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At least 307 records · Page 17

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NASA NTRS2004

Future of Fuel Savings

Using automation to free up controllers for more strategic management of air traffic is one approach being studied by NASA as it seeks to boost airspace system capacity and efficiency, thereby saving fuel. Heinz Erzberger, a NASA Ames Research Center senior scientist, says the Advanced Airspace Concept (AAC) has been studied for several years. It could increase efficiency 15% by providing optimal routes that cut airlines direct operating costs. A 25% increase in landings on existing runways could follow an important benefit. AAC is one of the efforts to be reviewed by the Joint Planning and Development Organization, an FAA-led initiative by six federal agencies to redesign the U.S. air transportation system by 2025. The main goal is to triple air traffic capacity within 20 years to avert the sort of gridlock that would make fuel consumption only one of many travel nightmares. The automated system approach would allow aircraft to fly optimal trajectories. A trajectory would be defined in the standard three dimensions and eventually include the fourth, time. The management of air traffic by the data-linked exchange of trajectories would start at high altitude and eventually move down to lower altitudes. The automated concept is an outgrowth of the type of tools developed by NASA for use by FAA controllers in managing traffic flows over the years, including ones that optimize routings for the best fuel burn. But AAC would push automation further to reduce workload so controllers can focus on "solving strategic control problems, managing traffic flow during changing weather and ... other unusal events." One key component, the automated trajectory server (ATS), is a ground systems that would rely on software to manage flight path requests from aircrews and controllers. But, Erzberger acknowledges, "The FAA's current plan for upgrades to air traffic services does not include [allowing] the future ground system to issue separation-critical clearances of trajectory changes autonomously to aircraft via data link without explicit approval of a controller," as the AAC proposes. The AAC enables pilots or controllers to data link requests for a trajectory change to the ATS for approval after they are deconflicted with the paths of other aircraft. To divert around storms, for example, pilots could data link their trajectory preference to the ATS. Since several aircraft might request similar routes, the computer would then have to suggest alternatives. This could be accomplished without pilot-controller radio calls, a big bottleneck now. The ATS would have a built-in conflict monitor to call for a resolution (turn, climb or descend), when loss of separation is likely in 1-20 min. The AAC system would reduce controller errors by 90%, according to NASA Ames estimates. The AAC would have a back-up program to assure separation-Tactical Separation Assurance (TSAFE). It s designed to detect short-term traffic conflicts within 3-4 min. of loss of separation. The last line of defense would still be provided by traffic alert & collision avoidance systems (TCAS).

Hughes, David↗

NASA NTRS1985

Post-flight BET products for the 2nd discovery entry, STS-19 (51-A)

The post-flight products for the second Discovery flight, STS-19 (51-A), are summarized. The inertial best estimate trajectory (BET), BT19D19/UN=169750N, was developed using spacecraft dynamic measurements from Inertial Measurement Unit 2 (IMU2) in conjunction with the best tracking coverage available for any of the earlier Shuttle entries. As a consequence of the latter, an anchor epoch was selected which conforms to an initial altitude of greater than a million feet. The Extended BET, ST19BET/UN=274885C, incorporated the previously mentioned inertial reconstructed state information and the Langley Atmospheric Information Retrieval System (LAIRS) atmosphere, ST19MET/UN=712662N, with some minor exceptions. Primary and back-up AEROBET reels are NK0165 and NK0201, respectively. This product was only developed over the lowermost 360 kft altitude range due to atmosphere problems but this relates to altitudes well above meaningful signal in the IMUs. Summary results generated from the AEROBET for this flight are presented with meaningful configuration and statistical comparisons from the previous thirteen flights. Modified maximum likelihood estimation (MMLE) files were generated based on IMU2 and the Rate Gyro Assembly/Accelerometer Assembly (RGA/AA), respectively. Appendices attached define spacecraft and physical constants utilized, show plots of the final tracking data residuals from the post-flight fit, list relevant parameters from the BET at a two second spacing, and retain for archival purpose all relevant input and output tapes and files generated.

Kelly, G. M.↗

NASA NTRS1983

Ions upstream of the earth's bow shock - A theoretical comparison of alternative source populations

The trajectories of ions reflected or leaked upstream from the earth's bow shock and subject solely to the Lorentz force in a steady interplanetary magnetic field B and the V x B electric field are studied theoretically. Expressions are obtained for the guiding center motion and gyromotion in a frame (the Hoffman-Teller frame) moving parallel to the shock surface with sufficient speed to transform the incident solar wind velocity into motion entirely along the interplanetary magnetic field. Equations are derived which transform these motions back to the observer's frame. The predicted upstream motions for four different source models for upstream ions are compared using these expressions: magnetic moment-conserving reflection of solar wind ions, specular reflection of solar wind ions, magnetic moment-conserving leakage of magnetosheath ions, and leakage of magnetosheath ions parallel to the shock normal.

Schwartz, S. J.↗

NASA NTRS2015

An Alternative Humans to Mars Approach: Reducing Mission Mass with Multiple Mars Flyby Trajectories and Minimal Capability Investments

Mars flyby trajectories and Earth return trajectories have the potential to enable lower- cost and sustainable human exploration of Mars. Flyby and return trajectories are true minimum energy paths with low to zero post-Earth departure maneuvers. By emplacing the large crew vehicles required for human transit on these paths, the total fuel cost can be reduced. The traditional full-up repeating Earth-Mars-Earth cycler concept requires significant infrastructure, but a Mars only flyby approach minimizes mission mass and maximizes opportunities to build-up missions in a stepwise manner. In this paper multiple strategies for sending a crew of 4 to Mars orbit and back are examined. With pre-emplaced assets in Mars orbit, a transit habitat and a minimally functional Mars taxi, a complete Mars mission can be accomplished in 3 SLS launches and 2 Mars Flyby's, including Orion. While some years are better than others, ample opportunities exist within a given 15-year Earth-Mars alignment cycle. Building up a mission cadence over time, this approach can translate to Mars surface access. Risk reduction, which is always a concern for human missions, is mitigated by the use of flybys with Earth return (some of which are true free returns) capability.

Whitley, Ryan J.↗

NASA NTRS2005

Development and Test Plans for the MSR EEV

The goal of the proposed Mars Sample Return mission is to bring samples from the surface of Mars back to Earth for thorough examination and analysis. The Earth Entry Vehicle is the passive entry body designed to protect the sample container from entry heating and deceleration loads during descent through the Earth s atmosphere to a recoverable location on the surface. This paper summarizes the entry vehicle design and outlines the subsystem development and testing currently planned in preparation for an entry vehicle flight test in 2010 and mission launch in 2013. Planned efforts are discussed for the areas of the thermal protection system, vehicle trajectory, aerodynamics and aerothermodynamics, impact energy absorption, structure and mechanisms, and the entry vehicle flight test.

Dillman, Robert↗

DOE OSTI2025

Single cell RNA sequencing reveals shifts in cell maturity and function of endogenous and infiltrating cell types in response to acute intervertebral disc injury

Intervertebral disc (IVD) degeneration contributes to disabling back pain. Degeneration can be initiated by injury and progressively leads to an irreversible loss of cells and function. IVD function restoration through cell replacement therapies have had limited success due to knowledge gaps in the critical cell populations important for repair. Here, in this study, we used single cell RNA sequencing to identify the transcriptional changes of IVD resident and infiltrating cell populations from Control and Injured coccygeal IVDs extracted from 12-week-old female C57BL/6J mice 7 days post injury. Clustering, gene ontology, and pseudotime trajectory analyses determined transcriptomic divergences with injury, flow cytometry identified they types of infiltrating immune cells, and immunofluorescence was utilized to define mesenchymal stem cell (MSC) localization. We identified 11 distinct clusters that included IVD, immune, vascular cells, and MSCs. Differential gene expression analysis determined that Outer Annulus Fibrosus, Neutrophils, Saa2-High MSCs, Macrophages, and Krt18 + Nucleus Pulposus (NP) cells were the major drivers of transcriptomic differences between Control and Injured cells. Gene ontology revealed that the most upregulated biological pathways were angiogenesis and T cell-related while wound healing and ECM regulation were downregulated. Pseudotime trajectory analyses revealed that IVD injury directed cells towards increased differentiation in all clusters, except for Krt18 + NP cells which remained in a less mature cell state. Saa2-High and Grem1-High MSCs populations shifted towards more differentiated IVD cells profiles with injury and localized distinctly within the IVD. This study revealed novel MSC populations with the potential to be leveraged for future IVD repair studies.

Cartilage↗

NASA NTRS1992

Design of a wing shape for study of hypersonic crossflow transition in flight

Computational fluid dynamics methods were used in the design of a wing shape for study of hypersonic crossflow transition in flight. The flight experiment is to be performed on the delta wing of the first stage of a Pegasus launch vehicle as a piggy-back experiment to support boundary-layer stability code development and validation. The design goal is to obtain crossflow-induced transition at 20-40 percent of the chord for a flight Mach number of approximately six. The present paper describes the design and analysis process utilized to obtain desired glove shape. A variety of schemes were used in the design, ranging from simple empirical crossflow correlations to three-dimensional Navier-Stokes codes in conjunction with linear stability/N-factor computations. The sensitivity to various parameters, such as trajectory variations, allowable wing thickness, leading-edge radius and surface temperature, is also discussed.

Godil, A.↗

NASA NTRS1995

Estimates of Comet Fragment Masses from Impact Crater Chains on Callisto and Ganymede

Chains of impact craters, or catenae, have been identified in Voyager images of Callisto and Ganymede. Although these resemble in some respects secondary crater chains, the source craters and basins for the catenae cannot be identified. The best explanation is a phenomenon similar to that displayed by former comet Shoemaker-Levy 9; tidal (or other) breakup close to Jupiter followed by gradual orbital separation of the fragments and collision with a Galilean satellite on the outbound leg of the trajectory. Because the trajectories must pass close to Jupiter, this constrains the impact geometry (velocity and impact angle) of the individual fragments. For the dominant classes of impactors, short period Jupiter-family comets and asteroids, velocities at Callisto and Ganymede are dominated by Jovian gravity and a satellite's orbital motion, and are insensitive to the pre-fragmentation heliocentric velocity; velocities are insensitive to satellite gravity for all impactor classes. Complex crater shapes on Callisto and Ganymede are determined from Voyager images and Schmidt-Holsapple scaling is used to back out individual fragment masses. We find that comet fragment radii are generally less than about 500 m (for ice densities) but can be larger. These estimates can be compared with those for the Shoemaker-Levy 9 impactors.

McKinnon, William B.↗

NASA NTRS2015

Orion Exploration Flight Test-1 (EFT-1) Absolute Navigation Performance

The Orion vehicle, being design to take men back to the Moon and beyond, successfully completed its first flight test, EFT-1 (Exploration Flight Test-1), on December 5th, 2014. The main objective of the test was to demonstrate the capability of re-enter into the Earth's atmosphere and safely splash-down into the pacific ocean. This un-crewed mission completes two orbits around Earth, the second of which is highly elliptical with an apogee of approximately 5908 km, higher than any vehicle designed for humans has been since the Apollo program. The trajectory was designed in order to test a high-energy re-entry similar to those crews will undergo during lunar missions. The mission overview is shown in Figure 1. The objective of this paper is to document the performance of the absolute navigation system during EFT-1 and to present its design.

Zanetti, Renato↗

NASA NTRS2010

Z-Pinch Pulsed Plasma Propulsion Technology Development

Fusion-based propulsion can enable fast interplanetary transportation. Magneto-inertial fusion (MIF) is an approach which has been shown to potentially lead to a low cost, small reactor for fusion break even. The Z-Pinch/dense plasma focus method is an MIF concept in which a column of gas is compressed to thermonuclear conditions by an axial current (I approximates 100 MA). Recent advancements in experiments and the theoretical understanding of this concept suggest favorable scaling of fusion power output yield as I(sup 4). This document presents a conceptual design of a Z-Pinch fusion propulsion system and a vehicle for human exploration. The purpose of this study is to apply Z-Pinch fusion principles to the design of a propulsion system for an interplanetary spacecraft. This study took four steps in service of that objective; these steps are identified below. 1. Z-Pinch Modeling and Analysis: There is a wealth of literature characterizing Z-Pinch physics and existing Z-Pinch physics models. In order to be useful in engineering analysis, simplified Z-Pinch fusion thermodynamic models are required to give propulsion engineers the quantity of plasma, plasma temperature, rate of expansion, etc. The study team developed these models in this study. 2. Propulsion Modeling and Analysis: While the Z-Pinch models characterize the fusion process itself, propulsion models calculate the parameters that characterize the propulsion system (thrust, specific impulse, etc.) The study team developed a Z-Pinch propulsion model and used it to determine the best values for pulse rate, amount of propellant per pulse, and mixture ratio of the D-T and liner materials as well as the resulting thrust and specific impulse of the system. 3. Mission Analysis: Several potential missions were studied. Trajectory analysis using data from the propulsion model was used to determine the duration of the propulsion burns, the amount of propellant expended to complete each mission considered. 4. Vehicle Design: To understand the applicability of Z-Pinch propulsion to interplanetary travel, it is necessary to design a concept vehicle that uses it -- the propulsion system significantly impacts the design of the electrical, thermal control, avionics and structural subsystems of a vehicle. The study team developed a conceptual design of an interplanetary vehicle that transports crew and cargo to Mars and back and can be reused for other missions. Several aspects of this vehicle are based on a previous crewed fusion vehicle study -- the Human Outer Planet Exploration (HOPE) Magnetized Target Fusion (MTF) vehicle. Portions of the vehicle design were used outright and others were modified from the MTF design in order to maintain comparability.

Polsgrove, Tara↗

NASA NTRS2010

Orion Optical Navigation for Loss of Communication Lunar Return Contingencies

The Orion Crew Exploration Vehicle (CEV) will replace the Space Shuttle and serve as the next-generation spaceship to carry humans back to the Moon for the first time since the Apollo program. For nominal lunar mission operations, the Mission Control Navigation team will utilize radiometric measurements to determine the position and velocity of Orion and uplink state information to support Lunar return. However, in the loss of communications contingency return scenario, Orion must safely return the crew to the Earth's surface. The navigation design solution for this loss of communications scenario is optical navigation consisting of lunar landmark tracking in low lunar orbit and star- horizon angular measurements coupled with apparent planetary diameter for Earth return trajectories. This paper describes the optical measurement errors and the navigation filter that will process those measurements to support navigation for safe crew return.

Getchius, Joel↗

NASA NTRS2011

First Stage Solid Propellant Multi Debris Thermal Analysis

The crew launch vehicle considered for the Constellation (Cx) Program utilizes a first stage solid rocket motor. If an abort is initiated in first stage flight the Crew Module (CM) will separate and be pulled away from the launch vehicle via a Launch Abort System (LAS) in order to safely and quickly carry the crew away from the malfunction launch vehicle. Having aborted the mission, the launch vehicle will likely be destroyed via a Flight Termination System (FTS) in order to prevent it from errantly traversing back over land and posing a risk to the public. The resulting launch vehicle debris field, composed primarily of first stage solid propellant, poses a threat to the CM. The harsh radiative thermal environment induced by surrounding burning propellant debris may lead to CM parachute failure. A methodology, detailed herein, has been developed to address this concern and quantify the risk of first stage propellant debris leading to radiative thermal demise of the CM parachutes. Utilizing basic thermal radiation principles, a software program was developed to calculate parachute temperature as a function of time for a given abort trajectory and debris piece trajectory set. Two test cases, considered worst-case aborts with regard to launch vehicle debris environments, were analyzed using the simulation: an abort declared at Mach 1 and an abort declared at maximum dynamic pressure (Max Q). For both cases, the resulting temperature profiles indicated that thermal limits for the parachutes were not exceeded. However, short duration close encounters by single debris pieces did have a significant effect on parachute temperature, with magnitudes on the order of 10 s of degrees Fahrenheit. Therefore while these two test cases did not indicate exceedance of thermal limits, in order to quantify the risk of parachute failure due to radiative effects from the abort environment, a more thorough probability-based analysis using the methodology demonstrated herein must be performed.

Toleman, Benjamin M.↗

NASA NTRS2011

First Stage Solid Propellant Multiply Debris Thermal Analysis

Destruction of a solid rocket stage of a launch vehicle can create a thermal radiation hazard for an aborting crew module. This hazard was assessed for the Constellation Program (Cx) crew and launch vehicle concept. For this concept, if an abort was initiated in first stage flight, the Crew Module (CM) will separate and be pulled away from the malfunctioning launch vehicle via a Launch Abort System (LAS). Having aborted the mission, the launch vehicle will likely be destroyed via a Flight Termination System (FTS) in order to prevent it from errantly traversing back over land and posing a risk to the public. The resulting launch vehicle debris field, composed primarily of first stage solid propellant, poses a threat to the CM. The harsh radiative thermal environment, caused by surrounding burning propellant debris, may lead to CM parachute failure. A methodology, detailed herein, has been developed to address this concern and to quantify the risk of first stage propellant debris leading to the thermal demise of the CM parachutes. Utilizing basic thermal radiation principles, a software program was developed to calculate parachute temperature as a function of time for a given abort trajectory and debris piece trajectory set. Two test cases, considered worst case aborts with regard to launch vehicle debris environments, were analyzed using the simulation: an abort declared at Mach 1 and an abort declared at maximum dynamic pressure (Max Q). For both cases, the resulting temperature profiles indicated that thermal limits for the parachutes were not exceeded. However, short duration close encounters by single debris pieces did have a significant effect on parachute temperature. Therefore while these two test cases did not indicate exceedance of thermal limits, in order to quantify the risk of parachute failure due to radiative effects from the abort environment, a more thorough probability-based analysis using the methodology demonstrated herein must be performed.

Toleman, Benjamin M.↗

NASA NTRS2004

Faster Processing for Inverting GPS Occultation Data

A document outlines a computational method that can be incorporated into two prior methods used to invert Global Positioning System (GPS) occultation data [signal data acquired by a low-Earth-orbiting satellite as either this or the GPS satellite rises above or falls below the horizon] to obtain information on altitude-dependent properties of the atmosphere. The two prior inversion methods, known as back propagation and canonical transform, are computationally expensive because for each occultation, they involve numerical evaluation of a large number of diffraction-like spatial integrals. The present method involves an angular-spectrum-based phase-extrapolation approximation in which each data point is associated with a plane-wave component that propagates in a unique direction from the orbit of the receiving satellite to intersect a straight line tangent to the orbit at a nearby point. This approximation enables the use of fast Fourier transforms (FFTs), which apply only to data collected along a straight-line trajectory. The computation of the diffraction-like integrals in the angular-spectrum domain by use of FFTs takes only seconds, whereas previously, it took minutes.

Ao, Chi↗

NASA NTRS2014

Human Outer Solar System Exploration via Q-Thruster Technology

Propulsion technology development efforts at the NASA Johnson Space Center continue to advance the understanding of the quantum vacuum plasma thruster (QThruster), a form of electric propulsion. Through the use of electric and magnetic fields, a Q-thruster pushes quantum particles (electrons/positrons) in one direction, while the Qthruster recoils to conserve momentum. This principle is similar to how a submarine uses its propeller to push water in one direction, while the submarine recoils to conserve momentum. Based on laboratory results, it appears that continuous specific thrust levels of 0.4 - 4.0 N/kWe are achievable with essentially no onboard propellant consumption. To evaluate the potential of this technology, a mission analysis tool was developed utilizing the Generalized Reduced Gradient non-linear parameter optimization engine contained in the Microsoft Excel® platform. This tool allowed very rapid assessments of "Q-Ship" minimum time transfers from earth to the outer planets and back utilizing parametric variations in thrust acceleration while enforcing constraints on planetary phase angles and minimum heliocentric distances. A conservative Q-Thruster specific thrust assumption (0.4 N/kWe) combined with "moderate" levels of space nuclear power (1 - 2 MWe) and vehicle specific mass (45 - 55 kg/kWe) results in continuous milli-g thrust acceleration, opening up realms of human spaceflight performance completely unattainable by any current systems or near-term proposed technologies. Minimum flight times to Mars are predicted to be as low as 75 days, but perhaps more importantly new "retro-phase" and "gravity-augmented" trajectory shaping techniques were revealed which overcome adverse planetary phasing and allow virtually unrestricted departure and return opportunities. Even more impressively, the Jovian and Saturnian systems would be opened up to human exploration with round-trip times of 21 and 32 months respectively including 6 to 12 months of exploration at the destinations. Finally, interstellar trip times are assessed at milli-g acceleration levels.

Joosten, B. Kent↗

DOE OSTI2026

Strong-Drive Limits in Josephson Circuits: From Chaos to an Unbound-Resonance Threshold

Strong microwave drives enable fast measurement and parametric control in superconducting circuits but can induce transitions out of the intended low-energy manifold. We develop a unified description of strong-drive limits in flux- and charge-driven Josephson circuits across drive frequency and dc flux bias. Using classical phase-space analysis and Floquet--Markov simulations, we identify distinct low- and high-frequency mechanisms. At low frequency, we characterize bound-state resonances and separatrix chaos and find that the flux-drive chaos threshold depends strongly on dc flux bias. At high frequency, these mechanisms are suppressed, and the dissipative steady state transfers from the central bound-state sector to outer resonances formed by above-barrier running trajectories. The resulting unbound-resonance threshold is nearly independent of drive frequency and circuit parameters over the regime studied and is controlled primarily by dc flux bias. Coherent simulations show that parametric operation persists beyond this threshold, but at a reduced rate, setting an effective upper bound on the achievable operation speed. We derive analytical criteria for both thresholds, validate them numerically, and experimentally confirm the predicted dc-bias dependence of the low-frequency threshold in a flux-driven SQUID. We also determine the timescales for transfer into the unbound-resonance regime and relaxation back to the bound-state manifold after the drive is removed. Finally, we relate the stability limits to a complementary picture based on the junction critical current and extend the framework to multitone drives and inductively shunted circuits. Together, these results identify the mechanisms limiting strong driving and suggest routes to extend the stable operating range of Josephson circuits.

You, Xinyuan [Fermilab] (ORCID:0000000291789419)↗

NASA NTRS1991

Trajectory control sensor engineering model detailed test objective

The concept employed in an existing Trajectory Control Sensor (TCS) breadboard is being developed into an engineering model to be considered for flight on the Shuttle as a Detailed Test Objective (DTO). The sensor design addresses the needs of Shuttle/SSF docking/berthing by providing relative range and range rate to 1500 meters as well as the perceived needs of AR&C by relative attitude measurement over the last 100 meters. Range measurement is determined using a four-tone ranging technique. The Doppler shift on the highest frequency tone will be used to provide direct measurement of range rate. Bearing rate and attitude rates will be determined through back differencing of bearing and attitude, respectively. The target consists of an isosceles triangle configuration of three optical retroreflectors, roughly one meter and one-half meter in size. After target acquisition, the sensor continually updates the positions of the three retros at a rate of about one hertz. The engineering model is expected to weigh about 25 pounds, consume 25-30 watts, and have an envelope of about 1.25 cubic feet. The following concerns were addressed during the presentation: are there any concerns with differentiating attitude and bearing to get attitude and bearing rates? Since the docking scenario has low data bandwidth, back differencing is a sufficient approximation of a perfect differentiator for this application. Could range data be obtained if there were no retroreflectors on the target vehicle? Possibly, but only at close range. It would be dependent on target characteristics.

Dekome, Kent↗

NASA NTRS1998

X-33 Integrated Test facility, Extended Range Simulation

In support of the X-33 Single Stage To Orbit program, NASA Dryden Flight Research Center was selected to provide continuous communications coverage of the X-33 vehicle from launch, through landing at Malmstrom Air Force Base, Montana and Michaels Army Air Field, Utah. An extensive real-time range simulation capability is being developed to ensure successful communications with the autonomous X-33 vehicle. This paper will provide an overview of the various levels of simulation being developed to support the X-33 extended range subsystems. These subsystems include the Flight Termination System, L-Band command uplink subsystem and the S-Band telemetry downlink subsystem. In addition, the radar model developed provides continuous azimuth, elevation and range information based on the flight trajectory. The Dynamic Ground Station Analysis model developed by NASA Goddard Space Flight Center, calculate the received signal strength at each ground station. This model takes into consideration Radio Frequency (RF) link parameters such as frequency, antenna gain, space loss, plasma effects and the vehicle's position and attitude at any point in time during the flight path. All three RF links are then attenuated based on this calculated level and the RF signals are sent into telemetry receivers to emulate remote sites, or the power incident on the vehicle from uplinked signals. The best source received telemetry data is then passed back to the Launch and Mission Control Monitoring System (LMCMS) resident in the Operations Control Center. The LMCMS also provides the range simulation system the uplink command combined with differential GPS corrections. Later stages will require the progressive integration of actual range hardware with this simulation effort, leading to communication between telemetry, uplink and FTS antennas at NASA Dryden Flight Research Center, with vehicle antennas mounted on the Walter C. Williams Research Aircraft Integration Facility (RAIF). Decommutated Pulse Code Modulated (PCM) data is displayed on one of the four monitors that comprise the Range Safety Officer's (RSO) station. Also displayed are instantaneous impact prediction models, and Federal Aviation Administration (FAA) data for notification of other traffic in the area. Aside from initiating the flight termination command and validating communication links, the RSO station with the range simulation will be used to provide both range control and range safety officers training. The training is necessary to perform their respective functions with greater levels of confidence prior to first flight.

Sharma, Ashley↗