Search NASA⌕ Search

SEARCH · Search NASA

Results for “PROPELLANT SENSITIVITY”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

CO2 (dry ice) cleaning system

Tomco Equipment Company has participated in the dry ice (solid carbon dioxide, CO2) cleaning industry for over ten years as a pioneer in the manufacturer of high density, dry ice cleaning pellet production equipment. For over four years Tomco high density pelletizers have been available to the dry ice cleaning industry. Approximately one year ago Tomco introduced the DI-250, a new dry ice blast unit making Tomco a single source supplier for sublimable media, particle blast, cleaning systems. This new blast unit is an all pneumatic, single discharge hose device. It meters the insertion of 1/8 inch diameter (or smaller), high density, dry ice pellets into a high pressure, propellant gas stream. The dry ice and propellant streams are controlled and mixed from the blast cabinet. From there the mixture is transported to the nozzle where the pellets are accelerated to an appropriate blasting velocity. When directed to impact upon a target area, these dry ice pellets have sufficient energy to effectively remove most surface coatings through dry, abrasive contact. The meta-stable, dry ice pellets used for CO2 cleaning, while labeled 'high density,' are less dense than alternate, abrasive, particle blast media. In addition, after contacting the target surface, they return to their equilibrium condition: a superheated gas state. Most currently used grit blasting media are silicon dioxide based, which possess a sharp tetrahedral molecular structure. Silicon dioxide crystal structures will always produce smaller sharp-edged replicas of the original crystal upon fracture. Larger, softer dry ice pellets do not share the same sharp-edged crystalline structures as their non-sublimable counterparts when broken. In fact, upon contact with the target surface, dry ice pellets will plastically deform and break apart. As such, dry ice cleaning is less harmful to sensitive substrates, workers and the environment than chemical or abrasive cleaning systems. Dry ice cleaning system components include: a dry ice pellet supply, a non-reactive propellant gas source, a pellet and propellant metering device, and a media transport and acceleration hose and nozzle arrangement. Dry ice cleaning system operating parameters include: choice of propellant gas, its pressure and temperature, dry ice mass flow rate, dry ice pellet size and shape, and acceleration nozzle configuration. These parameters may be modified to fit different applications. The growth of the dry ice cleaning industry will depend upon timely data acquisition of the effects that independent changes in these parameters have on cleaning rates, with respect to different surface coating and substrate combinations. With this data, optimization of cleaning rates for particular applications will be possible. The analysis of the applicable range of modulation of these parameters, within system component mechanical constraints, has just begun.

Barnett, Donald M.↗
NASA NTRS2015

Mars Hybrid Propulsion System Trajectory Analysis: Crew Missions - Part I

NASAs Human spaceflight Architecture team is developing a reusable hybrid transportation architecture in which both chemical and electric propulsion systems are used to send crew and cargo to Mars destinations such as Phobos, Deimos, the surface of Mars, and other orbits around Mars. By combining chemical and electrical propulsion into a single space- ship and applying each where it is more effective, the hybrid architecture enables a series of Mars trajectories that are more fuel-efficient than an all chemical architecture without significant increases in flight times. This paper provides the analysis of the interplanetary segments of the three Evolvable Mars Campaign crew missions to Mars using the hybrid transportation architecture. The trajectory analysis provides departure and arrival dates and propellant needs for the three crew missions that are used by the campaign analysis team for campaign build-up and logistics aggregation analysis. Sensitivity analyses were performed to investigate the impact of mass growth, departure window, and propulsion system performance on the hybrid transportation architecture. The results and system analysis from this paper contribute to analyses of the other human spaceflight architecture team tasks and feed into the definition of the Evolvable Mars Campaign.

Chai, Patrick R.↗
NASA NTRS2004

An aerobot for global in situ exploration of Titan

This paper describes the design and component testing of an aerobot that would be capable of global in situ exploration of Saturn's moon, Titan, over a 6 to 12 month mission lifetime. The proposed aerobot is a propeller-driven, buoyant vehicle that resembles terrestrial airships. However, the extremely cold Titan environment requires the use of cryogenic materials of construction and careful thermal design for protection of temperature-sensitive payload elements. Multiple candidate balloon materials have been identified based on extensive laboratory testing at 77 K. The most promising materials to date are laminates comprised of polyester fabrics and/or films with areal densities in the range of 40-100 g/m2. The aerobot hull is a streamlined ellipsoid 14 meters in length with a maximum diameter of 3 meters. The enclosed volume of 60 m3 is sufficient to float a mass of 234 kg at a maximum altitude of 8 km at Titan. Forward and aft ballonets are located inside the hull to enable the aerobot to descend to the surface while preserving a fully inflated streamlined shape. Altitude changes are effected primarily through thrust vectoring of the twin main propellers, with pressure modulated buoyancy change via the ballonets available as a slower backup option. A total of 100 W of electrical power is provided to the vehicle by a radioisotope power supply. Up to half of this power is available to the propulsion system to generate a top flight speed in the range of 1-2 m/s. This speed is expected to be greater than the near surface winds at Titan, enabling the aerobot to fly to and hover over targets of interest. A preliminary science payload has been devised for the aerobot to give it the capability for aerial imaging of the surface, atmospheric observations and sampling, and surface sample acquisition and analysis. Targeting, hovering, surface sample acquisition and vehicle health monitoring and automatic safing actions will all require significant on-board autonomy due to the over two hour round trip light time between Titan and Earth. An autonomy architecture and a core set of perception, reasoning and control technologies is under development using a free-flying airship testbed of approximately the same size as the proposed Titan aerobot. Data volume from the Titan science mission is expected to be on the order of 100-300 Mbit per day transmitted either direct to Earth through an 0.8 m high gain antenna or via an orbiter relay using an omni-directional antenna on the aerobot.

mobility↗
NASA NTRS2011

Solar Cycle Prediction

Solar cycle predictions are needed to plan long-term space missions; just like weather predictions are needed to plan your next vacation. Fleets of satellites circle the Earth collecting many types of science data, protecting astronauts, and relaying information. All of these satellites are sensitive at some level to solar cycle effects. Predictions of drag on LEO spacecraft are one of the most important. Launching a satellite with less propellant can mean a higher orbit, but unanticipated solar activity and increased drag can make that a Pyrrhic victory. Energetic events at the Sun can produce crippling radiation storms that endanger all assets in space. Testing solar dynamo theories by quantitative predictions of what will happen in 5-20 years is the next arena for solar cycle predictions. I will describe the current state of solar cycle predictions and anticipate how those predictions could be made more accurate in the future.

Pesnell, William Dean↗
NASA NTRS2012

Solar Cycle Predictions

Solar cycle predictions are needed to plan long-term space missions; just like weather predictions are needed to plan the launch. Fleets of satellites circle the Earth collecting many types of science data, protecting astronauts, and relaying information. All of these satellites are sensitive at some level to solar cycle effects. Predictions of drag on LEO spacecraft are one of the most important. Launching a satellite with less propellant can mean a higher orbit, but unanticipated solar activity and increased drag can make that a Pyrrhic victory as you consume the reduced propellant load more rapidly. Energetic events at the Sun can produce crippling radiation storms that endanger all assets in space. Solar cycle predictions also anticipate the shortwave emissions that cause degradation of solar panels. Testing solar dynamo theories by quantitative predictions of what will happen in 5-20 years is the next arena for solar cycle predictions. A summary and analysis of 75 predictions of the amplitude of the upcoming Solar Cycle 24 is presented. The current state of solar cycle predictions and some anticipations how those predictions could be made more accurate in the future will be discussed.

Pesnell, William Dean↗
NASA NTRS1980

Retrieval techniques: LVLH and inertially stabilized payloads

Procedures and techniques are discussed for retrieving payloads that are inertially or local vertical/local horizontal (LVLH) stabilized. Selection of the retrieval profile to be used depends on several factors: (1) control authority of the payload, (2) payload sensitivity to primary reaction control system (PRCS) plumes, (3) whether the payload is inertially or LVLH stabilized, (4) location of the grapple fixture, and (5) orbiter propellant consumption. The general retrieval profiles recommended are a V-bar approach for payloads that are LVLH or gravity-gradient stabilized, and the V-bar approach with one or two phase flyaround for inertially stabilized payloads. Once the general type of profile has been selected, the detailed retrieval profile and timeline should consider the various guidelines, groundrules, and constraints associated with a particular payload or flight. Reaction control system (RCS) propellant requirements for the recommended profiles range from 200 to 1500 pounds, depending on such factors as braking techniques, flyaround maneuvers (if necessary), and stationkeeping operations. The time required to perform a retrieval (starting from 1000 feet) varies from 20 to 130 minutes, depending on the complexity of the profile. The goals of this project are to develop a profile which ensures mission success; to make the retrieval profiles simple; and to keep the pilot workload to a minimum by making use of the automatic features of the orbiter flight software whenever possible.

Yglesias, J. A.↗
NASA NTRS2015

The Gravity Field of Mercury After the Messenger Low-Altitude Campaign

The final year of the MESSENGER mission was designed to take advantage of the remaining propellant onboard to provide a series of lowaltitude observation campaigns and acquire novel scientific data about the innermost planet. The lower periapsis altitude greatly enhances the sensitivity to the short-wavelength gravity field, but only when the spacecraft is in view of Earth. After more than 3 years in orbit around Mercury, the MESSENGER spacecraft was tracked for the first time below 200-km altitude on 5 May 2014 by the NASA Deep Space Network (DSN). Between August and October, periapsis passages down to 25-km altitude were routinely tracked. These periods considerably improved the quality of the data coverage. Before the end of its mission, MESSENGER will fly at very low altitudes for extended periods of time. Given the orbital geometry, however the periapses will not be visible from Earth and so no new tracking data will be available for altitudes lower than ~75 km. Nevertheless, the continuous tracking of MESSENGER in the northern hemisphere will help improve the uniformity of the spatial coverage at altitudes lower than ~150 km, which will further improve the overall quality of the Mercury gravity field.

Mercury↗
NASA NTRS1979

Advanced cooling techniques for high-pressure hydrocarbon-fueled engines

The regenerative cooling limits (maximum chamber pressure) for 02/hydrocarbon gas generator and staged combustion cycle rocket engines over a thrust range of 89,000 N (20,000lbf) to 2,669,000 N (600,000 lbf) for a reusable life of 250 missions were defined. Maximum chamber pressure limits were first determined for the three propellant combinations (O2/CH4, O2/C3H8, and O2/RP-1 without a carbon layer (unenhanced designs). Chamber pressure cooling enhancement limits were then established for seven thermal barriers. The thermal barriers evaluated for these designs were: carbon layer, ceramic coating, graphite liner, film cooling, transpiration cooling, zoned combustion, and a combination of two of the above. All fluid barriers were assessed a 3 percent performance loss. Sensitivity studies were then conducted to determine the influence of cycle life and RP-1 decomposition temperature on chamber pressure limits. Chamber and nozzle design parameters are presented for the unenahanced and enhanced designs. The maximum regenerative cooled chamber pressure limits were attained with the O2/CH4 propellant combination. The O2/RP-1 designs relied on a carbon layer and liquid gas injection chamber contours, short chamber, to be competitive with the other two propellant combinations. This was attributed to the low decomposition temperature of RP-1.

Cook, R. T.↗
NASA NTRS1973

Study of advanced techniques for determining the long term performance of components

The application of existing and new technology to the problem of determining the long-term performance capability of liquid rocket propulsion feed systems is discussed. The long term performance of metal to metal valve seats in a liquid propellant fuel system is stressed. The approaches taken in conducting the analysis are: (1) advancing the technology of characterizing components through the development of new or more sensitive techniques and (2) improving the understanding of the physical of degradation.

Source record↗
NASA NTRS1991

Satellite/spacecraft propulsion

Propulsion system performance has high leverage for many future missions because of large propellant mass requirements. Relatively small performance improvements can translate into large increases in payload and science return. Contamination control becomes more important as science instruments become more sensitive. This places more emphasis on exhaust plume contamination control. The need for reliable operation and long life places increased importance on health monitoring and control of spacecraft propulsion systems. The need for accurate spacecraft pointing and control increases the need for small impulse-bit thrusters. This presentation is represented by viewgraphs.

Dowdy, Mack W.↗
NASA NTRS2014

Space Launch System Upper Stage Technology Assessment

The Space Launch System (SLS) is envisioned as a heavy-lift vehicle that will provide the foundation for future beyond low-Earth orbit (LEO) exploration missions. Previous studies have been performed to determine the optimal configuration for the SLS and the applicability of commercial off-the-shelf in-space stages for Earth departure. Currently NASA is analyzing the concept of a Dual Use Upper Stage (DUUS) that will provide LEO insertion and Earth departure burns. This paper will explore candidate in-space stages based on the DUUS design for a wide range of beyond LEO missions. Mission payloads will range from small robotic systems up to human systems with deep space habitats and landers. Mission destinations will include cislunar space, Mars, Jupiter, and Saturn. Given these wide-ranging mission objectives, a vehicle-sizing tool has been developed to determine the size of an Earth departure stage based on the mission objectives. The tool calculates masses for all the major subsystems of the vehicle including propellant loads, avionics, power, engines, main propulsion system components, tanks, pressurization system and gases, primary structural elements, and secondary structural elements. The tool uses an iterative sizing algorithm to determine the resulting mass of the stage. Any input into one of the subsystem sizing routines or the mission parameters can be treated as a parametric sweep or as a distribution for use in Monte Carlo analysis. Taking these factors together allows for multi-variable, coupled analysis runs. To increase confidence in the tool, the results have been verified against two point-of-departure designs of the DUUS. The tool has also been verified against Apollo moon mission elements and other manned space systems. This paper will focus on trading key propulsion technologies including chemical, Nuclear Thermal Propulsion (NTP), and Solar Electric Propulsion (SEP). All of the key performance inputs and relationships will be presented and discussed in light of the various missions. For each mission there are several trajectory options and each will be discussed in terms of delta-v required and transit duration. Each propulsion system will be modeled, sized, and judged based on their applicability to the whole range of beyond LEO missions. Criteria for scoring will include the resulting dry mass of the stage, resulting propellant required, time to destination, and an assessment of key enabling technologies. In addition to the larger metrics, this paper will present the results of several coupled sensitivity studies. The ultimate goals of these tools and studies are to provide NASA with the most mass-, technology-, and cost-effective in-space stage for its future exploration missions.

Holladay, Jon↗
NASA NTRS1971

Shuttle cryogenic supply system optimization study

Technical information on different cryogenic supply systems is presented for selecting representative designs. Parametric data and sensitivity studies, and an evaluation of related technology status are included. An integrated mathematical model for hardware program support was developed. The life support system, power generation, and propellant supply are considered. The major study conclusions are the following: Optimum integrated systems tend towards maximizing liquid storage. Vacuum jacketing of tanks is a major effect on integrated systems. Subcritical storage advantages over supercritical storage decrease as the quantity of propellant or reactant decreases. Shuttle duty cycles are not severe. The operational mode has a significant effect on reliability. Components are available for most subsystem applications. Subsystems and components require a minimum amount of technology development.

Source record↗
NASA NTRS1979

An acoustical study of the XV-15 Tilt Rotor Research Aircraft

Acoustic data were obtained during a full-scale test of the XV-15 Tilt Rotor Research Aircraft in the Ames 40- by 80-Foot Wind Tunnel. The XV-15 has two 25-ft-diameter, three-bladed rotors at the tips of a 32-ft span wing. The rotors are used as lifting rotors, as propellers, and in various intermediate stages. Acoustic waveforms, dBA, and spectra as functions of different rotor conditions are presented and discussed. The noise level was found to be sensitive to change in rotor-disk angle of attack in the helicopter configuration. Much higher noise levels and harmonic contents were found in the helicopter mode than in the airplane mode. The measurements are compared with existing theoretical predictions which include wind-tunnel wall reflections.

Lee, A.↗
NASA NTRSDate not supplied

An Interim Set of TNT Curves for LOX/LNG Explosions

Space launch companies are actively developing, or in some cases have already developed, new vehicles that use large quantities of liquid oxygen (LOX) and liquid natural gas (LNG) propellants. This propellant mixture currently lacks sound/verified explosive safety standards for use in explosive siting and flight safety analysis. In the U.S., multiple government and commercial organizations have conducted limited testing, or are preparing to conduct tests, intended to provide a sound technical basis for LOX/LNG explosive standards applicable to space launch vehicle ground and flight operations. The limited testing and analysis conducted so far indicates the potential to produce extremely energetic explosions due to the miscibility of LOX/LNG (methane), a unique feature relative to previously used propellant mixtures such as hydrogen/oxygen or kerosene/oxygen. MOX is a homogenous mixture of methane and oxygen that is possible because of methane’s 100% miscibility in LOX. Preliminary testing by N2L, Inc. described in this paper shows that it is possible to form MOX with a variety of mixing methods, and that MOX is a sensitive high explosive capable of producing overpressures greater than comparable masses of C-4. The limited amount and quality of large-scale LOX/LNG explosive test data, and the potential formation of high-explosive MOX, create significant unknowns in the determination of hazard areas (cleared of the public) for ground operations (such as a static fire test) and flights of launch vehicles with LOX/LNG propellant. This paper will review the methods used by NASA to develop and validate the LOX/LH2 blast model using large-scale explosive test programs such as Project PYRO, the Hydrogen-Oxygen Vertical Impact (HOVI), and the Large-Scale Hydrogen-Oxygen Explosion (LSHOE) tests. This paper will describe how the same process was used to prepare an interim LOX/LNG blast model that applies to various LV failure scenarios and conservatively accounts for potential MOX formation. This paper also summarizes past and future testing and modeling efforts funded by a consortium of NASA, the US Federal Aviation Administration (FAA), and the US Space Force (USSF). These test programs should be concluded within the next 3 years and are intended to provide empirical data for model verification and validation.

Liquid Oxygen↗
NASA NTRSDate not supplied

An Interim Set of TNT Curves for LOX/LNG Explosions

Space launch companies are actively developing, or in some cases have already developed, new vehicles that use large quantities of liquid oxygen (LOX) and liquid natural gas (LNG) propellants. This propellant mixture currently lacks sound/verified LOX/LNG explosive safety standards for use in explosive siting and flight safety analysis. In the U.S., multiple government and commercial organizations have conducted limited testing, or are preparing to conduct tests, intended to provide a sound technical basis for explosive standards applicable to space launch vehicle ground and flight operations. The limited testing and analysis conducted so far indicates the potential to produce extremely energetic explosions due to the miscibility of LOX/LNG (methane), a unique feature relative to previously used propellant mixtures such as hydrogen/oxygen or kerosene/oxygen. MOX is a homogenous mixture of methane and oxygen that is possible because of methane’s 100% miscibility in LOX. Preliminary testing by N2L, Inc. described in this paper shows that it is possible to form MOX with a variety of mixing methods, and that MOX is a sensitive high explosive capable of producing overpressures greater than comparable masses of C-4. The limited amount and quality of large-scale LOX/LNG explosive test data, and the potential formation of high-explosive MOX, create significant unknowns in the determination of hazard areas (cleared of the public) for ground operations (such as a static fire test) and flights of launch vehicles with LOX/LNG propellant. This paper will review the methods used by NASA to develop and validate the LOX/LH2 blast model using large-scale explosive test programs such as Project PYRO, the Hydrogen-Oxygen Vertical Impact (HOVI), and the Large-Scale Hydrogen-Oxygen Explosion (LSHOE) tests. This paper will describe how the same process was used to prepare an interim LOX/LNG blast model that applies to various LV failure scenarios and conservatively accounts for potential MOX formation. This paper also summarizes past and future testing and modeling efforts funded by a consortium of NASA, the US Federal Aviation Administration (FAA), and the US Space Force (USSF). These test programs should be concluded within the next 3 years and are intended to provide empirical data for model verification and validation.

Liquid Oxygen↗
NASA NTRS2019

SPARROW: A Steam Propelled Autonomous Retrieval Robot for Ocean Worlds

This paper presents the results of a NIAC Phase I study into the use of a propulsively hopping robot for the exploration of Europa’s rugged, icy surface. Named the“Steam Propelled Retrieval Robot for Ocean Worlds,” SPARROW is a multi-thruster robot passively gimballed within a protective, spherical shell, which enables it to freely rotate, self-right, and tumble over chaotic terrains. SPARROW is envisioned as a soccerball-sized payload to a primary lander mission. Europa’s abundant surface ice would be harvested as an in situ propellant source. The principal objective of SPARROW would be to increase the science return of a Europa landed asset by enabling access to distal, spatially distributed geologic units. The design of mobility systems for Europa is challenging, due in part to its almost entirely unconstrained surface topography and strength. Images returned by Voyager and Galileo yielded resolutions on the order of hundreds of meters per pixel, with localized regions reaching 6 meters per pixel—still far larger than a typical rover. A key benefit of SPARROW’s hopping, impact-tolerant design, is that it eliminates the need for a piori information on the terrain topography and surface strength; no surface reaction forces are required for motion. In this context, SPARROW is entirely terrain agnostic. In this paper we detail the results of three study objectives: i) to quantify the energy required to collect surface ice, change its phase, and maintain propellant temperature, ii) to identify control and estimation strategies that enable SPARROW to successfully reach, and return from, regions of scientific interest, and iii) to characterize the impact of SPARROW’s range on likely science return. Five water-based propellant architectures are presented alongside their mass, power, and volume requirements. Monte Carlo simulations of SPARROW hopping and tumbling over 1 km of glacial ice are summarized, characterizing SPARROW’s sensitivity to uncertainty in: initial conditions, thrust control, and cage-terrain interaction. Finally, a science traceability matrix is presented, which details the effect of sortie range on three science goals: constraining Europa’s evolutionary morphology, assessing sub-surface ocean habitability, and searching for life and/or biosignatures.

Phillips, Cynthia↗
NASA NTRS1982

Development and utilization of a laser velocimeter system for a large transonic wind tunnel

The need for measurements of the velocity flow field about spinner propeller nacelle configurations at Mach numbers to 0.8 was met by a specially developed laser velocimeter system. This system, which uses an argon ion laser and 4 beam 2 color optics, was required to operate in the hostile environment associated with the operation of a large transonic wind tunnel. To overcome the conditions present in locating the sensitive optics in close proximity to the wind tunnel, an isolation system was developed. The system protects the velocimeter from the high vibrations, elevated temperatures, destructive acoustic pressures and low atmospheric pressures attendant with the operation of the wind tunnel. The system was utilized to map the flow field in front of, behind and in between the rotating blades of an advanced swept blade propeller model at a Mach number of 0.8. The data collected by the system will be used to correlate and verify computer analyses of propeller nacelle flow fields and propeller performance.

Freedman, R. J.↗
NASA NTRS2002

Sensitivity of Lunar Resource Economic Model to Lunar Ice Concentration

Lunar Prospector mission data indicates sufficient concentration of hydrogen (presumed to be in the form of water ice) to form the basis for lunar in-situ mining activities to provide a source of propellant for near-Earth and solar system transport missions. A model being developed by JPL, Colorado School of Mines, and CSP, Inc. generates the necessary conditions under which a commercial enterprise could earn a sufficient rate of return to develop and operate a LEO propellant service for government and commercial customers. A combination of Lunar-derived propellants, L-1 staging, and orbital fuel depots could make commercial LEO/GEO development, inter-planetary missions and the human exploration and development of space more energy, cost, and mass efficient.

Blair, Brad↗