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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Liquid Nitrogen Testing of an Integrated Reaction Control System

Integrated Reaction Control Systems (IRCS) are considered an enabling technology for future human exploration of the solar system. An IRCS uses the same fuel and oxidizer as the main propulsion system, allowing for increased performance and simplified cryogenic fluid management operations. The IRCS team at Marshall Spaceflight Center is currently evaluating an IRCS system that uses a combination of electric pumps and pressure-regulation devices to provide a constant flow of conditioned propellant to a thruster bank during operation. In contrast to accumulator and gasifier designs, the recirculation loop design is less massive and can provide large, sustained flowrates to the thruster inlet. In the summer 2021 test series presented herein, the feasibility of this concept was evaluated using liquid nitrogen as a simulant for liquid oxygen. The goals of this test series were to demonstrate the ability of the pump and pressure regulator to maintain adequate thruster inlet conditions during transient operation, and to demonstrate the ability for the system to operate in an “idle mode” that maintains system chill between activations. During test operations, the back-pressure regulator responded more slowly than the opening/closing of valves in the thruster simulant, but it was still able to attenuate the pressure fluctuation at the pump discharge and ensure a stable flowrate through the pump. Similarly, the pump was able to be run far below its nominal operating point and successfully flow sufficient fluid to maintain chill conditions.

integrated RCS↗
NASA NTRS2011

MLAS and LAS: Project Comparison of Lessons Learned

The development of new and safer manned space vehicles is a top priority at NASA. Recently two different approaches of how to accomplish this mission of keeping astronauts safe was successfully demonstrated. With work already underway on an Apollo-like launch abort system for the Orion Crew Exploration Vehicle (CEV), an alternative design concept named the Max Launch Abort System, or MLAS, was developed as a parallel effort. The Orion system, managed by the Constellation office, is based on the design of a single solid launch abort motor in a tower positioned above the capsule. The MLAS design takes a different approach placing the solid launch abort motor underneath the capsule. This effort was led by the NASA Engineering and Safety Center (NESC). Both escape systems were designed with the Ares I Rocket as the launch vehicle and had the same primary requirement to safely propel a crew module away from any emergency event either on the launch pad or during accent. Beyond these two parameters, there was little else in common between the two projects, except that they both concluded in successful launches that will further promote the development of crew launch abort systems. This presentation will compare both projects from the standpoint of technical requirements; program management and flight test objectives and highlight the synergistic lessons learned by two engineers who worked on each program.

Dittemore, Gary↗
NASA NTRS2018

NASA's Cryogenic Fluid Management Technology Development Roadmaps

The maturation of Cryogenic Fluid Management (CFM) Technologies is essential for achieving NASA's future long duration missions. Propulsion systems utilizing cryogens are necessary to achieve NASA's exploration missions to the moon, Mars, and beyond. Current State Of the Art (SOA) CFM technologies enable cryogenic propellants to be stored for several hours prior to their use. However, some envisioned mission architectures require that cryogens to be stored for two years or longer. The fundamental roles of CFM technologies are long term storage of cryogens, propellant tank pressure control and propellant delivery. In the presence of heat, the cryogens will "boil-off" over time resulting in excessive pressure buildup, off-nominal propellant conditions for engine consumption, and propellant loss. To achieve long term storage and tank pressure control, the CFM elements will intercept and/or remove any heat from the propulsion system. All functions are required to be performed both with and without the presence of a gravitational field. Which CFM technologies are required is a function of the cryogens used, mission architecture, vehicle design and propellant tank size. To enable NASA's crewed missions beyond Low Earth Orbit, a total of twenty-seven CFM technologies have been identified to support various In-Space Stages and Lander/Ascent Vehicles. A set of CFM Technology Development Roadmaps have been created identifying the current Technology Readiness Level (TRL) of each element, current technology "gaps", and existing technology development efforts. The roadmaps include a methodical approach and schedule to achieve a flight demonstration, hence maturing CFM technologies to TRL 6/7 for infusion into the NASA's exploration elements. Additionally, a survey of the aerospace industry was completed to understand their views on the various technologies and how they would be infused. This does not cover all possible technologies, but rather those that are of interest to NASA specifically.

Cryogenic fluid management↗
NASA NTRSDate not supplied

Ground-Based Cryogenic Leak Test of Fittings for Cryogenic Fluid Management

Mechanically connected fluid joints are virtually unavoidable in complex cryogenic system designs. In the case of spacecraft, the performance of these joints is critical to mission success. This is especially true for long-duration space missions where even very low leak rates can eventually lead to significant propellant losses or failure of vital cryogenic cooling systems. NASA has recently undertaken an effort to quantify the leak rate of Vacuum Coupling Radiation (VCR) fittings over the temperature range from ambient to 20 K, both before and after exposure to a launch vibration profile. A test apparatus employing a cryocooler and a calibrated helium mass spectrometer was developed, validated, and used to obtain quantifiable leak rates at a fitting test pressure of 31 bar (450 psig). Three different fitting sizes were tested, 25.4 mm (¼ inch), 12.7 mm (½ inch), and 6.35 mm (1 inch) and two gasket materials, stainless steel, and silver-plated nickel. Each fitting configuration (size/seal material) was subjected to two consecutive cryogenic thermal cycles/measurement tests, followed by a launch vibration test profile at ambient temperature, and then two additional cryogenic thermal cycles/measurement tests. The design and development of the test apparatus, and test data are presented and discussed in detail.

G J Tamasy↗
NASA NTRS1981

The Inertial Upper Stage - A space transportation system element nearing first flight

The Inertial Upper Stage (IUS) developed by the USAF and NASA is a highly reliable, cost-effective solid propellant upper stage, with inherent flexibility and adaptability for integration with the Space Shuttle. The propulsion system is simple, utilizing safe, solid rocket motors with extremely light-weight nonmetallic cases and nozzles. The IUS can deliver 2268 kg from the Shuttle to geosynchronous altitude; it consists of a 9700 kg propellant weight first stage, an interstage structure, a 2720 kg propellant weight second stage, and an equipment support section. The avionics system includes the electronic and electrical hardware used to perform all signal conditioning, data processing, and software formatting associated with navigation, guidance, control, data management, and redundancy management. The generic thermal design of the IUS is suited to a wide range of thermal environments; the software design provides for selectable thermal maneuvers (rotisserie, reciprocating, toasting, space facing, sun facing) to satisfy different payload thermal requirements. A 1982 launch with the Titan 34D and a 1983 launch with the Shuttle Orbiter are planned.

Rohrbaugh, D. J.↗
NASA NTRS2005

Capillary Flow Experiments Began on the International Space Station

The Capillary Flow Experiments (CFEs) are a suite of fluid physics flight experiments designed to investigate capillary flows and phenomena in low gravity. Data obtained from the CFEs will be crucial to NASA s Space Exploration Initiative, particularly pertaining to fluids management systems such as fuels and cryogen storage systems, thermal control systems (e.g., water recycling), and materials processing in the liquid state. NASA s current plans for exploration missions assume the use of larger liquid propellant masses than have ever flown on interplanetary missions. Under low-gravity conditions, capillary forces can be exploited to control fluid orientation so that such large mission-critical systems perform predictably. The first of the CFE experiments has been conducted on the International Space Station, and the data are being analyzed. The experiment suite is described briefly.

J. Mark Hickman↗
NASA NTRS1997

Systems Simulation of NASA Shooting Star Experiment Using Matlab/Simulink

The Shooting Star Experiment (SSE) is an experiment that incorporates advance propulsion technology. This project is being managed by the Marshall Space Flight Center, Huntsville, Alabama. Whenever spacecraft are launched from Low Earth Orbit (LEO), (typically 150 nautical miles) they are powered by a upper propulsive stage utilizing either a solid or liquid propellant engine. A typically mission for a spacecraft utilizing an upper stage would be a transfer from LEO to a Geostationary Orbit (GEO) or an interplanetary mission. These upper stages are heavy and bulky because they must carry propellants to provide sufficient energy to perform the mission. The SSE utilizes the energy of the Sun by focusing this energy by means of a Frensel lens into an engine where hydrogen (or nitrogen) gas is injected. The focusing of the solar energy heats the engine to very high temperatures. When the gas is injected into the hot engine, the gas is expelled at very high velocities. This process is extremely efficient. Because of the efficiency of the SSE type engine, more payload can be carried for a typical mission since the propulsive element is much smaller.

Reagan, Shawn↗
NASA NTRS2006

Low-Cost Propellant Launch From a Tethered Balloon

A document presents a concept for relatively inexpensive delivery of propellant to a large fuel depot in low orbit around the Earth, for use in rockets destined for higher orbits, the Moon, and for remote planets. The propellant is expected to be at least 85 percent of the mass needed in low Earth orbit to support the NASA Exploration Vision. The concept calls for the use of many small ( 10 ton) spin-stabilized, multistage, solid-fuel rockets to each deliver 250 kg of propellant. Each rocket would be winched up to a balloon tethered above most of the atmospheric mass (optimal altitude 26 2 km). There, the rocket would be aimed slightly above the horizon, spun, dropped, and fired at a time chosen so that the rocket would arrive in orbit near the depot. Small thrusters on the payload (powered, for example, by boil-off gases from cryogenic propellants that make up the payload) would precess the spinning rocket, using data from a low-cost inertial sensor to correct for small aerodynamic and solid rocket nozzle misalignment torques on the spinning rocket; would manage the angle of attack and the final orbit insertion burn; and would be fired on command from the depot in response to observations of the trajectory of the payload so as to make small corrections to bring the payload into a rendezvous orbit and despin it for capture by the depot. The system is low-cost because the small rockets can be mass-produced using the same techniques as those to produce automobiles and low-cost munitions, and one or more can be launched from a U.S. territory on the equator (Baker or Jarvis Islands in the mid-Pacific) to the fuel depot on each orbit (every 90 minutes, e.g., any multiple of 6,000 per year).

Wilcox, Brian↗
NASA NTRS2016

Liquefaction and Storage of In-Situ Oxygen on the Surface of Mars

The In-Situ production of propellants for Martian and Lunar missions has been heavily discussed since the mid 1990's. One portion of the production of the propellants is the liquefaction, storage, and delivery of the propellants to the stage tanks. Two key technology development efforts are required: large refrigeration systems (cryocoolers) to perform the liquefaction and high performance insulation within a soft vacuum environment. Several different concepts of operation may be employed to liquefy the propellants based on how and where these two technologies are implemented. The concepts that were investigated include: using an accumulator tank to store the propellant until it is needed, liquefying in the flow stream going into the tank, and liquefying in the flight propellant tank itself. The different concept of operations were studied to assess the mass and power impacts of each concept. Additionally, the trade between insulation performance and cryocooler mass was performed to give performance targets for soft vacuum insulation development. It was found that liquefying within the flight propellant tank itself adds the least mass and power requirements to the mission.

Cryogenic Fluid Management↗
NASA NTRS1990

Evolutionary Space Station fluids management

The demand for and management issues associated with fluids usage at the evolutionary Space Station are examined. A variety of fluids such as N2, He, methane and rare gases for research and development activities as well as massive quantities of cryogenic propellants for geosynchronous orbit and planetary exploration missions will need to be accommodated at the Space Station. A data base of fluid types, quantities, and projected usage schedules suggested, and potential accommodation concepts defined. Impacts to the Space Station operational configuration and necessary 'hooks and scars' to be included in the baseline design are addressed. Also presented is an operational scenario of the delivery of a cryogenic propellant tank set to the Space Station, attachment to the Space Station, and propellant transfer to and launch of a space transfer vehicle (STV). Results indicate that the cryogenic propellant requirements of STV and lunar missions may be met by tank sets attached to the Space Station, but those of Mars missions would require off-station methods.

Stevenson, Steve↗
NASA NTRS1994

Facility design consideration for continuous mix production of class 1.3 propellant

In November of 1989, NASA awarded the Advanced Solid Rocket Motor (ASRM) contract to Lockheed Missiles and Space Company (LMSC) for production of advanced solid rocket motors using the continuous mix process. Aerojet ASRM division (AAD) was selected as the facility operator and RUST International Corporation provided the engineering, procurement, and construction management services. The continuous mix process mandates that the mix and cast facilities be 'close-coupled' along with the premix facilities, creating unique and challenging requirements for the facility designer. The classical approach to handling energetic materials-division into manageable quantities, segregation, and isolation-was not available due to these process requirements and quantities involved. This paper provides a description of the physical facilities, the continuous mix process, and discusses the monitoring and detection techniques used to mitigate hazards and prevent an incident.

Williamson, K. L.↗
NASA NTRS1996

A Theoretical Study of Remobilizing Surfactant Retarded Fluid Particle Interfaces

Microgravity processes must rely on mechanisms other than bouyancy to move bubbles or droplets from one region to another in a continuous liquid phase. One suggested method is thermocapillary migration in which a temperature gradient is applied to the continuous phase. When a fluid particle contacts this gradient, one pole of the particle becomes warmer than the opposing pole. The interfacial tension between the drop or bubble phase and the continuous phase usually decreases with temperature. Thus the cooler pole is of higher interfacial tension than the warmer pole, and the interface is tugged in the direction of the cooler end. This thermocapillary or thermally induced Marangoni surface stress causes a fluid streaming in the continuous phase from which develops a viscous shear traction and pressure gradient which together propel the particle in the direction of the warmer fluid. In this paper, we provide a theoretical basis for remobilizing surfactant retarded fluid particle interfaces in an effort to make viable the use of thermocapillary migrations for the management of bubbles and drops in microgravity,

Wang, Yanping↗
NASA NTRS1989

Galileo Science Writers' Briefing

This NASA Kennedy video production presents Part 1 of a press conference held at JPL on August 8, 1989. The briefing in its entirety covers the Galileo Project's mission design from launch to completion in 1997 and is moderated by JPL Public Information Mgr. Robert Macmillan. Part 1 of the 3 part video series includes presentations by Richard J. Spehalski (Galileo Project Manager) and Clayne M. Yeates (Acting Science Mission Design Manager). Mr. Spehalski's presentation includes actual footage of spacecraft preparations at Kennedy Space Center and slides of mission timelines. Dr. Yeates discusses the Galileo mission in chronological order and includes slides of the interplanetary trajectory, encounter geometry, propellant margins vs. launch date, and planned earth images.

Source record↗
NASA NTRS2001

NASA's Space Launch Initiative Targets Toxic Propellants

When manned and unmanned space flight first began, the clear and overriding design consideration was performance. Consequently, propellant combinations of all kinds were considered, tested, and, when they lifted the payload a kilometer higher, or an extra kilogram to the same altitude, they became part of our operational inventory. Cost was not considered. And with virtually all of the early work being performed by the military, safety was hardly a consideration. After all, fighting wars has always been dangerous. Those days are past now. With space flight, and the products of space flight, a regular part of our lives today, safety and cost are being reexamined. NASA's focus turns naturally to its Shuttle Space Transportation System. Designed, built, and flown for the first time in the 1970s, this system remains today America's workhorse for manned space flight. Without its tremendous lift capability and mission flexibility, the International Space Station would not exist. And the Hubble telescope would be a monument to shortsighted management, rather than the clear penetrating eye on the stars it is today. But the Shuttle system fully represents the design philosophy of its period: it is too costly to operate, and not safe enough for regular long term access to space. And one of the key reasons is the utilization of toxic propellants. This paper will present an overview of the utilization of toxic propellants on the current Shuttle system.

Hurlbert, Eric↗
NASA NTRS1986

Evaluation of on-orbit cryogenic propellant depot options for the orbital transfer vehicle

An orbital cryogenic propellant storage facility will be required for a space-based Orbital Transfer Vehicle. The facility tanks will have features to permit fluid acquisition and transfer in low gravity and to limit cryogen boiloff caused by environmental heating. Boiloff management features will include thick multilayer insulation, vapor-cooled shields, low conductance structural supports and penetrations, and possibly refrigeration systems.

Schuster, J. R.↗
NASA NTRS2011

Modeling a Transient Pressurization with Active Cooling Sizing Tool

As interest in the area of in-space zero boil-off cryogenic propellant storage develops, the need to visualize and quantify cryogen behavior during ventless tank self-pressurization and subsequent cool-down with active thermal control has become apparent. During the course of a mission, such as the launch ascent phase, there are periods that power to the active cooling system will be unavailable. In addition, because it is not feasible to install vacuum jackets on large propellant tanks, as is typically done for in-space cryogenic applications for science payloads, instances like the launch ascent heating phase are important to study. Numerous efforts have been made to characterize cryogenic tank pressurization during ventless cryogen storage without active cooling, but few tools exist to model this behavior in a user-friendly environment for general use, and none exist that quantify the marginal active cooling system size needed for power down periods to manage tank pressure response once active cooling is resumed. This paper describes the Transient pressurization with Active Cooling Tool (TACT), which is based on a ventless three-lump homogeneous thermodynamic self-pressurization model1 coupled with an active cooling system estimator. TACT has been designed to estimate the pressurization of a heated but unvented cryogenic tank, assuming an unavailable power period followed by a given cryocooler heat removal rate. By receiving input data on the tank material and geometry, propellant initial conditions, and passive and transient heating rates, a pressurization and recovery profile can be found, which establishes the time needed to return to a designated pressure. This provides the ability to understand the effect that launch ascent and unpowered mission segments have on the size of an active cooling system. A sample of the trends found show that an active cooling system sized for twice the steady state heating rate would results in a reasonable time for tank pressure recovery with ZBO of a liquid oxygen propellant tank.

Guzik, Monica C.↗
NASA NTRS1997

Reusable Launch Vehicle Technology Program

Industry/NASA reusable launch vehicle (RLV) technology program efforts are underway to design, test, and develop technologies and concepts for viable commercial launch systems that also satisfy national needs at acceptable recurring costs. Significant progress has been made in understanding the technical challenges of fully reusable launch systems and the accompanying management and operational approaches for achieving a low cost program. This paper reviews the current status of the RLV technology program including the DC-XA, X-33 and X-34 flight systems and associated technology programs. It addresses the specific technologies being tested that address the technical and operability challenges of reusable launch systems including reusable cryogenic propellant tanks, composite structures, thermal protection systems, improved propulsion and subsystem operability enhancements. The recently concluded DC-XA test program demonstrated some of these technologies in ground and flight test. Contracts were awarded recently for both the X-33 and X-34 flight demonstrator systems. The Orbital Sciences Corporation X-34 flight test vehicle will demonstrate an air-launched reusable vehicle capable of flight to speeds of Mach 8. The Lockheed-Martin X-33 flight test vehicle will expand the test envelope for critical technologies to flight speeds of Mach 15. A propulsion program to test the X-33 linear aerospike rocket engine using a NASA SR-71 high speed aircraft as a test bed is also discussed. The paper also describes the management and operational approaches that address the challenge of new cost effective, reusable launch vehicle systems.

Freeman, Delma C., Jr.↗
NASA NTRS2006

AstroBiology Explorer Mission Concepts (ABE/ASPIRE)

The AstroBiology Explorer (ABE) and the Astrobiology Space InfraRed Explorer (ASPIRE) Mission Concepts are two missions designed to address the questions (1) Where do we come from? and (2) Are we alone? as outlined in NASA s Origins Program using infrared spectroscopy to explore the identity, abundance, and distribution of molecules of astrobiological importance throughout the Universe. The ABE mission s observational program is focused on six tasks to: (1) Investigate the evolution of ice and organics in dense clouds and star formation regions, and the young stellar/planetary systems that form in them; (2) Measure the evolution of complex organic molecules in stellar outflows; (3) Study the organic composition of a wide variety of solar system objects including asteroids, comets, and the planets and their satellites; (4) Identify organic compounds in the diffuse interstellar medium and determine their distribution , abundance, and change with environment; (5) Detect and identify organic compounds in other galaxies and determine their dependence on galactic type; and (6) Measure deuterium enrichments in interstellar organics and use them as tracers of chemical processes. The ASPIRE mission s observational program expands upon ABE's core mission and adds tasks that (7) Address the role of silicates in interstellar organic chemistry; and (8) Use different resolution spectra to assess the relative roles and abundances of gas- and solid-state materials. ABE (ASPIRE) achieves these goals using a highly sensitive, cryogenically-cooled telescope in an Earth drift-away heliocentric orbit, armed with a suite of infrared spectrometers that cover the 2.5-20(40) micron spectral region at moderate spectral resolution (R>2000). ASPIRE's spectrometer complement also includes a high-resolution (R>25,000) module over the 4-8 micron spectral region. Both missions target lists are chosen to observe a statistically significant sample of a large number of objects of varied types in support of the tasks outline above. The ABE and ASPIRE mission lifetimes are designed to be 14 months and 3 years, respectively, both with significant cryogen and propellant lifetime margins to support an extended observing campaign. The ABE/ASPIRE mission concepts and their supporting Science Teams are led by Principal Investigator Dr. Scott Sandford of NASA s Ames Research Center, with industry partner Ball Aerospace Technologies Ltd., and managed by Jet Propulsion Laboratory. The ABE/ASPIRE Science Operations will be carried out at NASA s Ames Research Center, and the ABE/ASPIRE database will be archived at Caltech/IPAC.

Sandford, Scott↗