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Results for “Propellant Management”

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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Low-thrust chemical orbit to orbit propulsion system propellant management study

Propellant requirements, tankage configurations, preferred propellant management techniques, propulsion systems weights, and technology deficiencies for low thrust expendable propulsion systems are examined. A computer program was utilized which provided a complete propellant inventory (including boil-off for cryogenic cases), pressurant and propellant tank dimensions for a given ullage, pressurant requirements, insulation requirements, and miscellaneous masses. The output also includes the masses of all tanks; the mass of the insulation, engines and other components; total wet system and burnout mass; system mass fraction; total impulse and burn time.

Dergance, R. H.↗
NASA NTRS2016

A Detailed Historical Review of Propellant Management Devices for Low Gravity Propellant Acquisition

This paper presents a comprehensive background and historical review of Propellant Management Devices (PMDs) used throughout spaceflight history. The purpose of a PMD is to separate liquid and gas phases within a propellant tank and to transfer vapor-free propellant from a storage tank to a transfer line en route to either an engine or receiver depot tank, in any gravitational or thermal environment. The design concept, basic flow physics, and principle of operation are presented for each type of PMD. The three primary capillary driven PMD types of vanes, sponges, and screen channel liquid acquisition devices are compared and contrasted. For each PMD type, a detailed review of previous applications using storable propellants is given, which include space experiments as well as space missions and vehicles. Examples of previous cryogenic propellant management are also presented.

Liquid acquisition device↗
NASA NTRS1991

Cryogenic propellant management system requirements for Space Station Freedom

Specific propellant management system requirements have been identified for each facility category of SSF. Distributed systems have been analyzed to indentify momentum management, guidance, and traffic management requirements associated with the guidance, navigation, and control system; space-to-space communications and enhanced tracking requirements associated with the communications and tracking system; and propellant management system utility requirements associated with the electrical power system. Flight element analyses determined attach structure, utility distribution, and structural integrity requirements for the pre-integrated truss and high mass manipulation and translation requirements for the mobile base system.

Saucillo, R. J.↗
NASA NTRS1979

Performance of a capillary propellant management device with hydrazine

The propellant management device that was successfully used in the Viking Orbiter spacecraft was selected for the main propulsion system of the Teleoperator Retrieval System (TRS). Due to differences in the missions and different propellants, the operation of this sheet metal vane device required reverification for the TRS application. An analytical investigation was performed considering the adverse acceleration environment and the high contract angle of the hydrazine propellant. Drop tower tests demonstrated that the device would provide propellant acquisition while the TRS was docked with Skylab, but its operation would have to be supplemented through propellant settling when free-flying.

Tegart, J. R.↗
NASA NTRS1972

Selection of a surface tension propellant management system for the Viking 75 Orbiter.

Discussion of the propellant management system requirements derived for the Viking 75 mission, and review of a series of surface tension propellant management system design concepts. The chosen concept is identified and its mission operation described. The ullage bubble and bulk liquid positioning characteristics are presented, along with propellant dynamic considerations entailed by thrust initiation/termination. Pressurization design considerations, required to assure minimum disturbance to the bulk propellant, are introduced as well as those of the tank ullage vent. Design provisions to assure liquid communication between tank ends are discussed. Results of a preliminary design study are presented, including mechanical testing requirements to assure structural integrity, propellant compatibility, and proper installation.

Dowdy, M. W.↗
NASA NTRS2010

NASA's Evolutionary Xenon Thruster (NEXT) Prototype Model 1R (PM1R) Ion Thruster and Propellant Management System Wear Test Results

The results of the NEXT wear test are presented. This test was conducted with a 36-cm ion engine (designated PM1R) and an engineering model propellant management system. The thruster operated with beam extraction for a total of 1680 hr and processed 30.5 kg of xenon during the wear test, which included performance testing and some operation with an engineering model power processing unit. A total of 1312 hr was accumulated at full power, 277 hr at low power, and the remainder was at intermediate throttle levels. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, was steady with no indications of performance degradation. The propellant management system performed without incident during the wear test. The ion engine and propellant management system were also inspected following the test with no indication of anomalous hardware degradation from operation.

VanNoord, Jonathan L.↗
NASA NTRS2011

NEXT Propellant Management System Integration With Multiple Ion Thrusters

As a critical part of the NEXT test validation process, a multiple-string integration test was performed on the NEXT propellant management system and ion thrusters. The objectives of this test were to verify that the PMS is capable of providing stable flow control to multiple thrusters operating over the NEXT system throttling range and to demonstrate to potential users that the NEXT PMS is ready for transition to flight. A test plan was developed for the sub-system integration test for verification of PMS and thruster system performance and functionality requirements. Propellant management system calibrations were checked during the single and multi-thruster testing. The low pressure assembly total flow rates to the thruster(s) were within 1.4 percent of the calibrated support equipment flow rates. The inlet pressures to the main, cathode, and neutralizer ports of Thruster PM1R were measured as the PMS operated in 1-thruster, 2-thruster, and 3-thruster configurations. It was found that the inlet pressures to Thruster PM1R for 2-thruster and 3-thruster operation as well as single thruster operation with the PMS compare very favorably indicating that flow rates to Thruster PM1R were similar in all cases. Characterizations of discharge losses, accelerator grid current, and neutralizer performance were performed as more operating thrusters were added to the PMS. There were no variations in these parameters as thrusters were throttled and single and multiple thruster operations were conducted. The propellant management system power consumption was at a fixed voltage to the DCIU and a fixed thermal throttle temperature of 75 C. The total power consumed by the PMS was 10.0, 17.9, and 25.2 W, respectively, for single, 2-thruster, and 3-thruster operation with the PMS. These sub-system integration tests of the PMS, the DCIU Simulator, and multiple thrusters addressed, in part, the NEXT PMS and propulsion system performance and functionality requirements.

Sovey, James S.↗
NASA NTRS1977

Design and development of a passive propellant management system

This paper describes the passive propellant management system program being conducted for the NASA-Marshall Space Flight Center. The objectives of this program are to design and develop a hydrazine propellant acquisition system which will acquire propellant in low- or zero-g environments and also retain this propellant under high axially directed accelerations that may be experienced during launch and orbit-to-orbit transfer. The system design requirements were established to satisfy requirements for a large number of potential NASA and military applications, such as orbit-to-orbit shuttles and satellite vehicles. The resulting concept is a multicompartmented tank with independent surface-tension acquisition channels in each compartment. The tank is designed to provide greater than 98% expulsion efficiency when subjected to the simultaneous requirements of acceleration, vibration, and outflow usage. The system design has the unique capability to demonstrate low-g performance in a one-g test environment, and the test program has been structured around this capability.

Regnier, W. W.↗
NASA NTRS2012

NASA Flexible Screen Propellant Management Device (PMD) Demonstration With Cryogenic Liquid

While evaluating various options for liquid methane and liquid oxygen propellant management for lunar missions, Innovative Engineering Solutions (IES) conceived the flexible screen device as a potential simple alternative to conventional propellant management devices (PMD). An apparatus was designed and fabricated to test flexible screen devices in liquid nitrogen. After resolution of a number of issues (discussed in detail in the paper), a fine mesh screen (325 by 2300 wires per inch) spring return assembly was successfully tested. No significant degradation in the screen bubble point was observed either due to the screen stretching process or due to cyclic fatigue during testing. An estimated 30 to 50 deflection cycles, and approximately 3 to 5 thermal cycles, were performed on the final screen specimen, prior to and between formally recorded testing. These cycles included some "abusive" pressure cycling, where gas or liquid was driven through the screen at rates that produced differential pressures across the screen of several times the bubble point pressure. No obvious performance degradation or other changes were observed over the duration of testing. In summary, it is felt by the author that these simple tests validated the feasibility of the flexible screen PMD concept for use with cryogenic propellants.

Wollen, Mark↗
NASA NTRS2010

Cryogenic Propellant Management Device: Conceptual Design Study

Concepts of Propellant Management Devices (PMDs) were designed for lunar descent stage reaction control system (RCS) and lunar ascent stage (main and RCS propulsion) missions using liquid oxygen (LO2) and liquid methane (LCH4). Study ground rules set a maximum of 19 days from launch to lunar touchdown, and an additional 210 days on the lunar surface before liftoff. Two PMDs were conceptually designed for each of the descent stage RCS propellant tanks, and two designs for each of the ascent stage main propellant tanks. One of the two PMD types is a traditional partial four-screen channel device. The other type is a novel, expanding volume device which uses a stretched, flexing screen. It was found that several unique design features simplified the PMD designs. These features are (1) high propellant tank operating pressures, (2) aluminum tanks for propellant storage, and (3) stringent insulation requirements. Consequently, it was possible to treat LO2 and LCH4 as if they were equivalent to Earth-storable propellants because they would remain substantially subcooled during the lunar mission. In fact, prelaunch procedures are simplified with cryogens, because any trapped vapor will condense once the propellant tanks are pressurized in space.

Wollen, Mark↗
NASA NTRS1981

Propellant management for low thrust chemical propulsion systems

Low-thrust chemical propulsion systems (LTPS) will be required for orbital transfer of large space systems (LSS). The work reported in this paper was conducted to determine the propellant requirements, preferred propellant management technique, and propulsion system sizes for the LTPS. Propellants were liquid oxygen (LO2) combined with liquid hydrogen (LH2), liquid methane or kerosene. Thrust levels of 100, 500, and 1000 lbf were combined with 1, 4, and 8 perigee burns for transfer from low earth orbit to geosynchronous earth orbit. This matrix of systems was evaluated with a multilayer insulation (MLI) or a spray-on-foam insulation. Vehicle sizing results indicate that a toroidal tank configuration is needed for the LO2/LH2 system. Multiple perigee burns and MLI allow far superior LSS payload capability. Propellant settling, combined with a single screen device, was found to be the lightest and least complex propellant management technique.

Hamlyn, K. M.↗
NASA NTRS1990

Cryogenic propellant management architectures to support the Space Exploration Initiative

The initial results of a current study to develop fuel system architectures to support the lunar requirements of the Space Exploration Initiative (SEI) are reported. The study includes the development and assessment of propellant management facility concepts, supporting infrastructure, operations analysis, and identification of impact on current programs, including Space Station Freedom, Earth-to-Orbit vehicles, and the Space Transfer Vehicle. The cryogenic propellant management architectures are evaluated using criteria that have been defined to provide for minimum subjective assessment and effective data reliability.

Cady, E. C.↗
NASA NTRSDate not supplied

Sensitivity Analysis of Heat Rejection and Propellant Management Technologies for Nuclear Thermal Propulsion Architectures

Cryogenic fluid management (CFM) technologies are very important for enabling a wider range of missions to utilize space nuclear propulsion (SNP) concepts such as nuclear thermal propulsion (NTP). Technologies for thermal and cryogenic propellant management allow for vehicles to take full advantage of the higher efficiency NTP systems for longer duration human interplanetary and deep space robotic missions. Currently, the CFM Portfolio Project at NASA’s Marshall Space Flight Center (MSFC) is developing thermal and propellant management technologies needed for SNP. When developing these technologies, it is important to understand the sensitivities of key performance parameters (KPPs) at the system and overall mission level due to the ways the technologies interact with each other, other subsystems, and influence the overall vehicle. The Advanced Concepts Office (ACO) at NASA-MSFC was tasked with building an integrated system model of a human Mars NTP mission to evaluate the impacts and sensitivities of CFM technologies on the overall vehicle and mission. This paper will cover the buildup of the model and highlight major sensitivities and breakpoints encountered, as well as future work in improving the existing models and sensitivities being evaluated.

Robert J. Hetterich↗
NASA NTRSDate not supplied

Sensitivity Analysis of Heat Rejection and Propellant Management Technologies for Nuclear Thermal Propulsion Architectures

Cryogenic fluid management (CFM) technologies are very important for enabling a wider range of missions to utilize space nuclear propulsion (SNP) concepts such as nuclear thermal propulsion (NTP). Technologies for thermal and cryogenic propellant management allow for vehicles to take full advantage of the higher efficiency NTP systems for longer duration human interplanetary and deep space robotic missions. Currently, the CFM Portfolio Project at NASA’s Marshall Space Flight Center (MSFC) is developing thermal and propellant management technologies needed for SNP. When developing these technologies, it is important to understand the sensitivities of key performance parameters (KPPs) at the system and overall mission level due to the ways the technologies interact with each other, other subsystems, and influence the overall vehicle. The Advanced Concepts Office (ACO) at NASA-MSFC was tasked with building an integrated system model of a human Mars NTP mission to evaluate the impacts and sensitivities of CFM technologies on the overall vehicle and mission. This paper will cover the buildup of the model and highlight major sensitivities and breakpoints encountered, as well as future work in improving the existing models and sensitivities being evaluated.

Robert J Hetterich↗
NASA NTRS1978

Design and performance verification of a passive propellant management system

This paper describes the design and verification testing of a reusable passive propellant management system. The system was designed to acquire propellant in low- or zero-g environments and also retain this propellant under high axially directed accelerations that may be experienced during launch and orbit-to-orbit transfer. The system design requirements were established to satisfy generally the requirements for a large number of potential NASA and military applications, such as orbit-to-orbit shuttles and satellite vehicles. The resulting concept was a multicompartmented tank with independent surface tension acquisition channels in each compartment. The tank was designed to provide a minimum expulsion efficiency of 98 percent when subjected to the simultaneous conditions of acceleration, vibration, and outflow. The system design has the unique capability to demonstrate low-g performance in a 1-g test environment, and the test program summarized was structured around this capability.

Hess, D. A.↗
NASA NTRS2023

Benchmark Performance Metrics of a Vane Propellant Management Device for a 0.15 m3 Liquid Hydrogen Tank

The use of cryogenic propellants has and will continue to play an integral role in manned-space exploration due the high specific impulses offered and its ubiquity through in-situ resource utilization. But guaranteeing vapor-free transfer of such low-surface tension liquids is difficult for traditional capillary-action propellant management devices (PMDs). Screen-channel liquid acquisition devices and compliant origami bladders are potential solutions, but to quantify the benefits these technologies offer, this paper present a cast study analyzing the performance metrics of an orthodox vane PMD as a benchmark for comparison. A 0.15 m3 liquid hydrogen tank at 20.3 K and 103 kPa was selected for study. Then assuming no body forces and no heat transfer (for simplicity), a steady-state 1-D differential equation was numerically solved in tandem with four possible wetted area configurations to yield an expulsion efficiency for a theoretically maximum inputted expulsion flow rate. Maximization of the flow rate was constrained by the onset of choked flow. Additional inputs, including vane height and vane number, were parametrically varied between 0.1 – 10 cm and 4 – 28 vanes, respectively.

Liquid Acquisition Device↗