White Sands Test Facility

Flame-retarded epoxy composites and phenolic composites containing fiberglass, aramid (Kevlar® 49), and graphite fiber-reinforcements were tested using the NASA upward flame propagation test, the controlled-atmosphere cone calorimeter test, and the liquid oxygen (LOX) mechanical impact test. The upward flame propagation test showed that phenolic/graphite had the highest flame resistance and epoxy/graphite had the lowest flame resistance. The controlled-atmosphere cone calorimeter was used to investigate the effect of oxygen concentration and fiber reinforcement on the burning behavior of composites. The LOX mechanical impact test showed that epoxy/fiberglass had the lowest ignition resistance and phenolic/aramid had the highest ignition resistance in LOX. The composites containing epoxy resin and/or aramid fiber reinforcement reacted very violently in LOX upon mechanical impact. © 1997 by John Wiley & Sons, Ltd.

Results of metals flammability tests performed on twenty-six metals in the NASA/White Sands Test Facility are discussed together with the test systems. The promoted combustion and ignition characteristics of these metals are described, and the metals are ranked according to their suitability for use in oxygen systems. In general, alloys with high copper and nickel contents and low iron content were found to rank higher than those that had high iron content, while alloys that had high aluminum content were ranked the lowest.

One of the major obstacles slowing the acceptance of parts made by additive manufacturing (AM) in NASA applications is the lack of a broadly accepted materials and process quality systems; and more specifically, the lack of adequate nondestructive evaluation (NDE) processes integrated into AM. Matching voluntary consensus standards are also needed to control the consistency of input materials, process equipment, process methods, finished part properties, and how those properties are characterized. As for nondestructive characterization, procedures are needed to interrogate features unique to parts made by AM, such as fine-scale porosity, deeply embedded flaws, complex part geometry, and intricate internal features. The NDE methods developed must be tailored to meet materials, design and test requirements encountered throughout the part life cycle, whether during process optimization, real-time process monitoring, finished part qualification and certification (especially of flight hardware), or in situ health monitoring. Restated, individualized process/product-specific NDE methods are needed to satisfy NASA's various quality assurance requirements. To date, only limited data have been acquired by NASA on parts made by AM. This paper summarizes the NASA AM effort, highlights available NDE data, and outlines the approach NASA is taking to apply NDE to its various AM efforts.

The ignition of carbon steel and 316 and 304 stainless steels caused by the impact of low-velocity particles (a standard mixture consisting of 2 g of iron and 3 g of inert materials) in gaseous oxygen was investigated using NASA/White Sands Test Facility for the ignition test, and a subsonic particle impact chamber to accelerate the particles that were injected into flowing oxygen upstream of the target specimen. It was found that the oxygen velocities required to ignite the three alloys were the same as that required to ignite the particle mixture. Ignition occurred at oxygen velocities greater than 45 m/sec at 20 to 24 MPa and was found to be independent of pressure between 2 and 30 MPa. Comparison of the present results and the past results from Wegener (1964) with the Compressed Gas Association (CGA) oxygen velocity limits for safe operations indicates that the CGA limits may be excessively conservative at high pressures and too liberal at low pressures.

Organic combustion products generated by the lunar module descent engine, which burns a 1:1 mixture of unsymmetrical dimethylhydrazine fuel and nitrogen tetroxide oxidizer, have been analyzed. The major gaseous combustion products found were ammonia, water, carbon monoxide, nitrous oxide, oxygen, carbon dioxide, and nitric oxide. The minor products were acetylene, hydrogen cyanide, ethylene, formaldehyde, propadiene, ketene, cyanous acid, hydrazoic acid, various methylamines, acetaldehyde, methyl nitrite, formic acid, nitrous acid, butadiyne, nitrilohydrazines, nitromethane, and nitrosohydrazines with other oxidized derivatives of unsymmetrical dimethylhydrazine and hydrazine. The ion intensities of the various species in all mass spectra were estimated as the following concentrations: the gases (NH(3), H(2)O, CO, NO, O(2), CO(2), and NO(2)), 87.7 percent; compounds of C, H, and O, 6.0 percent; and compounds of C, H, and N (with traces of O), 5.8 percent.

Tests were conducted with the RS-18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA's Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to ~122,000 ft (~37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. LO2 flow ranged from 5.9 - 9.5 lbm/sec (2.7 - 4.3 kg/sec), and LCH4 flow varied from 3.0 - 4.4 lbm/sec (1.4 - 2.0 kg/sec) during the RS-18 hot-fire test series. Propellant flow rate was measured using a coriolis mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup due to two-phase flow effects. Subsequent cold-flow testing demonstrated that the propellant manifolds must be adequately flushed in order for the coriolis flow meters to give accurate data. The coriolis flow meters were later shown to provide accurate steady-state data, but the turbine flow meter data should be used in transient phases of operation. Thrust was measured using three load cells in parallel, which also provides the capability to calculate thrust vector alignment. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes. All of these objectives were met with the RS-18 data and additional testing data from subsequent LO2/methane test programs in 2009 which included the first simulated-altitude pyrotechnic ignition demonstration of LO2/methane.

SpaceX SuperDraco combustion chamber for Dragon V2 made from Inconel using the DMLS process
“Through 3D printing, robust and high-performing engine parts can be created at a fraction of the cost and time of traditional manufacturing methods,”
It’s a very complex engine, and it was very difficult to form all the cooling channels, the injector head, and the throttling mechanism. Being able to print very high strength advanced alloys ... was crucial to being able to create the SuperDraco engine as it is.

Abstract Spacecraft materials selection is based on an upward flammability test conducted in a quiescent environment at the highest expected oxygen concentration. However, NASA’s advanced space exploration program is anticipating using various habitable environments. Because limited data are available to support current program requirements, a different test logic is suggested to address the expanded atmospheric environments through the determination of materials self-extinguishment limits. This paper provides additional pressure effects data on oxygen concentration and partial pressure self-extinguishment limits under quiescent conditions. For the range of total pressures tested, the oxygen concentration and oxygen partial pressure flammability thresholds show a near linear dependence on total pressure, and appear to increase with increasing oxygen concentration (and oxygen partial pressure) thresholds. For the Constellation Program, the flammability threshold information will allow NASA to identify materials with increased flammability risk from oxygen concentration and total pressure changes, minimize potential impacts, and allow for development of sound requirements for new spacecraft and extraterrestrial landers and habitats.

Depending on input material, process method, process parameters, and post-processing, the resulting defect state in as-built and finished additive manufactured (AM) parts can be highly variable and complex. To complicate matters further, the terminology used to describe specific defect types can be archaic or user specific and is in need of global harmonization. A common understanding of the root causes of defects and the effect of defects on relevant properties continues to evolve. In powder bed processing, for example, potential defects can be very small, down to the powder particle size. Defects also can occur because of single or multiple causes. Even when there are multiple causes, single defect types can be produced that fail by a single failure mode. Alternatively, a single defect type can have several different failure modes. The objective of this paper is to classify and identify types of technologically important defects that occur in AM parts produced by powder bed fusion (PBF) and directed energy deposition (DED). A breakdown of technologically important defects is presented in three sections: the cause, the defect, and detection by nondestructive evaluation (NDE). The effect-of-defect on relevant end-use properties is addressed whenever possible. For example, the effect of lack-of-fusion flaws on ultimate tensile properties and high cycle fatigue life is discussed, thus demonstrating the need to be able to detect such flaws. Thus, although the causes of the defects occurring in PBF and DED parts can be quite different, the actual defects can have some similarities. In general, reliable detection of defects by NDE does not depend on the process cause, but depends more on the size, geometry, and location (and, potentially, the morphology) of the defect as well as the complexity, density, and surface finish of the part.

A 1,500 lbf thrust-class liquid oxygen (LO2)/Liquid Methane (LCH4) rocket engine was developed and tested at both sea-level and simulated altitude conditions. The engine was fabricated by Armadillo Aerospace (AA) in collaboration with NASA Johnson Space Center. Sea level testing was conducted at Armadillo Aerospace facilities at Caddo Mills, TX. Sea-level tests were conducted using both a static horizontal test bed and a vertical take-off and landing (VTOL) test bed capable of lift-off and hover-flight in low atmosphere conditions. The vertical test bed configuration is capable of throttling the engine valves to enable liftoff and hover-flight. Simulated altitude vacuum testing was conducted at NASA Johnson Space Center White Sands Test Facility (WSTF), which is capable of providing altitude simulation greater than 120,000 ft equivalent. The engine tests demonstrated ignition using two different methods, a gas-torch and a pyrotechnic igniter. Both gas torch and pyrotechnic ignition were demonstrated at both sea-level and vacuum conditions. The rocket engine was designed to be configured with three different nozzle configurations, including a dual-bell nozzle geometry. Dual-bell nozzle tests were conducted at WSTF and engine performance data was achieved at both ambient pressure and simulated altitude conditions. Dual-bell nozzle performance data was achieved over a range of altitude conditions from 90,000 ft to 50,000 ft altitude. Thrust and propellant mass flow rates were measured in the tests for specific impulse (Isp) and C* calculations.

Materials selection for spacecraft is based on an upward flammability test conducted in a quiescent environment in the highest expected oxygen concentration environment. The test conditions and its pass/fail test logic do not provide sufficient quantitative materials flammability information for an advanced space exploration program. A modified approach has been suggested determination of materials self-extinguishment limits. The flammability threshold information will allow NASA to identify materials with increased flammability risk from oxygen concentration and total pressure changes, minimize potential impacts, and allow for development of sound requirements for new spacecraft and extraterrestrial landers and habitats. This paper provides data on oxygen concentration self-extinguishment limits under quiescent conditions for selected materials considered for the Constellation Program.

This article presents the results of an investigation of the Raman scattering effects of aging Kevlar strands. The goal of this research was to investigate the potential application of Raman spectroscopy as a nondestructive evaluation tool for the detection of aging effects of in-service composite materials. Kevlar composites used as overwrapping of metal-lined composite overwrapped pressure vessels (COPVs) have been analyzed. Raman spectra produced from the Kevlar fibers and the effects of resin materials, utilized to bind the fibers into strands to provide composite behavior, have been investigated. A series of Kevlar/epoxy strands exposed to elevated temperatures and sustained loading have been evaluated. It was found that this exposure had an effect on the width and intensity of certain bands in the Raman spectra of Kevlar. The potential application of these findings to the nondestructive evaluation of Kevlar composites is discussed.

The sources of variability in the pneumatic impact test commonly used to screen nonmetallic materials for oxygen service and to rank batches or lots of particular materials were investigated together with the mechanism of ignition. Tests were conducted at the NASA/White Sands Test Facility in which the standard test chamber was replaced by an instrumented chamber to obtain information on the rates of pressurization, specimen heating, and time to ignition. Results showed that much of the variability in test data could be traced to the variability in system cleanliness and in the opening time of the high-speed valve. The prinicipal mechanism of ignition in the test is considered to be the heating of the gas initially in the test chamber by adiabatic compression; the presence of voids in the test materials may also be important.

Abstract A Controlled‐atomosphere cone calorimeter was used to investigate the burning of pure and flame retardant‐treated cotton fabrics. The condensed‐phase flame retardants used were Morguard (containing ammonium dihydrogen phosphate and diammonium hydrogen phoisphate) and Nochar (containing ammonium sulfate and a sodium salt). The fabrics were tested at 25 kW m −2 incident heat flux in environments containing 15–30% oxygen. The flame retardants increased the time to ignition, residue yield, and CO and CO 2 yields. The flame retardants decreased the peak and average mass loss rates, the peak and average heat release rates, the effective heat of combustion at peak heat release rate, and the propensity to flashover. The effect of oxygen concentration on the burning of pure and flame retardant‐treated cotton fabrics has also been investigated. The flame retardants had better performance when the treated fabrics burned in the lower oxyge concentrations. The result of this study indicate that the controlled‐atmosphere cone calorimeter is a good tool for studying the effect of flame retardant and oxygen concentration on the burning of materials.

The effects of convective and absolute instabilities on the formation of drops formed from cylindrical liquid jets of glycerol/water issuing into still air were investigated. Medium-duration reduced gravity tests were conducted aboard NASA’s KC-135 and compared to similar tests performed under normal gravity conditions to aid in understanding the drop formation process. In reduced gravity, the Rayleigh–Chandrasekhar Equation was found to accurately predict the transition between a region of absolute and convective instability as defined by a critical Weber number. Observations of the physics of the jet, its breakup, and subsequent drop dynamics under both gravity conditions and the effects of the two instabilities on these processes are presented. All the normal gravity liquid jets investigated, in regions of convective or absolute instability, were subject to significant stretching effects, which affected the subsequent drop and associated geometry and dynamics. These effects were not displayed in reduced gravity and, therefore, the liquid jets would form drops which took longer to form (reduction in drop frequency), larger in size, and more spherical (surface tension effects). Most observed changes, in regions of either absolute or convective instabilities, were due to a reduction in the buoyancy force and an increased importance of the surface tension force acting on the liquid contained in the jet or formed drop. Reduced gravity environments allow better investigations to be performed into the physics of liquid jets, subsequently formed drops, and the effects of instabilities on these systems. In reduced gravity, drops form up to three times more slowly and as a consequence are up to three times larger in volume in the theoretical absolute instability region than in the theoretical convective instability region. This difference was not seen in the corresponding normal gravity tests due to the masking effects of gravity. A drop is shown to be able to form and detach in a region of absolute instability, and spanning the critical Weber number (from a region of convective to absolute instability) resulted in a marked change in dynamics and geometry of the liquid jet and detaching drops.

Variations in build process parameters, post-processing parameters, and feedstock have a significant impact on the structural integrity and performance of components made with additive manufacturing (AM). Effective nondestructive testing (NDT) is critical for ensuring the structural integrity of components. Complex geometries, nonequilibrium microstructures, new process variables, and lack of clear accept or reject criteria for AM components present new challenges to NDT. Quantitative, volumetric NDT methods that can detect material defects of interest in complex geometries are required. Process compensated resonance testing (PCRT) is an NDT method that uses a swept sine input to excite the component's resonance modes of vibration. The resonance frequencies are recorded, analyzed statistically, and compared to acceptability limits established using a database of training components. The swept sine input excites whole-body vibrational modes in nearly any geometry, and the component's resonance frequencies correlate directly to its structural integrity. In this study, PCRT evaluations were performed on titanium alloy (Ti-6Al-4V) populations made with electron beam PBF and aluminum alloy (AlSi10Mg) populations made with laser PBF. The evaluations were conducted in support of ASTM round-robin testing. In the Ti-6Al-4V population, PCRT showed clear resonance frequency differences between nominal specimens and off-nominal specimens with defective material states. PCRT also quantified the effects of hot isostatic pressing (HIP). PCRT pass/fail NDT of the Ti-6Al-4V population in the pre-HIP and post-HIP states demonstrated 100% accuracy. Computed tomography scans of the post-HIP specimens showed no clear indications of porosity. Follow-up tensile testing of a subset of nominal and off-nominal specimens in the post-HIP state showed that the off-nominal specimens had lower yield stresses and ultimate tensile stresses than nominal specimens. In the AlSi10Mg population, PCRT detected differences between recycled and virgin feedstock powder. PCRT pass/fail NDT of AlSi10Mg specimens exposed to nominal and off-nominal heat treatment demonstrated 100% accuracy.

Views Icon Views Article contents Figures & tables Video Audio Supplementary Data Peer Review Share Icon Share Twitter Facebook Reddit LinkedIn Tools Icon Tools Reprints and Permissions Cite Icon Cite Search Site Citation J. M. Waller, E. Andrade, R. L. Saulsberry; USE OF ACOUSTIC EMISSION TO MONITOR PROGRESSIVE DAMAGE ACCUMULATION IN KEVLAR® 49 COMPOSITES. AIP Conf. Proc. 22 February 2010; 1211 (1): 1111–1118. https://doi.org/10.1063/1.3362168 Download citation file: Ris (Zotero) Reference Manager EasyBib Bookends Mendeley Papers EndNote RefWorks BibTex toolbar search Search Dropdown Menu toolbar search search input Search input auto suggest filter your search All ContentAIP Publishing PortfolioAIP Conference Proceedings Search Advanced Search |Citation Search

McDonnell Douglas Aerospace (MDA) and NASA are investigating the compatibility of composite materials and liquid oxygen (LO2). The objective is to determine if these advanced lightweight materials can be used to fabricate LO2 propellant tanks for launch vehicles. Rele- vant background on MDA's development of composite materials for liquid hydrogen tanks is discussed. Mate- rial qualification standards for liquid oxygen service are reviewed. Hazards analyses on composite LO*, tank designs are discussed in which various ignition mecha- nisms are considered. Results are presented from a series of tests conducted to simulate potential ignition mechanisms in the presence of LO 2 and composite materials.

Broad‐band modal acoustic emission (AE) was used to characterize micromechanical damage progression in uniaxial IM7 and T1000 carbon fiber‐epoxy (C/Ep) tows, and a helical and hoop‐wrapped IM7 composite overwrapped pressure vessel (COPV). To expedite analysis, tows and the COPV were subjected to an intermittent load hold tensile stress profile. Damage progression in tow specimens was followed by analyzing the Fast Fourier Transforms (FFTs) associated with AE events. FFT analysis showed that damage was usually cooperative, consisting of several failure modes occurring at once, and was dominated by fiber breakage throughout the duration of the stress profile. Evidence was found for the existence of a universal damage parameter, referred to here as the critical Felicity ratio, or Felicity ratio at rupture (FR*), which had a value close to 0.96 for the tows and the COPV tested. The use of FR* to predict the burst pressure of the COPV is demonstrated.

A ground test article was prepared for RS-18 engine testing using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA Glenn Research Center’s Propulsion and Cryogenics Advanced Development (PCAD) project. Green propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propellant systems in lunar vehicle ascent engine applications. LO2/LCH4 testing capability at altitude conditions did not previously exist at WSTF for this size engine, and modifications were made to the Auxiliary Propulsion Systems Test Bed (APSTB) article. Altitude simulation is achieved using the WSTF Large Altitude Simulation System, which provides altitude conditions equivalent to ~90,000 ft (~27 km). For specific impulse calculations, engine thrust and propellant mass flow rates are measured. Propellant flow rate is measured using a coriolis-style mass-flow meter, and accuracy is compared with a serial turbine-style flow meter. Thrust is measured using three load cells in parallel. Igniter system capability is being developed to demonstrate two methods, a gaseous oxygen/methane spark torch igniter and solid propellant pyrotechnic igniter. Design, procurement and assembly are complete for the test article and test readiness is expected for hot-fires to begin pending completion of manifold buildup and system checkout. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/LCH4 lunar ascent engines, 2) obtain nozzle kinetics data to anchor twodimensional kinetics codes, and 3) obtain vacuum ignition data for the torch and pyrotechnic igniters.