Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “Propellant Tank”

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 1,459 records · Page 81
All sourcesNASA NTRSDOE OSTI
NASA NTRS2018

Helium Pressurization Modeling of the EDU Submerged Diffuser Using Thermal Desktop

NASA is currently working on developing its predictive modeling capabilities to support future missions that require long term storage of cryogenic propellants. Through the Evolvable Cryogenics project, Development and Validation of Analysis Tools (DVAT) task, both CFD and multi-node fluid modeling tools are being validated against to increase the following cryogenic fluid management technologies: Self-Pressurization, Pressure Control, Helium Pressurization, de-stratification, jet mixing, Tank and transfer line chill-down, tank filling and transfer. NASA is using previous and future micro-gravity experiments to validate both types of models. Recent activities include: Validation of the Tank Pressure Control Experiment (TPCE), Hydrogen and Helium pressurization, and Completion of the Zero Boil Off Tank (ZBOT) experiment on ISS. Future activities include validation of the ZBOT experiment with both CFD and multi-node analysis tools, validation of the Cryogenic Demonstration System of the Robotic Refueling Mission (RRM3).

low-gravity fluid modeling↗
NASA NTRS1972

Two phase pumping of cryogenic propellants

Two-phase cryogenic propellant pumping techniques developed during the past 5 years are discussed. During this period the theory of two-phase pumping of hydrogen was developed into an operational technique that can result in significant benefits to space vehicle propulsion systems. Using two-phase pumping techniques, zero LH2 tank net positive suction head (NPSH) was demonstrated on turbopump and rocket engine firings, and limited experimental results show that two-phase pumping of liquid oxygen is feasible. Design techniques were developed for improving the two-phase pumping capability of inducers.

H P Stinson↗
NASA NTRS1983

Modeling of space vehicle propellant mixing

An experimental program was conducted to examine the liquid flow patterns that result from the axial-jet mixing of ethanol in 10-cm-diameter spherical and cylindrical containers under zero-, reduced-, and normal-gravity conditions. Dimensionless parameters were developed that characterized the observed liquid flow patterns and the bulk-liquid mixing phenomena. The correlations developed, were used to analyze a typical liquid hydrogen tank and internal thermodynamic vent system for a shuttle-compatible space tug similar to current orbit transfer vehicle concepts.

Aydelott, J. C.↗
NASA NTRS2013

Atmospheric Processing Module for Mars Propellant Production

The multi-NASA center Mars Atmosphere and Regolith COllector/PrOcessor for Lander Operations (MARCO POLO) project was established to build and demonstrate a methane/oxygen propellant production system in a Mars analog environment. Work at the Kennedy Space Center (KSC) Applied Chemistry Laboratory is focused on the Atmospheric Processing Module (APM). The purpose of the APM is to freeze carbon dioxide from a simulated Martian atmosphere containing the minor components nitrogen, argon, carbon monoxide, and water vapor at Martian pressures (approx.8 torr) by using dual cryocoolers with alternating cycles of freezing and sublimation. The resulting pressurized CO2 is fed to a methanation subsystem where it is catalytically combined with hydrogen in a Sabatier reactor supplied by the Johnson Space Center (JSC) to make methane and water vapor. We first used a simplified once-through setup and later employed a HiCO2 recycling system to improve process efficiency. This presentation and paper will cover (1) the design and selection of major hardware items, such as the cryocoolers, pumps, tanks, chillers, and membrane separators, (2) the determination of the optimal cold head design and flow rates needed to meet the collection requirement of 88 g CO2/hr for 14 hr, (3) the testing of the CO2 freezer subsystem, and (4) the integration and testing of the two subsystems to verify the desired production rate of 31.7 g CH4/hr and 71.3 g H2O/hr along with verification of their purity. The resulting 2.22 kg of CH4/O2 propellant per 14 hr day (including O2 from electrolysis of water recovered from regolith, which also supplies the H2 for methanation) is of the scale needed for a Mars Sample Return mission. In addition, the significance of the project to NASA's new Mars exploration plans will be discussed.

Muscatello, Anthony↗
NASA NTRS2013

Atmospheric Processing Module for Mars Propellant Production

The multi-NASA center Mars Atmosphere and Regolith COllector/PrOcessor for Lander Operations (MARCO POLO) project was established to build and demonstrate a methane/oxygen propellant production system in a Mars analog environment. Work at the Kennedy Space Center (KSC) Applied Chemistry Laboratory is focused on the Atmospheric Processing Module (APM). The purpose of the APM is to freeze carbon dioxide from a simulated Martian atmosphere containing the minor components nitrogen, argon, carbon monoxide, and water vapor at Martian pressures (~8 torr) by using dual cryocoolers with alternating cycles of freezing and sublimation. The resulting pressurized CO2 is fed to a methanation subsystem where it is catalytically combined with hydrogen in a Sabatier reactor supplied by the Johnson Space Center (JSC) to make methane and water vapor. We first used a simplified once-through setup and later employed a HiCO2 recycling system to improve process efficiency. This presentation and paper will cover (1) the design and selection of major hardware items, such as the cryocoolers, pumps, tanks, chillers, and membrane separators, (2) the determination of the optimal cold head design and flow rates needed to meet the collection requirement of 88 g CO2/hr for 14 hr, (3) the testing of the CO2 freezer subsystem, and (4) the integration and testing of the two subsystems to verify the desired production rate of 31.7 g CH4/hr and 71.3 g H20/hr along with verification of their purity. The resulting 2.22 kg of CH4/O2 propellant per 14 hr day (including O2 from electrolysis of water recovered from regolith, which also supplies the H2 for methanation) is of the scale needed for a Mars Sample Return mission. In addition, the significance of the project to NASA's new Mars exploration plans will be discussed.

Muscatello, Anthony↗
DOE OSTI2025

Cold Spray Field Deployment Evaluation for Double-Shell Tank 241-AN-105

This report highlights the documentation that supported Cold Spray field deployment in double-shell tank (DST) 241-AN-105 (AN-105) in July of 2025. This work was a collaborative effort between VRC Metal Systems (VRC), Robotic Technologies of Tennessee (RTT), the Hanford Tank Waste Operation and Closure (H2C) Chief Technology Office, H2C Tank Farm Projects Engineering, H2C Tank and Pipeline Integrity (TAPI) and H2C Ultrasonic Testing (UT) Operations & Support. Since 2019, the Chief Technology Office has been developing Cold Spray technology for refurbishing Hanford’s DSTs (RPP-RPT-65015). This technology is particularly relevant for addressing localized corrosion within the annulus of the DSTs by propelling metal powder particles at supersonic speeds to targeted areas. This process results in high-quality coatings characterized by high bond strength and low porosity result.

12 MANAGEMENT OF RADIOACTIVE AND NON-RADIOACTIVE W↗
NASA NTRS1976

Exospheric density measurements from the drag-free satellite Triad

The Triad satellite is maintained within 1 mm of its proof mass (a small metal ball) in a purely gravitational orbit, using a drag-free control system, named Discos, as the disturbance compensation system. The Discos proof mass is surrounded by an outer shell which holds fuel tanks and cold gas jets. The shell shields the proof mass from such nongravitational forces as radiation pressure, atmospheric drag, and micrometeorite impact. Whenever these forces displace the outer shell relative to the proof mass, an error signal is generated by a capacitive bridge sensor, and the satellite is propelled by gas jets to remain centered on the proof mass when the error reaches 1 mm. Local atmospheric densities near 800 km were measured, using ball position data, and the observed accelerations were corrected for the effects of solar radiation pressure. The measured densities reveal a greater dependence on latitude than that indicated by balloon satellites, and also a dependence on longitude. Only a small dependence on Kp, however, was observed in the region where the measurements were made.

Moe, K.↗
NASA NTRS1980

A long-life bipropellant system demonstration: Viking Orbiter Propulsion System - 4 years in space and operating

On April 23, 1980, Viking Orbiter One (VO-1), operating in the blowdown mode, completed a ten second Mars orbit trim maneuver to position the spacecraft for its final science sequence in May-June 1980. This brought the number of propulsive maneuvers for VO-1 to 23. Total accumulated operating time for the rocket engine was 2896 seconds, representing a total impulse of 3.93 x 10 to the 6th N-sec. The estimated propellant remaining was sufficient to operate the rocket engine for an additional 30 seconds. VO-1 has completed more than 1700 days in space, 1400 days in orbit around Mars and more than two years of attitude control system operation with helium gas transferred from the propulsion system pressurant tank. The mass of helium remaining is expected to be sufficient for attitude control through June 1980. This paper presents the flight history of the Viking 75 Orbiter Propulsion Systems and summarizes the design and test philosophy which have contributed to their success.

Schmit, D. D.↗
NASA NTRS2014

A Crewed Mission to Apophis Using a Hybrid Bimodal Nuclear Thermal Electric Propulsion (BNTEP) System

A BNTEP system is a dual propellant, hybrid propulsion concept that utilizes Bimodal Nuclear Thermal Rocket (BNTR) propulsion during high thrust operations, providing 10's of kilo-Newtons of thrust per engine at a high specific impulse (Isp) of 900 s, and an Electric Propulsion (EP) system during low thrust operations at even higher Isp of around 3000 s. Electrical power for the EP system is provided by the BNTR engines in combination with a Brayton Power Conversion (BPC) closed loop system, which can provide electrical power on the order of 100's of kWe. High thrust BNTR operation uses liquid hydrogen (LH2) as reactor coolant propellant expelled out a nozzle, while low thrust EP uses high pressure xenon expelled by an electric grid. By utilizing an optimized combination of low and high thrust propulsion, significant mass savings over a conventional NTR vehicle can be realized. Low thrust mission events, such as midcourse corrections (MCC), tank settling burns, some reaction control system (RCS) burns, and even a small portion at the end of the departure burn can be performed with EP. Crewed and robotic deep space missions to a near Earth asteroid (NEA) are best suited for this hybrid propulsion approach. For these mission scenarios, the Earth return V is typically small enough that EP alone is sufficient. A crewed mission to the NEA Apophis in the year 2028 with an expendable BNTEP transfer vehicle is presented. Assembly operations, launch element masses, and other key characteristics of the vehicle are described. A comparison with a conventional NTR vehicle performing the same mission is also provided. Finally, reusability of the BNTEP transfer vehicle is explored.

Electric Propulsion↗
NASA NTRSDate not supplied

Modeling Self-Pressurization and Spray Bar Pressure Control of a Cryogenic Storage Tank in Normal Gravity

This paper presents Computational Fluid Dynamics (CFD) models for simulating self-pressurization and spray bar cooling processes in a large-scale liquid hydrogen storage tank under normal gravity. The self-pressurization model utilizes the kinetics-based Schrage equation with the Volume-Of-Fluid (VOF) model to account for the evaporative and condensing interfacial mass fluxes. It uses a laminar approach to model natural convection and compares the pressures and temperatures predicted by the CFD model with experimental data from the Multipurpose Hydrogen Test Bed (MHTB) self-pressurization experiment. Additionally, the predicted interfacial mass flux is presented. A CFD model for simulating pressure control using a spray bar is also presented. An Eulerian-Lagrangian approach models the interactions between the discrete droplets and the continuous ullage phase. The spray model is coupled with the VOF model by tracking particles in the ullage, removing particles from the ullage when they reach the interface, and then adding their contributions to the liquid. The T-sat model is used to calculate the droplet-ullage heat and mass transfer. In this model, droplets warm up to the saturation temperature corresponding to the ullage vapor pressure thereafter evaporating while remaining at the saturation temperature. The tank pressure, vapor, and liquid temperature evolutions are validated against data from the MHTB spray bar mixing experiment.

Self-Pressurization↗
NASA NTRS1990

Superfluid and normal fluid helium II in a rotating tank under low and microgravity environments

The physical behavior of superfluid and normal fluid helium II is discussed. It is shown that the centrifugal force due to a rotating container can never separate the superfluid from the normal fluid helium in a large rotating cylindrical spacecraft. A classical approach is appropriate for treating superfluid and normal fluid cryogenic helium as a mixture without separation in a large rotating cylinder in a spacecraft propellant systems.

Hung, R. J.↗
NASA NTRS1989

Space Shuttle RCS Oxidizer Leak Repair for STS-26

Following propellant loading of the Space Shuttle's reaction control system (RCS) for mission STS 26, an oxidizer leak was detected in the left orbital maneuvering system (OMS) pod, where the RCS is located. Subsequent investigation determined that the leak was isolated at a mechanical Dynatube fitting near the RCS nitrogen tetroxide tank. An intense effort was initiated to design, fabricate, and qualify a sealing device to stop the oxidizer leak externally so that the Space Shuttle launch could proceed. It was discovered that sealing devices called clamshells were widely used throughout the petrochemical and power generation industries to stop leaks developed in large diameter pipes which carry steam or other hazardous fluids. These clamshells are available in different diameters and strengths and are placed around the pipe at the location of the leak. A sealing compound is then injected under high pressure into the clamshell to stop the leak. This technology was scaled down and applied to the problem of stopping the leak on the Orbiter, which was on a half-inch diameter line in a nearly inaccessible location. Many obstacles had to be overcome such as determining that the sealing material would be compatible with the nitrogen tetroxide and ensuring that the clamshell would actually fit around the Dynatube fitting without interfering with other lines which were in close proximity. The effort at the NASA Johnson Space Center included materials compatibility testing of several sealants, design of a clamshell to fit in the confined compartment, and manufacture and qualification of the flight hardware. A clamshell was successfully placed around the Dynatube fitting on the Orbiter and the oxidizer leak was terminated. Then it was decided to apply this technology further and design clamshells for other mechanical fittings onboard the Orbiter and develop sealing compounds which will be compatible with fuels such as monomethyl hydrazine (MMH). The potential exists for using this type of sealing device in numerous other applications throughout the aerospace industry.

Delventhal, R. A.↗
NASA NTRS1974

Space Shuttle Reaction Control Subsystem propellant acquisition

A surface tension propellant acquisition/expulsion configuration was selected for the Space Shuttle Reaction Control Subsystem tankage to supply gas-free propellant during the low- and high-g operational environment (.00001 to 3.0 g) plus the numerous omnidirectional intermittent (pulse) demands of each mission. Candidate concepts to meet these stringent requirements were identified; this was followed by analysis and design sensitivity evaluations. A trade study resulted in the selection of a compartmented tank with individual flow channels as the preferred concept. Details of the analysis and design are presented.

Fester, D. A.↗
NASA NTRS2002

Oxygen Mass Flow Rate Generated for Monitoring Hydrogen Peroxide Stability

Recent interest in propellants with non-toxic reaction products has led to a resurgence of interest in hydrogen peroxide for various propellant applications. Because peroxide is sensitive to contaminants, material interactions, stability and storage issues, monitoring decomposition rates is important. Stennis Space Center (SSC) uses thermocouples to monitor bulk fluid temperature (heat evolution) to determine reaction rates. Unfortunately, large temperature rises are required to offset the heat lost into the surrounding fluid. Also, tank penetration to accomodate a thermocouple can entail modification of a tank or line and act as a source of contamination. The paper evaluates a method for monitoring oxygen evolution as a means to determine peroxide stability. Oxygen generation is not only directly related to peroxide decomposition, but occurs immediately. Measuring peroxide temperature to monitor peroxide stability has significant limitations. The bulk decomposition of 1% / week in a large volume tank can produce in excess of 30 cc / min. This oxygen flow rate corresponds to an equivalent temperature rise of approximately 14 millidegrees C, which is difficult to measure reliably. Thus, if heat transfer were included, there would be no temperature rise. Temperature changes from the surrounding environment and heat lost to the peroxide will also mask potential problems. The use of oxygen flow measurements provides an ultra sensitive technique for monitoring reaction events and will provide an earlier indication of an abnormal decomposition when compared to measuring temperature rise.

Ross, H. Richard↗
NASA NTRS2015

Reduced Boil-Off System Sizing

NASA is currently developing cryogenic propellant storage and transfer systems for future space exploration and scientific discovery missions by addressing the need to raise the technology readiness level of cryogenic fluid management technologies. Cryogenic propellants are baselined in many propulsion systems due to their inherently high specific impulse; however, their low boiling points can cause substantial boil-off losses over time. Recent efforts such as the Reduced Boil-off Testing and the Active Thermal Control Scaling Study provide important information on the benefit of an active cooling system applied to LH2 propellant storage. Findings show that zero-boil off technologies can reduce overall mass in LH2 storage systems when low Earth orbit loiter periods extend beyond two months. A significant part of this mass reduction is realized by integrating two stages of cooling: a 20 K stage to intercept heat at the tank surface, and a 90 K stage to reduce the heat entering the less efficient 20 K stage. A missing element in previous studies, which is addressed in this paper, is the development of a direct method for sizing the 90 K cooling stage. Such a method requires calculation of the heat entering both the 90 K and 20 K stages as compared to the overall system masses, and is reliant upon the temperature distribution, performance, and unique design characteristics of the system in question. By utilizing the known conductance of a system without active thermal control, the heat being intercepted by a 90 K stage can be calculated to find the resultant lift and mass of each active thermal control stage. Integral to this is the thermal conductance of the cooling straps and the broad area cooling shield, key parts of the 90 K stage. Additionally, a trade study is performed to show the ability of the 90 K cooling stage to reduce the lift on the 20 K cryocooler stage, which is considerably less developed and efficient than 90 K cryocoolers.

Cryogenic fluid management↗
NASA NTRSDate not supplied

Modeling Self-Pressurization and Spray Bar Pressure Control of A Cryogenic Storage Tank in Normal Gravity

This paper presents computational fluid dynamics (CFD) models for simulating self-pressurization and spray-bar pressure control processes in a large-scale liquid hydrogen storage tank under normal gravity conditions. For self-pressurization, the model employs the kinetics-based Schrage equation alongside the volume-of-fluid (VOF) method to account for interfacial mass transfer. The CFD predictions of pressure and temperature are compared with experimental data from the Multipurpose Hydrogen Test Bed (MHTB) experiment, and the predicted interfacial mass transfer rates are also presented. A CFD model simulating pressure control using a spray bar has also been developed. An Eulerian-Lagrangian approach models the interactions between discrete droplets and the continuous ullage (vapor) phase. The spray model is coupled with the VOF method by tracking droplets in the ullage and removing them when they reach the liquid interface. The T-sat model calculates droplet-ullage heat and mass transfer, wherein droplets warm up to the saturation temperature corresponding to the ullage vapor pressure before evaporating while remaining at the saturation temperature. The evolution of tank pressure, vapor temperature, and liquid temperature predicted by the CFD model is validated against data from the MHTB spray-bar mixing experiment. Overall, the CFD models agree with experimental data, demonstrating their capability to simulate self-pressurization and pressure control processes in large-scale cryogenic storage tanks. These models can be valuable tools for designing and optimizing cryogenic fluid management systems in future applications.

Self-Pressurization↗
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 NTRS1979

Analysis and test for space shuttle propellant dynamics (1/10th scale model test results). Volume 2: 1/10th scale model test data

Thirty sets of test data selected from the 89 low-g aircraft tests flown by NASA KC-135 zero-g aircraft are listed in tables with their accompanying test conditions. The data for each test consists of the time history plots of digitalized data (in engineering units) and the time history plots of the load cell data transformed to the tank axis system. The transformed load cell data was developed for future analytical comparisons; therefore, these data were transformed and plotted from the time at which the aircraft Z axis acceleration passed through l-g. There are 14 time history plots per test condition. The contents of each plot is shown in a table.

Berry, R. L.↗