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NASA NTRSDate not supplied

Numerical Simulation of No Vent Chill and Fill of a Large Liquid Hydrogen Tank

On orbit transfer of cryogenic liquids is necessary for many future space missions. The techniques required to transfer cryogens in micro gravity are quite different from ground system applications. In normal gravity, cryogenic tanks are usually filled from the bottom at nearly atmospheric pressure. The vapor generated due to heat transfer from warm tank walls, is allowed to vent from the top of the tank while the tank is being filled. If the normal gravity technique is used on orbit, the uncertainty of liquid and vapor distributions in unsettled condition may result in dumping of large amounts of liquid propellant overboard. The Vented Chill and No Vent Fill (VCNVF) process is a methodology [1] used to reduce propellant loss by holding the vent valve open during initial wall chilldown, preventing over-pressurization of the tank and stalling of the incoming flow, but closing the tank vent valve after chilling the tank wall sufficiently. An accurate simulation of no vent tank chill and fill is necessary to estimate the amount of propellant necessary to chill the tank, determine the maximum sustainable flowrate to fill the tank without stalling, and establish a timeline for the filling process. Numerical modeling of filling a cryogenic tank is complex because the flow is two-phase and the process involves both boiling and condensation with heat transfer between solid and fluid. In recent years, progress has been made to develop a numerical modeling technique [2,3,4] that has been verified by comparing with the test data from a small flight tank using liquid nitrogen [5]. The purpose of the present paper is to apply this technique in a larger tank with liquid hydrogen and compare with the test data of Chato [1]. The K-site Test Facility [1], shown in Figure 1, consists of a Test Tank, spray system, instrumentation, and the vacuum chamber. The test tank selected was ellipsoidal with an 87 inch major diameter and a 1.2 to 1 major to minor axis ratio. The two ends are joined by a short 1.5 inch cylindrical section. The tank is made of 2219 aluminum chemically milled to a nominal thickness of 0.087 inches. The tank weighs 329.25 pounds, and the tank volume is 175 ft 3 . The GFSSP model of the Test Tank is shown in Figure 2. Node 13 is a boundary node that represents the supply tank which is supplying liquid hydrogen at a given pressure and temperature. The total measured flowrate is evenly distributed through Branches 131 through 139. Nodes 1 through 9 represent the test tank and they are connected with metal solid nodes 14 through 22 through fluid-to-solid conductors that allow convective heat transfer between the solid and fluid nodes. The predicted tank pressure, filling rate, wall temperature and fluid temperature will be compared with test data.

Alok Majumdar↗
NASA NTRS1993

Mixing and transient interface condensation of a liquid hydrogen tank

Experiments were conducted to investigate the effect of axial jet-induced mixing on the pressure reduction of a thermally stratified liquid hydrogen tank. The tank was nearly cylindrical, having a volume of about 0.144 cu m with 0.559 m in diameter and 0.711 m long. A mixer/pump unit, which had a jet nozzle outlet of 0.0221 m in diameter was located 0.178 m from the tank bottom and was installed inside the tank to generate the axial jet mixing and tank fluid circulation. The liquid fill and jet flow rate ranged from 42 to 85 percent (by volume) and 0.409 to 2.43 cu m/hr, respectively. Mixing tests began with the tank pressure ranging from 187.5 to 238.5 kPa at which the thermal stratification results in 4.9 to 6.2 K liquid sub cooling. The mixing time and transient vapor condensation rate at the liquid-vapor interface are determined. Two mixing time correlations, based on the thermal equilibrium and pressure equilibrium, are developed. Both mixing time correlations are expressed as functions of system and buoyancy parameters and compared well with other experimental data. The steady state condensation rate correlation of Sonin et al. based on steam-water data is modified and expressed as a function of jet subcooling. The limited liquid hydrogen data of the present study shows that the modified steady state condensation rate correlation may be used to predict the transient condensation rate in a mixing process if the instantaneous values of jet sub cooling and turbulence intensity at the interface are employed.

Lin, C. S.↗
NASA NTRS1993

Mixing and transient interface condensation of a liquid hydrogen tank

Experiments were conducted to investigate the effect of axial jet-induced mixing on the pressure reduction of a thermally stratified liquid hydrogen tank. The tank was nearly cylindrical, having a volume of about 0.144 cu m with 0.559 m in diameter and 0.711 m length. A mixer/pump unit, which had a jet nozzle outlet of 0.0221 m in diameter was located 0.178 m from the tank bottom and was installed inside the tank to generate the axial jet mixing and tank fluid circulation. Mixing tests began with the tank pressures at which the thermal stratification results in 4.9-6.2 K liquid subcooling. The mixing time and transient vapor condensation rate at the liquid-vapor interface are determined. Two mixing time correlations, based on the thermal equilibrium and pressure equilibrium, are developed and expressed as functions of system and buoyancy parameters. The limited liquid hydrogen data of the present study shows that the modified steady state condensation rate correlation may be used to predict the transient condensation rate in a mixing process if the instantaneous values of jet sub cooling and turbulence intensity at the interface are employed.

Lin, C. S.↗
NASA NTRSDate not supplied

Liquid Hydrogen Tank Chill and No-Vent Fill Prediction using Computational Fluid Dynamics

Cryogenic tank chill and fill is an important cryogenic fluid management (CFM) technology that supports and enables many of NASA’s long-duration space missions. For no-vent fill, the receiver tank pressure remains below the supply tank pressure during the entire duration of the fill so that the tank does not require venting. This is especially advantageous for tank fill operations in low gravity where the position of the liquid is not always known and venting the tank may cause loss of propellant by venting liquid. In lieu of expensive tests conducted on-orbit, accurate computational models capable of predicting receiver tank pressure during cryogenic propellant tank fill may be used to reduce system and propellant mass as well as mission risk. However, these numerical models must be validated or anchored to test data. This study presents a computational fluid dynamics (CFD) model with conjugate heat transfer that is used to predict a liquid hydrogen tank chill and no-vent fill ground test conducted at Lewis Research Center (now Glenn Research Center) in 1991. The specific test case chosen for model validation implemented an upward-facing jet near the bottom of a cylindrical 34 liter tank. CFD predictions are compared to experimental measurements of tank pressure, fill level, fluid temperatures, and wall temperatures. The CFD results show reasonable agreement to the test data but overpredict the pressure collapse near the end of the fill despite the liquid jet penetrating the liquid-vapor interface. Several sensitivity studies are considered due to notable uncertainties in the experiment.

Computational Fluid Dynamics↗
NASA NTRSDate not supplied

Liquid Hydrogen Tank Chill and No-Vent Fill Prediction Using Computational Fluid Dynamics

Cryogenic tank chill and fill is an important cryogenic fluid management (CFM) technology that supports and enables many of NASA’s long-duration space missions. For no-vent fill, the receiver tank pressure remains below the supply tank pressure during the entire duration of the fill so that the tank does not require venting. This is especially advantageous for tank fill operations in low gravity where the position of the liquid is not always known and venting the tank may cause loss of propellant by venting liquid. In lieu of expensive tests conducted on-orbit, accurate computational models capable of predicting receiver tank pressure during cryogenic propellant tank fill may be used to reduce system and propellant mass as well as mission risk. However, these numerical models must be validated or anchored to test data. This study presents a computational fluid dynamics (CFD) model with conjugate heat transfer that is used to predict a liquid hydrogen tank chill and no-vent fill ground test conducted at Lewis Research Center (now Glenn Research Center) in 1991. The specific test case chosen for model validation implemented an upward-facing jet near the bottom of a cylindrical 34 liter tank. CFD predictions are compared to experimental measurements of tank pressure, fill level, fluid temperatures, and wall temperatures. The CFD results show reasonable agreement to the test data but overpredict the pressure collapse near the end of the fill despite the liquid jet penetrating the liquid-vapor interface. Several sensitivity studies are considered due to notable uncertainties in the experiment.

Computational Fluid Dynamics↗
NASA NTRSDate not supplied

Liquid Hydrogen Tank Chill and No-Vent Fill Prediction using Computational Fluid Dynamics

Cryogenic tank chill and fill is an important cryogenic fluid management technology that supports and enables many of NASA’s long-duration space missions. For no-vent fill, the receiver tank pressure remains below the supply tank pressure during the entire duration of the fill so that the tank does not require venting. This is especially advantageous for tank fill operations in low gravity where the position of the liquid is not always known and venting the tank may cause loss of propellant by venting liquid. In lieu of expensive tests conducted on-orbit, accurate computational models capable of predicting receiver tank pressure during cryogenic propellant tank fill may be used to reduce system and propellant mass as well as mission risk. However, these numerical models must be validated or anchored to test data. This study presents computational fluid dynamics (CFD) models with conjugate heat transfer that are used to predict a liquid hydrogen tank chill and no-vent fill ground test conducted at Lewis Research Center (now Glenn Research Center) in 1991. The specific test case chosen for model validation implemented an upward-facing jet near the bottom of a cylindrical 34-liter tank. CFD predictions are compared to experimental measurements of tank pressure, fill level, and wall temperature. The CFD results show reasonable agreement to the test data but overpredict the pressure collapse near the end of the fill despite the liquid jet penetrating the liquid-vapor interface.

Computational Fluid Dynamics↗
NASA NTRS2024

Helium Conservation by Diffusion Limited Purging of Liquid Hydrogen Tanks

Recently, the new 1.2-million-gallon liquid hydrogen (LH 2 ) storage tank at the Kennedy Space Center Launch Complex 39 was purged, replacing nitrogen with helium, using a piston purge process. This purging approach had been used on Space Shuttle External Tanks (ET) in the past, but not on one of the large LH 2 tanks. The result was a huge success, potentially saving more than 1 million cubic feet of helium and significant labor. To better understand this result, a diffusion-based purge model was developed. The model predictions accurately match the helium concentration data, providing understanding of the helium savings and suggesting how to better optimize the purging of this large tank for future operations. The model was then applied to the piston purging of the Space Shuttle ETs, showing that these purges were not optimal, but that the amount of wasted helium was not excessive and that additional helium might have been saved by small changes to the process. The Space Shuttle Program (SSP) has ended, but these results are applicable to the purging of future large cylindrical LH 2 tanks. Finally, data have been obtained on inerting the Space Launch System (SLS) Core Stage (CS) LH 2 tank, replacing hydrogen with helium after an aborted launch. This is a more complicated case, but the model predicts that substantial helium might be saved by modifying the current purging process.

Helium Conservation↗
NASA NTRS1972

European technology on composite materials

The following advanced composite materials were evaluated for future space shuttle applications: (1) Carbon and boron fiber composites using epoxy and polyimide as resins for stringer and skin panels: (2) high pressure steel tanks where the cylindrical part is reinforced by carbon fiber; and (3) an orbiter cargo door in boron fiber and polyimide resin.

Fray, J.↗
NASA NTRS2000

High Energy Flywheel Containment Evaluation

A flywheel testing facility is being constructed at the NASA Glenn Research Center. This facility is to be used for life cycle testing of various flywheel rotors. The lifecycle testing consists of spinning a rotor from a low rpm (approx. 20,000 ) to a high rpm (approx. 60,000) and then back to the low rpm. This spin cycle will model that which the rotor will see during use. To simulate the lifetime of the rotor, the spin cycle will be performed tens of thousands of times. A typical life cycle spin test is expected to last six months. During this time the rotor will be spun through a cycle every five minutes. The test will run continuously for the six month period barring a flywheel failure. Since it is not reasonable to have the surrounding area evacuated of personnel for the duration of the testing, the flywheel facility has to be designed to withstand a flywheel rotor failure and insure that there is no danger to any personnel in the adjacent buildings or surrounding areas. In order to determine if the facility can safely contain a flywheel rotor failure an analysis of the facility in conjunction with possible flywheel failure modes was performed. This analysis is intended as a worst case evaluation of the burst liner and vacuum tank's ability to contain a failure. The test chamber consists of a cylindrical stainless steel vacuum tank, two outer steel containment rings, and a stainless steel burst liner. The stainless steel used is annealed 302, which has an ultimate strength of 620 MPa (90,000 psi). A diagram of the vacuum tank configuration is shown. The vacuum tank and air turbine will be located below ground in a pit. The tank is secured in the pit with 0.3 m (12 in.) of cement along the base and the remaining portion of the tank is surrounded by gravel up to the access ports. A 590 kg (1300 lb.) bulkhead is placed on top of the pit during operation and the complete facility is housed within a concrete structure which has 7.5 cm (3 in.) thick walls. A cutaway of the facility is shown.

Colozza, Anthony J.↗
NASA NTRS1987

Flutter study of an advanced composite wing with external stores

A flutter test using a scaled model of an advanced composite wing for a Navy attack aircraft has been conducted in the NASA Langley Research Center Transonic Dynamics Tunnel. The model was a wall-mounted half-span wing with a semi-span of 6.63 ft. The wing had an aspect ratio of 5.31, taper ratio of 0.312, and quarter-chord sweep of 25 degrees. The model was supported in a manner that simulated the load path in the carry-through structure of the aircraft and the symmetric boundary condition at the fuselage centerline. The model was capable of carrying external stores from three pylon locations on the wing. Flutter tests were conducted for the wing with and without external stores. No flutter was encountered for the clean wing at test conditions which simulated the scaled airplane operating envelope. Flutter boundaries were obtained for several external store configurations. The flutter boundaries for the fuel tanks were nearly Mach number independent (occurring at constant dynamic pressure). To study the aerodynamic effect of the fuel tank stores, pencil stores (slender cylindrical rods) which had the same mass and pitch and yaw inertia as the fuel tanks were tested on the model. These pencil store configurations exhibited a transonic dip in the flutter dynamic pressure, indicating that the aerodynamic effect of the actual fuel tanks on flutter was significant. Several flutter analyses methods were used in an attempt to predict the flutter phenomenon exhibited during the wind-tunnel test. The analysis gave satisfactory predictions of flutter for the pencil store configurations, but unsatisfactory correlation for the actual fuel tank configurations.

Cole, Stanley R.↗
NASA NTRS1998

Liquid Motion in a Rotating Tank Experiment (LME)

The Liquid Motion Experiment (LME), which flew on STS 84 in May 1997, was an investigation of liquid motions in spinning, nutating tanks. LME was designed to quantify the effects of such liquid motions on the stability of spinning spacecraft, which are known to be adversely affected by the energy dissipated by the liquid motions. The LME hardware was essentially a spin table which could be forced to nutate at specified frequencies at a constant cone angle, independently of the spin rate. Cylindrical and spherical test tanks, partially filled with liquids of different viscosities, were located at the periphery of the spin table to simulate a spacecraft with off-axis propellant tanks; one set of tanks contained generic propellant management devices (PMDs). The primary quantitative data from the flight tests were the liquid-induced torques exerted on the tanks about radial and tangential axes through the center of the tank. Visual recordings of the liquid oscillations also provided qualitative information. The flight program incorporated two types of tests: sine sweep tests, in which the spin rate was held constant and the nutation frequency varied over a wide range; and sine dwell test, in which both the spin rate and the nutation frequency were held constant. The sine sweep tests were meant to investigate all the prominent liquid resonant oscillations and the damping of the resonances, and the sine dwell tests were meant to quantify the viscous energy dissipation rate of the liquid oscillations for steady state conditions. The LME flight data were compared to analytical results obtained from two companion IR&D programs at Southwest Research Institute. The comparisons indicated that the models predicted the observed liquid resonances, damping, and energy dissipation rates for many test conditions but not for all. It was concluded that improved models and CFD simulations are needed to resolve the differences. This work is ongoing under a current IR&D program.

Deffenbaugh, D. M.↗
NASA NTRS2017

Zero Boil-Off Methods for Large Scale Liquid Hydrogen Tanks Using Integrated Refrigeration and Storage

NASA has completed a series of tests at the Kennedy Space Center to demonstrate the capability of using integrated refrigeration and storage (IRAS) to remove energy from a liquid hydrogen (LH2) tank and control the state of the propellant. A primary test objective was the keeping and storing of the liquid in a zero boil-off state, so that the total heat leak entering the tank is removed by a cryogenic refrigerator with an internal heat exchanger. The LH2 is therefore stored and kept with zero losses for an indefinite period of time. The LH2 tank is a horizontal cylindrical geometry with a vacuum-jacketed, multi-layer insulation system and a capacity of 125,000 liters. The closed-loop helium refrigeration system was a Linde LR1620 capable of 390W cooling at 20K (without any liquid nitrogen pre-cooling). Three different control methods were used to obtain zero boil-off: temperature control of the helium refrigerant, refrigerator control using the tank pressure sensor, and duty cycling (on/off) of the refrigerator as needed. Summarized are the IRAS design approach, zero boil-off control methods, and results of the series of zero boil-off tests.

ZBO↗
NASA NTRS2000

Composite Tank

A composite tank for containing liquid oxygen and the method of making the same Wherein a water-soluble mandrel having ing the desired tank configuration and a cylindrical protuberance on at least one end is fitted with an inner boss conformance, to the configuration of the mandrel and in outer boss conforming to the configuration of the inner boss, the bosses each having a tubular portion for receiving the protuberance on the mandrel and a spherical portion. The mandrel and the bosses are first coated with a nickel coating. The mandrel is then wrapped with graphite fibers wetted with an epoxy resin and this resin is cured. A layer of insulating foam is then applied to the tank and cured. The insulating foam is machined to a desired concentration and a layer of aramid fibers wetted with a second epoxy resin is wrapped around the tank. The second resin is cured and the water soluble mandrel is washed from inside the tank.

Thomas K DeLay↗
NASA NTRS1990

Measuring Diameters Of Large Vessels

Computerized apparatus produces accurate results quickly. Apparatus measures diameter of tank or other large cylindrical vessel, without prior knowledge of exact location of cylindrical axis. Produces plot of inner circumference, estimate of true center of vessel, data on radius, diameter of best-fit circle, and negative and positive deviations of radius from circle at closely spaced points on circumference. Eliminates need for time-consuming and error-prone manual measurements.

Currie, James R.↗
NASA NTRS2005

Making a Metal-Lined Composite-Overwrapped Pressure Vessel

process has been devised for the fabrication of a pressure vessel that comprises a composite-material (matrix/fiber) shell with a metal liner on its inner surface. The use of the composite material makes it possible for the tank to be strong enough to withstand the anticipated operating pressure and yet weigh less than does an equivalent all-metal tank. The metal liner is used as a barrier against permeation: In the absence of such a barrier, the pressurized gas in the tank could leak by diffusing through the composite-material shell. The figure depicts workpieces at four key stages in the process, which consists of the following steps: 1. A mandrel that defines the size and shape of the pressure vessel is made by either molding or machining a piece of tooling wax. 2. Silver paint is applied to the surface of the mandrel to make it electrically conductive. 3. The ends of the mandrel are fitted with metal bosses. 4. The mandrel is put into a plating bath, wherein the metal liner is electrodeposited. Depending on the applications, the liner metal could be copper, nickel, gold, or an alloy. Typical liner thicknesses range from 1 to 10 mils (0.025 to 0.25 mm). 5. The wax is melted from within, leaving the thin metal liner. 6. A hollow shaft that includes holes and fittings through which the liner can be pressurized is sealed to both ends of the liner. The liner is pressurized to stiffen (and hence stabilize) it for the next step. 7. The pressurized liner is placed in a filament-winding machine, which is then operated to cover the liner with multiple layers of an uncured graphite-fiber/epoxy-matrix or other suitable composite material. 8. The composite-overwrapped liner is cured in an oven. 9. The pressure is relieved and the shaft is removed. The tank is then ready for use. The process as described above accommodates variations: a) The mandrel could be made of a wax that melts at a higher temperature and not removed until the tank is cured in the oven. b) The tank need not be cylindrical or axisymmetric, as long as the filament-winding machine can accommodate the chosen shape. c) Shallow grooves could be formed on the surface of the mandrel to give the liner a bellows-like character for reinforcement and/or to accommodate expansion and contraction.

DeLay, Tom↗
NASA NTRS2022

Passive Cooling in Aerogel-Based Insulation Systems for Liquid Hydrogen Upper Stage Launch Vehicle Tanks

Spray-On Foam Insulation (SOFI) is typically used to protect the windward-facing side of flight tanks for space launch vehicle liquid hydrogen (LH2) stages. SOFI is an excellent insulator in ambient pressure environments, however, its performance pales in comparison to reflective-type systems such as Multi-Layer Insulation in the thermal radiation dominated vacuum environment of space. If a windward insulation system employing radiation shields could be devised to replace or supplement SOFI, the on-orbit heat load could be drastically reduced, and the residual propellants could then be used to facilitate secondary missions. Such insulation systems, utilizing aerogel blanket insulation, have been explored by the Cryogenics Test Laboratory at NASA Kennedy Space Center. Owing to its nano-porous structure, aerogel is an excellent adsorbent as well as an insulator at cryogenic temperatures and will readily uptake condensable background gasses such as air when protecting a surface near LH2 temperatures. When an adsorbed blanket is rapidly exposed to a vacuum, such as during the ascent of a rocket, it will release the background gas, producing a passive cooling effect that could potentially reduce or eliminate the heat load on the propellant tank for a time. This work discusses the setup and results of a test program carried out to determine the effectiveness and impact of this cooling effect as it pertains to LH2 upper stage tanks. A vertical-cylindrical liquid nitrogen (LN2) calorimeter test instrument acted as the upper stage analog. The cold-mass was wrapped with two layers of 10 mm thick aerogel blanket material, with and without aluminum heat shields depending on the test, and interspersed with numerous thermocouple temperature sensors. Gaseous argon was used as a stand-in for air as the condensable background gas. The test procedure mimicked vehicle cryo-loading/stabilization, and ascent phases, with the ascent phase accomplished using parallel vacuum pumps. Results show that the desorption cooling effect was enough to not only eliminate the heat load on the LN2 cold-mass entirely, but further refrigerated the liquid below its boiling point.

A.M. Swanger↗
NASA NTRS2005

Structural Analysis of Helios Filament-Wound Tanks Subjected to Internal Pressure and Cooling

A finite-element stress analysis is performed on Helios filament-wound hydrogen tanks to examine the stress field and effect of end dome geometry on the stress field. Each tank is composed of a central circular cylindrical section with either geodesic or hemispherical end domes, which have metallic polar bosses. The tanks are subjected to combined and separate internal pressure and temperature loading conditions, and the stress contributions of each loading component are examined. The tank-wall-polar-boss interfacial meridional tensile stress in the hemispherical dome is found to be approximately 27 percent lower than that in the geodesic dome. The effects of both material anisotropy and the aluminum lining on the intensities of tensile meridional stress at the tank-wall-polar-boss bonding interface are examined.

Ko, William L↗