Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “Tank testing”

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 829 records · Page 46
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Boundary layer manipulators at high Reynolds numbers

Airfoil large-eddy breakup (LEBU) devices were tested on an axisymmetric body in the Langley Two Tank up to speeds of 50 ft/sec. NACA-0009, NACA-2412, E-193, and Clark Y contours were tested in single and tandem configurations. At the higher Reynolds numbers local skin friction downstream of the devices showed minimal reductions O (10 percent) and total body drag was increased 1 to 3 percent. At lower Reynolds numbers skin-friction reductions as large as 25 percent were measured and total body drag tended toward net reductions. The loss of effectiveness with increasing Reynolds number of conventional, outer layer devices suggests a decoupling of the outer and inner scales in high Reynolds number turbulent boundary layers.

Anders, J. B.↗
NASA NTRS1992

Ion engine endurance testing at high background pressures

Ion engine endurance testing at vacuum chamber pressures in the low 10 exp -3 Pa range is enabled through the use of a three-grid accelerator system with the decelerator grid biased 50 to 100 volts negative of neutralizer cathode potential. The negative decelerator grid serves to collect the facility induced charge exchange ion current which normally results in rapid erosion of the accelerator grid during testing at elevated vacuum chamber pressures. This screen, accelerator, negative decelerator (SAND) grid configuration enables an order of magnitude reduction in vacuum chamber pumping speeds relative to that required for endurance testing of ion engines with conventional two-grid accelerator systems. A 900-hr test of a 30-cm diameter engine at 6.5 kW and a tank pressure of 3.7 x 10 exp -3 Pa was performed to test the feasibility of the three-grid SAND accelerator system technology. Grid erosion rates from this test are compared to those from a 200-hr test performed with the same discharge chamber, in the same test facility, and at the same background pressure with a conventional two grid accelerator system. The SAND optics resulted in greater than a factor of 100 reduction in accelerator grid erosion rate relative to the two-grid system.

Brophy, John R.↗
NASA NTRS1993

Domain connectivity among systems of overset grids

This progress report describes research performed to achieve algorithmic improvements in the use of overset grids for the computation of unsteady flowfields about geometrically complex and moving component configurations. A hole-map computational method is detailed that reduces computational demand and corresponding demand on human resources. Results indicate that the method is both efficient and flexible for use in complex simulation cases. Tests were performed on shuttle orbiter/external tank, wing/store, and V-22 tilt-rotor configurations. The number of simulated points and the corresponding time on a SGI 4D-210 workstation are presented.

Meakin, Robert L.↗
NASA NTRS1980

Lightning Protection

An airplane's wingtip tank is being struck by simulated lightning in a test conducted by Lightning Technologies, Inc., a firm specializing in design features to protect aircraft from the hazard of electrical phenomena. Much of the technology employed in tests originated in NASA - sponsored studies focusing on effects of lightning on aircraft structures, electrical systems, and fuel tanks, and on means of protecting against hazardous effects.

Source record↗
NASA NTRS2003

MAP Propulsion System Thermal Design

The propulsion system of the Microwave Anisotropy Probe (MAP) had stringent requirements that made the thermal design unique. To meet instrument stability requirements the system had to be designed to keep temperatures of all components within acceptable limits without heater cycling. Although the spacecraft remains at a fixed 22 sun angle at L2, the variations in solar constant, property degradation, and bus voltage range all significantly affect the temperature. Large portions of the fuel lines are external to the structure and all components are mounted to non-conductive composite structure. These two facts made the sensitivity to the MLI effective emissivity and bus temperature very high. Approximately two years prior to launch the propulsion system was redesigned to meet MAP requirements. The new design utilized hardware that was already installed in order to meet schedule constraints. The spacecraft design and the thermal requirements were changed to compensate for inadequacies of the existing hardware. The propulsion system consists of fuel lines, fill and drain lines/valve, eight thrusters, a HXCM, and a propulsion tank. A voltage regulator was added to keep critical components within limits. Software was developed to control the operational heaters. Trim resistors were put in series with each operational heater circuits and the tank survival heater. A highly sophisticated test program, which included real time model correlation, was developed to determine trim resistors sizes. These trim resistors were installed during a chamber break and verified during thermal balance testing.

Mosier, Carol L.↗
NASA NTRS1948

Free-Spinning-Tunnel Tests of a 1/24-Scale Model of the Grumman XF9F-2 Airplane with Wing-Tip Tanks Installed

An investigation of the spin and recovery characteristics of a 1/24-scale model of the Grumman XF9F-2 airplane with wing-tip tanks installed has been conducted-in the Langley 20-foot free-spinning tunnel. The effects of control settings and movements on the erect spin and recovery characteristics of the model for a range of possible loadings of the tip tanks were determined. Spin and recovery characteristics without tanks were determined in a previous investigation. The model results indicated that the airplane spins will generally be oscillatory and that recoveries will be satisfactory for all loadings by normal recovery technique (full rudder reversal followed approximately one-half turn later by moving the elevator down). The rudder force necessary for recovery should be within the physical capability of the pilot but the elevator force may be excessive so that some type of balance or booster might be necessary, or it might be necessary to jettison the wing-tip tanks.

Berman, Theodore↗
NASA NTRS2010

NASA Ares I Launch Vehicle First Stage Roll Control System Cold Flow Development Test Program Overview

The Ares I launch vehicle is the selected design, chosen to return humans to the moon, Mars, and beyond. It is configured in two inline stages: the First Stage is a Space Shuttle derived five-segment Solid Rocket Booster and the Upper Stage is powered by a Saturn V derived J-2X engine. During launch, roll control for the First Stage (FS) is handled by a dedicated Roll Control System (RoCS) located on the connecting Interstage. That system will provide the Ares I with the ability to counteract induced roll torque while any induced yaw or pitch moments are handled by vectoring of the booster nozzle. This paper provides an overview of NASA s Ares I FS RoCS cold flow development test program including detailed test objectives, types of tests run to meet those objectives, an overview of the results, and applicable lessons learned. The test article was built and tested at the NASA Marshall Space Flight Center in Huntsville, AL. The FS RoCS System Development Test Article (SDTA) is a full scale, flight representative water flow test article whose primary objective was to obtain fluid system performance data to evaluate integrated system level performance characteristics and verify analytical models. Development testing and model correlation was deemed necessary as there is little historical precedent for similar large flow, pulsing systems such as the FS RoCS. The cold flow development test program consisted of flight-similar tanks, pressure regulators, and thruster valves, as well as plumbing simulating flight geometries, combined with other facility grade components and structure. Orifices downstream of the thruster valves were used to simulate the pressure drop through the thrusters. Additional primary objectives of this test program were to: evaluate system surge pressure (waterhammer) characteristics due to thruster valve operation over a range of mission duty cycles at various feed system pressures, evaluate temperature transients and heat transfer in the pressurization system, including regulator blowdown and propellant ullage performance, measure system pressure drops for comparison to analysis of tubing and components, and validate system activation and re-activation procedures for the helium pressurant system. Secondary objectives included: validating system processes for loading, unloading, and purging, validating procedures and system response for multiple failure scenarios, including relief valve operation, and evaluating system performance for contingency scenarios. The test results of the cold flow development test program are essential in validating the performance and interaction of the Roll Control System and anchoring analysis tools and results to a Critical Design Review level of fidelity.

Butt, Adam↗
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 NTRS1979

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

Space shuttle propellant dynamics during ET/Orbiter separation in the RTLS (return to launch site) mission abort sequence were investigated in a test program conducted in the NASA KC-135 "Zero G" aircraft using a 1/10th-scale model of the ET LOX Tank. Low-g parabolas were flown from which thirty tests were selected for evaluation. Data on the nature of low-g propellant reorientation in the ET LOX tank, and measurements of the forces exerted on the tank by the moving propellent will provide a basis for correlation with an analytical model of the slosh phenomenon.

Berry, R. L.↗
NASA NTRS2001

Innovative Method for Developing a Helium Pressurant Tank Suitable for the Upper Stage Flight Experiment

The AFRL USFE project is an experimental test bed for new propulsion technologies. It will utilize ambient temperature fuel and oxidizers (Kerosene and Hydrogen peroxide). The system is pressure fed, not pump fed, and will utilize a helium pressurant tank to drive the system. Mr. DeLay has developed a method for cost effectively producing a unique, large pressurant tank that is not commercially available. The pressure vessel is a layered composite structure with an electroformed metallic permeation barrier. The design/process is scalable and easily adaptable to different configurations with minimal cost in tooling development 1/3 scale tanks have already been fabricated and are scheduled for testing. The full-scale pressure vessel (50" diameter) design will be refined based on the performance of the sub-scale tank. The pressure vessels have been designed to operate at 6,000 psi. a PV/W of 1.92 million is anticipated.

DeLay, Tom K.↗
NASA NTRS1979

Test evaluation of shock buffering concept for hydrodynamic ram induced by yawing projectile impacting a simulated integral fuel tank

A concept for containing the shock inputs due to hydrodynamic ram caused by an impacting projectile within a fuel cell is discussed. This is to provide a buffering layer of foam, flexible, rigid or a combination thereof, which is sealed from the liquid. A program is described in which this buffering concept was tested. The effectiveness of a novel muzzle-mounted, 'tumble', test device is shown.

Zabel, P. H.↗
NASA NTRS1987

Design, development and test of shuttle/Centaur G-prime cryogenic tankage thermal protection systems

The thermal protection systems for the shuttle/Centaur would have had to provide fail-safe thermal protection during prelaunch, launch ascent, and on-orbit operations as well as during potential abort. The thermal protection systems selected used a helium-purged polyimide foam beneath three rediation shields for the liquid-hydrogen tank and radiation shields only for the liquid-oxygen tank (three shields on the tank sidewall and four on the aft bulkhead). A double-walled vacuum bulkhead separated the two tanks. The liquid-hydrogen tank had one 0.75-in-thick layer of foam on the forward bulkhead and two layers on the larger area sidewall. Full scale tests of the flight vehicle in a simulated shuttle cargo bay that was purged with gaseous nitrogen gave total prelaunch heating rates of 88,500 Btu/hr and 44,000 Btu/hr for the liquid-hydrogen and -oxygen tanks, respectively. Calorimeter tests on a representative sample of the liquid-hydrogen tank sidewall thermal protection system indicated that the measured unit heating rate would rapidly decrease from the prelaunch rate of approx 100 Btu/hr/sq ft to a desired rate of less than 1.3 Btu/hr/sq ft once on orbit.

Knoll, Richard H.↗
NASA NTRS2005

Millimeter Wave Detection of Localized Anomalies in the Space Shuttle External Fuel Tank Insulating Foam and Acreage Heat Tiles

The Space Shuttle Columbia's catastrophic accident emphasizes the growing need for developing and applying effective, robust and life-cycle oriented nondestructive testing (NDT) methods for inspecting the shuttle external fuel tank spray on foam insulation (SOFI) and its protective acreage heat tiles. Millimeter wave NDT techniques were one of the methods chosen for evaluating their potential for inspecting these structures. Several panels with embedded anomalies (mainly voids) were produced and tested for this purpose. Near-field and far-field millimeter wave NDT methods were used for producing millimeter wave images of the anomalies in SOFI panel and heat tiles. This paper presents the results of an investigation for the purpose of detecting localized anomalies in two SOFI panels and a set of heat tiles. To this end, reflectometers at a relatively wide range of frequencies (Ka-band (26.5 - 40 GHz) to W-band (75 - 110 GHz)) and utilizing different types of radiators were employed. The results clearly illustrate the utility of these methods for this purpose.

Kharkovsky, S.↗
NASA NTRS2003

Cryogenic Temperature-Gradient Foam/Substrate Tensile Tester

The figure shows a fixture for measuring the tensile strength of the bond between an aluminum substrate and a thermally insulating polymeric foam. The specimen is meant to be representative of insulating foam on an aluminum tank that holds a cryogenic liquid. Prior to the development of this fixture, tensile tests of this type were performed on foam/substrate specimens immersed in cryogenic fluids. Because the specimens were cooled to cryogenic temperatures throughout their thicknesses, they tended to become brittle and to fracture at loads below true bond tensile strengths. The present fixture is equipped to provide a thermal gradient from cryogenic temperature at the foam/substrate interface to room temperature on the opposite foam surface. The fixture includes an upper aluminum block at room temperature and a lower aluminum block cooled to -423 F (approx. -253 C) by use of liquid helium. In preparation for a test, the metal outer surface (the lower surface) of a foam/substrate specimen is bonded to the lower block and the foam outer surface (the upper surface) of the specimen is bonded to the upper block. In comparison with the through-the-thickness cooling of immersion testing, the cryogenic-to-room-temperature thermal gradient that exists during testing on this fixture is a more realistic approximation of the operational thermal condition of sprayed insulating foam on a tank of cryogenic liquid. Hence, tensile tests performed on this fixture provide more accurate indications of operational bond tensile strengths. In addition, the introduction of the present fixture reduces the cost of testing by reducing the amount of cryogenic liquid consumed and the time needed to cool a specimen.

Vailhe, Christophe↗
NASA NTRSDate not supplied

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

To enable the design of future in-space cryogenic propellant transfer vehicles, high accuracy models of the cryogenic propellant transfer process are desired. For in-space refueling applications, accurate simulations are desired of the propellant tank chilldown and no-vent fill (NVF) process in order to estimate the amount of propellant required, to determine the maximum sustainable flow rate to fill the tank, and to establish a timeline for the filling process. NVF is a complex process that involves numerous two-phase heat and mass transfer phenomena, such as boiling at the wall and evaporation and condensation at the liquid/vapor interface. In recent years, progress has been made in developing a numerical modeling technique using the Generalized Fluid System Simulation Program (GFSSP) that has been validated against several small-scale liquid nitrogen NVF tests. This paper presents GFSSP numerical model validation against three liquid hydrogen NVF tests on a large-scale propellant tank. Comparing the GFSSP model to the data for pressure, fill level, and wall temperature, the mean absolute percentage error is below 20% across the range of simulated conditions, indicating very good predictive agreement for this large-scale hydrogen NVF test series.

Alok Majumdar↗
NASA NTRSDate not supplied

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

To enable the design of future in-space cryogenic propellant transfer vehicles, high accuracy models of the cryogenic propellant transfer process are desired. For in-space refueling applications, accurate simulations are desired of the propellant tank chilldown and no-vent fill (NVF) process in order to estimate the amount of propellant required, to determine the maximum sustainable flow rate to fill the tank, and to establish a timeline for the filling process. NVF is a complex process that involves numerous two-phase heat and mass transfer phenomena, such as boiling at the wall and evaporation and condensation at the liquid/vapor interface. In recent years, progress has been made in developing a numerical modeling technique using the Generalized Fluid System Simulation Program (GFSSP) that has been validated against several small-scale liquid nitrogen NVF tests. This paper presents GFSSP numerical model validation against three liquid hydrogen NVF tests on a large-scale propellant tank. Comparing the GFSSP model to the data for pressure, fill level, and wall temperature, the mean absolute percentage error is below 20% across the range of simulated conditions, indicating very good predictive agreement for this large-scale hydrogen NVF test series.

Alok Majumdar↗