Free convection in a container of cryogenic fluid.
Semiempirical description of stratification in wall heated containers of cryogenic fluids, noting decrease due to bottom heating
PROPELLANT STORAGE↗
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1000 Hours of Testing Completed on 10-kW Hall Thruster
Between the months of April and August 2000, a 10-kW Hall effect thruster, designated T- 220, was subjected to a 1000-hr life test evaluation. Hall effect thrusters are propulsion devices that electrostatically accelerate xenon ions to produce thrust. Hall effect propulsion has been in development for many years, and low-power devices (1.35 kW) have been used in space for satellite orbit maintenance. The T-220, shown in the photo, produces sufficient thrust to enable efficient orbital transfers, saving hundreds of kilograms in propellant over conventional chemical propulsion systems. This test is the longest operation ever achieved on a high-power Hall thruster (greater than 4.5 kW) and is a key milestone leading to the use of this technology for future NASA, commercial, and military missions.
Mason, Lee S.↗
Electron Beam Welding of Pure Tungsten Hex Cans for Nuclear Thermal Propulsion Engines
Nuclear thermal propulsion (NTP) is an in-space propulsion method currently being developed at the NASA Marshall Space Flight Center (MSFC). NTP systems are a high specific impulse (750–1,100 s), high thrust (15,000–250,000 lbf ) method of propulsion which have the potential to allow for faster transit times when optimizing for high ΔV. In the nuclear rocket engine, the heat from the nuclear fission reaction is transferred to a low molecular mass propellant (such as hydrogen). Hot propellant is expanded through a nozzle to generate thrust. Development of ceramic metal (cermet) fuel systems for NTP applications is currently ongoing at MSFC. In cermet fuel systems, ceramic fissile fuel particles such as uranium nitride or uranium dioxide are dispersed within a net-shaped, high-density structural matrix. The composite material is cladded by a protective metal structure to make up an NTP fuel element. Cladding materials must be able to withstand the demanding operating conditions required of the engine as well as retain a hermetic seal to allow for retention of fuel element structural integrity, prevent hydrogen attack or migration of the ceramic fuel, and limit release of fission products during operation. For NTP applications, tungsten is a prime material for both the metal matrix and cladding in cermet fuel systems because of its high melting point, high temperature strength, and compatibility with hot hydrogen. If a weld in tungsten with the capability of holding a hermetic seal is achievable, tungsten becomes a strong candidate for NTP applications. This Technical Memorandum focuses on determining the weldability of pure tungsten using electron beam welding (EBW). Tungsten appears well suited for NTP applications, but it has a high ductile to brittle transition temperature (DBTT) dependent upon chemical composition, structure/stress distribution, and mechanical conditions. Therefore, it is highly subject to brittle fracture. Because of its high susceptibility to brittle fracture, it is very difficult to weld. EBW was chosen for joining pure tungsten because of its low heat input compared to gas tungsten arc welding. Reduced heat input can be directly correlated with an increase in ductility of a tungsten weld. EBW is a high energy density welding process in which a stream of electrons penetrates a weld joint in a deep, narrow spike in contrast to a broad gas tungsten arc weld pool. The investigation initially focused on EBW of tungsten plates of both 0.01 in and 0.03 in thickness to determine if EBW could weld pure tungsten without the presence of visual defects—particularly cracking—in the welds. Variation in the weld procedure and post-weld heat treatment (PWHT) was used to improve the surface appearance of flat EBWs on a pure tungsten sheet. The investigation moved on to weld 0.05-in-thick hexagonal tungsten cans with a weld joint thickness of 0.025 in. The goal for welding the pure tungsten hex cans was to avoid any visual surface defects and generate a weld capable of a hermetic seal. This proved difficult. Cold welds commonly exhibited porosity that leaked air. Hot welds exhibited cracks, typically observed immediately after welding. Later welds were preheated to increase ductility and decrease the likelihood of through-thickness cracking. PWHT was used to arrest microcrack growth both in the flat weld samples and hexagonal weld samples.
Courtright, Z. S.↗
Crowd-Sourced Technology Challenge for Improving Visual Color Detection of Hydrazine and Monomethylhydrazine Vapors in Spacecraft Environments
NASA currently uses a visual colorimetric detection method for potential hydrazine, monomethylhydrazine (MMH), or unsymmetrical dimethylhydrazine (UDMH) contamination in the International Space Station. Astronauts exposed to propellants or their residues during extravehicular activities may transfer contaminants into the airlock. The colorimetric detection method employs the Contamination Detection Kit (CDK), which uses a potassium tetrachloroaurate redox reaction with the propellant hydrazine vapors and a color comparison card to determine airborne concentrations. Seeking ideas for improvement, the NASA Tournament Lab (NTL) crowdsourced a way to tackle the challenge of detecting hydrazine and MMH vapors using colorimetric detection methods. This Rid the Rocket competition drew over 200 participants and 20 submissions from around the world proposing innovative ways to develop a new chemical colorimetric detection method for hydrazine and MMH vapors on spacecraft. Using a phased approach to evaluate contestants, NASA eventually narrowed the field to five finalists from the United States, Romania, Taiwan, and India. Concept papers and hardware submissions were judged on feasibility, creativity, and ability to detect hydrazine and MMH vapors before being sent to the NASA White Sands Test Facility for laboratory evaluation. Finalists employed variations of sampling methods and color-detection chemistry using a variety of sampling pumps and indicator pads or solutions—including those employing potassium or hydrogen tetrachloroaurate, para -dimethylaminobenzaldehyde (PDAB), and modifiers including sodium metasilicate and cetyltrimethylammonium bromide—to enhance gold nanoparticle formation and surface plasmon resonance (SPR) resulting in visual blue to purple color development. This paper presents a summary of the crowdsourced submissions and results of laboratory testing.
Crowd-Sourced↗
In situ propellant production: Alternatives for Mars exploration
Current planning for the Space Exploration Initiative (SEI) recognizes the need for extraterrestrial resources to sustain long-term human presence and to attain some degree of self-sufficiency. As a practical matter, reducing the need to carry large supplies of propellant from Earth will make space exploration more economical. For nearly every round trip planned with conventional propulsion, the actual payload is only a small fraction - perhaps 10-15 percent - of the mass launched from Earth. The objective of this study was to analyze the potential application for SEI missions of propellants made exclusively from lunar or martian resources. Using such propellants could minimize or eliminate the cost of carrying propellant for surface excursion vehicles and return transfers through two high-energy maneuvers: Earth launch and trans-Mars injection. Certain chemical mono- and bipropellants are candidates for this approach; they could be recovered entirely from in situ resources on the Moon and Mars, without requiring a continuing Earth-based resupply of propellant constituents (e.g., fuel to mix with a locally obtained oxidizer) and, perhaps, with minimal need to resupply consumables (e.g., reagents or catalyst for process reactions). A complete assessment of the performance potential of these propellants must include the requirements for installation, operations, maintenance, and resupply of the chemical processing facility.
Stancati, Michael L.↗
Nodal Modeling of Submerged Helium Injection Pressurization of a Cryogenic Propellant Tank
Subcooling of cryogenic propellant by helium injection is one of the most effective methods for suppressing bulk boiling and keeping subcooled propellant conditions for pre-launch, launch, and post-launch pressurization applications. For tank pressurization, submerged helium injection can substantially reduce helium consumption by infusing gaseous propellant into the tank ullage. This paper describes a thermodynamic model of the helium bubbling process in liquid oxygen to estimate the amount of oxygen vapor absorbed by the rising helium bubbles and the amount of subcooling of liquid oxygen due to evaporative heat and mass transfer. The physics of helium dissolution during the pressurization process is also modeled, primarily for liquid hydrogen propellant where the dissolution is more significant. The analyses were performed in a simulation model of tank pressurization built with Generalized Fluid System Simulation Program (GFSSP), a general-purpose flow network code developed at NASA/Marshall Space Flight Center. The numerical predictions of subcooling have been compared with the experimental data of Cho et al. which investigated the propellant subcooling effect as a function of system pressure, helium injection temperature, and flowrate for a non-drained submerged injection system. The numerical predictions of helium consumption have been compared with the test data from a NASA Centaur test vehicle which included both direct and submerged injection with draining of propellants. Hydrogen propellant testing of the Cryogenic Propellant Storage and Transfer Engineering Developmental Unit (CPST EDU) conducted at NASA/Glenn Research Facility was also analyzed. The data used for the model validation were taken in 1-g, but the model was developed to be applicable in both multi-g and micro-g environments.
tank pressurization↗
Tethered versus attached cryogenic propellant storage at Space Station
The space-based orbital transfer vehicle will require a large cryogenic fuel storage facility at the Space Station. An alternative to fuel storage on-board the Space Station is on a tethered orbital refueling facility (TORF) which is separated from the Space Station by a sufficient distance to induce a gravity gradient to settle the propellants. Overall costs and benefits of a particular tethered facility design have been defined relative to a representative zero-gravity facility on the Space Station. Results indicate that the TORF hardware and operations costs are roughly 40 percent higher than the comparable zero-g facility costs. The cost difference is negligible when compared to the launch cost of the fuel over the facility 10-year lifetime.
Fester, Dale A.↗
Nuclear reactor power for an electrically powered orbital transfer vehicle
To help determine the systems requirements for a 300-kWe space nuclear reactor power system, a mission and spacecraft have been examined which utilize electric propulsion and this nuclear reactor power for multiple transfers of cargo between low earth orbit (LEO) and geosynchronous earth orbit (GEO). A propulsion system employing ion thrusters and xenon propellant was selected. Propellant and thrusters are replaced after each sortie to GEO. The mass of the Orbital Transfer Vehicle (OTV), empty and dry, is 11,000 kg; nominal propellant load is 5000 kg. The OTV operates between a circular orbit at 925 km altitude, 28.5 deg inclination, and GEO. Cargo is brought to the OTV by Shuttle and an Orbital Maneuvering Vehicle (OMV); the OTV then takes it to GEO. The OTV can also bring cargo back from GEO, for transfer by OMV to the Shuttle. OTV propellant is resupplied and the ion thrusters are replaced by the OMV before each trip to GEO. At the end of mission life, the OTV's electric propulsion is used to place it in a heliocentric orbit so that the reactor will not return to earth. The nominal cargo capability to GEO is 6000 kg with a transit time of 120 days; 1350 kg can be transferred in 90 days, and 14,300 kg in 240 days. These capabilities can be considerably increased by using separate Shuttle launches to bring up propellant and cargo, or by changing to mercury propellant.
Jaffe, L.↗
Nuclear-electric reusable orbital transfer vehicle
To help determine the systems requirements for a 300-kWe space nuclear reactor power system, a mission and spacecraft have been examined that utilize electric propulsion supported by the nuclear reactor's power for multiple transfers of cargo between low earth orbit (LEO) and geosynchronous earth orbit (GEO). A propulsion system employing ion thrusters and xenon propellant was selected. Propellant and thrusters are replaced after each sortie to GEO. The mass of the orbital transfer vehicle (OTV), empty and dry, is 11,000 kg; nominal propellant load is 5000 kg. The OTV operates between a circular orbit at 925-km altitude, 28.5-deg inclination, and GEO. Cargo is brought to the OTV by Shuttle and an orbital maneuvering vehicle (OMV); the OTV then takes it to GEO. The OTV can also bring cargo back from GEO for transfer by OMV to the Shuttle. OTV propellant is resupplied, and the ion thrusters are replaced, by the OMV before each sortie to GEO. At the end of mission life, the OTV's electric propulsion is used to place it in a heliocentric orbit so that the reactor will not return to earth. The nominal cargo capability to GEO is 6000 kg, with a transit time of 120 days; 1350 kg can be transferred in 90 days, and 14,300 kg in 240 days. These capabilities can be considerably increased by using separate Shuttle launches to bring up propellant and cargo or by changing to mercury propellant.
Jaffe, Leonard D.↗
Transportation mode performance comparison for a sustained manned Mars base
The results of a study performed to characterize the propellant mass requirements of two new types of orbit transfers between earth and Mars are discussed. These new orbit types, called VISIT and Up/Down Escalators, cycle continuously between the two planets, allowing a large crew facility, or CASTLE, to remain in this orbit with smaller crew transfer vehicles, or Taxis, used to shuttle between planetary orbits and this vehicle. Trajectory options and infrastructure elements are discussed along with the assumptions made in this study. The latter include the existence of a mature Martian base and the production and use of extraterrestrial propellants. Performance results from each of the three orbit transfer options presented here are discussed.
Hoffman, S. J.↗
Thermodynamic Vent System Applied as Propellant Delivery System for Air Force
Responding to a request from the Air Force, NASA Lewis Research Center engineers designed a combination pressure control and propellant delivery system based on thermodynamic vent system (TVS) technology. The Air Force is designing a new type of orbit transfer vehicle that uses energy from sunlight to both propel and power the vehicle. Because this vehicle uses propellant at a substantially slower rate than higher-energy rockets, it needed the Lewis-developed TVS technology for long-duration storage of cryogen propellants. Lewis engineers, in conjunction with industry partners, showed how this TVS technology could also be used to deliver propellant to the thruster. The Air Force has now begun the ground test demonstration phase. After successful completion of ground testing, the Air Force plans to use this technology in a space flight as early as 1999.
Source record↗
Comparison of Low-Thrust Control Laws for Application in Planetocentric Space
Recent interest at NASA for the application of solar electric propulsion for the transfer of significant payloads in cislunar space has led to the development of high-fidelity simulations of such missions. With such transfers involving transfer times on the order of months, simulation time can be significant. In the past, the examination of such missions typically began with the use of lower-fidelity trajectory optimization tools such as SEPSPOT to develop and tune guidance laws which delivered optimal or near- optimal trajectories, where optimal is generally defined as minimizing propellant expenditure or time of flight. The transfer of these solutions to a high-fidelity simulation is typically an iterative process whereby the initial solution may nearly, but not precisely, meet mission objectives. Further tuning of the guidance algorithm is typically necessary when accounting for high-fidelity perturbations such as those due to more detailed gravity models, secondary-body effects, solar radiation pressure, etc. While trajectory optimization is a useful method for determining optimal performance metrics, algorithms which deliver nearly optimal performance with minimal tuning are an attractive alternative.
SOLAR ELECTRIC PROPULSION↗
Numerical Investigation of Radiative Heat Transfer in Laser Induced Air Plasmas
Radiative heat transfer is one of the most important phenomena in the laser induced plasmas. This study is intended to develop accurate and efficient methods for predicting laser radiation absorption and plasma radiative heat transfer, and investigate the plasma radiation effects in laser propelled vehicles. To model laser radiation absorption, a ray tracing method along with the Beer's law is adopted. To solve the radiative transfer equation in the air plasmas, the discrete transfer method (DTM) is selected and explained. The air plasma radiative properties are predicted by the LORAN code. To validate the present nonequilibrium radiation model, several benchmark problems are examined and the present results are found to match the available solutions. To investigate the effects of plasma radiation in laser propelled vehicles, the present radiation code is coupled into a plasma aerodynamics code and a selected problem is considered. Comparisons of results at different cases show that plasma radiation plays a role of cooling plasma and it lowers the plasma temperature by about 10%. This change in temperature also results in a reduction of the coupling coefficient by about 10-20%. The present study indicates that plasma radiation modeling is very important for accurate modeling of aerodynamics in a laser propelled vehicle.
Liu, J.↗
Rocket Propulsion (RP) 21 Steering Committee Meeting - NASA Spacecraft Propulsion Update
Lander Tech is three separate but synergistic efforts: Lunar CATALYST (Lunar Cargo Transportation and Landing by Soft Touchdown) Support U.S. industry led robotic lunar lander development via three public-private efforts. Support U.S. industry led robotic lunar lander development via three public-private partnerships. Infuse or transfer landing technologies into these public private partnerships. Advanced Exploration Systems-Automated Propellant Loading (APL) -Integrated Ground Operations. Demonstrate LH2 zero loss storage, loading and transfer operations via testing on a large scale in a relevant launch vehicle servicing environment. (KSC, GRC). Game Changing Technology-20 Kelvin -20 Watt Cryocooler Development of a Reverse Turbo-Brayton Cryocooler operating at 20 Kelvin with 20 Watts of refrigeration lift.
Propellant↗
A Status of the Advanced Space Transportation Program from Planning to Action
A Technology Plan for Enabling Commercial Space Business was presented at the 48th International Astronautical Congress in Turin, Italy. This paper presents a status of the program's accomplishments. Technology demonstrations have progressed in each of the four elements of the program; (1) Low Cost Technology, (2) Advanced Reusable Technology, (3) Space Transfer Technology and (4) Space Transportation Research. The Low Cost Technology program element is primarily focused at reducing development and acquisition costs of aerospace hardware using a "design to cost" philosophy with robust margins, adapting commercial manufacturing processes and commercial off-the-shelf hardware. The attributes of this philosophy for small payload launch are being demonstrated at the component, sub-system, and system level. The X-34 "Fastrac" engine has progressed through major component and subsystem demonstrations. A propulsion system test bed has been implemented for system-level demonstration of component and subsystem technologies; including propellant tankage and feedlines, controls, pressurization, and engine systems. Low cost turbopump designs, commercial valves and a controller are demonstrating the potential for a ten-fold reduction in engine and propulsion system costs. The Advanced Reusable Technology program element is focused on increasing life through high strength-to-weight structures and propulsion components, highly integrated propellant tanks, automated checkout and health management and increased propulsion system performance. The validation of rocket based combined cycle (RBCC) propulsion is pro,-,ressing through component and subsystem testing. RBCC propulsion has the potential to provide performance margin over an all rocket system that could result in lower gross liftoff weight, a lower propellant mass fraction or a higher payload mass fraction. The Space Transfer Technology element of the program is pursuing technology that can improve performance and dramatically reduce the propellant and structural mass of orbit transfer and deep space systems. Flight demonstration of ion propulsion is progressing towards launch. Ion propulsion is the primary propulsion for Deep Space 1; a flyby of comet West-kohoutek-lkemura and asteroid 3352 McAuliffe. Testing of critical solar-thermal propulsion subsystems have been accomplished and planning is continuing for the flight demonstration of an electrodynamic tether orbit transfer system. The forth and final element of the program, Space Transportation Research, has progressed in several areas of propulsion research. This element of the program is focused at long-term (25 years) breakthrough concepts that could bring launch costs to a factor of one hundred below today's cost or dramatically expand planetary travel and enable interstellar travel.
Lyles, Garry↗
Energy transfer and scale dynamics in 2D and 3D laser-driven jets
We demonstrate a methodology for diagnosing the multiscale dynamics and energy transfer in complex HED flows with realistic driving and boundary conditions. The approach separates incompressible, compressible, and baropycnal contributions to energy scale-transfer and quantifies the direction of these transfers in (generalized) wavenumber space. We use this to compare the kinetic energy (KE) transfer across scales in simulations of 2D axisymmetric vs fully 3D laser-driven plasma jets. Using the FLASH code, we model a turbulent jet ablated from an aluminum cone target in the configuration outlined. We show that, in addition to its well known bias for underestimating hydrodynamic instability growth, 2D modeling suffers from significant spurious energization of the bulk flow by a turbulent upscale cascade. In 2D, this arises as vorticity and strain from instabilities near the jet's leading edge transfer KE upscale, sustaining a coherent circulation that helps propel the axisymmetric jet farther ( ≈25% by 3.5 ns) and helps keep it collimated. In 3D, the coherent circulation and upscale KE transfer are absent. Here, the methodology presented here may also help with inter-model comparison and validation, including future modeling efforts to alleviate some of the 2D hydrodynamic artifacts highlighted in this study.
70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
Solid-State Thermodynamic Vent System for Control of Cryogenic Propellants
A Solid-State Thermodynamic Vent System (STVS) is a novel Cryogenic Fluid Management (CFM) technology for spacecraft cryogenic propellant tanks that may reduce boiloff while providing greater control over the propellent condition. By exploiting the vacuum-induced cryocooling potential of cryogen-saturated silica aerogel material, an internal STVS heat exchanger expels some sacrificial cryogenic propellant to the vacuum of space to produce cooling within the storage tank. This cooling is transferred directly to the stored fluid, thereby reducing boiloff and increasing hold times. The primary STVS project objective was to design and build a bespoke STVS heat exchanger that employed aerogel blanket material and perform a ground demonstration of the technology using liquid nitrogen (LN2) as the cryogenic propellant. This demonstration aimed to prove that an STVS can have a positive influence on the tank boiloff, shown through a reduction in mass flow rate out of the system during operation, and exercise control over the tank pressure. Testing proved the effectiveness of the concept by reducing the LN2 boiloff rate and tank pressure by roughly 70% and 77% respectively during a single pump-down cycle, which sacrificed around 1.8 kg of propellant.
Propellant Production↗
In-Space Cryogenic Propellant Depot Potential Commercial and Exploration Applications
The key goals and objectives for an In-Space Cryogenic Propellant Depot are to support a safe, reliable, affordable and effective future human and robotic space exploration initiative. Previous studies have been conducted at the NASA Marshall Space Flight Center to determine technical requirements and feasibility for exploration and commercial potential of an in-space cryogenic propellant depot in low-Earth-orbit (LEO), low-Lunar orbit (LLO) and/or on the lunar surface. Results indicate that in-space cryogenic propellant depots are technically feasible given continued technology development and that there is a substantial growing market that depots could support. Systems studies showed that the most expensive part of transferring payloads to geo-synchronous-orbit (GEO) is the fuel. A cryogenic propellant production and storage depot stationed in LEO could lower the cost of missions to GEO and beyond. Propellant production separates water into hydrogen and oxygen through electrolysis. This process requires large amounts of power which is enabled by Space Solar Power technologies. Recent analysis indicate that in the coming decades there could be a significant demand for water-based propellants from Earth, moon, or asteroid resources if in-space transfer vehicles (upper stages) transitioned to reusable systems using water based propellants. This type of strategic planning move could create a substantial commercial market for space resources development, and ultimately lead toward significant commercial infrastructure development within the Earth-Moon system.
Howell, Joe T.↗