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At least 289 records · Page 16

Direct simulation Monte Carlo and Navier-Stokes simulations of blunt body wake flows

Numerical results obtained with direct simulation Monte Carlo and Navier-Stokes methods are presented for a Mach-20 nitrogen flow about a 70-deg blunted cone. The flow conditions simuulated are those that can be obtained in existing low-density hypersonic wind tunnels. Three sets of flow conditions are considered with freestream Knudsen numbers ranging from 0.03 to 0.001. The focus is on the wake structure: how the wake structure changes as a function of rarefaction, what the afterbody levels of heating are, and to what limits the continuum models are realistic as rarefaction in the wake is progressively increased. Calculations are made with and without an afterbody sting. Results for the after body sting are emphasizes in anticipation of an experimental study for the current flow conditions and model configuration. The Navier-Stokes calculations were made with and without slip boundary conditions. Comparisons of the results obtained with the two simulation methodologies are made for both flowfield structure and surface quantities.

Moss, James N.↗

Chromatic Image Analysis For Quantitative Thermal Mapping

Chromatic image analysis system (CIAS) developed for use in noncontact measurements of temperatures on aerothermodynamic models in hypersonic wind tunnels. Based on concept of temperature coupled to shift in color spectrum for optical measurement. Video camera images fluorescence emitted by phosphor-coated model at two wavelengths. Temperature map of model then computed from relative brightnesses in video images of model at those wavelengths. Eliminates need for intrusive, time-consuming, contact temperature measurements by gauges, making it possible to map temperatures on complex surfaces in timely manner and at reduced cost.

Buck, Gregory M.↗

Hypersonic flows as related to the National Aerospace Plane

The object of Cooperative Agreement NCC2-452 was to identify, develop, and document reliable turbulence models for incorporation into CFD codes, which would then subsequently be incorporated into numerical design procedures for the NASP and any other hypersonic vehicles. In a two-pronged effort, consisting of an experimental and a theoretical approach, several key features of flows over complex vehicles were identified, and test bodies were designed which were composed of simple geometric shapes over which these flow features were measured. The experiments were conducted in the 3.5' Hypersonic Wind Tunnel at NASA Ames Research Center, at nominal Mach numbers from 7 to 8.3 and Re/m from 4.9 x 10(exp 6) to 5.8 x 10(exp 6). Boundary layers approaching the interaction region were 2.5 to 3.7 cm thick. Surface and flow field measurements were conducted, and the initial boundary conditions were experimentally documented.

Kussoy, Marvin↗

The Use of Steady and Pulsed Detonations for Propulsion Systems

Objectives of the ODWE concept studies are: demonstrate the feasibility of the oblique detonation wave engine (ODWE) for hypersonic propulsion; demonstrate the existance and stability of an oblique detonation wave in hypersonic wind tunnels; develop engineering codes which predict the performance characteristics of the ODWE including specific impulse and thrust coefficients for various operating conditions; develop multi-dimensional computer codes which can model all aspects of the ODWE including fuel injection, mixing, ignition, combustion and expansion with fully detailed chemical kinetics and turbulence models; and validate the codes with experimental data use the simulations to predict the ODWE performance for conditions not easily obtained in wind tunnels.

Adelman, Henry G.↗

Direct Monte Carlo Simulations of Hypersonic Low-Density Flows about an ASTV Including Wake Structure

Results of a numerical study concerning flow past a 70-deg blunted cone in hypersonic low-density flow environments are presented using the direct simulation Monte-Carlo method. The flow conditions simulated are those that can be obtained in existing low-density hypersonic wind tunnels. Results indicate that a stable vortex forms in the near wake at and below a freestream Knudsen number (based on cone diameter) of 0.01 and the size of the vortex increases with decreasing Knudsen number. The base region of the flow remains in thermal nonequilibrium for all cases considered herein.

Dogra, V. K.↗

Rapid Model Fabrication and Testing for Aerospace Vehicles

Advanced methods for rapid fabrication and instrumentation of hypersonic wind tunnel models are being developed and evaluated at NASA Langley Research Center. Rapid aeroheating model fabrication and measurement techniques using investment casting of ceramic test models and thermographic phosphors are reviewed. More accurate model casting techniques for fabrication of benchmark metal and ceramic test models are being developed using a combination of rapid prototype patterns and investment casting. White light optical scanning is used for coordinate measurements to evaluate the fabrication process and verify model accuracy to +/- 0.002 inches. Higher-temperature (<210C) luminescent coatings are also being developed for simultaneous pressure and temperature mapping, providing global pressure as well as global aeroheating measurements. Together these techniques will provide a more rapid and complete experimental aerodynamic and aerothermodynamic database for future aerospace vehicles.

Buck, Gregory M.↗

Hypersonic Nozzle/Afterbody Experiment: Flow Visualization and Boundary Layer Experiments

This study was conducted to experimentally characterize the flow field created by the interaction of a single-expansion-ramp-nozzle (SERN) flow with a hypersonic external stream Data were obtained from a generic nozzle/afterbody model in the 3.5-Foot Hypersonic Wind Tunnel of the NASA Ames Research Center in a cooperative experimental program involving Ames and the McDonnell Douglas Aerospace. The model design and test planning were performed in close cooperation with members of the Ames computational fluid dynamics (CFD) team for the National Aero-Space Plane (NASP) program. This paper presents experimental results consisting of oil-flow and shadowgraph flow-visualization photographs, afterbody surface-pressure distributions, boundary-layer rake measurements, and Preston-tube skin-friction measurements.

Keener, Earl R.↗

Aerothermodynamics at NASA-Langley Research Center

The Aerothermodynamics Branch at NASA - Langley Research Center is tasked with developing, assessing and applying aerothermodynamic technologies to enable the development of hypersonic aircraft, launch vehicles, and planetary/earth entry systems. To accomplish this mission, the Branch capitalizes on the synergism between the experimental and computational facilities/tools which reside in the branch and a staff that can draw on five decades of experience in aerothermodynamics. The Aerothermodynamics Branch is staffed by 30 scientists/engineers. The staff, of which two-thirds are less than 40 years old, is split evenly between experimentalists and computationalists. Approximately 90 percent of the staff work on space transportation systems while the remainder work on planetary missions. The Branch manages 5 hypersonic wind tunnels which are staffed by 14 technicians, numerous high end work stations and a SGI Origin 2000 system. The Branch also utilizes other test facilities located at Langley as well as other national and international test sites. Large scale computational requirements are met by access to Agency resources.

Weilmuenster, K. James↗

Improvement of Flow Quality in NAL Chofu Mach 10 Nozzle

As a result of CFD analysis and remachining of the nozzle, the flow quality of the Mach 10 Hypersonic Wind Tunnel at NAL Chofu, Japan was improved. The subsequent test results validated the CFD analytical predictions by NASA and MHL.

Lacey, John↗

Experimental Hypersonic Aerodynamic Characteristics of the Space Shuttle Orbiter for a Range of Damage Scenarios

Aerodynamic tests in support of the Columbia accident investigation were conducted in two hypersonic wind tunnels at the NASA Langley Research Center, the 20-Inch Mach 6 Air Tunnel and the 20-Inch Mach 6 CF4 Tunnel. The primary purpose of these tests was to measure the forces and moments generated by a variety of outer mold line alterations (damage scenarios) using 0.0075-scale models of the Space Shuttle Orbiter (approximately 10 inches in length). Simultaneously acquired global heat transfer mappings were obtained for a majority of the configurations tested. Test parameters include angles of attack from 38 to 42 deg, unit Reynolds numbers from 0.26 to 3.0 x10^6 per foot, and normal shock density ratios of 5 (Mach 6 air) and 12 (Mach 6 CF4). The damage scenarios evaluated included asymmetric boundary layer transition, gouges in the windward surface acreage thermal protection system tiles, wing leading edge damage (partially and fully missing reinforced carbon-carbon (RCC) panels), holes through the wing from the windward surface to the leeside, deformation of the wing windward surface, and main landing gear door and/or gear deployment. The aerodynamic data were compared to the magnitudes and directions observed in flight, and the heating images were evaluated in terms of the location of the generated disturbances and how these disturbance might relate to the response of discrete gages on the Columbia Orbiter vehicle during entry. The measured aerodynamic increments were generally small in magnitude, as were the flight-derived values during most of the entry. Asymmetric boundary layer transition (ABLT) results were consistent with the flight-derived Shuttle ABLT model, but not with the observed flight trends for STS-107. The partially missing leading edge panel results best matched both the early aerodynamic and heating trends observed in flight. A progressive damage scenario is presented that qualitatively matches the flight observations for the full entry.

Brauckman, Gregory J.↗

Experimental Hypersonic Aerodynamic Characteristics of the Space Shuttle Orbiter for a Range of Damage Scenarios

Aerodynamic tests in support of the Columbia accident investigation were conducted in two hypersonic wind tunnels at the NASA Langley Research Center, the 20-Inch Mach 6 Air Tunnel and the 20-Inch CF4 Tunnel. The primary purpose of these tests was to measure the forces and moments generated by a variety of outer mold line alterations (damage scenarios) using 0.0075-scale models of the Space Shuttle Orbiter. Simultaneously acquired global heat transfer mappings were obtained for a majority of the configurations tested. Test parametrics included angles of attack from 38 to 42 deg, unit Reynolds numbers from 0.3 x 10(exp 6) to 3.0 x 10(exp 6) per foot, and normal shock density ratios of 5 (Mach 6 air) and 12 (CF4). The damage scenarios evaluated included asymmetric boundary layer transition, gouges in the windward surface thermal protection system tiles, wing leading edge damage (partially and fully missing reinforced carbon-carbon (RCC) panels), deformation of the wing windward surface, and main landing gear and/or door deployment. The measured aerodynamic increments for the damage scenarios examined were generally small in magnitude, as were the flight-derived values during most of the entry prior to loss of communication. A progressive damage scenario is presented that qualitatively matches the flight observations for the STS-107 entry.

Brauckmann, Gregory J.↗

Reduction and Analysis of Phosphor Thermography Data With the IHEAT Software Package

Detailed aeroheating information is critical to the successful design of a thermal protection system (TPS) for an aerospace vehicle. This report describes NASA Langley Research Center's (LaRC) two-color relative-intensity phosphor thermography method and the IHEAT software package which is used for the efficient data reduction and analysis of the phosphor image data. Development of theory is provided for a new weighted two-color relative-intensity fluorescence theory for quantitatively determining surface temperatures on hypersonic wind tunnel models; an improved application of the one-dimensional conduction theory for use in determining global heating mappings; and extrapolation of wind tunnel data to flight surface temperatures. The phosphor methodology at LaRC is presented including descriptions of phosphor model fabrication, test facilities and phosphor video acquisition systems. A discussion of the calibration procedures, data reduction and data analysis is given. Estimates of the total uncertainties (with a 95% confidence level) associated with the phosphor technique are shown to be approximately 8 to 10 percent in the Langley's 31-Inch Mach 10 Tunnel and 7 to 10 percent in the 20-Inch Mach 6 Tunnel. A comparison with thin-film measurements using two-inch radius hemispheres shows the phosphor data to be within 7 percent of thin-film measurements and to agree even better with predictions via a LATCH computational fluid dynamics solution (CFD). Good agreement between phosphor data and LAURA CFD computations on the forebody of a vertical takeoff/vertical lander configuration at four angles of attack is also shown. In addition, a comparison is given between Mach 6 phosphor data and laminar and turbulent solutions generated using the LAURA, GASP and LATCH CFD codes. Finally, the extrapolation method developed in this report is applied to the X-34 configuration with good agreement between the phosphor extrapolation and LAURA flight surface temperature predictions. The phosphor process outlined in the paper is believed to provide the aerothermodynamic community with a valuable capability for rapidly obtaining (4 to 5 weeks) detailed heating information needed in TPS design.

Merski, N. Ronald↗

Overset Techniques for Hypersonic Multibody Configurations with the DPLR Solver

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

Hyatt, Andrew James↗

Calibration of PCB-132 Sensors in a Shock Tube

While PCB-132 sensors have proven useful for measuring second-mode instability waves in many hypersonic wind tunnels, they are currently limited by their calibration. Until now, the factory calibration has been all that was available, which is a single-point calibration at an amplitude three orders of magnitude higher than a second-mode wave. In addition, little information has been available about the frequency response or spatial resolution of the sensors, which is important for measuring high-frequency instability waves. These shortcomings make it difficult to compare measurements at different conditions and between different sensors. If accurate quantitative measurements could be performed, comparisons of the growth and breakdown of instability waves could be made in different facilities, possibly leading to a method of predicting the amplitude at which the waves break down into turbulence, improving transition prediction. A method for calibrating the sensors is proposed using a newly-built shock tube at Purdue University. This shock tube, essentially a half-scale version of the 6-Inch shock tube at the Graduate Aerospace Laboratories at Caltech, has been designed to attain a moderate vacuum in the driven section. Low driven pressures should allow the creation of very weak, yet still relatively thin shock waves. It is expected that static pressure rises within the range of second-mode amplitudes should be possible. The shock tube has been designed to create clean, planar shock waves with a laminar boundary layer to allow for accurate calibrations. Stronger shock waves can be used to identify the frequency response of the sensors out to hundreds of kilohertz.

Berridge, Dennis C.↗

NASA's Hydrogen Outpost: The Rocket Systems Area at Plum Brook Station

"There was pretty much a general knowledge about hydrogen and its capabilities," recalled former researcher Robert Graham. "The question was, could you use it in a rocket engine? Do we have the technology to handle it? How will it cool? Will it produce so much heat release that we can't cool the engine? These were the questions that we had to address." The National Aeronautics and Space Administration's (NASA) Glenn Research Center, referred to historically as the Lewis Research Center, made a concerted effort to answer these and related questions in the 1950s and 1960s. The center played a critical role transforming hydrogen's theoretical potential into a flight-ready propellant. Since then NASA has utilized liquid hydrogen to send humans and robots to the Moon, propel dozens of spacecraft across the universe, orbit scores of satellite systems, and power 135 space shuttle flights. Rocket pioneers had recognized hydrogen's potential early on, but its extremely low boiling temperature and low density made it impracticable as a fuel. The Lewis laboratory first demonstrated that liquid hydrogen could be safely utilized in rocket and aircraft propulsion systems, then perfected techniques to store, pump, and cleanly burn the fuel, as well as use it to cool the engine. The Rocket Systems Area at Lewis's remote testing area, Plum Brook Station, played a little known, but important role in the center's hydrogen research efforts. This publication focuses on the activities at the Rocket Systems Area, but it also discusses hydrogen's role in NASA's space program and Lewis's overall hydrogen work. The Rocket Systems Area included nine physically modest test sites and three test stands dedicated to liquid-hydrogen-related research. In 1962 Cleveland Plain Dealer reporter Karl Abram claimed, "The rocket facility looks more like a petroleum refinery. Its test rigs sprout pipes, valves and tanks. During the night test runs, excess hydrogen is burned from special stacks in the best Oklahoma oil field tradition." Besides the Rocket Systems Area, Plum Brook Station also included a nuclear test reactor, a large vacuum tank, a hypersonic wind tunnel, and a full-scale upper-stage rocket stand. The Rocket Systems Area operated from 1961 until NASA shut down all of Plum Brook in 1974. The center reopened Plum Brook in the late 1980s and continues to use several test facilities. The Rocket Systems Area, however, was not restored. Today Plum Brook resembles a nature preserve more than an oil refinery. Lush fields and forests separate the large test facilities. Until recently, the abandoned Rocket Systems Area structures and equipment were visible amongst the greenery. These space-age ruins, particularly the three towers, stood as silent sentinels over the sparsely populated reservation. Few knew the story of these mysterious facilities when NASA removed them in the late 2000s.

Rocket Systems↗

Direct Numerical Simulation of Hypersonic Turbulent Boundary Layer inside an Axisymmetric Nozzle

As a first step toward a study of acoustic disturbance field within a conventional, hypersonic wind tunnel, direct numerical simulations (DNS) of a Mach 6 turbulent boundary layer on the inner wall of a straight axisymmetric nozzle are conducted and the results are compared with those for a flat plate. The DNS results for a nozzle radius to boundary-layer thickness ratio of 5:5 show that the turbulence statistics of the nozzle-wall boundary layer are nearly unaffected by the transverse curvature of the nozzle wall. Before the acoustic waves emanating from different parts of the nozzle surface can interfere with each other and undergo reflections from adjacent portions of the nozzle surface, the rms pressure fluctuation beyond the boundary layer edge increases toward the nozzle axis, apparently due to a focusing effect inside the axisymmetric configuration. Spectral analysis of pressure fluctuations at both the wall and the freestream indicates a similar distribution of energy content for both the nozzle and the flat plate, with the peak of the premultiplied frequency spectrum at a frequency of [(omega)(delta)]/U(sub infinity) approximately 6.0 inside the free stream and at [(omega)(delta)]/U(sub infinity) approximately 2.0 along the wall. The present results provide the basis for follow-on simulations involving reverberation effects inside the nozzle.

Huang, Junji↗