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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Propulsive effect of spacecraft propellant tank rupture following hypervelocity impact by a micrometeoroid

The impulse due to propellant tank micrometeroid damage is estimated for the Galileo spacecraft. The modes of tank failure considered are tank burst, or rupture, and puncture, which results in a hole through which propellant escapes. Other factors relevant to this calculation include spacecraft geometry, propellant volume and properties, and available heat sources. Tank burst was found to result in a spacecraft Delta V of less than 5 m/s, while the net impulse due to puncture is less than 10,000 Newton-seconds.

Jan, Darrell L.↗
NASA NTRS2001

Testing of Densified Liquid Hydrogen Stratification in a Scale Model Propellant Tank

This paper describes a test program that was conducted at NASA to demonstrate the ability to load densified LH2 into a subscale propellant tank. This work was done through a collaborative effort between NASA Glenn Research Center and the Lockheed Martin Michoud Space Systems (LMMSS). The Multilobe tank, which was made from composite materials similar to that to be used on X-33, was formed from two lobes with a center septum. Test results are shown for data that was collected on filling the subscale tank with densified liquid hydrogen (DLH2) propellant that was produced at the NASA Plum Brook Station. Data is compared to analytical predictions. Data collected for this test series agrees well with analytical predictions of the environmental heat leak into the tank and the thermal stratification characteristics of the hydrogen propellant in the tank as it was filled with DLH2.

Jurns, John M.↗
NASA NTRS2012

Concepts for a Shroud or Propellant Tank Derived Deep Space Habitat

Long duration human spaceflight missions beyond Low Earth Orbit will require much larger spacecraft than capsules such as the Russian Soyuz or American Orion Multi-Purpose Crew Vehicle. A concept spacecraft under development is the Deep Space Habitat, with volumes approaching that of space stations such as Skylab, Mir, and the International Space Station. This paper explores several concepts for Deep Space Habitats constructed from a launch vehicle shroud or propellant tank. It also recommends future research using mockups and prototypes to validate the size and crew station capabilities of such a habitat. Keywords: Exploration, space station, lunar outpost, NEA, habitat, long duration, deep space habitat, shroud, propellant tank.

Howard, Robert L.↗
NASA NTRS1967

Propellant tank pressurization analysis program

Computer program for the analysis of a single propellant tank pressurization system includes many pertinent physical phenomena previously ignored in other mathematical models. This program can be used for analysis, simulation, and design of propellant pressurization systems.

Epstein, M.↗
NASA NTRS1983

Test program to demonstrate the stability of hydrazine in propellant tanks

The suitability of stainless steels and Inconel for long-term hydrazine propellant-storage tanks is investigated. Rectangular coupon samples cut from propellent tanks were sealed with a measured amount of hydrazine in glass capsules, stored at 43 or 60 C, and removed after 6 to 24 months, when corrosion of the coupon and decomposition of the hydrazine was determined, and SEM and electron spectroscopy were performed on some coupons. Corrosion was found to be unmeasurably low for all the coupons, and hydrazine decomposition produced less than 1.0 cu cm of gas per sq cm of wetted surface per year, except in those few cases when catalysis or contamination were detected. Especially good stability was observed for type 304L stainless steel. The decomposition rates determined in the coupon tests are confirmed by preliminary results of actual tank storage trials.

Moran, C. M.↗
NASA NTRS2014

Pressure-Volume-Temperature (PVT) Gauging of an Isothermal Cryogenic Propellant Tank Pressurized with Gaseous Helium

Results are presented for pressure-volume-temperature (PVT) gauging of a liquid oxygen/liquid nitrogen tank pressurized with gaseous helium that was supplied by a high-pressure cryogenic tank simulating a cold helium supply bottle on a spacecraft. The fluid inside the test tank was kept isothermal by frequent operation of a liquid circulation pump and spray system, and the propellant tank was suspended from load cells to obtain a high-accuracy reference standard for the gauging measurements. Liquid quantity gauging errors of less than 2 percent of the tank volume were obtained when quasi-steady-state conditions existed in the propellant and helium supply tanks. Accurate gauging required careful attention to, and corrections for, second-order effects of helium solubility in the liquid propellant plus differences in the propellant/helium composition and temperature in the various plumbing lines attached to the tanks. On the basis of results from a helium solubility test, a model was developed to predict the amount of helium dissolved in the liquid as a function of cumulative pump operation time. Use of this model allowed correction of the basic PVT gauging calculations and attainment of the reported gauging accuracy. This helium solubility model is system specific, but it may be adaptable to other hardware systems.

VanDresar, Neil T.↗
NASA NTRS1984

Evaluation of propellent tank insulation concepts for low-thrust chemical propulsion systems: Executive summary

Cryogenic propellant tank insulations or liquid oxygen/liquid hydrogen low-thrust 2224N (500 lbf) propulsion systems (LTPS) were assessed. The insulation studied consisted of combinations of N2-purged foam and multilayer insulation (MLI) as well as He-purged MLI-only. Heat leak and payload performance predictions were made for three shuttle-launched LTPS designed for shuttle bay packaged payload densities of 56 kg cu/m (3.5 lbm/cu ft), 40 kg/cu m (2.5 lbm/cu ft) and 24 kg/cu m (1.5 lbm/cu ft). Foam/MLI insulations were found to increase LTPS payload delivery capability when compared with He-purged MLI-only. An additional benefit of foam/MLI was reduced operational complexity because orbiter cargo bay N2 purge gas could be used for MLI purging. Maximum payload mass benefit occurred when an enhanced convection, rather than natural convection, heat transfer was specified for the insulation purge enclosure. The enhanced convection environment allowed minimum insulation thickness to be used for the foam/MLI interface temperature selected to correspond to the moisture dew point in the N2 purge gas. Experimental verification of foam/MLI benefits was recommended. A conservative program cost estimate for testing a MLI-foam insulated tank was 2.1 million dollars. This cost could be reduced significantly without increasing program risk.

Kramer, T.↗