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At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Ground testing for the no-vent fill of cryogenic tanks - Results of tests for a 71 cubic foot tank

NASA Lewis Research has been investigating the no-vent till method, since it is a promising approach to transfer liquid while handling the problems of low-g venting. This paper reports the results of a test series for filling a 71 cu ft tank with liquid hydrogen without venting. Twenty two tests were conducted, ten with a bottom orifice as the inlet and 12 with a spray bar. Parameters investigated included inlet saturation pressures of approximately 5, 15, and 25 psia; transfer pressures of 20, 30, and 45 psia; and various starting wall temperatures. Of the tests, only the one run at the highest wall temperature (238 R) failed to fill the tank. Test results are compared to a thermodynamic equilibrium model. Overall model-data agreement was good except for the tendency of the model to overshoot during the initial wall cool-down of the higher starting wall temperature fills.

Chato, David J.↗
NASA NTRS1993

Ground testing for the no-vent fill of cryogenic tanks: Results of tests for a 71 cubic foot tank

NASA Lewis Research has been investigating the no-vent fill method, since it is a promising approach to transfer liquid while handling the problems of low-g venting. This paper reports the results of a test series for filling a 71 cu ft tank with liquid hydrogen without venting. 22 tests were conducted, 10 with a bottom orifice as the inlet and 12 with a spray bar. Parameters investigated included inlet saturation pressures of approximately 5, 15, and 25 psia, transfer pressures of 20, 30, and 45 psia, and various starting wall temperatures. Of the tests, only the one run at the highest wall temperature (238 R) failed to fill the tank. Test results are compared to a thermodynamic equilibrium model. Overall model-data agreement was good except for the tendency of the model to overshoot during the initial wall cool down of the higher starting wall temperature fills.

Chato, David J.↗
NASA NTRS1975

An investigation of the 0.0091 scale External Tank ogive nose (MSFC model 470) in the MSFC 14 inch TWT to determine the pressure distribution around the External Tank Nose (TA3F), volume 1

The model tested was a 0.0091 scale representation of the ogive nose section of the External Tank with nose cap and lightning rod and protuberances. Pressure distribution around the nose cap was determined. Pressure data were also obtained along the ogive nose. Data were obtained over an angle of attack range of + or -5 degrees and over a Mach number range of .6 to 4.96. The Reynolds number per unit length (ft.) ranged from 4.1 million to 4.96 million. There were 22 pressure ports in a single row. Circumferential positions of 0, 22.5, 45, 67.5 and 90 degrees were simulated by rotating the model. The LO2 feed line and LO2 recirculation line were simulated. The effects of the nose spike were investigated over a range of Mach numbers. The data will be useful for venting analyses. Photographs showing flow visualization of the test configuration are given.

Ramsey, P. E.↗
NASA NTRS1946

Tank Tests of a 1/7-Size Powered Dynamic Model of the Grumman XJR2F-1 Amphibian: Spray Characteristics, Take-Off and Landing Stability in Smooth Water - Langley Tank Model 212, TED No. NACA 2378

Tests of a model of the XJR2F-Y amphibian were made in Langley tank no. to determine the spray characteristics and the take-off and landing stability. At a gross load of 22,000 pounds full size, spray entered the propeller disk only at a very narrow range of speeds. The spray striking the flaps was not excessive and no appreciable wetting of the tail surfaces was noted. The trim limits of stability appeared to be satisfactory and the upper-limit porpoising was not violent. The stable range of center-of-gravity locations with flaps set 20deg was well aft of the desired operating range. However, with flaps up, the forward limit was about 18 percent mean aerodynamic chord and the aft limit about 28.5 percent mean aerodynamic chord at a load of 26,000 pounds and with elevators deflected -10deg. Under these conditions the location of the step is considered satisfactory. Tests showed that the effect of water in the nose-wheel well would be to move the forward limit aft about 2-percent mean aerodynamic chord. Without ventilation of the main step, the model skipped during landing at most trims, but this skipping was not violent. With the ventilation, the model skipped lightly only at trims where the afterbody keel was approximately parallel to the water (around 7.5 deg).

Land, Norman S.↗
NASA NTRS1946

Tank Tests of 1/5.5-Scale Forward Dynamic Model of the Columbia XJL-1 Amphibian - Langley Tank Model 208, TED No. NACA 2336

Tests of a powered dynamic model of the Columbia XJL-1 amphibian were made in Langley tank no.1 to determine the hydrodynamic stability and spray characteristics of the basic hull and to investigate the effects of modifications on these characteristics. Modifications to the forebody chime flare, the step, and the afterbody, and an increase in the angle of incidence of the wing were included in the test program. The seaworthiness and spray characteristics were studied from simulated taxi runs in smooth and rough water. The trim limits of stability, the range of stable positions of the enter of gravity for take-off, and the landing stability were determined in smooth water. The aerodynamic lift, pitching moment, and thrust were determined at speeds up to take-off speed.

Havens, Robert F.↗
NASA NTRS2010

Radio-Frequency Tank Eigenmode Sensor for Propellant Quantity Gauging

Although there are several methods for determining liquid level in a tank, there are no proven methods to quickly gauge the amount of propellant in a tank while it is in low gravity or under low-settling thrust conditions where propellant sloshing is an issue. Having the ability to quickly and accurately gauge propellant tanks in low-gravity is an enabling technology that would allow a spacecraft crew or mission control to always know the amount of propellant onboard, thus increasing the chances for a successful mission. The Radio Frequency Mass Gauge (RFMG) technique measures the electromagnetic eigenmodes, or natural resonant frequencies, of a tank containing a dielectric fluid. The essential hardware components consist of an RF network analyzer that measures the reflected power from an antenna probe mounted internal to the tank. At a resonant frequency, there is a drop in the reflected power, and these inverted peaks in the reflected power spectrum are identified as the tank eigenmode frequencies using a peak-detection software algorithm. This information is passed to a pattern-matching algorithm, which compares the measured eigenmode frequencies with a database of simulated eigenmode frequencies at various fill levels. A best match between the simulated and measured frequency values occurs at some fill level, which is then reported as the gauged fill level. The database of simulated eigenmode frequencies is created by using RF simulation software to calculate the tank eigenmodes at various fill levels. The input to the simulations consists of a fairly high-fidelity tank model with proper dimensions and including internal tank hardware, the dielectric properties of the fluid, and a defined liquid/vapor interface. Because of small discrepancies between the model and actual hardware, the measured empty tank spectra and simulations are used to create a set of correction factors for each mode (typically in the range of 0.999 1.001), which effectively accounts for the small discrepancies. These correction factors are multiplied to the modes at all fill levels. By comparing several measured modes with the simulations, it is possible to accurately gauge the amount of propellant in the tank. An advantage of the RFMG approach of applying computer simulations and a pattern-matching algorithm is that the Although there are several methods for determining liquid level in a tank, there are no proven methods to quickly gauge the amount of propellant in a tank while it is in low gravity or under low-settling thrust conditions where propellant sloshing is an issue. Having the ability to quickly and accurately gauge propellant tanks in low-gravity is an enabling technology that would allow a spacecraft crew or mission control to always know the amount of propellant onboard, thus increasing the chances for a successful mission. The Radio Frequency Mass Gauge (RFMG) technique measures the electromagnetic eigenmodes, or natural resonant frequencies, of a tank containing a dielectric fluid. The essential hardware components consist of an RF network analyzer that measures the reflected power from an antenna probe mounted internal to the tank. At a resonant frequency, there is a drop in the reflected power, and these inverted peaks in the reflected power spectrum are identified as the tank eigenmode frequencies using a peak-detection software algorithm. This information is passed to a pattern-matching algorithm, which compares the measured eigenmode frequencies with a database of simulated eigenmode frequencies at various fill levels. A best match between the simulated and measured frequency values occurs at some fill level, which is then reported as the gauged fill level. The database of simulated eigenmode frequencies is created by using RF simulation software to calculate the tank eigenmodes at various fill levels. The input to the simulations consists of a fairly high-fidelity tank model with proper dimensions and including internal tank harare, the dielectric properties of the fluid, and a defined liquid/vapor interface. Because of small discrepancies between the model and actual hardware, the measured empty tank spectra and simulations are used to create a set of correction factors for each mode (typically in the range of 0.999 1.001), which effectively accounts for the small discrepancies. These correction factors are multiplied to the modes at all fill levels. By comparing several measured modes with the simulations, it is possible to accurately gauge the amount of propellant in the tank. An advantage of the RFMG approach of applying computer simulations and a pattern-matching algorithm is that the

Zimmerli, Gregory A.↗
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 NTRS2016

Vented Chill / No-Vent Fill of Cryogenic Propellant Tanks

Architectures for extended duration missions often include an on-orbit replenishment of the space vehicle's cryogenic liquid propellants. Such a replenishment could be accomplished via a tank-to-tank transfer from a dedicated tanker or a more permanent propellant depot storage tank. Minimizing the propellant loss associated with transfer line and receiver propellant tank thermal conditioning is essential for mass savings. A new methodology for conducting tank-to-tank transfer while minimizing such losses has been demonstrated. Charge-Hold-Vent is the traditional methodology for conducting a tank-to-tank propellant transfer. A small amount of cryogenic liquid is introduced to chill the transfer line and propellant tank. As the propellant absorbs heat and undergoes a phase change, the tank internal pressure increases. The tank is then vented to relieve pressure prior to another charge of cryogenic liquid being introduced. This cycle is repeated until the transfer lines and tank are sufficiently chilled and the replenishment of the propellant tank is complete. This method suffers inefficiencies due to multiple chill and vent cycles within the transfer lines and associated feed system components. Additionally, this system requires precise measuring of cryogenic fluid delivery for each transfer, multiple valve cycling events, and other complexities associated with cycled operations. To minimize propellant loss and greatly simplify on-orbit operations, an alternate methodology has been designed and demonstrated. The Vented Chill / No Vent Fill method is a simpler, constant flow approach in which the propellant tank and transfer lines are only chilled once. The receiver tank is continuously vented as cryogenic liquid chills the transfer lines, tank mass and ullage space. Once chilled sufficiently, the receiver tank valve is closed and the tank is completely filled. Interestingly, the vent valve can be closed prior to receiver tank components reaching liquid saturation temperature. An incomplete fill results if insufficient energy is removed from the tank's thermal mass and ullage space. The key to successfully conducting the no vent fill is to assure that sufficient energy is removed from the system prior to closing the receiver tank vent valve. This paper will provide a description of the transfer methodology and test article, and will provide a discussion of test results.

Rhys, Noah O.↗
NASA NTRS1999

Conformal Cryogenic Tank Trade Study for Reusable Launch Vehicles

Future reusable launch vehicles may be lifting bodies with non-circular cross section like the proposed Lockheed-Martin VentureStar(tm). Current designs for the cryogenic tanks of these vehicles are dual-lobed and quad-lobed tanks which are packaged more efficiently than circular tanks, but still have low packaging efficiencies with large gaps existing between the vehicle outer mold line and the outer surfaces of the tanks. In this study, tanks that conform to the outer mold line of a non-circular vehicle were investigated. Four structural concepts for conformal cryogenic tanks and a quad-lobed tank concept were optimized for minimum weight designs. The conformal tank concepts included a sandwich tank stiffened with axial tension webs, a sandwich tank stiffened with transverse tension webs, a sandwich tank stiffened with rings and tension ties, and a sandwich tank stiffened with orthogrid stiffeners and tension ties. For each concept, geometric parameters (such as ring frame spacing, the number and spacing of tension ties or webs, and tank corner radius) and internal pressure loads were varied and the structure was optimized using a finite-element-based optimization procedure. Theoretical volumetric weights were calculated by dividing the weight of the barrel section of the tank concept and its associated frames, webs and tension ties by the volume it circumscribes. This paper describes the four conformal tank concepts and the design assumptions utilized in their optimization. The conformal tank optimization results included theoretical weights, trends and comparisons between the concepts, are also presented, along with results from the optimization of a quad-lobed tank. Also, the effects of minimum gauge values and non-optimum weights on the weight of the optimized structure are described in this paper.

Rivers, H. Kevin↗
NASA NTRS2007

Runtime and Pressurization Analyses of Propellant Tanks

Multi-element unstructured CFD has been utilized at NASA SSC to carry out analyses of propellant tank systems in different modes of operation. The three regimes of interest at SSC include (a) tank chill down (b) tank pressurization and (c) runtime propellant draw-down and purge. While tank chill down is an important event that is best addressed with long time-scale heat transfer calculations, CFD can play a critical role in the tank pressurization and runtime modes of operation. In these situations, problems with contamination of the propellant by inclusion of the pressurant gas from the ullage causes a deterioration of the quality of the propellant delivered to the test article. CFD can be used to help quantify the mixing and propellant degradation. During tank pressurization under some circumstances, rapid mixing of relatively warm pressurant gas with cryogenic propellant can lead to rapid densification of the gas and loss of pressure in the tank. This phenomenon can cause serious problems during testing because of the resulting decrease in propellant flow rate. With proper physical models implemented, CFD can model the coupling between the propellant and pressurant including heat transfer and phase change effects and accurately capture the complex physics in the evolving flowfields. This holds the promise of allowing the specification of operational conditions and procedures that could minimize the undesirable mixing and heat transfer inherent in propellant tank operation. It should be noted that traditional CFD modeling is inadequate for such simulations because the fluids in the tank are in a range of different sub-critical and supercritical states and elaborate phase change and mixing rules have to be developed to accurately model the interaction between the ullage gas and the propellant. We show a typical run-time simulation of a spherical propellant tank, containing RP-1 in this case, being pressurized with room-temperature nitrogen at 540 R. Nitrogen, shown in blue on the right-hand side of the figures, enters the tank from the diffuser at the top of the figures and impinges on the RP-1, shown in red, while the propellant is being continuously drained at the rate of 1050 lbs/sec through a pipe at the bottom of the tank. The sequence of frames in Figure 1 shows the resultant velocity fields and mixing between nitrogen and RP-1 in a cross-section of the tank at different times. A vortex is seen to form in the incoming nitrogen stream that tends to entrain propellant, mixing it with the pressurant gas. The RP-1 mass fraction contours in Figure 1 are also indicative of the level of mixing and contamination of the propellant. The simulation is used to track the propagation of the pure propellant front as it is drawn toward the exit with the evolution of the mixing processes in the tank. The CFD simulation modeled a total of 10 seconds of run time. As is seen from Figure 1d, after 5.65 seconds the propellant front is nearing the drain pipe, especially near the center of the tank. Behind this pure propellant front is a mixed fluid of compromised quality that would require the test to end when it reaches the exit pipe. Such unsteady simulations provide an estimate of the time that a high-quality propellant supply to the test article can be guaranteed at the modeled mass flow rate. In the final paper, we will discuss simulations of the LOX and propellant tanks at NASA SSC being pressurized by an inert ullage. Detailed comparisons will be made between the CFD simulations and lower order models as well as with test data. Conditions leading to cryo collapse in the tank will also be identified.

Field, Robert E.↗