Results for “plume impingement” — Page 14

NASA NTRS2004

Exhaust Simulation Testing of a Hypersonic Airbreathing Model at Transonic Speeds

An experimental study was performed to examine jet-effects for an airframe-integrated, scramjet-rocket combined-cycle vehicle configuration at transonic test conditions. This investigation was performed by testing an existing exhaust simulation wind tunnel model, known as Model 5B, in the NASA Langley 16-Ft. Transonic Tunnel. Tests were conducted at freestream Mach numbers from 0.7 to 1.2, at angles of attack from 2 to +14 degrees, and at up to seven nozzle static pressure ratio values for a set of horizontal-tail and body-flap deflections. The model aftbody, horizontal tails, and body flaps were extensively pressure instrumented to provide an understanding of jet-effects and control-surface/plume interactions, as well as for the development of analytical methodologies and calibration of computational fluid dynamic codes to predict this type of flow phenomenon. At all transonic test conditions examined, the exhaust flow at the exit of the internal nozzle was over-expanded, generating an exhaust plume that turned toward the aftbody. Pressure contour plots for the aftbody of Model 5B are presented for freestream transonic Mach numbers of 0.70, 0.95, and 1.20. These pressure data, along with shadowgraph images, indicated the impingement of an internal plume shock and at least one reflected shock onto the aftbody for all transonic conditions tested. These results also provided evidence of the highly three-dimensional nature of the aftbody exhaust flowfield. Parametric testing showed that angle-of-attack, static nozzle pressure ratio, and freestream Mach number all affected the exhaust-plume size, exhaust-flowfield shock structure, and the aftbody-pressure distribution, with Mach number having the largest effect. Integration of the aftbody pressure data showed large variations in the pitching moment throughout the transonic regime.

Huebner, Lawrence D.↗

NASA NTRS2008

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility: Steady Predictions - Part 1

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. Advanced numerical tools assume greater significance in supporting testing and design of high altitude testing facilities and plume induced testing environments of high thrust engines because of the greater inter-dependence and synergy in the functioning of the different sub-systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Facility designs in this case will require a complex network of diffuser ducts, steam ejector trains, fast operating valves, cooling water systems and flow diverters that need to be characterized for steady state performance. In this paper, we will demonstrate with the use of CFD analyses s advanced capability to evaluate supersonic diffuser and steam ejector performance in a sub-scale A-3 facility at NASA Stennis Space Center (SSC) where extensive testing was performed. Furthermore, the focus in this paper relates to modeling of critical sub-systems and components used in facilities such as the A-3 facility. The work here will address deficiencies in empirical models and current CFD analyses that are used for design of supersonic diffusers/turning vanes/ejectors as well as analyses for confined plumes and venting processes. The primary areas that will be addressed are: (1) supersonic diffuser performance including analyses of thermal loads (2) accurate shock capturing in the diffuser duct; (3) effect of turning duct on the performance of the facility (4) prediction of mass flow rates and performance classification for steam ejectors (5) comparisons with test data from sub-scale diffuser testing and assessment of confidence levels in CFD based flowpath modeling of the facility. The analyses tools used here expand on the multi-element unstructured CFD which has been tailored and validated for impingement dynamics of dry plumes, complex valve/feed systems, and high pressure propellant delivery systems used in engine and component test stands at NASA SSC. The analyses performed in the evaluation of the sub-scale diffuser facility explored several important factors that influence modeling and understanding of facility operation such as (a) importance of modeling the facility with Real Gas approximation, (b) approximating the cluster of steam ejector nozzles as a single annular nozzle, (c) existence of mixed subsonic/supersonic flow downstream of the turning duct, and (d) inadequacy of two-equation turbulence models in predicting the correct pressurization in the turning duct and expansion of the second stage steam ejectors. The procedure used for modeling the facility was as follows: (i) The engine, test cell and first stage ejectors were simulated with an axisymmetric approximation (ii) the turning duct, second stage ejectors and the piping downstream of the second stage ejectors were analyzed with a three-dimensional simulation utilizing a half-plane symmetry approximation. The solution i.e. primitive variables such as pressure, velocity components, temperature and turbulence quantities were passed from the first computational domain and specified as a supersonic boundary condition for the second simulation. (iii) The third domain comprised of the exit diffuser and the region in the vicinity of the facility (primary included to get the correct shock structure at the exit of the facility and entrainment characteristics). The first set of simulations comprising the engine, test cell and first stage ejectors was carried out both as a turbulent real gas calculation as well as a turbulent perfect gas calculation. A comparison for the two cases (Real Turbulent and Perfect gas turbulent) of the Ma Number distribution and temperature distributions are shown in Figures 1 and 2 respectively. The Mach Number distribution shows small yet distinct differences between the two cases such as locations of shocks/shock reflections and a slightly different impingement point on the wall of the diffuser from the expansion at the exit of the nozzle. Similarly the temperature distribution indicates different flow recirculation patterns in the test cell. Both cases capture all the essential flow phenomena such as the shock-boundary layer interaction, plume expansion, expansion of the first stage ejectors, mixing between the engine plume and the first stage ejector flow and pressurization due to the first stage ejectors. The final paper will discuss thermal loads on the walls of the diffuser and cooling mechanisms investigated.

Allgood, Daniel C.↗

NASA NTRS2010

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility: Steady Predictions - Part 1

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. Advanced numerical tools assume greater significance in supporting testing and design of high altitude testing facilities and plume induced testing environments of high thrust engines because of the greater inter-dependence and synergy in the functioning of the different sub-systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Facility designs in this case will require a complex network of diffuser ducts, steam ejector trains, fast operating valves, cooling water systems and flow diverters that need to be characterized for steady state performance. In this paper, we will demonstrate with the use of CFD analyses s advanced capability to evaluate supersonic diffuser and steam ejector performance in a sub-scale A-3 facility at NASA Stennis Space Center (SSC) where extensive testing was performed. Furthermore, the focus in this paper relates to modeling of critical sub-systems and components used in facilities such as the A-3 facility. The work here will address deficiencies in empirical models and current CFD analyses that are used for design of supersonic diffusers/turning vanes/ejectors as well as analyses for confined plumes and venting processes. The primary areas that will be addressed are: (1) supersonic diffuser performance including analyses of thermal loads (2) accurate shock capturing in the diffuser duct; (3) effect of turning duct on the performance of the facility (4) prediction of mass flow rates and performance classification for steam ejectors (5) comparisons with test data from sub-scale diffuser testing and assessment of confidence levels in CFD based flowpath modeling of the facility. The analyses tools used here expand on the multi-element unstructured CFD which has been tailored and validated for impingement dynamics of dry plumes, complex valve/feed systems, and high pressure propellant delivery systems used in engine and component test stands at NASA SSC. The analyses performed in the evaluation of the sub-scale diffuser facility explored several important factors that influence modeling and understanding of facility operation such as (a) importance of modeling the facility with Real Gas approximation, (b) approximating the cluster of steam ejector nozzles as a single annular nozzle, (c) existence of mixed subsonic/supersonic flow downstream of the turning duct, and (d) inadequacy of two-equation turbulence models in predicting the correct pressurization in the turning duct and expansion of the second stage steam ejectors. The procedure used for modeling the facility was as follows: (i) The engine, test cell and first stage ejectors were simulated with an axisymmetric approximation (ii) the turning duct, second stage ejectors and the piping downstream of the second stage ejectors were analyzed with a three-dimensional simulation utilizing a half-plane symmetry approximation. The solution i.e. primitive variables such as pressure, velocity components, temperature and turbulence quantities were passed from the first computational domain and specified as a supersonic boundary condition for the second simulation. (iii) The third domain comprised of the exit diffuser and the region in the vicinity of the facility (primary included to get the correct shock structure at the exit of the facility and entrainment characteristics). The first set of simulations comprising the engine, test cell and first stage ejectors was carried out both as a turbulent real gas calculation as well as a turbulent perfect gas calculation. A comparison for the two cases (Real Turbulent and Perfect gas turbulent) of the Ma Number distribution and temperature distributions are shown in Figures 1 and 2 respectively.

Allgood, Daniel C.↗

NASA NTRS1989

Assessment of impingement effects in the isentropic core of a small satellite control thruster plume

Two computational techniques commonly employed in the calculation of rocket and thruster expansion plumes are assessed. These are the method of characteristics (MOC), which is derived from the continuum Euler equations, and the direct simulation Monte Carlo (DSMC) method, which adopts a discrete particle approach. These techniques vary both in the computational expense and in the accuracy and detail of the solutions that they provide, depending upon the regime of application. The assessment is made with reference to the plume expanding from a small monopropellant hydrazine thruster and concentrates on the isentropic core of the jet for the flow regime lying between the continuum and free molecular limits. It is found that the more numerically intensive DSMC method offers the better correspondence to the available experimental data. In addition, large differences in typical impingement effects such as drag force and heat transfer are found at the free molecular limit of the plume expansion for the two predictive techniques. It is concluded that accurate estimation of impingement potential may only be achieved through application of the discrete particle method.

Boyd, I. D.↗

NASA NTRS2026

Investigation of Bipropellant Plume-Induced Contamination Effects on Coverglass Materials

Contamination and degradation of external spacecraft materials by unburned and partially combusted species from bipropellant thruster plumes has long been observed as a key component of the induced space environment. Space shuttle flight experiments and returned flight hardware from the International Space Station (ISS) have both experienced microscopic impact features induced by high-velocity thruster plume droplets. Analytical results have shown that droplet impingement angle relative to a receiving surface plays a key role in the surface damage. Although impacts with normal impingement angles contribute more severely to surface degradation than highly oblique angles, surface effects at higher impingement angles should not be dismissed. Thruster plume-induced materials degradation is a complex phenomenon that depends on a variety of parameters, including but not limited to material type, system temperature and pressure, plume composition, and thruster firing specifications such as number of pulses, pulse duration, and sample distance from the thruster. For space applications, attaining the vacuum pressure and temperature conditions necessary for flight-like plume expansion and exposure conditions is not a trivial task. The German Aerospace Center (Deutsches Zentrum für Luft- und Raumfahrt, DLR) is a facility uniquely capable of simulating such conditions. Test coupons were exposed to bipropellant thruster firings under high vacuum at the DLR facility. Percent area coverage (PAC) and droplet size distributions were evaluated for the uncoated and coated solar array coverglass materials over a range of impingement angles (0̊ to 75̊). A post-test imaging workflow was developed that aimed to quantify changes in sample surface morphology obtained from scanning electron microscopy (SEM) images using the Image Processing and Analysis in Java (ImageJ) tool; an opensource image processing software. The goal was to create a framework through which to evaluate the effect of bipropellant-induced PAC and droplet size distribution on solar array coverglass optical transmission losses. Understanding this relationship is important because optical transmission losses are known to lead to current reduction in solar power generation systems. In addition to the development of surface characterization workflows, valuable lessons learned as they pertain to future investigations and experiments will be discussed. The authors hope that sharing these lessons will facilitate more utilization of DLR’s unique capabilities as well as open the conversation for how best to address experimental characterization of flight-like plume expansion and its impacts on materials surface degradation effects.

Gateway↗

NASA NTRS1980

Shuttle program standard maneuver sequences for orbiter/upper-stage separation SSUS-A, SSUS-D, and IUS

Descriptions of standard post-ejection maneuver sequences for the deployment of IUS, SSUS-A, and SSUS-D upper stages from the space shuttle orbiter are presented. The sequences were designed to satisfy requirements for limiting the damage inflicted on the orbiter by upper-stage exhaust particles, subject to a further requirement for minimizing the impingement of orbiter thruster plumes on the deployed payload. In all cases it was assumed that the orbital maneuvering system engines would be used to apply the orbiter's major separation velocity increment.

Wilson, S. W.↗

NASA NTRS1995

Pressure And Thermal Modeling Of Rocket Launches

Report presents mathematical model for use in designing rocket-launching stand. Predicts pressure and thermal environment, as well as thermal responses of structures to impinging rocket-exhaust plumes. Enables relatively inexperienced analyst to determine time-varying distributions and absolute levels of pressure and heat loads on structures.

Sheldon D Smith↗

NASA NTRS1998

Thruster Injector Faceplate Testing in Support of the Aerojet Rocket-Based Combined Cycle (RBCC) Concept

To satisfy RBCC rocket thruster requirements of high performance and a minimum amount of free hydrogen at plume boundary, a new impinging injector element using gaseous hydrogen and gaseous oxygen as the propellants has been designed. Analysis has shown that this injector design has potential to provide a high specific impulse (Isp) while minimizing the amount of free hydrogen that is available to be burned with incoming secondary flow. Past studies and test programs have shown that gas/gas-impinging elements typically result in high injector face temperatures due to combustion occurring close to the face. Since this design is new, there is no hot fire experience with this element. Objectives of this test program were to gain experience and hot fire test data on this new rocket thruster element design and injector faceplate pattern. Twenty-two hot fire tests were run with maximum mixture ratio (MR) and chamber pressure (Pc) obtained at 7.25 and 1,822 psia, respectively. Post-test scanning microscope (SEM) images show only slight faceplate erosion during testing. This injector element design performed well and can be operated at design conditions: (1) Pc of 2,000 psia and MR of 7.0 and (2) Pc of 1,000 psia and MR of 5.0.

Fazah, M. M.↗

NASA NTRS2016

Improving Fidelity of Launch Vehicle Liftoff Acoustic Simulations

Launch vehicles experience high acoustic loads during ignition and liftoff affected by the interaction of rocket plume generated acoustic waves with launch pad structures. Application of highly parallelized Computational Fluid Dynamics (CFD) analysis tools optimized for application on the NAS computer systems such as the Loci/CHEM program now enable simulation of time-accurate, turbulent, multi-species plume formation and interaction with launch pad geometry and capture the generation of acoustic noise at the source regions in the plume shear layers and impingement regions. These CFD solvers are robust in capturing the acoustic fluctuations, but they are too dissipative to accurately resolve the propagation of the acoustic waves throughout the launch environment domain along the vehicle. A hybrid Computational Fluid Dynamics and Computational Aero-Acoustics (CFD/CAA) modeling framework has been developed to improve such liftoff acoustic environment predictions. The framework combines the existing highly-scalable NASA production CFD code, Loci/CHEM, with a high-order accurate discontinuous Galerkin (DG) solver, Loci/THRUST, developed in the same computational framework. Loci/THRUST employs a low dissipation, high-order, unstructured DG method to accurately propagate acoustic waves away from the source regions across large distances. The DG solver is currently capable of solving up to 4th order solutions for non-linear, conservative acoustic field propagation. Higher order boundary conditions are implemented to accurately model the reflection and refraction of acoustic waves on launch pad components. The DG solver accepts generalized unstructured meshes, enabling efficient application of common mesh generation tools for CHEM and THRUST simulations. The DG solution is coupled with the CFD solution at interface boundaries placed near the CFD acoustic source regions. Both simulations are executed simultaneously with coordinated boundary condition data exchange.

Liever, Peter↗

NASA NTRSDate not supplied

Large Vehicle Lunar Landing Surface Interaction and In-Situ Resource Based Risk Mitigation: Landing & Launch Pads

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/GCH4rocket engine fired on a 1m2area 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.

Lunar↗

NASA NTRS1985

Radial rift and block tectonics around the Tharsis bulge: Introductional postulation

Although the Vallis Marineris canyon and radial fossae grabens have been extensively studied, their origin and formation mechanism is still the subject of numerous questions. Possible rift formation is only one point of view and does not explain the rifting mechanism and the radial pattern of these structures around the Tharsis bulge. Both active and passive rifting must be taken into account. According to the active mechanism the building of a colcanic complex like the Tharsis bulge is caused by the rising of a huge mantle plume and adjoining extrusions which tap the generated magma. The surrounding fossae and valley structures are caused by more passive crustal rifting due to tensional failure of the surface layers. The main rising mantle plume activated and regenerated failure patterns radial to the centre of activity. These radial zones of weakness are then most easily utilized by the rising mantle plume. Deep zones of weakness regulate the penetration and distribution of hot mantle rock into upper levels while, contrarily, the effective impingement of the hot mantle plume into the lithosphere opens up new weakness zones.

Raitala, J.↗

NASA NTRS1993

Contamination control and plume assessment of low-energy thrusters

A method for evaluating the effects of gases originating from plumes or other direct sources impinging on sensitive surfaces located within their field of view is presented. The method is applied to an Xe ion plasma generator with a thrust of a few mN located near a cryogenic surface. The method is based on experimentally derived plume descriptions of emitted low-pressure vented gases and test results of plume distributions from a similar ion engine. It is concluded that the plume of the Xe neutrals represented by a cosine squared function of the angle is a possible contamination source for certain surfaces of the spacecraft.

Scialdone, John J.↗

NASA NTRS2000

Applying Contamination Modelling to Spacecraft Propulsion Systems Designs and Operations

Molecular and particulate contaminants generated from the operations of a propulsion system may impinge on spacecraft critical surfaces. Plume depositions or clouds may hinder the spacecraft and instruments from performing normal operations. Firing thrusters will generate both molecular and particulate contaminants. How to minimize the contamination impact from the plume becomes very critical for a successful mission. The resulting effect from either molecular or particulate contamination of the thruster firing is very distinct. This paper will discuss the interconnection between the functions of spacecraft contamination modeling and propulsion system implementation. The paper will address an innovative contamination engineering approach implemented from the spacecraft concept design, manufacturing, integration and test (I&T), launch, to on- orbit operations. This paper will also summarize the implementation on several successful missions. Despite other contamination sources, only molecular contamination will be considered here.

Chen, Philip T.↗

NASA NTRS2001

Unic Unstructured CFD Methodology Development

Base heating characteristics is crucial to the success and the overall performance of the X-33 engine. Base heating is important throughout the entire flight trajectory due to the aerospike engine design of X-33. The base region is surrounded by the hot-gas plume, which expands, circulates and impinges on the base. An advanced computation fluid dynamics method is employed in an effort to develop a robust, accurate and efficient tool for the X-33 base heating performance predictions. This computational tool is developed based on a Navier-Stokes flow solver, which is suitable for general complex geometry and includes turbulence, finite-rate chemistry, and radiation models. To fulfill the fast turnaround requirement as a design analysis tool, adaptive mesh refinement method and parallel-computing algorithm are also incorporated in the present model. Case study for the X-33 base-region fluid dynamics and heat transfer characteristics are presented.

Chen, Y. S.↗

NASA NTRS2014

Constraining the Composition of the Subcontinental Lithospheric Mantle Beneath the East African Rift: FTIR Analysis of Water in Spinel Peridotite Mantle Xenoliths

The East African Rift System was initiated by the impingement of the Afar mantle plume on the base of the non-cratonic continental lithosphere (assembled during the Pan-African Orogeny), producing over 300,000 kmof continental flood basalts approx.30 Ma ago. The contribution of the subcontinental lithospheric mantle (SCLM) to this voluminous period of volcanism is implied based on basaltic geochemical and isotopic data. However, the role of percolating melts on the SCLM composition is less clear. Metasomatism is capable of hybridizing or overprinting the geochemical signature of the SCLM. In addition, models suggest that adding fluids to lithospheric mantle affects its stability. We investigated the nature of the SCLM using Fourier transform infrared spectrometry (FTIR) to measure water content in mantle xenoliths entrained in young (1 Ma) basaltic lavas from the Ethiopian volcanic province. The mantle xenoliths consist dominantly of spinel lherzolites and are composed of nominally anhydrous minerals, which can contain trace water as H in mineral defects. Eleven mantle xenoliths come from the Injibara-Gojam region and two from the Mega-Sidamo region. Water abundances of olivines in six samples are 1-5ppm H2O while the rest are below the limit of detection (<0.5 ppm H2O); orthopyroxene and clinopyroxene contain 80-238 and 111-340 ppm wt H2O, respectively. Two xenoliths have higher water contents - a websterite (470 ppm) and dunite (229 ppm), consistent with involvement of ascending melts. The low water content of the upper SCLM beneath Ethiopia is as dry as the oceanic mantle except for small domains represented by percolating melts. Consequently, rifting of the East African lithosphere may not have been facilitated by a hydrated upper mantle.

Erickson, Stephanie Gwen↗

NASA NTRS2026

Moon Base Transportation - Deliveries to the Lunar Surface

Development of the Moon Base will enable a home away from home for astronauts who will live and work at humanity’s first lunar outpost. In this effort, NASA’s Moon Transportation Office is responsible for enabling the transformational missions required to deliver habitats, supplies, science payloads, and all other elements needed to cultivate a permanent presence on the Lunar Surface. The Mission Concept (MC) is characterized through evaluation of an end-to-end architecture that can successfully deliver a generalized heavy large-volume payload, in excess of 4000 kg, to a precision landing and touchdown on the lunar surface. The mission architecture utilizes a single launch configuration of a Lunar Lander (LL) with a unique propellant system. The LL has an integral orbital transfer capability and features jettisonable elements. The design circumvents the need for prop transfer on orbit and multiple launch configurations. The launch vehicle (LV) for this work will assume the capability to deliver a payload in excess of 40,000 kg to orbit, affording multiple LV solutions. Considerations for the LL and payload deployment from the fairing are assumed to be handled through compliance with a launch providers’ Interface Requirements Document (IRD). The MC will span from launch at Kennedy Space Center (KSC) to terminal descent and touchdown on the lunar surface, requiring a total ΔV on the order of 6 km/s beyond what is required to get the vehicle stack to a 200 km circular Low Earth Orbit (LEO). Major mission phases include: launch and launch vehicle separation, transfer operations, pre-landing navigation, and lunar descent and touchdown. A Concept of Operations (ConOps) is used as the primary design driver for defining architecture of the vehicles necessary to achieve final payload delivery. Numerous ground rules and assumptions will be provided for each phase of the mission. Concept designs for the LL is presented. An emphasis of the design maximizes a feasible path for maturation, manufacturing, and operation. A self-imposed practical consideration for this effort is the incorporation of legacy designed hardware to minimize expensive, time-intensive, and high-risk hardware development cycles. The propulsion system of the LL adopts a conventional storable bipropellant configuration of monomethyl hydrazine (MMH) and mixed oxides of nitrogen (MON3). This effort will showcase a unique propellant delivery system to minimize the reliance on propellant management devices (PMDs) during descent. Numerous key constraints have been considered, across the multiple segments of the mission. These include the unique aspects of center of gravity (CG) management, thruster plume effects including self-impingement, propulsion system hardware limitations, navigation during multiple mission phases, and landing gear geometry for uneven terrain. These constraints shape the trades necessary for precision landing of heavy cargo and ensure compatibility with broader Moon Base Transportation concepts. The resulting insights inform future transportation strategies for the Moon and beyond; directly contributing to the development of cargo‑delivery standards that will support the long‑term buildup of a sustainable, continuously inhabited Moon Base.

Lunar Habitat↗

NASA NTRS1991

Hot-spot evolution and the global tectonics of Venus

The global tectonics of Venus may be dominated by plumes rising from the mantle and impinging on the lithosphere, giving rise to hot spots. Global sea-floor spreading does not take place, but direct convective coupling of mantle flow fields to the lithosphere leads to regional-scale deformation and may allow lithospheric transport on a limited scale. A hot-spot evolutionary sequence comprises (1) a broad domal uplift resulting from a rising mantle plume, (2) massive partial melting in the plume head and generation of a thickened crust or crustal plateau, (3) collapse of dynamic topography, and (4) creep spreading of the crustal plateau. Crust on Venus is produced by gradual vertical differentiation with little recycling rather than by the rapid horizontal creation and consumption characteristic of terrestrial sea-floor spreading.

Phillips, Roger J.↗

DOE OSTI2020

Gas density structure of supersonic flows impinged on by thin blades for laser-plasma accelerator targets

Density transition injection is an effective technique for controllably loading electrons into a trapped phase for laser plasma accelerators. One common technique to achieve the required fluid structure is to impinge a thin blade on the plume of a supersonic nozzle. Density transitions induced in this way are often assumed to be bow shocks and therefore sharp, but simulations and fluorescence measurements presented in this work show that in many cases of interest, the density transition accessible to a laser propagating transverse to the shock is an intercepting shock, and therefore, shock thickness and density vary with pressure, laser height, and blade position. The fluid dynamics of a supersonic nozzle impinged on by a thin, flat object are explored through simulations and relevant features are verified via planar laser-induced fluorescence measurements. Overall, the implications of the results for tuning electron beam injectors in laser plasma accelerators are discussed.

42 ENGINEERING↗