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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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At least 505 records · Page 28
All sourcesNASA NTRSDOE OSTI
DOE OSTI2024

Modification of Jet Velocities in an Explosively Loaded Copper Target with a Conical Cavity

Here, in this work, the design and execution of an experiment with the goal of demonstrating control over the evolution of a copper jet is described. Simulations show that when using simple multi-material buffers placed between a copper target with a conical cavity and a cylinder of high-explosive, a variety of jetting behaviors occur based on material placement, including both jet velocity augmentation and mitigation. A parameter sweep was performed to determine optimal buffer designs in two configurations. Experiments using the optimal buffer designs verified the effectiveness of the buffers at altering jet velocities. Similar trends were shown between the experimental results and the modeling.

75 CONDENSED MATTER PHYSICS, SUPERCONDUCTIVITY AND↗
NASA NTRS1991

RMS active damping augmentation

The topics are presented in viewgraph form and include: RMS active damping augmentation; potential space station assembly benefits to CSI; LaRC/JSC bridge program; control law design process; draper RMS simulator; MIMO acceleration control laws improve damping; potential load reduction benefit; DRS modified to model distributed accelerations; accelerometer location; Space Shuttle aft cockpit simulator; simulated shuttle video displays; SES test goals and objectives; and SES modifications to support RMS active damping augmentation.

Michael G Gilbert↗
DOE OSTI2024

Challenges and Solutions for Fast Neutron Irradiation of Bulk Material Specimens

Reactor developers continue to recognize opportunities for further enhancing fast spectrum reactor designs with advanced core materials, but all the material test reactors currently available to the United States are thermal spectrum designs. Fortunately, the Advanced Test Reactor and High Flux Isotope Reactor are versatile high flux facilities where spectral modification strategies can be used to reduce undesirable thermal neutron capture transmutation damage and augment fast flux delivered to specimens. New opportunities to leverage high flux regions and specially designed fast flux boosting experiment configurations can be used to achieve meaningful fast fluences on large specimens in ATR. New optimization potentials can be employed to achieve even higher fluences, albeit for smaller specimens, using thermal neutron filters in HFIR test positions. These capabilities, while not true fast reactors, can provide highly relevant environments for researchers needing to study the effects of fast neutron damage in bulk material specimens.

11 NUCLEAR FUEL CYCLE AND FUEL MATERIALS↗
NASA NTRS2018

Additive Manufacturing of Liquid Rocket Engine Combustion Devices: A Summary of Process Developments and Hot-Fire Testing Results

Additive Manufacturing (AM) of metals is a processing technology that has significantly matured over the last decade. For liquid propellant rocket engines, the advantages of AM for replacing conventional manufacturing of complicated and expensive metallic components and assemblies are very attractive. AM can significantly reduce hardware cost, shorten fabrication schedules, increase reliability by reducing the number of joints, and improve hardware performance by allowing fabrication of designs not feasible by conventional means. The NASA Marshall Space Flight Center (MSFC) has been involved with various forms of metallic additive manufacturing for use in liquid rocket engine component design, development, and testing since 2010. The AM technique most often used at the NASA MSFC has been powder-bed fusion or selective laser melting (SLM), although other techniques including laser directed energy deposition (DED), arc-based deposition, and laser-wire cladding techniques have also been used to develop several components. The purpose of this paper is to discuss the various internal programs at the NASA MSFC using AM to develop combustion devices hardware. To date at the NASA MSFC, combustion devices component hardware ranging in size from 100 lbf to 35,000 lbf have been designed and manufactured using SLM and deposition-based AM processes, and many of these pieces have been hot-fire tested. Combustion devices component hardware have included thrust chamber injectors, injector components such as faceplates, regeneratively-cooled combustion chambers, regeneratively-cooled nozzles, gas generator and preburner hardware, and augmented spark igniters. Ongoing and future developments for combustion devices have also included design of components sized for boost-class engines. Several design and hot-fire test iterations have been completed on these subscale and larger scale components, and a summary of these results will be presented as well.

Gradl, Paul↗
NASA NTRS2007

Technical Findings, Lessons Learned, and Recommendations Resulting from the Helios Prototype Vehicle Mishap

The Helios Prototype was originally planned to be two separate vehicles, but because of resource limitations only one vehicle was developed to demonstrate two missions. The vehicle consisted of two configurations, one for each mission. One configuration, designated HP01, was designed to operate at extremely high altitudes using batteries and high-efficiency solar cells spread across the upper surface of its 247-foot wingspan. On August 13, 2001, the HP01 configuration reached an altitude of 96,863 feet, a world record for sustained horizontal flight by a winged aircraft. The other configuration, designated HP03, was designed for long-duration flight. The plan was to use the solar cells to power the vehicle's electric motors and subsystems during the day and to use a modified commercial hydrogen-air fuel cell system for use during the night. The aircraft design used wing dihedral, engine power, elevator control surfaces, and a stability augmentation and control system to provide aerodynamic stability and control. At about 30 minutes into the second flight of HP03, the aircraft encountered a disturbance in the way of turbulence and morphed into an unexpected, persistent, high dihedral configuration. As a result of the persistent high dihedral, the aircraft became unstable in a very divergent pitch mode in which the airspeed excursions from the nominal flight speed about doubled every cycle of the oscillation. The aircraft s design airspeed was subsequently exceeded and the resulting high dynamic pressures caused the wing leading edge secondary structure on the outer wing panels to fail and the solar cells and skin on the upper surface of the wing to rip away. As a result, the vehicle lost its ability to maintain lift, fell into the Pacific Ocean within the confines of the U.S. Navy's Pacific Missile Range Facility, and was destroyed. This paper describes the mishap and its causes, and presents the technical recommendations and lessons learned for improving the design, analysis, and testing methods and techniques required for this class of vehicle.

Noll, Thomas E.↗
NASA NTRSDate not supplied

Space-Based Solar Power: An Enabler for Expanded Lunar Surface Exploration and Mobility Operations

The achievement of industrial and scientific goals set by the National Aeronautics and Space Administration (NASA) Artemis missions within the Moon to Mars (M2M) architecture greatly relies on the breadth and efficacy of lunar surface exploration. The lunar surface environment presents various challenges and complexities to sustainable mission operations and activities. Power generation and distribution capabilities are vital for exploration on the lunar surface, which can enable activities such as subsurface sample collection, permanently shadowed region (PSR) prospecting, and lunar terrain mapping. Innovative power technologies that demonstrate flexibility, reliability, and high capacity are desirable to achieve sustainable operations and infrastructure on the lunar surface. Studies have proposed solar-powered spacecraft in cislunar orbit capable of beaming power to lunar surface assets as a solution to this problem. This study analyzes the benefit that space-based solar power (SBSP) assets could deliver to the M2M architecture by measuring its potential to augment lunar surface exploration opportunity. Qualitative measures of SBSP’s potential impact will include, but not be limited to, power available across surface elements, surface mobility range, and delivered landed mass. These measures are assessed for a notional surface architecture measured against its baseline configuration (no external power augmentation). Additionally, this presentation will provide an overview of the SBSP system design configuration trade space, subsystem parameter trade space, and results from preliminary spacecraft concept of operations (ConOps) and sizing.

Space-based solar power↗
NASA NTRS1986

Passive damping augmentation for flexible structures

The present work concentrates on the application and extension of absorber design and optimization techniques to a multimode, multi-DOF, large space structure, namely the NASA space station. The principal issue addressed is the optimal tuning of several absorbers for the transient response of a multi-DOF system, including the effects of modal coupling, existing structural damping, absorber placement, and adsorber mass. The space station is subject to many transient disturbances such as docking, orbit reboost, crew motion, and payload slewing. A notable steady-state excitation source is the Science Research Centrifuge, which rotates at a frequency in the bandwidth of the primary structural modes. Because of the relatively advanced state of development of steady-state absorber design techniques, only the transient cases are considered in this study.

Sesak, J. R.↗
NASA NTRS1994

Development of CFD model for augmented core tripropellant rocket engine

The Space Shuttle era has made major advances in technology and vehicle design to the point that the concept of a single-stage-to-orbit (SSTO) vehicle appears more feasible. NASA presently is conducting studies into the feasibility of certain advanced concept rocket engines that could be utilized in a SSTO vehicle. One such concept is a tripropellant system which burns kerosene and hydrogen initially and at altitude switches to hydrogen. This system will attain a larger mass fraction because LOX-kerosene engines have a greater average propellant density and greater thrust-to-weight ratio. This report describes the investigation to model the tripropellant augmented core engine. The physical aspects of the engine, the CFD code employed, and results of the numerical model for a single modular thruster are discussed.

Jones, Kenneth M.↗
NASA NTRS1980

Control-system techniques for improved departure/spin resistance for fighter aircraft

This paper summarizes some fundamental information on control-system effects on controllability of highly maneuverable aircraft at high angles of attack and techniques for enhancing fighter aircraft departure/spin resistance using control-system design. The discussion includes (1) a brief review of pertinent high-angle-of-attack phenomena including aerodynamics, inertia coupling, and kinematic coupling, (2) effects of conventional stability augmentation systems at high angle of attack, (3) high-angle-of-attack control-system concepts designed to enhance departure/spin resistance, and (4) the outlook for applications of these concepts to future fighters, particularly those designs which incorporate relaxed static stability.

High angle of attack↗
NASA NTRS2009

Motion Cueing Algorithm Modification for Improved Turbulence Simulation

Atmospheric turbulence cueing produced by flight simulator motion systems has been less than satisfactory because the turbulence profiles have been attenuated by the motion cueing algorithms. Cardullo and Ellor initially addressed this problem by directly porting the turbulence model output to the motion system. Reid and Robinson addressed the problem by employing a parallel aircraft model, which is only stimulated by the turbulence inputs and adding a filter specially designed to pass the higher turbulence frequencies. There have been advances in motion cueing algorithm development at the Man-Machine Systems Laboratory, at SUNY Binghamton. In particular, the system used to generate turbulence cues has been studied. The Reid approach, implemented by Telban and Cardullo, was employed to augment the optimal motion cueing algorithm installed at the NASA LaRC Simulation Laboratory, driving the Visual Motion Simulator. In this implementation, the output of the primary flight channel was added to the output of the turbulence channel and then sent through a non-linear cueing filter. The cueing filter is an adaptive filter; therefore, it is not desirable for the output of the turbulence channel to be augmented by this type of filter. The likelihood of the signal becoming divergent was also an issue in this design. After testing on-site it became apparent that the architecture of the turbulence algorithm was generating unacceptable cues. As mentioned above, this cueing algorithm comprised a filter that was designed to operate at low bandwidth. Therefore, the turbulence was also filtered, augmenting the cues generated by the model. If any filtering is to be done to the turbulence, it will utilize a filter with a much higher bandwidth, above the frequencies produced by the aircraft response to turbulence. The authors have developed an implementation wherein only the signal from the primary flight channel passes through the nonlinear cueing filter. This paper discusses three new algorithms. Testing shows that the new methods provide the pilot with a more realistic sensation of turbulence; the cues are not attenuated by algorithm. Results of offline testing show the credibility of the models. Offline test verification was based primarily on the evaluation of the power spectral density of the outputs and the time response.

Ercole, Anthony V.↗
NASA NTRS2005

The Effect of Pulse Length and Ejector Radius on Unsteady Ejector Performance

The thrust augmentation of a set of ejectors driven by a shrouded Hartmann-Sprenger tube has been measured at four different frequencies. Each frequency corresponded to a different length to diameter ratio of the pulse of air leaving the driver shroud. Two of the frequencies had length to diameter ratios below the formation number, and two above. The formation number is the value of length to diameter ratio below which the pulse converts to a vortex ring only, and above which the pulse becomes a vortex ring plus a trailing jet. A three level, three parameter Box-Behnken statistical design of experiment scheme was performed at each frequency, measuring the thrust augmentation generated by the appropriate ejectors from the set. The three parameters were ejector length, radius, and inlet radius. The results showed that there is an optimum ejector radius and length at each frequency. Using a polynomial fit to the data, the results were interpolated to different ejector radii and pulse length to diameter ratios. This showed that a peak in thrust augmentation occurs when the pulse length to diameter ratio equals the formation number, and that the optimum ejector radius is 0.87 times the sum of the vortex ring radius and the core radius.

Wilson, Jack↗
NASA NTRS2019

Improving Accuracy of Structural Dynamic Modification with Augmented Residual Vectors

It is often important to perform sensitivity analysis to determine how a structural model will be impacted by design changes. Often, the structural analysts will manually make changes to the finite element model (FEM) to determine the effects. But when dealing with a large FEM with millions of degrees of freedom these manual changes can be cumbersome and calculation of the effects can computationally expensive. Therefore, it is desirable to determine the effects of model changes through approximation methods. One common technique is to determine the analytical sensitivity of the FEM model with respect to the given change. These analytical sensitivities are valid when small changes are made to the structural model, but invalid if large changes need to be assessed. Another approach is to use Structural Dynamic Modification (SDM) to create a surrogate model to analyze model changes. SDM is a widely-used sensitivity method and is used in applications of model updating, uncertainty quantification, and model design studies. SMD is valid for moderate (10-20 percent) changes in the structural model, but model approximations are often needed for large parameter changes (greater than 20 percent). Structural Dynamic Modification can be improved by using residual vectors to augment the surrogate model formulation from SDM. Adding the residual modes increases the fidelity of the surrogate model while keeping the computational cost low. This paper discusses the application and limitations of the augmented residual modes method to two structures: the Integrated Spacecraft and Payload Element (ISPE) of the Space Launch System (SLS) and the full SLS as it is configured during its Integrated Modal Test (IMT).

Stewart, Eric C.↗
NASA NTRS1999

Multivariable Techniques for High-Speed Research Flight Control Systems

This report describes the activities and findings conducted under contract with NASA Langley Research Center. Subject matter is the investigation of suitable multivariable flight control design methodologies and solutions for large, flexible high-speed vehicles. Specifically, methodologies are to address the inner control loops used for stabilization and augmentation of a highly coupled airframe system possibly involving rigid-body motion, structural vibrations, unsteady aerodynamics, and actuator dynamics. Design and analysis techniques considered in this body of work are both conventional-based and contemporary-based, and the vehicle of interest is the High-Speed Civil Transport (HSCT). Major findings include: (1) control architectures based on aft tail only are not well suited for highly flexible, high-speed vehicles, (2) theoretical underpinnings of the Wykes structural mode control logic is based on several assumptions concerning vehicle dynamic characteristics, and if not satisfied, the control logic can break down leading to mode destabilization, (3) two-loop control architectures that utilize small forward vanes with the aft tail provide highly attractive and feasible solutions to the longitudinal axis control challenges, and (4) closed-loop simulation sizing analyses indicate the baseline vane model utilized in this report is most likely oversized for normal loading conditions.

Newman, Brett A.↗
NASA NTRSDate not supplied

Development of an Additive Manufactured Cryogenic TVS Augmented Injector

Cryogenic fluid transfer is critical in ground and space systems. Transfer typically requires venting a receiver vessel during the chill and fill process to maintain a pressure favorable to fluid flow. Venting in space is problematic due to low-gravity where fluid position is indeterminate; possibly allowing fluid to flow out the vent. To control receiver tank venting the vented chill/no-vent fill methodology is attractive. An additively manufactured Injector coupled with a Thermodynamic Venting System (TVS) augmentation can enable the vented chill/no-vent fill approach with simplified operations. The TVS augmented injector includes an internal heat exchanger, fluid injector spray head, and external surface condensation heat exchange; combined with integrated flow paths in one part. Operations using the injector allow the tank vent to be closed very early in the transfer and not reopened. The TVS circuit eliminates the potential for stalled fill by maintaining cooling flow on the injector and condenser elements, reducing tank pressure. The injector can also be employed for long-term cryo fluid storage as a means of controlling tank pressure if integrated with a recirculation pump. A cryo-cooler can be used in place of/or augmenting the TVS flow circuit for a zero loss pressure control system. The Injector was developed using the principles of design for additive manufacture and printed as one part, resulting in part count reduction, improved reproducibility, shorter lead times, and reduced cost. The TVS augmented injector is compact and highly adaptable simplifying its integration. A key benefit is that this methodology enables simplified operations in reaching 100% fill conditions.

Omar Mireles↗
NASA NTRS2017

Low and High-Power Inductive Pulsed Plasma Thruster Development Testing at NASA-MSFC

The inductive pulsed plasma thruster (IPPT) is an electromagnetic plasma accelerator that has been identified in NASA roadmaps as an enabling propulsion technology for some niche low-power missions and for high-power in-space propulsion needs. The IPPT is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged producing a high current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. Thrusters of this type possess many demonstrated and potential benefits that make them worthy of continued investigation. The electrodeless nature of these thrusters eliminates the lifetime and contamination issues associated with electrode erosion in conventional electric thrusters. Also, a wider variety of propellants are accessible when compatibility with metallic electrodes in no longer an issue. IPPTs have been successfully operated using propellants like ammonia, hydrazine, and CO2, and there is no fundamental reason why they would not operate on other in situ propellants like H2O. It is well-known that pulsed accelerators can maintain constant specific impulse (I(sub sp)) and thrust efficiency (eta(sub t)) over a wide range of input power levels by adjusting the pulse rate to hold the discharge energy per pulse constant. It has also been demonstrated that an inductive pulsed plasma thruster can operate in a regime where eta(sub t) is relatively constant over a wide range of I(sub sp) values (3000-8000 s). Finally, thrusters in this class have operated in single-pulse mode at high energy per pulse, and by increasing the pulse rate they offer the potential to process very high levels of power using a single thruster. There has been significant previous research on IPPTs designed around a planar-coil (flat-plate) geometry. The most notable of these was the Pulsed Inductive Thruster (PIT), with the PIT MkV presently representing the state-of- the-art in pulsed high-power IPPT technological development. In this paper, we focus on two planar-geometry devices that operate at significantly different power levels. Most work performed at NASA-Marshall Space Flight Center (MSFC) has, to date, focused on lower power thruster operation (approx. = 10s to 100s of J/pulse, up to 2-2.5 kW average power throughput) and previously described. The most recent work aimed to assemble a device that could be tested in cyclic mode on a thrust-stand, and which could augment the existing data set for IPPTs. In addition, the thruster was designed to serve as a test-bed for solid state switching circuitry and pulsed gas valves, with the modular design of the device allowing for variation in or upgrades to test configuration. Recently, MSFC obtained on loan from the Georgia Institute of Technology (Atlanta, GA) the PIT MkVI, successor to the PIT MkV. The MkV and MkVI are similar in design with much of the hardware from the former, specifically the capacitors and spark-gap switches, being reused in the latter. The coil is similar in geometry but has bent copper rods used in the latest iteration in place of the Litz wire windings found in the MkV. The MkVI master switch for the spark gaps is located in the vacuum chamber contained within a sealed, pressurized vessel fastened to the back of the thruster. This is different from the MkV where many capacitor charging lines and spark gap-triggering delay lines ran to the thruster from a master trigger located outside the vacuum chamber. The MkVI was damaged during testing soon after its fabrication was completed. The thruster arrived at MSFC still-damaged and mostly disassembled into many individual pieces. The device has been repaired, with a few additional design changes implemented after discussions with the late Prof. Lovberg regarding the initial testing results and issues encountered. In the present work, we present results from testing of both the small IPPT and the larger MkVI thruster. The smaller device (Fig. 1) is tested on a thrust stand on multiple gases to demonstrate its capability to operate in a repetition-rate mode and serve as a IPPT technology-development testbed. The larger MkVI (Fig. 2) is operated for the first time in its newly reconstituted state, demonstrating full-power pulsed operation and, for the first time, repetition-rate operation of a high-power IPPT. The additional upgrades required for synchronous operation of all the pulsed systems in single-pulse and repetition-rate mode are described in detail.

Polzin, Kurt A.↗
NASA NTRS2018

Inductive Pulsed Plasma Thruster Development Testing at NASA-MSFC

The inductive pulsed plasma thruster (IPPT) is an electromagnetic plasma accelerator that has been identified in NASA roadmaps as an enabling propulsion technology for some niche low-power missions and for high-power in-space propulsion needs. The IPPT is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged producing a high current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. Thrusters of this type possess many demonstrated and potential benefits that make them worthy of continued investigation. The electrodeless nature of these thrusters eliminates the lifetime and contamination issues associated with electrode erosion in conventional electric thrusters. Also, a wider variety of propellants are accessible when compatibility with metallic electrodes in no longer an issue. IPPTs have been successfully operated using propellants like ammonia, hydrazine, and CO2, and there is no fundamental reason why they would not operate on other in situ propellants like H2O. It is well-known that pulsed accelerators can maintain constant specific impulse (I(sub sp)) and thrust efficiency (η(sub t)) over a wide range of input power levels by adjusting the pulse rate to hold the discharge energy per pulse constant. It has also been demonstrated that an inductive pulsed plasma thruster can operate in a regime where η(sub t) is relatively constant over a wide range of I(sub sp) values (3000-8000 s). Finally, thrusters in this class have operated in single-pulse mode at high energy per pulse, and by increasing the pulse rate they offer the potential to process very high levels of power using a single thruster. There has been significant previous research on IPPTs designed around a planar-coil (flat-plate) geometry. The most notable of these was the Pulsed Inductive Thruster (PIT), with the PIT MkV presently representing the state of- the-art in pulsed high-power IPPT technological development. In this paper, we focus on two planar-geometry devices that operate at significantly different power levels. Most work performed at NASA-Marshall Space Flight Center (MSFC) has, to date, focused on lower power thruster operation (≈ 10s to 100s of J/pulse, up to 2-2.5 kW average power throughput) and previously described in Refs. [5,6]. The most recent work aimed to assemble a device that could be tested in cyclic mode on a thrust-stand, and which could augment the existing data set for IPPTs. In addition, the thruster was designed to serve as a test-bed for solid state switching circuitry and pulsed gas valves, with the modular design of the device allowing for variation in or upgrades to test configuration. Recently, MSFC obtained on loan from the Georgia Institute of Technology (Atlanta, GA) the PIT MkVI, successor to the PIT MkV. The MkV and MkVI are similar in design with much of the hardware from the former, specifically the capacitors and spark-gap switches, being reused in the latter. The coil is simliar in geometry but has bent copper rods used in the latest iteration in place of the Litz wire windings found in the MkV. The MkVI master switch for the spark gaps is located in the vacuum chamber contained within a sealed, pressurized vessel fastened to the back of the thruster. This is different from the MkV where many capacitor charging lines and spark gap-triggering delay lines ran to the thruster from a master trigger located outside the vacuum chamber. The MkVI was damaged during testing soon after its fabrication was completed. The thruster arrived at MSFC still-damaged and mostly disassembled into many individual pieces. The device has been repaired, with a few additional design changes implemented after discussions with the late Prof. Lovberg regarding the initial testing results and issues encountered.

Polzin, Kurt A.↗
NASA NTRS1981

Comparison of closed loop model with flight test results

An analytic technique capable of predicting the landing characteristics of proposed aircraft configurations in the early stages of design was developed. In this analysis, a linear pilot-aircraft closed loop model was evaluated using experimental data generated with the NT-33 variable stability in-flight simulator. The pilot dynamics are modeled as inner and outer servo loop closures around aircraft pitch attitude, and altitude rate-of-change respectively. The landing flare maneuver is of particular interest as recent experience with military and other highly augmented vehicles shows this task to be relatively demanding, and potentially a critical design point. A unique feature of the pilot model is the incorporation of an internal model of the pilot's desired flight path for the flare maneuver.

George, F. L.↗
DOE OSTI2020

Virtual and Augmented Realities to Support Development, Training, Deployment, and Operations of Remote Systems - 20516

Throughout the U.S. Department of Energy complex, there is a need to deploy new, unique equipment and systems for cleanup activities, waste processing, infrastructure repair, and decommissioning and demolition in radioactive and hazardous environments. The new equipment often consists of custom systems created by integrating and modifying commercial technologies. In addition, the systems are deployed in unique environments due to limited access, waste properties, and the presence of previously installed equipment or infrastructure that has been left in place. The deployment of custom systems that are 'turnkey' ready for first-time installation and application has its risks. Use of test platforms and mockups for cold (i.e., non-radioactive) testing is one method to mitigate that risk: however, full-scale test mockups for operations such as waste retrieval and tank inspection can be expensive and must be modified to incorporate the uniqueness of each application. An alternative to full-scale mockups is to use three-dimensional (3D) virtual environments and virtual reality (VR) technology to develop interactive virtual test beds. These can be used throughout the down-selection, development, system integration, operational planning, operator training, and mission execution phases of a task. Virtual test beds are readily reconfigured, can be used for multiple tasks, support multiple users in parallel, and can accommodate enhancements and changes not originally planned or conceived in the initial design. Virtual test beds also provide a platform for developing and testing augmented reality (AR) tools for improving operator efficiency. A VR experience has been generated for waste retrieval from a Hanford waste tank to present the potential applications of a virtual test bed. Features of the VR experience are presented, potential applications of VR and AR for remote applications are discussed, and development areas for improving future application are considered. (authors)

12 MANAGEMENT OF RADIOACTIVE AND NON-RADIOACTIVE W↗