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This paper describes the results of a preliminary evaluation of a 1.35 kW anode layer thruster. Data on performance and plume characteristics will be presented. Data on thruster characteristics such as thrust, plume current densities, discharge voltage, discharge current, and mass flow rate along with the results of a 500-hour endurance test wear will be presented.
In the last several years, astronomers have discovered several exoplanets with masses less than 10 times that of the Earth [1]. Despite the likely abundance of Earth-sized planets, little is known about the pathways through which these planets evolve to become habitable or uninhabitable. Venus and Earth have similar planetary radii and solar orbital distance, and therefore offer a chance to study in detail the divergent evolution of two objects that now have radically different climates. Understanding the extent, duration, and types of volcanism present on Venus is an important step towards understanding how volatiles released from the interior of Venus have influenced the development of the atmosphere. Placing constraints on the extent of explosive volcanism on Venus can provide boundary conditions for timing, volumes, and altitudes for atmospheric injection of volatiles. In addition, atmospheric properties such as near-surface temperature and density affect how interior heat and volatiles are released. Radar image data for Venus can be used to determine the physical properties of volcanic deposits, and in particular, they can be used to search for evidence of pyroclastic deposits that may result from explosive outgassing of volatiles. For explosive volcanism to occur with the current high atmospheric pressure, magma volatile contents must be higher than is typical on Earth (at least 2-4% by weight) [2,3]. In, addition, pyroclastic flows should be more prevalent on Venus than convective plumes and material may not travel as far from the vent source as it would on Earth [3]. Areas of high radar backscatter with wispy margins that occur near concentric fractures on Sapho Patera [4] and several coronae in Eastern Eistla Regio [5] have been attributed to collapse of eruption columns and runout of rough materials.
The temperature and pressure conditions experienced by rocks during an impact event can be constrained using petrologic and microstructural analysis and is crucial to providing ground truth to the impact cratering process. Suevite is a polymict, impact melt-bearing breccia, specific to Ries crater in Germany. There are competing models for suevite formation and emplacement, such as clastic flows pushed out of the crater rim or ejecta plume fallback. Knowledge of the temperature and pressure pathways recorded by grains within the suevite can help distinguish between these and other models. The accessory phase zircon (ZrSiO4) and its high-pressure polymorph reidite are particularly useful in such circumstances as they are highly refractory minerals that can record the high-temperature and/or high-pressure conditions of an impact event. Here we present evidence for a wide array of temperature and pressure conditions recorded in zircon grains within a single thin section of suevite. Zircons in this study range from unshocked to highly shocked (>53 GPa), and record temperatures more than 1673 °C. These findings confirm previous studies concluding that suevites contain material exposed to very diverse pressure and temperature conditions during initial shock compression and excavation but do not, as a whole, experience extreme temperatures (>1673 °C) or pressures (>30 GPa).
A low-boom supersonic transport was designed for a cruise Mach of 1.7 and 40 passengers. This low-boom aircraft, referred to as the Mach 1.7 40-PAX concept, was generated using computational fluid dynamics (CFD) based sonic boom analysis at the start of overland cruise (SOC). The engine for the Mach 1.7 40-PAX concept was designed using the Numerical Propulsion System Simulation (NPSS) and modeled as a flow-through nacelle for CFD analysis. To understand how the engine plume affects the undertrack ground signature, the Mach 1.7 40-PAX concept is redesigned using an aeropropulsive CFD simulation. A process for approximation of the NPSS engine at SOC by a CFD engine is developed. The generated CFD engine has the identical nozzle boundary conditions and approximately the same mass flow and thrust as those of the NPSS engine at SOC. Then, the configuration with the CFD engines is optimized to approximately restore the low-boom characteristics of the Mach 1.7 40-PAX concept. Finally, the CFD simulation data for the optimized concept with the CFD engines is used to calibrate the low-fidelity aerodynamic analyses for mission analysis of this concept. The cyclic dependency of the involved disciplinary analyses is resolved using an iteration method for a consistent coupling of the mission analysis, NPSS engine analysis, and low-boom redesign using the aeropropulsive CFD simulation. This low-boom redesign study is used as an example to demonstrate how the propulsion-airframe integration could be implemented for conceptual design of low-boom supersonic transports.
The interaction of two nozzles exhausting into vacuum generates a complex three-dimensional shock structure. The shock structure and resulting plume flowfield is characterized by the nozzle separation distance. For the appropriate range of penetration Knudsen numbers, the analysis of this shock structure can be suitably accomplished through a Monte Carlo simulation. This paper describes the application of a general three-dimensional Monte Carlo simulation on the Connection Machine CM-2 to the analysis of the plume self-interaction shock in the near field. Results are presented for two cases, corresponding to a small and a large nozzle separation distance. The results correctly reproduce the expected flow features and demonstrate the ability of this method to properly simulate the start of the plume self-interaction shock. This has significance not only for allowing analysis of the self-interacting plume in the near field, but also for allowing the subsequent simulation of the far field flow through the use of a continuation downstream exit boundary.
The primary test objective was to define the base pressure environment of the first and second stage mated vehicle in a supersonic flow field from Mach 2.60 through 3.50 with simulated rocket engine exhaust plumes. The secondary objective was to obtain the pressure environment of the Orbiter at various vent port locations at these same freestream conditions. Data were obtained at angles of attack from -4 deg through +4 deg at zero yaw, and at yaw angles from -4 deg through +4 deg at zero angle of attack, with rocket plume sizes varying from smaller than nominal to much greater than nominal. Failed Orbiter engine data were also obtained. Elevon hinge moments and wing panel load data were obtained during all runs. Photographs of test equipment and tested configurations are shown.
Sound sources in the plumes of unheated round jets, in the Mach number range 0.6 to 1.8, were investigated experimentally using "casuality" approach, where air density fluctuations in the plumes were correlated with the far field noise. The air density was measured using a newly developed Molecular Rayleigh scattering based technique, which did not require any seeding. The reference at the end provides a detailed description of the measurement technique.
Abstract Coal mining and reclamation can have a profound influence on hydrogeologic systems, with clear consequences for groundwater quality, yet their long-term influence on downgradient water quality over time following reclamation is less well documented. Geochemical trends were evaluated in water quality downgradient of a fully reclaimed landscape at the former Big Sky Mine in the Rosebud Creek watershed (southeastern Montana, USA), over a 3-year period (2020–2022), including bond release in 2022. Within 6 km downgradient from the reclaimed area, sulfate concentrations decreased from approximately 3500 to 1800 mg l −1 within the Miller Coulee alluvial aquifer. Major ions, δ 34 S SO4 values, and residence time tracers suggest that the observed decreases in sulfate concentration result from a combination of dilution by mixed-age inflows and incomplete transit of the high salinity plume from the mine boundary. Both bedrock and alluvial aquifers of the Rosebud Creek corridor contained contributions of millennia-old regional groundwater, which may serve to mitigate mine-derived high salinity waters. Rosebud Creek, which traverses the outflow zone of Miller Coulee in the study area, exhibited high sulfate concentrations during low flows and consistent downgradient increases in sulfate concentration. The possibility of plume dynamics in Miller Coulee suggests that the greatest water quality impacts may not yet have reached Rosebud Creek.
A low-boom supersonic transport was designed for a cruise Mach of 1.7 and 40 passengers. This low-boom aircraft, referred to as the Mach 1.7 40-PAX concept, was generated using computational fluid dynamics (CFD) based sonic boom analysis at the start of overland cruise (SOC). The engine for the Mach 1.7 40-PAX concept was designed using the Numerical Propulsion System Simulation (NPSS) and modeled as a flow-through nacelle for CFD analysis. To understand how the engine plume affects the undertrack ground signature, the Mach 1.7 40-PAX concept is redesigned using an aeropropulsive CFD simulation. A process is developed for approximation of the NPSS engine at SOC by a CFD engine for aeropropulsive CFD simulation. The generated CFD engine has the identical nozzle boundary conditions and approximately the same mass flow and thrust as those of the NPSS engine at SOC. Then, the configuration with the CFD engines is optimized to approximately restore the low-boom characteristics of the Mach 1.7 40-PAX concept. Finally, the CFD simulation data for the optimized concept with the CFD engines is used to calibrate the low-fidelity aerodynamic analyses for mission analysis of this concept. The cyclic dependency of the involved disciplinary analyses is resolved using an iteration method for a consistent coupling of the mission analysis, NPSS engine analysis, and low-boom redesign using the aeropropulsive CFD simulation. This low-boom redesign study is used as an example to demonstrate how the propulsion-airframe integration could be implemented for conceptual design of low-boom supersonic transports.
Computational Fluid Dynamics (CFD) technology has been used to assess the exhaust plume pollutant environment of the RD-170 engine hot-firing on the F1 Test Stand at Marshall Space Flight Center. Researchers know that rocket engine hot-firing has the potential for forming thermal nitric oxides (NO(x)), as well as producing carbon monoxide (CO) when hydrocarbon fuels are used. Because of the complicated physics involved, however, little attempt has been made to predict the pollutant emissions from ground-based engine testing, except for simplified methods which can grossly underpredict and/or overpredict the pollutant formations in a test environment. The objective of this work, therefore, has been to develop a technology using CFD to describe the underlying pollutant emission physics from ground-based rocket engine testing. This resultant technology is based on a three-dimensional (3D), viscous flow, pressure-based CFD formulation, where wet CO and thermal NO finite-rate chemistry mechanisms are solved with a Penalty Function method. A nominal hot-firing of a RD-170 engine on the F1 stand has been computed. Pertinent test stand flow physics such as the multiple-nozzle clustered engine plume interaction, air aspiration from base and aspirator, plume mixing with entrained air that resulted in contaminant dilution and afterburning, counter-afterburning due to flame bucket water-quenching, plume impingement on the flame bucket, and restricted multiple-plume expansion and turning have been captured. The predicted total emission rates compared reasonably well with those of the existing hydrocarbon engine hot-firing test data.
With the development of new manned spaceflight capabilities including NASA's Orion capsule and the Space-X Dragon capsule, there is a renewed importance of understanding the dynamics of ablative thermal protection systems. To this end, a new inductively coupled plasma torch facility is being developed at UT-Austin. The torch operates on argon and/or air at plasma powers up to 50 kW. In the present configuration the flow issues from a low-speed subsonic nozzle and the hot plume is characterized using slug calorimetry and emission spectroscopy. Preliminary measurements using emission spectroscopy have indicated that the torch is capable of producing an air plasma with a temperature between 6,000 K and 8,000 K depending on the power and flow settings and an argon plasma with a temperature of approximately 12,000 K. The operation envelope was measured, and heat flux measured for every point within the envelope using both a slug calorimeter and a Gardon gauge heat flux sensor. The torch was found to induce a stagnation point heat flux of between 90 and 225 W/sq cm.
A key capability required for the exploration of planetary bodies is the ability to land on the surface. Previous work performed by NASA and other institutions has primarily focused on landing small spacecraft on planetary surfaces and the associated small-to-medium thrusters required for the soft landing. In the case of human exploration—particularly the establishment of long duration exploration and habitation outposts—the ability to land large landers, such as the SpaceX Starship, is necessary. These larger landing systems require the use of more powerful engines, with higher engine exhaust temperatures and higher landing loads. Understanding the excavation of material by the engines, as well as the potential for the landing legs to sink into the subsurface, is key in ensuring reliable and safe landings. A further improvement in landing reliability can be achieved by constructing landing / launch pads, especially with in-situ resources. Some material excavation by the plume is inevitable, leaving at least a portion of the surface scoured and uneven under the lander and ejecting regolith particles and rocks at very high velocities. One possible solution would be to robotically build landing / launch pads (ideally autonomously) at the destination using in-situ materials. In this case, the first one or few landers will need to land on unimproved surfaces at higher risk; however, they would bring the required equipment to build the landing pads with mostly local resources, thus increasing the reliability of safe landing for subsequent larger landers. A number of methods to build in-situ landing and launch pads have already been developed. These methods include no or some addition of required binder additives to the local regolith material, different processing approaches and result in varying landing pad strengths. A sub-scale rocket engine plume, was used to simulate some of the conditions of a landing on the Moon to assess the effectiveness of various materials for an in-situ built landing pad, The GO2/GCH4 rocket engine fired on 1m2 area coupons of representative pad materials. The results will allow continued development towards materials that satisfy the landing pad properties required for the effective risk reduction and increased reliability for landing people and equipment on the lunar surface. This work contained two parts: (1) computer modeling of a large rocket engine plume interacting with regolith on the Moon, using the Granular Gas Flow Solver (GGFS) provided by CFD Research Corporation as well as other computational fluid dynamics codes (CFD) such as Loci/CHEM. (2) Developing landing/launch pad materials that could be used for in-situ construction on the lunar surface in the future, to mitigate the calculated effects of a large vehicle rocket engine landing and launching on the Moon. The resulting computed values of plume impingement surface temperature, stagnation pressure, gas velocity, shear stress and heat flux were then matched as closely as possible in the Earth’s atmosphere in a sub-scale rocket engine GO2/CH4 test which was provided by Masten Space Systems in Mojave, California. The rocket engine was mounted on a test stand with vertical translation capabilities so that the landing operations of a lander could be simulated (Figure 1).The pad materials test coupons were placed at the surface of a large bin containing simulated lunar basalt regolith and subjected to a test firing as shown in Figure 2. The results of this testing will be presented with related findings, analysis and discussion.
Tritium processing facilities may release tritium oxide (HTO) to the atmosphere which poses potential health risks to exposed co-located workers and to offsite individuals. Most radiological consequence analyses determine HTO dose by applying Gaussian plume models to simulate the transport of HTO. Within these models, deposition velocity is used to assess the sum of all deposition processes acting on the plume. While this may account for vegetative and soil uptake or respiration processes, it may currently lack inclusion of the complex interactions within heterogeneous forested environments. In this complex morphology, dispersion patterns are significantly altered by changing flow regimes above and below the forest canopy and by the transfer of plume material across the canopy boundary. To determine the effects of a heterogeneous forest canopy on an airborne HTO plume, a Gaussian plume model coupled with an advection-diffusion plume model was applied to estimate transport in the free atmosphere above the forest and within the forest canopy and understory. During 2012, wind speed and wind direction measurements taken at 5 heights, ranging from 2-m to 28-m, on an instrumented meteorological tower located in a loblolly pine forest at the Department of Energy (DOE) Savannah River Site (SRS), near Aiken, SC. From these measurements, model predictions were made over a full spectrum of meteorological conditions. Deposition and resuspension velocities were calculated based on the model-predicted flux of plume material across the top of the forest canopy. Additionally, net deposition velocity of the plume material was calculated as the difference between the deposition and resuspension velocities. The 1st and 5th percentile net deposition velocities were estimated to be 0.7 cm s-1 and 1.2 cm s-1, respectively.
The near field plum of a 1 kW class arcjet thruster was investigated using electrostatic probes of various geometries. The electron number densities and temperatures were determined in a simulated hydrazine plume at axial distances between 3 cm (1.2 in) and 15 cm (5.9 in) and radial distances extending to 10 cm (3.9 in) off centerline. Values of electron number densities obtained using cylindrical and spherical probes of different geometries agreed very well. The electron density on centerline followed a source flow approximation for axial distances as near as 3 cm (1.2 in) from the nozzle exit plane. The model agreed well with previously obtained data in the far field. The effects of propellant mass flow rate and input power level were also studied. Cylindrical probes were used to obtain ion streamlines by changing the probe orientation with respect to the flow. The effects of electrical configuration on the plasma characteristics of the plume were also investigated by using a segmented anode/nozzle thruster. The results showed that the electrical configuration in the nozzle affected the distribution of electrons in the plume.
The near field plume of a 1 kW class arcjet thruster was investigated using electrostatic probes of various geometries. The electron number densities and temperatures were determined in a simulated hydrazine plume at axial distances between 3 cm (1.2 in.) and 15 cm (5.9 in.) and radial distances extending to 10 cm (3.9 in.) off centerline. Values of electron number densities obtained using cylindrical and spherical probes of different geometries agreed very well. The electron density on centerline followed a source flow approximation for axial distances as near as 3 cm (1.2 in.) from the nozzle exit plane. The model agreed well with previously obtained data in the far field. The effects of propellant mass flow rate and input power level were also studied. Cylindrical probes were used to obtain ion streamlines by changing the probe orientation with respect to the flow. The effects of electrical configuration on the plasma characteristics of the plume were also investigated by using a segmented anode/nozzle thruster. The results showed that the electrical configuration in the nozzle affected the distribution of electrons in the plume.
Plasma properties in the plume produced by a 3+1 NEXT thruster array operating at full power were mapped using a series of planar Langmuir probes. The Langmuir probes were swept across the diameters of each thruster as well as the centerline of the array at multiple downstream axial locations to produce a plasma map of the plume produced by the array. Such maps yielded the spatial distribution of the plasma density, the electron temperature and the plasma potential in the near field of the array. This spatial information provides insight into local plasma particle flow. Flow direction is particularly important from both an array lifetime and spacecraft-plume plasma interaction standpoint. The variation in the plume plasma parameters tended to vary in a manner consistent with both plume shape and overlap of adjacent plumes.
Reynolds-Averaged Navier-Stokes (RANS) simulations were performed for a commercial supersonic transport aircraft concept and experimental hardware models designed to represent the installed propulsion system of the conceptual aircraft in an upcoming test campaign. The purpose of the experiment is to determine the effects of jet-surface interactions from supersonic aircraft on airport community noise. RANS simulations of the commercial supersonic transport aircraft concept were performed to relate the representative experimental hardware to the actual aircraft. RANS screening simulations were performed on the proposed test hardware to verify that it would be free from potential rig noise and to predict the aerodynamic forces on the model hardware to assist with structural design. The simulations showed a large region of separated flow formed in a junction region of one of the experimental configurations. This was dissimilar with simulations of the aircraft and could invalidate the noise measurements. This configuration was modified and a subsequent RANS simulation showed that the size of the flow separation was greatly reduced. The aerodynamic forces found on the experimental models were found to be relatively small when compared to the expected loads from the model’s own weight.Reynolds-Averaged Navier-Stokes (RANS) simulations were completed for two configurations of a three-stream inverted velocity profile (IVP) nozzle and a baseline single-stream round nozzle (mixed-flow equivalent conditions). For the Sideline and Cutback flow conditions, while the IVP nozzles did not reduce the peak turbulent kinetic energy on the lower side of the jet plume, the IVP nozzles did significantly reduce the size of the region of peak turbulent kinetic energy when compared to the jet plume of the baseline nozzle cases. The IVP nozzle at Sideline conditions did suffer a region of separated flow from the inner stream nozzle splitter that did produce an intense, but small, region of turbulent kinetic energy in the vicinity of the nozzle exit. When viewed with the understanding that jet noise is directly related to turbulent kinetic energy, these IVP nozzle simulations show the potential to reduce noise to observers located below the nozzle. However, these RANS simulations also show that some modifications may be needed to prevent the small region of separated flow-induced turbulent kinetic energy from the inner stream nozzle splitter at Sideline conditions.
Inlets and exhaust nozzles are common place in the world of flight. Yet, many aerodynamic simulation packages do not provide a method of modelling such high energy boundaries in the flow field. For the purposes of aerodynamic simulation, inlets and exhausts are often fared over and it is assumed that the flow differences resulting from this assumption are minimal. While this is an adequate assumption for the prediction of lift, the lack of a plume behind the aircraft creates an evacuated base region thus effecting both drag and pitching moment values. In addition, the flow in the base region is often mis-predicted resulting in incorrect base drag. In order to accurately predict these quantities, a method for specifying inlet and exhaust conditions needs to be available in aerodynamic simulation packages. A method for a first approximation of a plume without accounting for chemical reactions is added to the Cartesian mesh based aerodynamic simulation package CART3D. The method consists of 3 steps. In the first step, a components approach where each triangle is assigned a component number is used. Here, a method for marking the inlet or exhaust plane triangles as separate components is discussed. In step two, the flow solver is modified to accept a reference state for the components marked inlet or exhaust. In the third step, the flow solver uses these separated components and the reference state to compute the correct flow condition at that triangle. The present method is implemented in the CART3D package which consists of a set of tools for generating a Cartesian volume mesh from a set of component triangulations. The Euler equations are solved on the resulting unstructured Cartesian mesh. The present methods is implemented in this package and its usefulness is demonstrated with two validation cases. A generic missile body is also presented to show the usefulness of the method on a real world geometry.