Liquid Oxygen Tank Research Articles

Propellant Sloshing Parameter Extraction from Computational-Fluid-Dynamics Analysis

Propellant slosh is a potential source of disturbance critical to the stability of space vehicles. The sloshing dynamics are typically represented by a mechanical model of a spring–mass–damper system. This mechanical model is then included in the equation of motion of the entire vehicle for guidance, navigation and control analysis. The typical parameters required by the mechanical model include the natural frequency of the sloshing, sloshing mass, sloshing mass center coordinates, and critical damping coefficient. During the 1960s in the U.S. space program, these parameters were computed either from analytical solutions for simple geometries or by experimental testing for the subscale configurations. The purpose of this work is to demonstrate the soundness of a computational-fluid-dynamics approach in modeling the detailed fluid dynamics of tank sloshing and the excellent accuracy in extracting mechanical properties for different tank configurations and at different fill levels. As the first attempt, the work focuses mainly on the identification of the natural frequency and the equivalent slosh mass from the simulations. The paper presents verification against the analytical solution of natural frequency for two- and three-dimensional straight cylinders, and validation against experimental results for subscale Centaur Liquid Oxygen and Liquid Hydrogen tanks with and without baffles. The results show that computational-fluid-dynamics technology can provide accurate mechanical parameters for any tank configuration, and is especially valuable to the future design of propellant tanks.

Experimental study of moisture uptake of polyurethane foam subjected to a heat sink below 30 K

Rigid closed-cell foam is widely used to thermally insulate liquid hydrogen and oxygen tanks of space launch vehicles due to its lightweight, mechanical strength and thermal-insulating performance. Up to now, little information is available on the intrusion of moisture into the foam that subjects one side to liquid hydrogen temperatures and the other side to room temperatures and high relative humidity. A novel cryogenic moisture uptake apparatus has been designed and fabricated to measure the moisture uptake into the polyurethane foam. For safety and convenience, two identical single-stage pulse tube cryocoolers instead of liquid hydrogen are used to cool one side of the foam specimen to the lowest temperature of 26K. Total of eight specimens in three groups, according to whether there is a butt-joint or weathering period, are tested respectively for both 5h and 9h. The additional weight due to moisture uptake of the foam for the 26K cases is compared to previous measurements at 79K. The results are instructive for the applications of foam to the insulation of liquid hydrogen tanks in space launch vehicles.

The study on the compatibility of epoxy resin with liquid oxygen

Lightweight liquid oxygen (LOX) tank has attracted much attention recently, and epoxy resin is considered to be one of the most likely candidate materials. In this work, polymer wafer made from bisphenol-A epoxy resin was firstly treated with LOX, and then its macroscopic morphology, microscopic structure, and chemical composition were analyzed by atomic force microscopy and Fourier transformed infrared spectroscopy. Finally, its compatibility with LOX was studied by mechanical impact test. The results showed that LOX treatment has little influence on the macroscopic morphology and chemical composition of the epoxy resin wafer, but has a large influence on its microstructure and compatibility with LOX. With the increase of immersion time, the surface roughness, the number, length, and depth of the cracks appeared on the sample surface were all increased. The reaction caused by mechanical impact became more frequent and intense. Finally, possible damage mechanism was discussed.

Explosion Hazard from a Propellant-Tank Breach in Liquid Hydrogen-Oxygen Rockets

An engineering risk assessment of the conditions for massive explosions of cryogenic liquid hydrogen-oxygen rockets during launch accidents is presented. The assessment is based on the analysis of the data of purposeful rupture experiments with liquid oxygen and hydrogen tanks and on an interpretation of these data via analytical semiquantitative estimates and numerical simulations of simplified models for the whole range of the physical phenomena governing the outcome of a propellant-tank breach. The following sequence of events is reconstructed: rupture of fuel tanks, escape of the fluids from the ruptured tanks, liquid film boiling, fragmentation of liquid flow, formation of aerosol oxygen and hydrogen clouds, mixing of the clouds, droplet evaporation, self-ignition of the aerosol clouds, and aerosol combustion. The power of the explosion is determined by a small fraction of the escaped cryogens that become well mixed within the aerosol cloud during the delay time between rupture and ignition. Several scenarios of cavitation-induced self-ignition of the cryogenic hydrogen/oxygen mixture are discussed. The explosion parameters in a particular accident are expected to be highly varied and unpredictable due to randomness of the processes of formation, mixing, and ignition of oxygen and hydrogen clouds. Under certain conditions rocket accidents may result in very strong explosions with blast pressures from a few atm up to 100 atm. The most dangerous situations and the foreseeable risks for space missions are uncovered.

Spray-on foam insulations for launch vehicle cryogenic tanks

Spray-on foam insulation (SOFI) has been developed for use on the cryogenic tanks of space launch vehicles beginning in the 1960s with the Apollo program. The use of SOFI was further developed for the Space Shuttle program. The External Tank (ET) of the Space Shuttle, consisting of a forward liquid oxygen tank in line with an aft liquid hydrogen tank, requires thermal insulation over its outer surface to prevent ice formation and avoid in-flight damage to the ceramic tile thermal protection system on the adjacent Orbiter. The insulation also provides system control and stability throughout the lengthy process of cooldown, loading, and replenishing the tank. There are two main types of SOFI used on the ET: acreage (with the rind) and closeout (machined surface). The thermal performance of the seemingly simple SOFI system is a complex array of many variables starting with the large temperature difference of 200–260K through the typical 25-mm thickness. Environmental factors include air temperature and humidity, wind speed, solar exposure, and aging or weathering history. Additional factors include manufacturing details, launch processing operations, and number of cryogenic thermal cycles. The study of the cryogenic thermal performance of SOFI under large temperature differentials is the subject of this article. The amount of moisture taken into the foam during the cold soak phase, termed Cryogenic Moisture Uptake, must also be considered. The heat leakage rates through these foams were measured under representative conditions using laboratory standard liquid nitrogen boiloff apparatus. Test articles included baseline, aged, and weathered specimens. Testing was performed over the entire pressure range from high vacuum to ambient pressure. Values for apparent thermal conductivity and heat flux were calculated and compared with prior data. As the prior data of record was obtained for small temperature differentials on non-weathered foams, analysis of the different methods is provided. Recent advancements and applications of SOFI systems on future launch vehicles and spacecraft are also addressed.

S192022 極低温に冷却されたCAMUI型固体燃料の点火特性

During development of a large size CAMUI-type hybrid rocket motor in 2006, anomalous combustion associated with large pressure spike frequently occurred. After troubleshooting, low temperature of the fuel was found to have caused anomalous combustion. Because a cryogenic liquid oxygen tank is around the motor, polyethylene fuel grain is cooled to liquid oxygen temperature. The anomalous combustion was duplicated with a downsized motor. The size of the fuel grain was 2/5 of large size motor. Because Damkohler number is proportional to the ratio of pressure and oxygen mass flux, the ratio for the subscale motor was conformed to that for the full-scale motor. In a reproductive experiment, pressure spike also appeared. This result suggests that blow-off occurs before the pressure spike. After the blow-off, gas mixture in the chamber may ignite to cause the pressure spike.

MRI of adsorbed water in solid foams at 21.1 T

In a number of practical situations there is a critical need to evaluate the distribution of small amounts of water absorbed throughout a solid foam sample. One of these pertains to Spray On Foam Insulation (SOFI) NCFI 24124, a thermal insulation material used on the liquid hydrogen and oxygen tanks of the Space Shuttle at Kennedy Space Center (KSC). However, several problems including infinitesimal amounts of water and inevitable water binding to the foam makes the MR signal weak especially for high resolution MRI. The 900 MHz ultra wide bore (UWB) NMR spectrometer at the National High Magnetic Field Laboratory (NHMFL) provides a unique opportunity to perform this evaluation due to its high MRI sensitivity at 21.1 T and its ability to examine large volume samples. The recent upgrade of the UWB900 MHz MRI scanner (August, 2010), allows for ultra-short echo time (UTE) 3D MRI performance. The standard MR imaging technique in this case is not suitable, as it will yield a zero signal. Here we report the first 3D MR images of bound water and water content in solid foam NCFI 24124 samples conditioned to match launch pad conditions.

Cavitation-induced ignition of cryogenic hydrogen-oxygen fluids

The Challenger disaster and purposeful experiments with liquid hydrogen (H2) and oxygen (Ox) tank breaches demonstrated that cryogenic H2/Ox fluids always self-ignite in the process of their sudden mixing. Here, we propose a cavitation-induced self-ignition mechanism that may be realized under these conditions. In one possible scenario, self-ignition is caused by the strong shock waves generated by the collapse of pure Ox vapor bubble near the surface of the Ox liquid that may initiate detonation of the gaseous H2/Ox mixture next to the gas-liquid interface. This effect is further enhanced by H2/Ox combustion inside the collapsing bubble in the presence of admixed H2 gas.

Effect of Return Inlet on Thermal Stratification in a Rocket Tank

Thermal stratification in a rocket liquid-oxygen tank with a natural circulation precooling loop is numerically analyzed under different return inlet locations and influx angles. It is found that the heat convection is dominant in the bulk above the return inlet location and that the heat conduction becomes primary below the location. The axial thermal stratification exists in the whole range of the tank, and the radial stratification appears mainly in the bottom. The stratifications are influenced by the selection of the return inlet location and influx angle, and the optimal values of the return inlet location at hi/L = 0.3 and the influx angle of 30 deg are determined.

Optimizing an Intermediate Ozone System Used for Primary Disinfection at a 55 MGD Surface Water Treatment Plant

The Henrico County VA Water Treatment Facility has been in operation since April 2004, making use of intermediate ozone for primary disinfection to take advantage of both the disinfection and DBP benefits. Since the summer of 2007 the staff at the Henrico WTF has embarked on an optimizing journey with their ozone system. This paper overviews the significant milestones accomplished on this journey as well as ongoing improvements. Optimizing efforts to reduce operating costs while achieving disinfection goals included: Staff training workshops; Automating ozone data collection and access; Operating system at optimum ozone concentration; Selecting optimum contactor sampling locations and number of contactors; Overcoming minimum gas flow limitations; Solenoid contactor sampling system issues and improvements; Disinfection calculation constraints; Excessive buildup of liquid oxygen (LOX) tank pressure. Each of these items is discussed following an overview of this 55 MGD surface water treatment facility.

Optimisation methodology for cryotanks

Liquid Oxygen (LOX) tanks represent an important percentage of net weight and cost of any reusable launch vehicle (RLV). Moreover they play structural and thermal and propulsive roles at the same time. The transition from metal to composite material is then considered an obliged choice. In this study, a FEM optimisation procedure has been developed in order to both increase payload (by decreasing tank mass) and improve aerodynamic and controllability (by increasing global flexural stiffness). The optimisation of these two contradictory objectives have been obtained using a standard FE numerical tool within an innovative optimisation procedure, as a possible alternative to genetic algorithm approaches [Jaunky N, Knight NF, Ambur DR. Optimal design of general stiffened composite circular cylinders for global buckling with strength constraints. Compos Struct 1998;41:243–52. [1]]. The key features of this procedure are: the starting point determination methodology; a sensitivity analysis based on an integration between the classical laminate theory and the netting theory to validate the FE optimiser results; the strategy of choice of the best compromise solution in terms of design constraints and objective function.

Numerical Analysis for Propellant Management in Rocket Tanks

To track the three-dimensional behavior of liquid surfaces in storage tanks on orbit, a numerical scheme based on the level set method was developed. For the verification of the numerical methods, experimental data were acquired through the observation of the unsteady deformations of liquid surfaces in cylindrical containers under low-gravity conditions in a drop tower. Primary emphasis was placed on the dynamic behavior of the liquid surface and the displacement of the contact line on the solid wall. Based on the results, the boundary condition for surface tension was discussed and the model of wetting phenomena was adequately introduced into the computation. Compared with the experimental data, the corresponding numerical results showed good agreement. With the developed code, the flowfields at the draining process in the liquid oxygen (LO X ) tank for the upper-stage propulsion system of an H-IIA launch vehicle under low-gravity conditions were also investigated. It was found that the buoyancy induced by a slight acceleration was efficient to prevent dip growth and that the serviceable propellant in the launch-vehicle tank could thereby be increased in a realistic situation.

An uncertainty analysis of the PVT gauging method applied to sub-critical cryogenic propellant tanks

The PVT (pressure, volume, temperature) method of liquid quantity gauging in low-gravity is based on gas law calculations assuming conservation of pressurant gas within the propellant tank and the pressurant supply bottle. There is interest in applying this method to cryogenic propellant tanks since the method requires minimal additional hardware or instrumentation. To use PVT with cryogenic fluids, a non-condensable pressurant gas (helium) is required. With cryogens, there will be a significant amount of propellant vapor mixed with the pressurant gas in the tank ullage. This condition, along with the high sensitivity of propellant vapor pressure to temperature, makes the PVT method susceptible to substantially greater measurement uncertainty than is the case with less volatile propellants. A conventional uncertainty analysis is applied to example cases of liquid hydrogen and liquid oxygen tanks. It appears that the PVT method may be feasible for liquid oxygen. Acceptable accuracy will be more difficult to obtain with liquid hydrogen.

P31-02 航行型海中ロボット用閉鎖式ディーゼル機関

This paper presents details of an air-independent power (AIP) system of a Closed Cycle Diesel Engine System (CCDE) for a cruising type autonomous underwater vehicle (AUV) called "R-One Robot" The CCDE is controlled fully automatically, and produces 5kW output and its capacity is about 60kWh, which corresponds to 24 hour and 100 km range continuous dive of the robot. The CCDE consists of a diesel engine and peripheral equipment for closed cycle operation such as a chemical absorber for carbon dioxide removal from exhaust gas using potassium hydro-oxide, an oxygen replenishment system including oxygen sensor in the control system and a liquid oxygen tank. Vibration and noise of the engine are sufficiently isolated for the vehicle hull. The R-One Robot is 8 m length and 4.8 tons in weight, and its depth rate is 400 m. The first autonomous dive of the robot was carried out in 1996,and has been used for scientific research of the ocean so far.

Effects of initial geometric imperfections on the non-linear response of the Space Shuttle superlightweight liquid-oxygen tank

The results of an analytical study of the elastic buckling and non-linear behavior of the liquid-oxygen tank for the new Space Shuttle superlightweight external fuel tank are presented. Selected results that illustrate three distinctly different types of non-linear response phenomena for thin-walled shells which are subjected to combined mechanical and thermal loads are presented. These response phenomena consist of a bifurcation-type buckling response, a short-wavelength non-linear bending response, and a non-linear collapse or “snap-through” response associated with a limit point. The effects of initial geometric imperfections on the response characteristics are emphasized. The results illustrate that the buckling and non-linear response of a geometrically imperfect shell structure subjected to complex loading conditions may not be adequately characterized by an elastic linear bifurcation buckling analysis, and that the traditional industry practice of applying a buckling-load knock-down factor can result in an ultra-conservative design. Results are also presented that show that a fluid-filled shell can be highly sensitive to initial geometric imperfections, and that the use a buckling-load knock-down factor is needed for this case.

Effects of Welding-Induced Imperfections on Behavior of Space Shuttle Superlightweight Tank

Results of linear-bifurcation buckling and nonlinear analyses of the Space Shuttle superlightweight external liquid-oxygen tank are presented for an important prelaunch loading condition. These results show the effects of actual, measured welding-induced initial geometric imperfections on an important response mode for thin-walled shellsthataresubjectedtocombinedmechanicalandthermalloads.Thistypeofinitialgeometricimperfectionmay be encountered in the design of other liquid-fuel launch vehicles. Results are presented that show that the liquidoxygen tank will buckle in the barrel section, but at load levels nearly four times the magnitude of the operational load level, and will exhibit stable postbuckling behavior. The actual measured imperfections are located in this sectionofthetank. Resultsofimperfection sensitivity analysesarepresented thatshowthat thelargestdegradation in the apparent membrane stiffnesses of the liquid-oxygen tank barrel section is caused by an imperfection shape that is in the form of the linear-bifurcation buckling mode with a relatively small amplitude. These results also show that the effect of the relatively large-amplitude measured imperfection is benign.

Nonlinear Behavior of Space Shuttle Superlightweight Tank Under End-of-Flight Loads

Results of elastic, linear-bifurcation buckling and nonlinear analyses of the new Space Shuttle superlightweight external liquid-oxygen tank are presented for an important end-of-e ight loading condition. These results illustrate an important type of response mode for thin-walled shells subjected to combined mechanical and thermal loads that may be encountered in the design of other liquid-fuel launch vehicles. Linear-bifurcation buckling analyses are presented that predict several nearly equal eigenvalues that correspond to local buckling modes in the aft dome of the liquid-oxygen tank. In contrast, the nonlinear response phenomenon is shown to consist of a shortwavelength bending deformation in the aft elliptical dome of the liquid-oxygen tank that grows in amplitude in a stable manner with increasing load. Imperfection sensitivity analyses are presented that show that the presence of several nearly equal eigenvalues does not lead to a premature general instability mode for the aft dome. For the linear-bifurcation and nonlinear analyses, the results show that accurate predictions of the response of the shell generally require a large-scale, high-e delity, e nite element model, and that a design based on a linear-bifurcation buckling analysis and a buckling-load knockdown factor is overly conservative. Results are also presented that showthatthesuperlightweightliquid-oxygentankcansupportloadsin excessofapproximately1.9 timesthevalues of the operational loads considered.

Nonlinear Behavior of Space Shuttle Superlightweight Tank Under Booster Ascent Loads

Results of linear-bifurcation and nonlinear analyses of the Space Shuttle superlightweight (SLWT) external liquid-oxygen (LO2) tank for an important early booster ascent loading condition are presented. These results for thin-walled linear elastic shells that are subjected to combined mechanical and thermal loads illustrate an important type of response mode that may be encountered in the design of other liquid-fuel launch vehicles. Linear-bifurcation analyses are presented that predict several nearly equal eigenvalues that correspond to local buckling modes in the forward ogive section of the LO 2 tank. In contrast, the nonlinear response phenomenon is shown to consist of short-wavelength bending deformations in the forward-ogive and barrel sections of the LO2 tank that grow in amplitude in a stable manner with increasing load. Imperfection sensitivity analyses are presented that show that the presence of several nearly equal eigenvalues does not lead to a premature general instability mode for the forward-ogive section. For the linear-bifurcation and nonlinear analyses, the results show that accurate predictions of the response of the shell generally require a large-scale, high-e delity, e nite element model, and that a design based on a linear-bifurcation buckling analysis and a buckling-load knockdown factor is overly conservative. Results are presented that show that the SLWT LO 2 tank can support loads in excess of approximately 2.6 timesthevaluesof theoperationalloads considered. In addition, results are presented that show that local bending deformations may cause failure of the thermal protection system (TPS) at load levels less than the load level corresponding to structural collapse. Results are presented that can be used to estimate the load level at which TPS failure is likely to occur.

Control strategies for space boosters using air collection systems

A simple and efficient performance analysis method is developed for evaluating vehicle pitch and engine throttle controls to minimize booster fuel required to fill second-stage liquid oxygen (LOX) tanks and deliver the vehicle to the staging point. An optimization methodology finds a throttle schedule that controls both the air-breathing engine and the LOX collection rate. The altitude-velocity profile is derived from a variational calculus/energy management contouring method. Automatic adaptive-gain pitch-rate and throttle controls are developed. Results from a parametric study show that collecting on the run for an optimum schedule results in a 17% fuel savings over collection at a constant Mach number.

Lunar habitat concept employing the Space Shuttle external tank

The space shuttle external tank, which consists of a liquid oxygen tank, an intertank structure, and a liquid hydrogen tank, is an expendable structure used for approximately 8.5 min during each launch. A concept for outfitting the liquid oxygen tank-intertank unit for a 12-person lunar habitat is described. The concept utilizes existing structures and openings for both man and equipment access without compromising the structural integrity of the tank. Living quarters, instrumentation, environmental control and life support, thermal control, and propulsion systems are installed at Space Station Freedom. The unmanned habitat is then transported to low lunar orbit and autonomously soft landed on the lunar surface. Design studies indicate that this concept is feasible by the year 2000 with concurrent development of a space transfer vehicle and manned cargo lander for crew changeover and resupply.