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NASA NTRS1993

Capabilities and constraints of NASA's ground-based reduced gravity facilities

The ground-based reduced gravity facilities of NASA have been utilized to support numerous investigations addressing various processes and phenomina in several disciplines for the past 30 years. These facilities, which include drop towers, drop tubes, aircraft, and sounding rockets are able to provide a low gravity environment (gravitational levels that range from 10(exp -2)g to 10(exp -6)g) by creating a free fall or semi-free fall condition where the force of gravity on an experiment is offset by its linear acceleration during the 'fall' (drop or parabola). The low gravity condition obtained on the ground is the same as that of an orbiting spacecraft which is in a state of perpetual free fall. The gravitational levels and associated duration times associated with the full spectrum of reduced gravity facilities including spaced-based facilities are summarized. Even though ground-based facilities offer a relatively short experiment time, this available test time has been found to be sufficient to advance the scientific understanding of many phenomena and to provide meaningful hardware tests during the flight experiment development process. Also, since experiments can be quickly repeated in these facilities, multistep phenomena that have longer characteristic times associated with them can sometimes be examined in a step-by-step process. There is a large body of literature which has reported the study results achieved through using reduced-gravity data obtained from the facilities.

Lekan, Jack↗

NASA NTRS2010

A Study of Fluid Interface Configurations in Exploration Vehicle Propellant Tanks

The equilibrium shape and location of fluid interfaces in spacecraft propellant tanks while in low-gravity is of interest to system designers, but can be challenging to predict. The propellant position can affect many aspects of the spacecraft such as the spacecraft center of mass, response to thruster firing due to sloshing, liquid acquisition, propellant mass gauging, and thermal control systems. We use Surface Evolver, a fluid interface energy minimizing algorithm, to investigate theoretical equilibrium liquid-vapor interfaces for spacecraft propellant tanks similar to those that have been considered for NASA's new class of Exploration vehicles. The choice of tank design parameters we consider are derived from the NASA Exploration Systems Architecture Study report. The local acceleration vector employed in the computations is determined by estimating low-Earth orbit (LEO) atmospheric drag effects and centrifugal forces due to a fixed spacecraft orientation with respect to the Earth or Moon, and rotisserie-type spacecraft rotation. Propellant/vapor interface positions are computed for the Earth Departure Stage and Altair lunar lander descent and ascent stage tanks for propellant loads applicable to LEO and low-lunar orbit. In some of the cases investigated the vapor ullage bubble is located at the drain end of the tank, where propellant management device hardware is often located.

Zimmerli, Gregory A.↗

NASA NTRS2012

Development and Testing of Molecular Adsorber Coatings

The effect of on-orbit molecular contamination has the potential to degrade the performance of spaceflight hardware and diminish the lifetime of the spacecraft. For example, sensitive surfaces, such as optical surfaces, electronics, detectors, and thermal control surfaces, are vulnerable to the damaging effects of contamination from outgassed materials. The current solution to protect these surfaces is through the use of zeolite coated ceramic adsorber pucks. However, these pucks and its additional complex mounting hardware requirements result in several disadvantages, such as size, weight, and cost related concerns, that impact the spacecraft design and the integration and test schedule. As a result, a new innovative molecular adsorber coating was developed as a sprayable alternative to mitigate the risk of on-orbit molecular contamination. In this study, the formulation for molecular adsorber coatings was optimized using various binders, pigment treatment methods, binder to pigment ratios, thicknesses, and spray application techniques. The formulations that passed coating adhesion and vacuum thermal cycling tests were further tested for its adsorptive capacity. Accelerated molecular capacitance tests were performed in an innovatively designed multi-unit system containing idealized contaminant sources. This novel system significantly increased the productivity of the testing phase for the various formulations that were developed. Work performed during the development and testing phases has demonstrated successful application of molecular adsorber coatings onto metallic substrates, as well as, very promising results for the adhesion performance and the molecular capacitance of the coating. Continued testing will assist in the qualification of molecular adsorber coatings for use on future contamination sensitive spaceflight missions.

Abraham, Nithin↗

NASA NTRS1996

Thermal Control Subsystem Design for the Avionics of a Space Station Payload

A case study of the thermal control subsystem development for a space based payload is presented from the concept stage through preliminary design. This payload, the Space Acceleration Measurement System 2 (SAMS-2), will measure the acceleration environment at select locations within the International Space Station. Its thermal control subsystem must maintain component temperatures within an acceptable range over a 10 year life span, while restricting accessible surfaces to touch temperature limits and insuring fail safe conditions in the event of loss of cooling. In addition to these primary design objectives, system level requirements and constraints are imposed on the payload, many of which are driven by multidisciplinary issues. Blending these issues into the overall system design required concurrent design sessions with the project team, iterative conceptual design layouts, thermal analysis and modeling, and hardware testing. Multiple tradeoff studies were also performed to investigate the many options which surfaced during the development cycle.

Moran, Matthew E.↗

NASA NTRSDate not supplied

Towards Autonomous Lunar Resource Excavation via Reinforcement Learning

To continue on a sustainable and flexible path, NASA needs to address the challenge of collecting and moving large amounts of regolith at the destination. NASA’s Regolith Advanced Surface Systems Operations Robot (RASSOR) is principally designed to mine and deliver regolith for In-Situ Resource Utilization (ISRU) processing. RASSOR’s design enables it to efficiently collect and deposit regolith, return collected material for processing, and myriad related ISRU activities. To reliably perform these operations on the lunar surface, RASSOR software and sensory systems need to be robust and maximize the information extracted from a reduced sensor payload. Herein, we present preliminary findings from the Intelligent Capabilities Enhanced RASSOR project. We created reduced-order simulation environments to develop autonomous trenching controllers via reinforcement learning and prototype state estimation architectures. The goal of reinforcement learning is for an agent to learn a policy (task strategy) through interactions with an environment. When the agent performs an action, a change occurs in environment state and a numerical reward is received which informs the agent whether the action performed was good or not. Since reinforcement learning algorithms learn through trial-and-error, a simulation is a desirable first environment for development and learning. We developed two simulations, the first is a 2D excavation simulation developed to facilitate parameter selection, and a 3D simulation developed using a game physics engine, to simulate simplified soil interactions and increase the fidelity of the dynamic models of the robotic agents. The development of this 3D simulation has enabled the training of additional sensing capabilities and research both at the granular mechanics and operations levels. We experimented with various virtual sensor payloads to identify a combination that enabled efficient excavation operation and learning. Our reward function is based on how much material is excavated per step. A penalty is also received for leaving the dig site and to smooth the acceleration of the drum arms. We implemented pseudo time-of-flight sensors to report distance from each drum to ground and the height above ground which was found to be more efficient than existing solutions. Our findings suggest that reinforcement learning for autonomous operations has learned viable trenching strategies within 3000 training episodes in our simplified 2D environment and helped identify desirable sensing capabilities, arrangements, and considerations such as the positioning of time-of-flight sensors. Future work includes expanding our simulation to more complex environments and scenarios, and transfer learning from simulation to RASSOR 2.0 hardware for deployment in the Regolith Test Bin at NASA's Kennedy Space Center.

rassor↗

NASA NTRS2012

Advanced Stirling Convertor Durability Testing: Plans and Interim Results

The U.S. Department of Energy (DOE), Lockheed Martin Corporation (LM), and NASA Glenn Research Center (GRC) have been developing the Advanced Stirling Radioisotope Generator (ASRG) for use as a power system for space science missions. In support of this program, GRC has been involved in testing Stirling convertors, including the Advanced Stirling Convertor (ASC), for use in the ASRG. This testing includes electromagnetic interference/compatibility (EMI/EMC), structural dynamics, advanced materials, organics, and unattended extended operation. The purpose of the durability tests is to experimentally demonstrate the margins in the ASC design. Due to the high value of the hardware, previous ASC tests focused on establishing baseline performance of the convertors within the nominal operating conditions. The durability tests present the first planned extension of the operating conditions into regions beyond those intended to meet the product spec, where the possibility exists of lateral contact, overstroke, or over-temperature events. These tests are not intended to cause damage that would shorten the life of the convertors, so they can transition into extended operation at the conclusion of the tests. This paper describes the four tests included in the durability test sequence: 1) start/stop cycling, 2) exposure to constant acceleration in the lateral and axial directions, 3) random vibration at increased piston amplitude to induce contact events, and 4) overstroke testing to simulate potential failures during processing or during the mission life where contact events could occur. The paper also summarizes the analysis and simulation used to predict the results of each of these tests.

Meer, David W.↗

NASA NTRS2012

Advanced Stirling Convertor Durability Testing: Plans and Interim Results

The U.S. Department of Energy (DOE), Lockheed Martin Corporation (LM), and NASA Glenn Research Center (GRC) have been developing the Advanced Stirling Radioisotope Generator (ASRG) for use as a power system for space science missions. In support of this program, NASA?s Glenn Research Center (GRC) has been involved in testing Stirling convertors, including the Advanced Stirling Convertor (ASC), for use in the ASRG. This testing includes electromagnetic interference/compatibility (EMI/EMC), structural dynamics, advanced materials, organics, and unattended extended operation. The purpose of the durability tests is to experimentally demonstrate the margins in the ASC design. Due to the high value of the hardware, previous ASC tests focused on establishing baseline performance of the convertors within the nominal operating conditions. The durability tests present the first planned extension of the operating conditions into regions beyond those intended to meet the product spec, where the possibility exists of lateral contact, overstroke, or over-temperature events. These tests are not intended to cause damage that would shorten the life of the convertors, so they can transition into extended operation at the conclusion of the tests. This paper describes the four tests included in the durability test sequence: 1) start/stop cycling, 2) exposure to constant acceleration in the lateral and axial directions, 3) random vibration at increased piston amplitude to induce contact events, and 4) overstroke testing to simulate potential failures during processing or during the mission life where contact events could occur. The paper also summarizes the analysis and simulation used to predict the results of each of these tests.

Meer, Dave↗

NASA NTRS2012

High-Performance Modeling and Simulation of Anchoring in Granular Media for NEO Applications

NASA is interested in designing a spacecraft capable of visiting a near-Earth object (NEO), performing experiments, and then returning safely. Certain periods of this mission would require the spacecraft to remain stationary relative to the NEO, in an environment characterized by very low gravity levels; such situations require an anchoring mechanism that is compact, easy to deploy, and upon mission completion, easy to remove. The design philosophy used in this task relies on the simulation capability of a high-performance multibody dynamics physics engine. On Earth, it is difficult to create low-gravity conditions, and testing in low-gravity environments, whether artificial or in space, can be costly and very difficult to achieve. Through simulation, the effect of gravity can be controlled with great accuracy, making it ideally suited to analyze the problem at hand. Using Chrono::Engine, a simulation pack age capable of utilizing massively parallel Graphic Processing Unit (GPU) hardware, several validation experiments were performed. Modeling of the regolith interaction has been carried out, after which the anchor penetration tests were performed and analyzed. The regolith was modeled by a granular medium composed of very large numbers of convex three-dimensional rigid bodies, subject to microgravity levels and interacting with each other with contact, friction, and cohesional forces. The multibody dynamics simulation approach used for simulating anchors penetrating a soil uses a differential variational inequality (DVI) methodology to solve the contact problem posed as a linear complementarity method (LCP). Implemented within a GPU processing environment, collision detection is greatly accelerated compared to traditional CPU (central processing unit)- based collision detection. Hence, systems of millions of particles interacting with complex dynamic systems can be efficiently analyzed, and design recommendations can be made in a much shorter time. The figure shows an example of this capability where the Brazil Nut problem is simulated: as the container full of granular material is vibrated, the large ball slowly moves upwards. This capability was expanded to account for anchors of different shapes and penetration velocities, interacting with granular soils.

Quadrelli, Marco B.↗

NASA NTRS2013

Inductive Pulsed Plasma Thruster Development and Testing at NASA-MSFC

THE inductive pulsed plasma thruster (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. In the present work, we present a summary of the IPPT research and development conducted at NASA's Marshall Space Flight Center (MSFC). As a higher-power, still relatively low readiness level system, there are many issues associated with the eventual deployment and use of the IPPT as a primary propulsion system on spacecraft that remain to be addressed. The present program aimed to fabricate and test hardware to explore how these issues could be addressed. The following specific areas were addressed within the program and will be discussed within this paper. a) Conical theta-pinch IPPT geometry thruster configuration. b) Repetition-rate multi-kW thruster pulsing. c) Long-lifetime pulsed gas valve. d) Fast pulsed gas valve driver and controller. e) High-voltage, repetitive capacitor charging power processing unit. During the course of testing, a number of specific tests were conducted, including several that, to our knowledge, have either never been previously conducted (such as multi-KW repetition-rate operation) or have not been performed since the early 1990s (direct IPPT thrust measurements).2 Conical theta-pinch IPPT thrust stand measurements are presented in Fig. 1 while various time-integrated and time

Polzin, Kurt A.↗

NASA NTRS2016

Hot-Fire Test Results of an Oxygen/RP-2 Multi-Element Oxidizer-Rich Staged-Combustion Integrated Test Article

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. For the thrust chamber assembly of the test article, several configurations of new main injectors, using relatively conventional gas-centered swirl coaxial injector elements, were designed and fabricated. The design and fabrication of these main injectors are described in a companion paper at this JANNAF meeting. New ablative combustion chambers were fabricated based on hardware previously used at NASA for testing at similar size and pressure. An existing oxygen/RP-1 oxidizer-rich subscale preburner injector from a previous NASA-funded program, along with existing and new inter-connecting hot gas duct hardware, were used to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. Results from independent hot-fire tests of the preburner injector in a combustion chamber with a sonic throat are described in companion papers at this JANNAF conference. The resulting integrated test article - which includes the preburner, inter-connecting hot gas duct, main injector, and ablative combustion chamber - was assembled at Test Stand 116 at the East Test Area of the NASA Marshall Space Flight Center. The test article was well instrumented with static and dynamic pressure, temperature, and acceleration sensors to allow the collected data to be used for combustion analysis model development. Hot-fire testing was conducted with main combustion chamber pressures ranging from 1400 to 2100 psia, and main combustion chamber mixture ratios ranging from 2.4 to 2.9. Different levels of fuel film cooling injected from the injector face were examined ranging from none to about 12% of the total fuel flow. This paper presents the hot-fire test results of the integrated test article. Combustion performance, stability, thermal, and compatibility characteristics of both the preburner and the thrust chamber are described. Another companion paper at this JANNAF meeting includes additional and more detailed test data regarding the combustion dynamics and stability characteristics.

Hulka, J. R.↗

NASA NTRS1996

Medical and surgical applications of space biosensor technology

Researchers in space life sciences are rapidly approaching a technology impasse. Many of the critical questions on the impact of spaceflight on living systems simply cannot be answered with the limited available technologies. Research subjects, particularly small animal models like the rat, must be allowed to function relatively untended and unrestrained for long periods to fully reflect the impact of microgravity and spaceflight on their behavior and physiology. These requirements preclude the use of present hard-wired instrumentation techniques and limited data acquisition systems. Implantable sensors and miniaturized biotelemetry are the only means of capturing the fundamental and critical data. This same biosensor and biotelemetry technology has direct application to Earth-based medicine and surgery. Continuous, on-line data acquisition and improved measurement capabilities combined with the ease and flexibility offered by automated, wireless, and portable instruments and data systems, should provide a boon to the health care industry. Playing a key role in this technology revolution is the Sensors 2000! (S2K!) Program at NASA Ames Research Center. S2K!, in collaboration with space life sciences researchers and managers, provides an integrated capability for sensor technology development and applications, including advanced biosensor technology development, spaceflight hardware development, and technology transfer and commercialization. S2K! is presently collaborating on several spaceflight projects with dual-use medical applications. One prime example is a collaboration with the Fetal Treatment Center (FTC) at the University of California at San Francisco. The goal is to develop and apply implantable chemical sensor and biotelemetry technology to continuously monitor fetal patients during extra-uterine surgery, replacement into the womb, through birth and beyond. Once validated for ground use, the method will be transitioned to spaceflight applications to remotely monitor key biochemical parameters in flight animals. Successful application of NASA implantable biosensor and biotelemetry technologies should accelerate the advancement of this and other modern medical procedures while furthering the exploration of life in space.

NASA Discipline General Space Life Sciences↗

NASA NTRS2002

Occupant Responses in a Full-Scale Crash Test of the Sikorsky ACAP Helicopter

A full-scale crash test of the Sikorsky Advanced Composite Airframe Program (ACAP) helicopter was performed in 1999 to generate experimental data for correlation with a crash simulation developed using an explicit nonlinear, transient dynamic finite element code. The airframe was the residual flight test hardware from the ACAP program. For the test, the aircraft was outfitted with two crew and two troop seats, and four anthropomorphic test dummies. While the results of the impact test and crash simulation have been documented fairly extensively in the literature, the focus of this paper is to present the detailed occupant response data obtained from the crash test and to correlate the results with injury prediction models. These injury models include the Dynamic Response Index (DRI), the Head Injury Criteria (HIC), the spinal load requirement defined in FAR Part 27.562(c), and a comparison of the duration and magnitude of the occupant vertical acceleration responses with the Eiband whole-body acceleration tolerance curve.

Jackson, Karen E.↗

NASA NTRS2002

Occupant Responses in a Full-Scale Crash Test of the Sikorsky ACAP Helicopter

Jackson, Karen E.↗

NASA NTRS2019

A Comparison of ARTEMIS Data with the Lunar Plasma Design Environment for NASA Crewed Missions

NASA’s Gateway will provide the capability for sustaining a human presence in cis-lunar space. Operations of the Gateway will include spacecraft dockings, extra vehicular activities (EVA), and high-power solar arrays. NASA’s experience with the International Space Station highlighted the importance of evaluating spacecraft charging effects for such operations. For crewed spacecraft, which tend to employ the use of dielectric surfaces in this dynamic plasma environment, reliance on spacecraft charging simulation packages, such as the NASA/Air Force Spacecraft Charging Analyzer Program (Nascap-2k) [Mandell et al., 2006] and Spacecraft Plasma Interaction System (SPIS) [Roussel et al., 2008], is required to understand the risks to hardware and humans. The variability in the lunar plasma environment as the Moon revolves around the Earth, lunar wake effects, and a strong dependency on photoemission and secondary electron emission creates challenges for spacecraft charging analysis. The Design Specification for Natural Environments (DSNE) [NASA, MSFC] is the primary resource for space environments affecting NASA’s crewed missions, and the DSNE provides plasma environments in a standard form for input into simulation packages. NASA developed the existing lunar plasma environment using data from Geotail [Nishida, 1994] along with published lunar plasma wake models [Halekas et al., 2005] based on Lunar Prospector. Since 2011, NASA’s twin Acceleration Reconnection Turbulence & Electrodynamics of Moon’s Interaction with the Sun (ARTEMIS) satellites [Angelopoulos, 2010] have been collecting high resolution plasma and fields observations within the lunar plasma environment providing a much larger dataset of the plasma properties in cislunar space. This research compares the existing lunar plasma environment definition with ARTEMIS data and makes recommendations on the refinement of the environment definition for future lunar missions.

ARTEMIS mission↗

NASA NTRS1987

Neutralization of beam-emitting spacecraft by plasma injection

An impulsive plasma injection has been used to study charge neutralization of the Space Shuttle Orbiter while it was emitting an electron beam into space. This investigation was performed by Space Experiments with Particle Accelerators on Spacelab-1. A plasma consisting of 10 to the 19th argon ion-electron pairs was injected into space for 1 ms while an electron beam was also being emitted into space. The electron beam energy and current were as high as 5 keV and 300 mA. While the orbiter potential was positive before the plasma injection and began to decrease during the plasma injection, it was near zero for 6 to 20 ms after the plasma injection. The recovery time to the initial level of charging varied from 10 to 100 ms. In a laboratory test in a large space chamber using the same flight hardware, the neutralization time was 8-17 ms and the recovery time was 11-20 ms. The long duration of the neutralization effect in space can be explained by a model of diffusion of the cold plasma which is produced near the Orbiter by charge exchange between the neutral argon atoms and the energetic argon ions during plasma injection.

Sasaki, S.↗

NASA NTRS1996

Energy Beam Highways Through the Skies

The emergence of Energy Beam Flight Transportation Systems could dramatically change the way we travel in the 21st Century. A framework for formulating 'Highways of Light' and the top level architectures that invoke radically new Space Power Grid infrastructure, are introduced. Basically, such flight systems, hereafter called Lightcraft, would employ off-board energy beam sources (either laser or microwave) to energize on-board dependent 'motors' -- instead of the traditional autonomous 'engines' with their on-board energy sources (e.g., chemical fuels). Extreme reductions in vehicle dry mass appear feasible with the use of off-board power and a high degree of on-board artificial intelligence. Such vehicles may no longer need airports for refueling (since they require no propellant), and could possibly pick up travelers at their homes -- before motoring over to one of many local boost stations, for the flight out. With off-board power, hyper-energetic acceleration performance and boost-glide trajectories become feasible. Hypersonic MS airbreathing propulsion can enable boosts up to twice escape velocity, which will cut trip times to the moon down to 5.5 hours. The predominant technological, environmental and social factors that will result from such transportation systems will be stressed. This presentation first introduces the remote source siting options for the space power system infrastructure, and then provides three representative laser/microwave Lightcraft options (derived from historical Case Studies): i.e., 'Acorn', 'Toy Top', and 'Disc.' Next the gamut of combined-cycle engine options developed for these Lightcraft are examined -- to illuminate the 'emerging technologies' that must be harnessed to produce flight hardware. Needed proof-of concept experiments are identified, along with the Macro-Level Issues that can springboard these revolutionary concepts into hardware reality.

Myrabo, Leik N.↗

NASA NTRS2005

Results of a 2000 hour wear test of the NEXIS ion engine

The Nuclear Electric Xenon Ion System (NEXIS) ion thruster was developed for potential outer planet robotic missions under NASA's Prometheus program. This engine was designed to operate at power levels ranging from 16 to over 20 kWe at specific impulses of 6000 to 7500 s for burn times of up to 10 years, satisfying the requirements of nuclear electric propulsion systems such as that on the proposed Prometheus 1 mission to explore the icy moons of Jupiter. State-of-the-art performance and life assessment tools were used to design the thruster. Following the successful performance validation of a Laboratory Model thruster, Development Model hardware was fabricated and subjected to vibration and wear testing. The results of a 2000-hour wear test are reported herein. Thruster performance achieved the target requirements and was steady for the duration of the test. Ion optics performance was similarly stable. Discharge loss increases of 6 eV/ion were observed in the first 500 hours of the test and were attributed to primary electron energy decreases due to cathode insert conditioning. Relatively high recycle rates were observed and were identified to be high-voltage-to-ground arcs in the back of the thruster caused by wire insulation outgassing and electron penetration through the plasma screen. Field emission of electrons between the accelerator and screen grids was observed and attributed to evolution of field emitter sites at accelerator grid aperture edges caused by ion bombardment. Preliminary modeling and analysis indicates that the NEXIS engine can meet mission performance requirements over the required lifetime. Finally, successful validation of the NEXIS design methodology, design tools, and technologies with the results of the wear test and companion performance and vibration tests presents significant applicability of the NEXIS development effort to missions of near-term as well as long-term interest for NASA.

Nuclear Electric Xenon Ion System (NEXIS)Thruster ↗

NASA NTRS2005

Results of a 2000 Hour Wear Tof the NEXIS Ion Engine

Thruster performance↗