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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2010

Development of a Contingency Capillary Wastewater Management Device

The Personal Body .Attached Liquid Liquidator (PBALL) is conceived as a passive, capillary driven contingency wastewater disposal device. In this contingency scenario, the airflow system on the NASA Crew Exploration Vehicle (CEV) is assumed to have failed, leaving only passive hardware and vacuum vent to dispose of the wastewater. To meet these needs, the PBALL was conceived to rely on capillary action and urine wetting design considerations. The PBALL is designed to accommodate a range of wetting conditions, from 0deg < (theta)adv approx. 90deg, be adaptable for both male and female use, collect and retain up to a liter of urine, minimize splash-back, and allow continuous drain of the wastewater to vacuum while minimizing cabin air loss. A sub-scale PBALL test article was demonstrated on NASA's reduced gravity aircraft in April, 2010.

Thomas, Evan A.↗
NASA NTRS2004

Novel Horn Designs for Power Ultrasonics

Ultrasonic horns are used in a variety of industrial and medical applications. At JPL a rock-sampling tool based on an ultrasonic horn was developed to drill, abrade and core rock samples including hard basalts. This device is an impact device, which uses ultrasonic vibratlons that occur at the horn tip to produce a sonic resonance with the aid of a loosely connected mass. Although standard horns are found in many current industrial designs they suffer from a few key limitations when used for USDC applications. Manufacturing a horn requires turning down stock material (e.g. Titanium) from the larger outer diameter to the horn tip diameter, and this process is both time consuming and wasteful. In this paper, we present novel horn designs that are specifically designed for impact applications as the USDC. One such design addressed the excasive length that is involved with the use of the horn limiting its applications when system dimensions are constrained. For this purpose, a folded horn design was conceived that reduces the overall length of the resonator (physical length) but maintains or increases the acoustic length. Initial experiments with horns having such P design indicate that the tip displacement can be further adjusted by phasing the bending displacements and the extensional displacements. Another conceived horn design is the 'dog bone' horn that uses an end mass on the horn tip io increase the impact efficiency of the horn. In this paper, the experimental results for these novel born designs are presented and compared to the results predicted by theory.

acutators↗
NASA NTRS2011

SRMS History, Evolution and Lessons Learned

Early in the development of the Space Shuttle, it became clear that NASA needed a method of deploying and retrieving payloads from the payload bay. The Shuttle Remote Manipulator System (SRMS) was developed to fill this need. The 50 foot long robotic arm is an anthropomorphic design consisting of three electromechanical joints, six degrees of freedom, and two boom segments. Its composite boom construction provided a light weight solution needed for space operations. Additionally, a method of capturing payloads with the arm was required and a unique End Effector was developed using an electromechanical snare mechanism. The SRMS is operated using a Displays and Controls Panel and hand controllers located within the aft crew compartment of the shuttle. Although the SRMS was originally conceived to deploy and retrieve payloads, its generic capabilities allowed it to perform many other functions not originally conceived of. Over the years it has been used for deploying and retrieving constrained and free flying payloads, maneuvering and supporting EVA astronauts, satellite repair, International Space Station construction, and as a viewing aid for on-orbit International Space Station operations. After the Columbia accident, a robotically compatible Orbiter Boom Sensor System (OBSS) was developed and used in conjunction with the SRMS to scan the Thermal Protection System (TPS) of the shuttle. These scans ensure there is not a breach of the TPS prior to shuttle re-entry. Ground operations and pre mission simulation, analysis and planning played a major role in the success of the SRMS program. A Systems Engineering Simulator (SES) was developed to provide a utility complimentary to open loop engineering simulations. This system provided a closed-loop real-time pilot-driven simulation giving visual feedback, display and control panel interaction, and integration with other vehicle systems, such as GN&C. It has been useful for many more applications than traditional training. Evolution of the simulations, guided by the Math Model Working Group, showed the utility of input from multiple modeling groups with a structured forum for discussion.There were many unique development challenges in the areas of hardware, software, certification, modeling and simulation. Over the years, upgrades and enhancements were implemented to increase the capability, performance and safety of the SRMS. The history and evolution of the SRMS program provided many lessons learned that can be used for future space robotic systems.

Jorgensen, Glenn↗
NASA NTRS2011

Teleoperated Marsupial Mobile Sensor Platform Pair for Telepresence Insertion Into Challenging Structures

A platform has been developed for two or more vehicles with one or more residing within the other (a marsupial pair). This configuration consists of a large, versatile robot that is carrying a smaller, more specialized autonomous operating robot(s) and/or mobile repeaters for extended transmission. The larger vehicle, which is equipped with a ramp and/or a robotic arm, is used to operate over a more challenging topography than the smaller one(s) that may have a more limited inspection area to traverse. The intended use of this concept is to facilitate the insertion of a small video camera and sensor platform into a difficult entry area. In a terrestrial application, this may be a bus or a subway car with narrow aisles or steep stairs. The first field-tested configuration is a tracked vehicle bearing a rigid ramp of fixed length and width. A smaller six-wheeled vehicle approximately 10 in. (25 cm) wide by 12 in. (30 cm) long resides at the end of the ramp within the larger vehicle. The ramp extends from the larger vehicle and is tipped up into the air. Using video feedback from a camera atop the larger robot, the operator at a remote location can steer the larger vehicle to the bus door. Once positioned at the door, the operator can switch video feedback to a camera at the end of the ramp to facilitate the mating of the end of the ramp to the top landing at the upper terminus of the steps. The ramp can be lowered by remote control until its end is in contact with the top landing. At the same time, the end of the ramp bearing the smaller vehicle is raised to minimize the angle of the slope the smaller vehicle has to climb, and further gives the operator a better view of the entry to the bus from the smaller vehicle. Control is passed over to the smaller vehicle and, using video feedback from the camera, it is driven up the ramp, turned oblique into the bus, and then sent down the aisle for surveillance. The demonstrated vehicle was used to scale the steps leading to the interior of a bus whose landing is 44 in. (.1.1 m) from the road surface. This vehicle can position the end of its ramp to a surface over 50 in. (.1.3 m) above ground level and can drive over rail heights exceeding 6 in. (.15 cm). Thus configured, this vehicle can conceivably deliver the smaller robot to the end platform of New York City subway cars from between the rails. This innovation is scalable to other formulations for size, mobility, and surveillance functions. Conceivably the larger vehicle can be configured to traverse unstable rubble and debris to transport a smaller search and rescue vehicle as close as possible to the scene of a disaster such as a collapsed building. The smaller vehicle, tethered or otherwise, and capable of penetrating and traversing within the confined spaces in the collapsed structure, can transport imaging and other sensors to look for victims or other targets.

Krasowski, Michael J.↗
NASA NTRS2012

An Intelligent Strain Gauge with Debond Detection and Temperature Compensation

The harsh rocket propulsion test environment will expose any inadequacies associated with preexisting instrumentation technologies, and the criticality for collecting reliable test data justifies investigating any encountered data anomalies. Novel concepts for improved systems are often conceived during the high scrutiny investigations by individuals with an in-depth knowledge from maintaining critical test operations. The Intelligent Strain Gauge concept was conceived while performing these kinds of activities. However, the novel concepts are often unexplored even if it has the potential for advancing the current state of the art. Maturing these kinds of concepts is often considered to be a tangential development or a research project which are both normally abandoned within the propulsion-oriented environment. It is also difficult to justify these kinds of projects as a facility enhancement because facility developments are only accepted for mature and proven technologies. Fortunately, the CIF program has provided an avenue for bringing the Intelligent Strain Gauge to fruition. Two types of fully functional smart strain gauges capable of performing reliable and sensitive debond detection have been successfully produced. Ordinary gauges are designed to provide test article data and they lack the ability to supply information concerning the gauge itself. A gauge is considered to be a smart gauge when it provides supplementary data relating other relevant attributes for performing diagnostic function or producing enhanced data. The developed strain gauges provide supplementary signals by measuring strain and temperature through embedded Karma and nickel chromium (NiCr) alloy elements. Intelligently interpreting the supplementary data into valuable information can be performed manually, however, integrating this functionality into an automatic system is considered to be an intelligent gauge. This was achieved while maintaining a very low mass. The low mass enables debond detection and temperature compensation to be performed when the gauge is utilized on small test articles. It was also found that the element's mass must be relatively small to avoid overbearing the desired thermal dissipation characteristics. Detecting the degradation of a gauge s bond was reliably achieved by correlating thermal dissipation with the bond s integrity. This was accomplished by precisely coupling a NiCr element with a Karma element for accurately interjecting and quantifying thermal energy. A finite amount of thermal energy is consistently placed in the gauge by electrically powering the NiCr element. The energy will only be temporarily stored before it begins to dissipate into the surrounding structure through the gauge bond. The ability to transmit the energy into the structure becomes greatly inhibited by any discontinuity in the bond s substrate. Therefore, the way the thermal dissipation occurs will reveal even the slightest change in the integrity of the bond.

Jensen, Scott L.↗
NASA NTRS2012

Determination of Survivable Fires

At NASA, there exists no standardized design or testing protocol for spacecraft fire suppression systems (either handheld or total flooding designs). An extinguisher's efficacy in safely suppressing any reasonable or conceivable fire is the primary benchmark. That concept, however, leads to the question of what a reasonable or conceivable fire is. While there exists the temptation to over-size' the fire extinguisher, weight and volume considerations on spacecraft will always (justifiably) push for the minimum size extinguisher required. This paper attempts to address the question of extinguisher size by examining how large a fire a crew member could successfully survive and extinguish in the confines of a spacecraft. The hazards to the crew and equipment during an accidental fire include excessive pressure rise resulting in a catastrophic rupture of the vehicle skin, excessive temperatures that burn or incapacitate the crew (due to hyperthermia), carbon dioxide build-up or other accumulation of other combustion products (e.g. carbon monoxide). Estimates of these quantities are determined as a function of fire size and mass of material burned. This then becomes the basis for determining the maximum size of a target fire for future fire extinguisher testing.

Dietrich, D. L.↗
NASA NTRS2013

US Rocket Propulsion Industrial Base Health Metrics

The number of active liquid rocket engine and solid rocket motor development programs has severely declined since the "space race" of the 1950s and 1960s center dot This downward trend has been exacerbated by the retirement of the Space Shuttle, transition from the Constellation Program to the Space launch System (SLS) and similar activity in DoD programs center dot In addition with consolidation in the industry, the rocket propulsion industrial base is under stress. To Improve the "health" of the RPIB, we need to understand - The current condition of the RPIB - How this compares to past history - The trend of RPIB health center dot This drives the need for a concise set of "metrics" - Analogous to the basic data a physician uses to determine the state of health of his patients - Easy to measure and collect - The trend is often more useful than the actual data point - Can be used to focus on problem areas and develop preventative measures The nation's capability to conceive, design, develop, manufacture, test, and support missions using liquid rocket engines and solid rocket motors that are critical to its national security, economic health and growth, and future scientific needs. center dot The RPIB encompasses US government, academic, and commercial (including industry primes and their supplier base) research, development, test, evaluation, and manufacturing capabilities and facilities. center dot The RPIB includes the skilled workforce, related intellectual property, engineering and support services, and supply chain operations and management. This definition touches the five main segments of the U.S. RPIB as categorized by the USG: defense, intelligence community, civil government, academia, and commercial sector. The nation's capability to conceive, design, develop, manufacture, test, and support missions using liquid rocket engines and solid rocket motors that are critical to its national security, economic health and growth, and future scientific needs. center dot The RPIB encompasses US government, academic, and commercial (including industry primes and their supplier base) research, development, test, evaluation, and manufacturing capabilities and facilities. center dot The RPIB includes the skilled workforce, related intellectual property, engineering and support services, and supply chain operations and management. This definition touches the five main segments of the U.S. RPIB as categorized by the USG: defense, intelligence community, civil government, academia, and commercial sector.

Doreswamy, Rajiv↗
NASA NTRS2017

Synchronous Separation, Seaming, Sealing and Sterilization (S4) using Brazing for Sample Containerization and Planetary Protection

The potential return of samples back to Earth from other planetary bodies would be based on planetary protection requirements that vary depending on the type of body [1]. Potential Mars Sample Return would require the protection of our planet from backward contamination. To fulfill this requirement, it would be necessary to implement “break the chain of contact (BTC)” process, where any material reaching Earth would have to be inside a container that is sealed with an extremely high level of confidence. In order to accomplish this, it would be necessary to contain the acquired samples and destroy any potential biological materials that may have contaminated the external surface of the container, while protecting the samples for further analysis. Using brazing, a novel synchronous separation, seaming, sealing and sterilization (S4) process for sample containerization and planetary protection has been conceived and demonstrated. A prototype double-wall container with inner and outer shells and Earth clean interstitial space was used for this demonstration. For potential Mars sample return, the double wall container would be consist of two halves and prepared on Earth. The on-orbit execution would consist of inserting the sample into one of the halves and then mating to the other half and melt the braze material to perform the S4 process. The use of brazing material that melts at temperatures higher than 500OC would assure sterilization of the exposed areas due to pyrolysis since carbon bonds are broken at this temperature. The process consists of two-steps, Step-1: the double wall container halves are fabricated and brazed on Earth; and Step-2: Assembly and brazing the samples on orbit. To prevent potential jamming during the process of mating the two halves of the double-wall container and the extraction of the brazed inner container, a double cone-within-cone approach has been conceived. The results of this study are described and discussed in this manuscript.

Bar-Cohen, Yoseph↗
NASA NTRS2016

MarCO: Interplanetary Mission Development on a CubeSat Scale

Shortly after JPL’s Interior Exploration using Seismic Investigations, Geodesy and Heat Transport (InSight) mission launches, separates, and commences its cruise phase, two CubeSats will deploy from the launch vehicle’s upper stage and begin independent flight to Mars (Fig. 1). During InSight’s entry, descent, and landing (EDL) sequence, these twin Mars Cube One (MarCO) spacecraft will fly 3,500 km above the Martian surface, recording and relaying InSight UHF radio data to the Deep Space Network (DSN) on Earth1. MarCO is a twin CubeSat mission developed by the NASA Jet Propulsion Laboratory (JPL) to accompany the InSight (Interior Exploration using Seismic Investigations, Geodesy and Heat Transport) Mars mission lander. MarCO's primary mission objective is to launch with InSight and independently fly to Mars to serve as a communications relay during InSight's entry, descent, and landing (EDL) phase. MarCO represents a new type of deep space mission: CubeSats at Mars. Building on the development of JPL's first interplanetary CubeSat project, the Interplanetary Nano-Spacecraft Pathfinder in Relevant Environment (INSPIRE), MarCO further refined the approach to hardware, software, and ground architecture development to solve the challenges of quickly building low-budget spacecraft to fly to Mars. The greatest constraint, beyond others typical of CubeSat missions, was time. The duration between MarCO's conception to completion of spacecraft assembly was less than two years - an unprecedented schedule for any planetary mission to date. Through necessity, MarCO has built on previous experience, procedures, systems, and development methodologies, defining a new niche for supporting larger primary missions. The MarCO spacecraft are poised to write a new chapter in deep space exploration. Originally slated to launch and reach Mars in 2016, the InSight mission schedule subsequently slipped to 2018. During the original landing of InSight, Earth would not be in view, and no orbiter around Mars would have been in position to both receive UHF EDL data and simultaneously relay it back to Earth. It was from this obstacle that MarCO was conceived. Regardless of any changes to InSight’s 2018 EDL configuration geometry, MarCO is still expected to fly and serve in the same capacity as originally designed: the first CubeSat mission to Mars. CubeSats have historically been firmly in the domain of universities and small companies. As first conceived, they served as a platform upon which to teach all aspects of the space mission lifecycle. JPL took on this mission type with Interplanetary Nano-Spacecraft Pathfinder in Relevant Environment2 (INSPIRE), moving the concept into a new domain: deep space. Building from the INSPIRE platform and lessons learned, MarCO addressed new challenges in the domain of planetary missions: independent interplanetary flight and navigation, integration with a large-scale mission, long-distance and long-delay communication, short development time, and a small development team. Of these, the greatest constraint was schedule: only 18 months passed from conception of mission concept until delivery of fully assembled and tested flight hardware. Careful selection of mission team, along with extensive use of off-the-shelf equipment, and streamlining automated processes, was essential. This achievement represents the next step in the evolution of CubeSats beyond low-Earth orbit.

Werne, Thomas↗
NASA NTRSDate not supplied

Observing System Simulations for the AOS Mission

The Earth System Observatory (ESO) is NASA’s response to the recommendations of the 2017 Earth Sciences Decadal Survey conducted by the US National Academy of Sciences, Engineering and Medicine. The ESO is being conceived as a set of fully integrated missions addressing 4 main Earth science focus areas including aerosols, clouds, convection and precipitation (jointly re-ferred to as AOS, the Atmosphere Observing System). ESO ground breaking observations will provide critical measurements to address societally relevant problems in climate change, natural hazard mitiga-tion, fighting forest fires, and improving real-time agricultural processes. A critical element of the AOS observing strategy is to make extensive use of new passive and active sen-sors as well as of the so-called Program-of-Record (PoR), complemented by a fully integrated sub-orbital component. In order to achieve maximum benefit, all these observations need to be integrated into comprehensive observing and modeling/data assimilation systems. Such an approach requires compre-hensive model-data synthesis capabilities that needs to be conceived in conjunction with the space-based and suborbital components of AOS. In this presentation we will summarize the major science goals of AOS including cloud feedbacks, at-mospheric convection, emphasizing aerosol processes and aerosol radiative effects, and the synergistic aspects of clouds-precipitation-aerosol interactions. We will describe examples of the observing system simulation capabilities being developed for AOS, including global storm resolving nature runs, detailed instrument and retrieval simulators, as well as fast retrieval emulators for instrumenting climate models. This simulation environment, being developed under NASA’s open-source science initiative, will permit us to explore how AOS data will be used across space and time to better initialize forecasts and train modeling systems, and to infuse models and data assimilation systems with AOS data, well before launch.

Arlindo da Silva↗
NASA NTRS1962

Some Abort Techniques and Procedures for Manned Spacecraft

This paper is concerned with our present state of knowledge on the procedures and effectiveness of aborts from manned orbital and lunar missions. By abort is meant a deliberate or unintentional termination of the primary mission followed by an expedient return to earth of the payload (spacecraft and crew). A deliberate abort may be initiated if a critical component failed and thus impaired the probability of mission success or the crew's safety. Examples of this kind of abort are failure of the launch guidance system or the oxygen and cabin pressure system. Also, an unintentional abort may occur during the launch phase due to the complete or partial failure of one of the launch-vehicle stages to operate. For each mission every conceivable set of circumstances must be considered and evaluated to see if an abort of the mission is called for and to determine what sequence of events should occur to bring the spacecraft and crew safely back to earth with the highest probability of success. Factors which must be considered are: how to specify and sense an abort condition, how much fuel should be put onboard the vehicle for abort purposes, what size and kind of rocket engines are best suited for each mission, and what compromise is incurred due to other mission requirements? If an abort occurs, the optimum time of application as well as the magnitude and direction of the velocity changes must be determined beforehand and the crew must be clearly informed of what action is required or what the situation is so that they can decide the proper action.

Maneuver↗
NASA NTRS1962

A CONSIDERATION OF THE U.S. NAVY STRATO-LAB BALLOON PROGRAM AND ITS CONTRIBUTIONS TO MANNED SPACE FLIGHT

As the coined name Strato-Lab implies, it is a program which provides a stratospheric laboratory utilizing a plastic balloon platform for manned research investigations. The project was originated in 1954; actual work began early in 1955 and the first manned stratospheric flights were made in 1956. It was conceived and has been continued with the following three objectives: (1) Provide a high-altitude research system, a scientific laboratory, permitting man's active scientific participation (2) Provide an opportunity for studies of man's reactions to an actual environment of physiological and psychological stress (3) Allow an opportunity for man to conduct tests and evaluate specific components and/or techniques in an extremely high altitude environment.

Stratosphere↗
NASA NTRS1962

THE COMBINATION OF ACTIVE AND PASSIVE SHIELDING

I is conceivable that improved shielding of space vehicles against high-energy charged-particle radiation can be obtained by combining active and passive shielding. Methods of shielding calculations and some preliminary observations pertaining to active-passive shielding are given. An analysis based upon the field of a magnetic dipole indicates that weight savings in bulk shielding can be accomplished.

SPACECRAFT SHIELDING↗
NASA NTRS1962

Development of the Saturn Space Vehicle

The Saturn launch vehicle program has become a major element in our national space effort within little more than 3 years. At the beginning of the program, in 1958, various Saturn configurations were planned to lift instrumented payloads of 20,000 to 40,000 lb. into Earth orbit. As research and development progressed, it became apparent that the vehicle design approach had far more potential than anticipated that it was, in fact, a breakthrough into a new class of heavy launch vehicles. Thus, within the space of a few years, it was possible to consider even more advanced missions. In addition to placing multiton payloads in orbit, we can seriously conceive sending over 90,000 lb. of instruments on escape missions, using advanced Saturn-class vehicles. With these vehicles, we plan to accomplish lunar reconnaissance and manned lunar landings scheduled for the latter part of this decade. Likewise, these vehicles will permit flight tests of advanced nuclear power plants and stages.

SATURN LAUNCH VEHICLE↗
NASA NTRS1959

The Ten Year Plan of the National Aeronautics and Space Administration

The long term national objectives of the United States in aeronautical and space activities are stated in general terms in the enabling legislation establishing NASA. In operational terms, these objectives, are instructions to explore and to utilize both the atmosphere and the regions outside the earth's atmosphere for peaceful and scientific purposes, while at the same time providing research support to the Department of Defense. These objectives can be attained only by means of a broad and soundly conceived program of research, development and operations in space. In the long run, such activities should make feasible the manned exploration of the moon and the nearby planets, and this exploration may thus be taken as a long-term goal of NASA activities. To assure steady and rapid progress toward these objectives, a NASA Ten Year Plan has been developed and is presented in this document.

Source record↗
NASA NTRS1965

Space vacuum investigations

Inaccessible-Non Repairable- Costly: The meaning of these words is absolute in the event of a complete failure of an orbiting spacecraft. Adequate design, or rather, minimizing chance of failure, challenges all scientists and engineers. Where experience has not been developed, the risk involved in a new application must be weighed against the consequence of a design failure. In space work, the spectrum of consequence is extremely wide. It ranges from loss of scientific data on one hand, to loss of life on the other.. There is little question that knowledge of the space environment and its effects must be determined to produce reliable spacecraft. Tests to prove the design must be performed with equipment capable of simulating the effects of the environment in which the design is to operate. This does not necessarily mean that the natural environment must be duplicated. The degree of environmental simulation that should be achieved depends primarily upon the significance of the physical effects it causes. The purpose of space vacuum simulation in an environmental test program should be accurately conceived and the testing phased to obtain maximum information from the limited samples available for test. This discussion is confined to some of the factors involved in simulating the effects of space vacuum environment. This survey is by no means complete, but will touch on some of the highlights in this field: 1. A description of the environment 2. The level of simulation necessary to evaluate its effects 3. Test equipment capability 4. Experience in test 5. Experience in orbit 6. Observations & future emphasis

VACUUM EFFECT↗
NASA NTRS1962

Mechanical Design of Spacecraft

In the spring of 1962, engineers from the Engineering Mechanics Division of the Jet Propulsion Laboratory gave a series of lectures on spacecraft design at the Engineering Design seminars conducted at the California Institute of Technology. Several of these lectures were subsequently given at Stanford University as part of the Space Technology seminar series sponsored by the Department of Aeronautics and Astronautics. Presented here are notes taken from these lectures. The lectures were conceived with the intent of providing the audience with a glimpse of the activities of a few mechanical engineers who are involved in designing, building, and testing spacecraft. Engineering courses generally consist of heavily idealized problems in order to allow the more efficient teaching of mathematical technique. Students, therefore, receive a somewhat limited exposure to actual engineering problems, which are typified by more unknowns than equations. For this reason it was considered valuable to demonstrate some of the problems faced by spacecraft designers, the processes used to arrive at solutions, and the interactions between the engineer and the remainder of the organization in which he is constrained to operate. These lecture notes are not so much a compilation of sophisticated techniques of analysis as they are a collection of examples of spacecraft hardware and associated problems. They will be of interest not so much to the experienced spacecraft designer as to those who wonder what part the mechanical engineer plays in an effort such as the exploration of space.

STRUCTURAL MATERIAL↗