Mach disk in underexpanded exhaust plumes
Mach disk in underexpanded exhaust plume predicted by dividing flow field into subregions
Abbett, M.↗
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Results for “plume flow”
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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Vaporization of graphite in the temperature range of 4000 to 4500 K
The vaporization of graphite under intense laser radiation is considered both theoretically and experimentally. Under intense radiation, the mass-loss rate can be high enough to cause the flow in the laser plume to be supersonic. It is shown that under these conditions the vaporization process is coupled to the plume gasdynamics and the mass-loss rate for graphite is 62% of the free vaporization rate. Experimental results are presented for surface temperatures from 3985 to 4555 K and mass-loss rates from 0.56 to 27.0 g per sq cm sec. The results are used to determine the vapor pressure of graphite in a pressure range of 2 to 11 atm, and the values are shown to be in agreement with the JANAF vapor pressure curve.
Lundell, J. H.↗
Test data from solid propellant plume aerodynamics test program in Ames 6 x 6 foot supersonic wind tunnel (shuttle test FA7) (Ames test 033-66)
The aerodynamic effects of plumes from hot combustion gases in the presence of a transonic external flow field were measured to advance plumes simulation technology, extend a previously acquired data base, and provide data to compare with the effects observed using cold gas plumes. A variety of underexpanded plumes issuing from the base of a strut-mounted ogive-cylinder body were produced by combusting solid propellant gas generators. The gas generator fired in a short-duration mode (200 to 300 msec). Propellants containing 16 percent and 2 percent A1 were used, with chamber pressures from 400 to 1800 psia. Conical nozzles of 15 deg half-angle were tested with area ratios of 4 and 8. Pressures were measured in the gas generator combustion chamber, along the nozzle wall, on the base, and along the body rear exterior. Schlieren photographs were taken for all tests. Test data are presented along with a description of the test setup and procedures.
Hair, L. M.↗
Interactive analysis of thermal imagery
Necessary knowledge is presented on data acquisition and preparation for analysis of thermal imagery of power plant heated discharges remotely sensed from an aircraft, with special emphasis on analog to digital conversion of analog tapes acquired during scanning and to geometrical scaling. The central element in the interactive analysis of thermal imagery is an interactive graphics computer terminal which allows an interpreter to effectively interact with a large-scale computer, providing decisions or data as computations are carried out. A temperature calibration is performed, which the interpreter may test anywhere on the image. When satisfied that calibration is correct, the portion of the image to be analyzed is outlined. Printed and microfiche analyses of the plume are produced. The flow chart of programs for analysis of thermal imagery is presented and discussed in some detail.
Madding, R. P.↗
The direct observation of unidirectional solidification as a function of gravity level
A saturated solution of NH4 Cl-H2O was solidified unidirectionally on earth at 1 g and in a suborbital rocket at 0.001 g. In the 1 g experiment, extensive fluid flow was apparent, creating plumes which carry numerous crystallites vertically above the interface in a manner associated with freckling in castings. During growth the interface grew into the liquid at the same rate as the individual dendrites. When solidified in low g, the interface grew at a slower rate than the individual dendrites. No plumes of liquid carrying crystals were seen, but several crystals did appear and grow ahead of the interface. Growth rates, dendrite morphologies and preferred crystallographic orientations are compared for the two experiment conditions. Results indicate that solidification in low gravity produces more uniform dendrite arm spacings, and tends to be controlled by diffusion and crystallographic orientation.
Johnston, M. H.↗
Microwave radiometer measurement of tidally induced salinity changes off the Georgia coast
A quasi-synoptic survey of tidally induced salinity changes off the Georgia coast was performed by using a L band microwave radiomater onboard a NASA aircraft. Salinity maps were obtained for ebb and flood conditions in order to define the salinity distributions near rivers and sounds and major changes that occur from ebb flow to flood flow. The Savannah River plume dominated the salinity regime and extended out from the Savannah River mouth about 12 km during ebb tidal conditions. The plume merged into a band of low salinity water extending along the Georgia-South Carolina coast which was produced by the many river sources of freshwater entering the coastal waters. The changes in salinity observed offshore of the river plume area were consistent with estimates of the changes that would occur over a typical tidal excursion perpendicular to the observed gradient.
Kendall, B. M.↗
Evaluation of the communications impact of a low power arcjet thruster
The interaction of a 1 kW arcjet thruster plume with a communications signal is evaluated. A two-parameter, source flow equation has been used to represent the far flow field distribution of the arcjet plume in a realistic spacecraft configuration. Modelling the plume as a plasma slab, the interaction of the plume with a 4 GHz communications signal is then evaluated in terms of signal attenuation and phase shift between transmitting and receiving antennas. Except for propagation paths which pass very near the arcjet source, the impacts to transmission appear to be negligible. The dominant signal loss mechanism is refraction of the beam rather than absorption losses due to collisions. However, significant reflection of the signal at the sharp vacuum-plasma boundary may also occur for propagation paths which pass near the source.
Carney, Lynnette M.↗
Evaluation of the communications impact of a low power arcjet thruster
The interaction of a 1 kW arcjet thruster plume with a communications signal is evaluated. A two-parameter, source flow equation has been used to represent the far flow field distribution of the arcjet plume in a realistic spacecraft configuration. Modelling the plume as a plasma slab, the interaction of the plume with a 4 GHz communications signal is then evaluated in terms of signal attenuation and phase shift between transmitting and receiving antennas. Except for propagation paths which pass very near the arcjet source, the impacts to transmission appear to be negligible. The dominant signal loss mechanism is refraction of the beam rather than absorption losses due to collisions. However, significant reflection of the signal at the sharp vacuum-plasma boundary may also occur for propagation paths which pass near the source.
Carney, Lynnette M.↗
Exhaust plumes and their interaction with missile airframes - A new viewpoint
The present, novel treatment of missile airframe-exhaust plume interactions emphasizes their simulation via a formal solution of the Reynolds-averaged Navier-Stokes (RNS) equation and is accordingly able to address the simulation requirements of novel missiles with nonconventional/integrated propulsion systems. The method is made possible by implicit RNS codes with improved artificial dissipation models, generalized geometric capabilities, and improved two-equation turbulence models, as well as by such codes' recent incorporation of plume thermochemistry and multiphase flow effects.
Dash, S. M.↗
Numerical Simulation of the Generation of Axisymmetric Mode Jet Screech Tones
An imperfectly expanded supersonic jet, invariably, radiates both broadband noise and discrete frequency sound called screech tones. Screech tones are known to be generated by a feedback loop driven by the large scale instability waves of the jet flow. Inside the jet plume is a quasi-periodic shock cell structure. The interaction of the instability waves and the shock cell structure, as the former propagates through the latter, is responsible for the generation of the tones. Presently, there are formulas that can predict the tone frequency fairly accurately. However, there is no known way to predict the screech tone intensity. In this work, the screech phenomenon of an axisymmetric jet at low supersonic Mach number is reproduced by numerical simulation. The computed mean velocity profiles and the shock cell pressure distribution of the jet are found to be in good agreement with experimental measurements. The same is true with the simulated screech frequency. Calculated screech tone intensity and directivity at selected jet Mach number are reported in this paper. The present results demonstrate that numerical simulation using computational aeroacoustics methods offers not only a reliable way to determine the screech tone intensity and directivity but also an opportunity to study the physics and detailed mechanisms of the phenomenon by an entirely new approach.
Shen, Hao↗
Radiative Vaporization of Graphite in the Temperature Range of 4000 to 4500 deg K
The vaporization of graphite under intense laser radiation is considered both theoretically and experimentally. Under intense radiation, the mass-loss rate can be high enough to cause the flow in the laser plume to be supersonic. Under these conditions, the vaporization process is coupled to the plume gasdynamics. Experimental results are presented for surface temperatures of 3985 to 4555 K and mass-loss rates from 0.52 to 27.0 g/sq cm sec. The data are used to determine the vapor pressure of graphite in a range of 2 to 11 atm, and the results are shown to be in good agreement with the JANAF vapor pressure curve, if the vaporization coefficients are unity. The assumption of unity vaporization coefficients is shown to be reasonable by a comparison of the present results with other recent vapor pressure results for graphite.
Lundell, John H.↗
Orion Launch Abort Vehicle Attitude Control Motor Testing
Current Orion Launch Abort Vehicle (LAV) configurations use an eight-jet, solid-fueled Attitude Control Motor (ACM) to provide required vehicle control for all proposed abort trajectories. Due to the forward position of the ACM on the LAV, it is necessary to assess the effects of jet-interactions (JI) between the various ACM nozzle plumes and the external flow along the outside surfaces of the vehicle. These JI-induced changes in flight control characteristics must be accounted for in developing ACM operations and LAV flight characteristics. A test program to generate jet interaction aerodynamic increment data for multiple LAV configurations was conducted in the NASA Ames and NASA Langley Unitary Plan Wind Tunnels from August 2007 through December 2009. Using cold air as the simulant gas, powered subscale models were used to generate interaction data at subsonic, transonic, and supersonic test conditions. This paper presents an overview of the complete ACM JI experimental test program for Orion LAV configurations, highlighting ACM system modeling, nozzle scaling assumptions, experimental test techniques, and data reduction methodologies. Lessons learned are discussed, and sample jet interaction data are shown. These data, in conjunction with computational predictions, were used to create the ACM JI increments for all relevant flight databases.
Murphy, Kelly J.↗
Inviscid and Viscous CFD Analysis of Booster Separation for the Space Launch System Vehicle
This paper presents details of Computational Fluid Dynamic (CFD) simulations of the Space Launch System during solid-rocket booster separation using the Cart3D inviscid and Overflow viscous CFD codes. The discussion addresses the use of multiple data sources of computational aerodynamics, experimental aerodynamics, and trajectory simulations for this critical phase of flight. Comparisons are shown between Cart3D simulations and a wind tunnel test performed at NASA Langley Research Center's Unitary Plan Wind Tunnel, and further comparisons are shown between Cart3D and viscous Overflow solutions for the flight vehicle. The Space Launch System (SLS) is a new exploration-class launch vehicle currently in development that includes two Solid Rocket Boosters (SRBs) modified from Space Shuttle hardware. These SRBs must separate from the SLS core during a phase of flight where aerodynamic loads are nontrivial. The main challenges for creating a separation aerodynamic database are the large number of independent variables (including orientation of the core, relative position and orientation of the boosters, and rocket thrust levels) and the complex flow caused by exhaust plumes of the booster separation motors (BSMs), which are small rockets designed to push the boosters away from the core by firing partially in the direction opposite to the motion of the vehicle.
Viscous↗
Model Design and Pre-Test CFD Analysis for a Supersonic Retropropulsion Wind Tunnel Test
Future Mars human lander missions will require using powered descent beginning at supersonic conditions, something which has never been done before at Mars. Significant aerosciences challenges exist due to interactions between the retrorocket exhaust plumes and the freestream flow that alter the aerodynamic behavior of the powered descent vehicle. Historically, wind tunnel tests have been used to study supersonic retropropulsion with inert gas exhaust simulants. Also, SpaceX has successfully decelerated the Falcon 9 first stage numerous times by using powered flight at supersonic conditions. On the computational side, Reynolds-Averaged Navier-Stokes flowfield simulations are regularly completed at full-scale conditions. However, the available ground and flight data do not provide a basis for calibrating the computational uncertainties for aerodynamic interference forces and moments on proposed Mars descent vehicles, either because of insufficient data or dissimilar vehicles geometries and/or conditions. Thus, additional ground testing is needed to continue addressing the aerosciences risks for large-scale human Mars landers. To that end, a retropropulsion wind tunnel test will be conducted in the NASA Langley Unitary Plan Wind Tunnel in 2020. The test is designed to improve upon similar past tests, both in terms of model design and measured data. The test campaign will use subscale model geometries derived from the two current NASA powered descent reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry. This paper covers the reference vehicles, test objectives, model design, scaling parameters, test matrix, and computational analysis to date.
Karl T Edquist↗
Validation of Artemis I Aerothermal Design Models Using Developmental Flight Instrumentation
The inaugural flight of the Space Launch System (SLS) Block 1 launch vehicle, Artemis I, occurred on November 16, 2022, and featured a full suite of Developmental Flight Instrumentation (DFI) that provided aerothermodynamic measurements to assess thermal design and substantiate aerothermodynamic models. The Block 1 launch vehicle aerothermal instrumentation consisted of approximately 277 aerothermal gauges mounted throughout the Orion Multi-Purpose Crew Vehicle (MPCV), Integrated Spacecraft and Payload Element (ISPE), Core Stage (CS) and Solid Rocket Boosters (SRB) and an additional 179 thermal gauges on the Orion Crew Module (CM). Instrumentation included calorimeters, radiometers, pressure transducers, gas temperature probes, and thermocouples. Data was collected from lift-off through CS Main Engine Cut-Off (MECO). The flight data was invaluable for determining aerothermal model performance and developing flight-derived aerothermal environments for flight reconstruction thermal analysis and future SLS aerothermal models. The data offered critical insights into the aerothermodynamic conditions experienced during the launch and ascent of the SLS vehicle. This study compares the flight derived environments to pre-existing design models. The aerothermal models were constructed using MINIVER, the aerothermal engineering code which predicts aerodynamic heating and acts as an integration tool for incorporating databases from computational fluid dynamics (CFD) simulations and wind tunnel test data. The comparisons reveal the fidelity of the design models, highlighting areas where the design models accurately predicted flight conditions and instances where deviations were observed. Preliminary results suggest that while the design models largely aligned with the observed flight data, there were unique observations that reflected needed areas of model refinement. Aerothermal flight data from Artemis I for the SLS Block 1 vehicle will be further utilized to enhance the accuracy of Block 1B and Block 2 aerothermal models, ensuring improved safety and performance for subsequent Artemis missions.
aerothermodynamics↗
Validation of Artemis I Aerothermal Design Models Using Developmental Flight Instrumentation
The inaugural flight of the Space Launch System (SLS) Block 1 launch vehicle, Artemis I, occurred on November 16, 2022, and featured a full suite of Developmental Flight Instrumentation (DFI) that provided aerothermodynamic measurements to assess thermal design and substantiate aerothermodynamic models. The Block 1 launch vehicle aerothermal instrumentation consisted of approximately 277 aerothermal gauges mounted throughout the Orion Multi-Purpose Crew Vehicle (MPCV), Integrated Spacecraft and Payload Element (ISPE), Core Stage (CS) and Solid Rocket Boosters (SRB) and an additional 179 thermal gauges on the Orion Crew Module (CM). Instrumentation included calorimeters, radiometers, pressure transducers, gas temperature probes, and thermocouples. Data was collected from lift-off through CS Main Engine Cut-Off (MECO). The flight data was invaluable for determining aerothermal model performance and developing flight-derived aerothermal environments for flight reconstruction thermal analysis and future SLS aerothermal models. The data offered critical insights into the aerothermodynamic conditions experienced during the launch and ascent of the SLS vehicle. This study compares the flight derived environments to pre-existing design models. The aerothermal models were constructed using MINIVER, the aerothermal engineering code which predicts aerodynamic heating and acts as an integration tool for incorporating databases from computational fluid dynamics (CFD) simulations and wind tunnel test data. The comparisons reveal the fidelity of the design models, highlighting areas where the design models accurately predicted flight conditions and instances where deviations were observed. Preliminary results suggest that while the design models largely aligned with the observed flight data, there were unique observations that reflected needed areas of model refinement. Aerothermal flight data from Artemis I for the SLS Block 1 vehicle will be further utilized to enhance the accuracy of Block 1B and Block 2 aerothermal models, ensuring improved safety and performance for subsequent Artemis missions.
aerothermodynamics↗
NASA Space Launch System Artemis I & II Post Flight Ascent Aerothermal Environments Overview
Since 2011 the Aerosciences Branch/EV33 at NASA Marshall Space Flight Center has been involved with the development of ascent external aerothermal environments for the NASA Space Launch System (SLS) Block 1 launch vehicle for the purposes of supporting thermal analysis and the design of thermal protection systems. The SLS Block 1 Artemis I and II launch vehicles successfully launched from Pad39B at NASA Kennedy Space Center on November 16th, 2022 and April 1st, 2026, respectively. Over 70 aerothermal islands, consisting of over 265 operational instruments captured aerodynamic heating and plume induced environments throughout the launch vehicles. Gauges consisted of calorimeters, radiometers, gas temperature probes, pressure transducers, bi-directional pressure probes and thermocouples. Prior to launch, aerothermal design environment models were generated to predict ascent aerodynamic heating and plume induced environments over a design space that covered a range of vehicle trajectories that varied atmospheric, vehicle performance, and off-nominal, engine-out conditions. Post flight reconstruction models were developed for each flight island using the Day-of-Launch (DOL) Best Equivalent Trajectory (BET) that provided freestream conditions and propulsion system boundary conditions. This paper discusses a summary of the ascent aerothermal environments observed during the flights and the respective modelling approaches and the performance of them through comparisons of flight data and predictions.
aerothermodynamics↗
Estuarine and coastal water dynamics controlling sediment movement and plume development in Long Island Sound
The author has identified the following significant results. As the Connecticut River flows into Long Island Sound, large plumes develop during the mixing of ocean and estuarine waters. Plumes were delineated for July 28, October 8, October 27, and December 2, 1972, by analyzing ERTS-1 imagery with the SRI Electronic Satellite Image Analysis Console (ESIAC). Because the chemical and physical composition of the plume and ocean water were not too different, the ESIAC was utilized to expand the scenes and subject the transparencies to varying combinations of viewing techniques to identify and delineate the plumes. Best results were obtained when band 5 transparencies were used. Indications are, when the scene being analyzed is predominantly in the first two steps of the gray scale, it is best to use the negative transparencies. When the analysis is being done above the first two steps of the gray scale, it is best to use the positive transparencies.
Ruggles, F. H., Jr.↗