High-energy >30 keV inflight shell radiography recorded 215 km/s implosion velocities and show that implosion Legendre mode P 2 asymmetry can be tuned via inner-to-outer beam wavelength separation, similar to other implosions. Shell radiographs and neutron core images show similar P 2 asymmetry, suggesting no symmetry swings between peak implosion velocity and stagnation times. Despite the modest implosion velocities, gas-filled deuterium–tritium capsule implosions generate 10 15 neutron yields at relatively modest core ion temperatures of 2.75 keV, indicating that in spite of the high-density inner layer, the implosions have been stabilized by the design density gradient. When compared with hydrodynamic simulations, the measured yield-over-simulated is 35% due to fuel–pusher mix and other perturbations such as the capsule fill tube. Finally, simple analytical scalings of hot spot pressure and neutron yield show that PSS implosions reach similar performance at lower implosion velocities and higher shell densities to low-Z ICF capsules.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗

NASA NTRS1977

Aerodynamics of horizontally-opposed aircraft engine installations

The goal of cooling installation design is to provide adequate cooling air flow to the engine with a minimum drag penalty. To achieve this goal, attention must be given to certain aerodynamic factors involved in the design problem. In the present paper, the factors which influence the aerodynamic design of horizontally-opposed engine cooling installations are discussed with particular reference to the performance of the aircraft in terms of the dynamic pressure generated for the power output of the engine, the gains or losses through the propeller, proper inlet design, altitude effects on the orifice characteristics of the engine, and exit area sizing. Results of a flight research program showed that the inlet design has a significant influence on cooling effectiveness and drag; the inlet should be designed for low velocity ratios and should avoid small radii of curvature in its geometry. The influence of the propeller on the inlet appears to cause the stagnation point on the lip to move toward the inside.

Miley, S. J.↗