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NASA NTRS1995

High-Density Amorphous Ice, the Frost on Interstellar Grains

Most water ice in the universe is in a form which does not occur naturally on Earth and of which only minimal amounts have been made in the laboratory. We have encountered this 'high-density amorphous ice' in electron diffraction experiments of low-temperature (T less than 30 K) vapor-deposited water and have subsequently modeled its structure using molecular dynamics simulations. The characteristic feature of high-density amorphous ice is the presence of 'interstitial' oxygen pair distances between 3 and 4 A. However, we find that the structure is best described as a collapsed lattice of the more familiar low-density amorphous form. These distortions are frozen in at temperatures below 38 K because, we propose, it requires the breaking of one hydrogen bond, on average, per molecule to relieve the strain and to restructure the lattice to that of low-density amorphous ice. Several features of astrophysical ice analogs studied in laboratory experiments are readily explained by the structural transition from high-density amorphous ice into low-density amorphous ice. Changes in the shape of the 3.07 gm water band, trapping efficiency of CO, CO loss, changes in the CO band structure, and the recombination of radicals induced by low-temperature UV photolysis all covary with structural changes that occur in the ice during this amorphous to amorphous transition. While the 3.07 micrometers ice band in various astronomical environments can be modeled with spectra of simple mixtures of amorphous and crystalline forms, the contribution of the high-density amorphous form nearly always dominates.

Jenniskens, P.↗

NASA NTRS1994

Influence of Two-Phase Thermocapillary Flow on Cryogenic Liquid Retention in Microscopic Pores

Previous experiments indicate that the bubble point pressure of spacecraft liquid hydrogen acquisition devices is reduced substantially when the ullage is pressurized with heated hydrogen vapor. The objective is to determine whether the two-phase thermocapillary convection arising from thermodynamic non-equilibrium along the porous surfaces of such devices could lead to this observed degradation in retention performance. We also examine why retention capability appears to be unaffected by pressurization with heated helium or direct heating through the porous structure. Computational assessments based on coupled solution of the flowfield and liquid free surface indicate that for highly wetting fluids in small pores, dynamic pressure and vapor recoil dictate surface morphology and drive meniscus deformation. With superheating, the two terms exert the same influence on curvature and promote mechanical equilibrium, but with subcooling, the pressure distribution produces a suction about the pore center-line that degrades retention. This result points to thermocapillary-induced deformation arising from condensation as the cause for retention loss. It also indicates that increasing the level of non-equilibrium by reducing accommodation coefficient restricts deformation and explains why retention failure does not occur with direct screen heating or helium pressurization.

Schmidt, G. R.↗

NASA NTRS1994

Ways to Increase Launch Velocities of 2-Stage Gas Guns

The amount of space debris is rapidly increasing and the debris is distributed over a wide variety of orbits. Satellites, manned space vehicles and space stations will have to pay increasing attention to the dangers of impacts with space debris. Various armoring techniques (i.e., double or triple layer armor) will have to tested extensively to determine the most effective armor per unit weight. Intersecting near-earth orbits can lead to impact velocities up to 15 km/sec. Conventional two-stage light gas guns can launch intact, controlled-shape projectiles with a density of 1.2 gm/cc and length- to-diameter ratios of 0.5-1.0 at velocities up to 8-9 km/sec. Higher velocities (10-11 km/sec) can be obtained' for very light projectiles. The higher launch velocities tend to be very severe on the high pressure coupling and barrel of the gun and lead to short component lifetimes. Clearly, the ability to raise the launch velocity of a gun (for reasonably massive projectile shapes) from 8-9 km/sec to 11-13 km/sec (or higher), without reduction of component lifetimes, would have significant benefits. This would allow much better simulation of the higher velocity debris impacts as well as better simulation of high speed re-entry into planetary atmospheres. Several techniques for increasing the launcher muzzle velocity above 8-9 km/sec have been studied using CFD simulations and appear to offer the potential for significant gains. The first technique is to use multiple compressions, instead of a single compression, in the pump tube of the light gas gun. In a sense, this is a kind of pre-heating of the gas in the pump tube; other types of pre-heating have yielded disappointing results in the past. The dynamics of the multiple compression pump tube is very different, however, from the earlier techniques, where the pump tube was typically heated ohmically before the gun 2 cycle was started. In this paper, we present CFD calculations that show that significant increases in muzzle velocity can be obtained with multiple compressions in the pump tube. With a conventional two-stage gun, an important limitation to obtaining higher velocities is friction and heat transfer to the barrel, which typically has a length- to- diameter ratio of 200-400. These viscous losses greatly reduce the effectiveness of the regions of the barrel far removed from the second stage breech. We have studied computationally the effect of adding an additional breech (or breeches) along the barrel to reduce these viscous losses. Velocity increases from 6.5 to 7.2 km/sec have been obtained using the main breech and one additional breech. In these results, both breeches were operated with hydrogen, heated electrothermally. We have also studied a gun geometry where the main breech is operated in the conventional manner, using piston compression. The additional breech is operated either with electrothermal heating or heating by using a high explosive charge in a novel geometry. The latter option provides very effective compression, heating and acceleration of the hydrogen working gas and is fully reusable. Calculations are presented which show that very substantial increases in muzzle velocity can be obtained this way, without overstressing the projectile or the 'gun. The third technique studied is to add a section of ram accelerator tube after the barrel to further accelerate the projectile. The ram accelerator used here is not the conventional premixed gas ram accelerator, but a new technique using high explosive as the energy source and pure hydrogen as the working gas in a geometry which can be made fully reusable. Preliminary results with this new rain accelerator geometry were presented and showed that stable ram accelerator drive can be established. Herein, detailed calculations axe presented which show that substantial velocity increases can be obtained using this ram accelerator technique in tandem with a conventional light gas gun.

Bogdanoff, David W.↗

NASA NTRS2004

Characterization of Space Shuttle External Tank Thermal Protection System (TPS) Materials in Support of the Columbia Accident Investigation

NASA suffered the loss of the seven-member crew of the Space Shuttle Columbia on February 1, 2003 when the vehicle broke apart upon re-entry to the Earth's atmosphere. The final report of the Columbia Accident Investigation Board (CAIB) determined that the accident was caused by a launch ascent incident-a suitcase-sized chunk of insulating foam on the Shuttle's External Tank (ET) broke off, and moving at almost 500 mph, struck an area of the leading edge of the Shuttle s left wing. As a result, one or more of the protective Reinforced Carbon-Carbon (RCC) panels on the wing leading edge were damaged. Upon re-entry, superheated air approaching 3,000 F breached the wing damage and caused the vehicle breakup and loss of crew. The large chunk of insulating foam that broke off during the Columbia launch was determined to come from the so-called bipod ramp area where the Shuttle s orbiter (containing crew) is attached to the ET. Underneath the foam in the bipod ramp area is a layer of TPS that is a cork-filled silicone rubber composite. In March 2003, the NASA Marshall Space Flight Center (MSFC) in Huntsville, Alabama received cured samples of the foam and composite for testing from the Michoud Assembly Facility (MAF) in New Orleans, Louisiana. The MAF is where the Shuttle's ET is manufactured. The foam and composite TPS materials for the ET have been well characterized for mechanical property data at the super-cold temperatures of the liquid oxygen and hydrogen fuels used in the ET. However, modulus data on these materials is not as well characterized. The TA Instruments 2980 Dynamic Mechanical Analyzer (DMA) was used to determine the modulus of the two TPS materials over a range of -145 to 95 C in the dual cantilever bending mode. Multi-strain, fixed frequency DMA tests were followed by multi-frequency, fixed strain tests to determine the approximate bounds of linear viscoelastic behavior for the two materials. Additional information is included in the original extended abstract.

Wingard, Charles D.↗

NASA NTRS1989

Time-dependent computational studies of flames in microgravity

The research performed at the Center for Reactive Flow and Dynamical Systems in the Laboratory for Computational Physics and Fluid Dynamics, at the Naval Research Laboratory, in support of the NASA Microgravity Science and Applications Program is described. The primary focus was on investigating fundamental questions concerning the propagation and extinction of premixed flames in Earth gravity and in microgravity environments. The approach was to use detailed time-dependent, multispecies, numerical models as tools to simulate flames in different gravity environments. The models include a detailed chemical kinetics mechanism consisting of elementary reactions among the eight reactive species involved in hydrogen combustion, coupled to algorithms for convection, thermal conduction, viscosity, molecular and thermal diffusion, and external forces. The external force, gravity, can be put in any direction relative to flame propagation and can have a range of values. A combination of one-dimensional and two-dimensional simulations was used to investigate the effects of curvature and dilution on ignition and propagation of flames, to help resolve fundamental questions on the existence of flammability limits when there are no external losses or buoyancy forces in the system, to understand the mechanism leading to cellular instability, and to study the effects of gravity on the transition to cellular structure. A flame in a microgravity environment can be extinguished without external losses, and the mechanism leading to cellular structure is not preferential diffusion but a thermo-diffusive instability. The simulations have also lead to a better understanding of the interactions between buoyancy forces and the processes leading to thermo-diffusive instability.

Oran, Elaine S.↗

NASA NTRS1995

Unsteady numerical simulations of the stability and dynamics of flames

In this report we describe the research performed at the Naval Research Laboratory in support of the NASA Microgravity Science and Applications Program over the past three years (from Feb. 1992) with emphasis on the work performed since the last microgravity combustion workshop. The primary objective of our research is to develop an understanding of the differences in the structure, stability, dynamics and extinction of flames in earth gravity and in microgravity environments. Numerical simulations, in which the various physical and chemical processes can be independently controlled, can significantly advance our understanding of these differences. Therefore, our approach is to use detailed time-dependent, multi-dimensional, multispecies numerical models to perform carefully designed computational experiments. The basic issues we have addressed, a general description of the numerical approach, and a summary of the results are described in this report. More detailed discussions are available in the papers published which are referenced herein. Some of the basic issues we have addressed recently are (1) the relative importance of wall losses and gravity on the extinguishment of downward-propagating flames; (2) the role of hydrodynamic instabilities in the formation of cellular flames; (3) effects of gravity on burner-stabilized flames, and (4) effects of radiative losses and chemical-kinetics on flames near flammability limits. We have also expanded our efforts to include hydrocarbon flames in addition to hydrogen flames and to perform simulations in support of other on-going efforts in the microgravity combustion sciences program. Modeling hydrocarbon flames typically involves a larger number of species and a much larger number of reactions when compared to hydrogen. In addition, more complex radiation models may also be needed. In order to efficiently compute such complex flames recent developments in parallel computing have been utilized to develop a state-of-the-art parallel flame code. This is discussed below in some detail after a brief discussion of the numerical models.

Kailasanath, K.↗

NASA NTRS1992

Turbine blade tip and seal clearance excitation forces

Experimental and theoretical work done as Phase 3 of a program sponsored by MSFC to investigate the magnitude, origin, and parametric variations of destabilizing forces which arise in high power turbines due to blade-tip leakage effects are described. The two facilities which were built for this purpose are first described. The larger one is a closed, 2 atm pressurized Freon-12 flow loop in which is installed a 1:1 replica of the SSME first stage hydrogen turbine, which can be driven by the flow, and which generates about 14 KW of power into a load-absorbing DC generator. The smaller facility is used to measure the forces on labyrinth seals of the same type as those used in our turbine tests with a shrouded turbine. The seals can be kinematically whirled and spun (independently), and the inlet swirl can be set to a variety of values. Air is the working fluid (with atmospheric discharge) and the data are real-time pressure distributions in the seal glands. The five different unshrouded turbine configurations were tested with static offsets, plus one with a shroud band and a two-ridge seal. Theoretical models of various degrees of complexity were developed to help interpreting and extrapolating the data. The notion of partial work done by the fluid leaking through the tip gaps was put on a quantitative basis by examining the leakage vortex roll-up dynamics. This was used to obtain a theory of the work loss due to a uniform gap. Perturbation and multiple scale arguments were then used to extend this to the case of an eccentric turbine. This yields an unsteady, 3-D theory which can predict the distribution of the approach flow, and its effect on work defect, cross-forces, pressure patterns, and dynamic damping. The predictions agree qualitatively with the data and exhibit the correct trends, but the cross-forces are generally under-predicted.

Martinez-Sanchez, M.↗

NASA NTRS2001

Pulse Detonation Engine Modeled

Pulse Detonation Engine Technology is currently being investigated at Glenn for both airbreathing and rocket propulsion applications. The potential for both mechanical simplicity and high efficiency due to the inherent near-constant-volume combustion process, may make Pulse Detonation Engines (PDE's) well suited for a number of mission profiles. Assessment of PDE cycles requires a simulation capability that is both fast and accurate. It should capture the essential physics of the system, yet run at speeds that allow parametric analysis. A quasi-one-dimensional, computational-fluid-dynamics-based simulation has been developed that may meet these requirements. The Euler equations of mass, momentum, and energy have been used along with a single reactive species transport equation, and submodels to account for dominant loss mechanisms (e.g., viscous losses, heat transfer, and valving) to successfully simulate PDE cycles. A high-resolution numerical integration scheme was chosen to capture the discontinuities associated with detonation, and robust boundary condition procedures were incorporated to accommodate flow reversals that may arise during a given cycle. The accompanying graphs compare experimentally measured and computed performance over a range of operating conditions for a particular PDE. Experimental data were supplied by Fred Schauer and Jeff Stutrud from the Air Force Research Laboratory at Wright-Patterson AFB and by Royce Bradley from Innovative Scientific Solutions, Inc. The left graph shows thrust and specific impulse, Isp, as functions of equivalence ratio for a PDE cycle in which the tube is completely filled with a detonable hydrogen/air mixture. The right graph shows thrust and specific impulse as functions of the fraction of the tube that is filled with a stoichiometric mixture of hydrogen and air. For both figures, the operating frequency was 16 Hz. The agreement between measured and computed values is quite good, both in terms of trend and magnitude. The error is under 10 percent everywhere except for the thrust value at an equivalence ratio of 0.8 in the left figure, where it is 14 percent. The simulation results shown were made using 200 numerical cells. Each cycle of the engine, approximately 0.06 sec, required 2.0 min of CPU time on a Sun Ultra2. The simulation is currently being used to analyze existing experiments, design new experiments, and predict performance in propulsion concepts where the PDE is a component (e.g., hybrid engines and combined cycles).

Paxson, Daniel E.↗

NASA NTRS2014

Nuclear Thermal Propulsion Ground Test History

Nuclear Thermal Propulsion (NTP) was started in ~1955 under the Atomic Energy Commission as project Rover and was assigned to Los Alamos National Laboratory. The Nevada Test Site was selected in 1956 and facility construction began in 1957. The KIWI-A was tested on July 1, 1959 for 5 minutes at 70MW. KIWI-A1 was tested on July 8, 1960 for 6 minutes at 85MW. KIWI-A3 was tested on October 10, 1960 for 5 minutes at 100MW. The National Aeronautics and Space Administration (NASA) was formed in 1958. On August 31, 1960 the AEC and NASA established the Space Nuclear Propulsion Office and named Harold Finger as Director. Immediately following the formation of SNPO, contracts were awarded for the Reactor In Flight Test (RIFT), master plan for the Nuclear Rocket Engine Development Station (NRDS), and the Nuclear Engine for Rocket Vehicle Application (NERVA). From December 7, 1961 to November 30, 1962, the KIWI-B1A, KIWI-B1B, and KIWI-B4A were tested at test cell A. The last two engines were only tested for several seconds before noticeable failure of the fuel elements. Harold Finger called a stop to any further hot fire testing until the problem was well understood. The KIWI-B4A cold flow test showed the problem to be related to fluid dynamics of hydrogen interstitial flow causing fuel element vibrations. President Kennedy visited the NTS one week after the KIWI-B4A failure and got to see the engine starting to be disassembled in the maintenance facility. The KIWI-B4D and KIWI-B4E were modified to not have the vibration problems and were tested in test cell C. The NERVA NRX program started testing in early 1964 with NRX-A1 cold flow test series (unfueled graphite core), NRX-A2 and NRX-A3 power test series up to 1122 MW for 13 minutes. In March 1966, the NRX-EST (Engine System Test) was the first breadboard using flight functional relationship and total operating time of 116 minutes. The NRX-EST demonstrated the feasibility of a hot bleed cycle. The NRX-A5 had multiple start-ups in May-June 1966 with 30.75 minutes accumulative operating time at or above 1GW. The NRX-A6 was tested in December 1969 and ran for 62 minutes at 1100 MW. Each engine had post-test examination and found various structure anomalies which were identified for correction and the fuel element corrosion rate was reduced. The Phoebus series of research reactors began testing at test cell C, in June 1965 with Phoebus 1A. Phoebus 1A operated for 10.5 minutes at 1100 MW before unexpected loss of propellant and leading to an engine breakdown. Phoebus 1B ran for 30 minutes in February of 1967. Phoebus 2A was the highest steady state reactor built at 5GW. Phoebus 2A ran for 12 minutes at 4100 MW demonstrating sufficient power is available. The Peewee test bed reactor was tested November- December 1968 in test cell C for 40 minutes at 500MW with overall performance close to pre-run predictions. The XE' engine was the only engine tested with close to a flight configuration and fired downward into a diffuser at the Engine Test Stand (ETS) in 1969. The XE' was 1100 MW and had ~28 start-ups. The nuclear furnace NF-1 was operated at 44 MW with multiple test runs at 90 minutes in the summer of 1972. The NF-1 was the last NTP reactor tested. The Rover/NERVA program was cancelled in 1973. However, before cancellation, a lot of other engineering work was conducted by Aerojet on a 75, 000 lbf prototype flight engine and by Los Alamos on a ~16,000 lbf "Small Engine" nuclear rocket design. The ground test history of NTP at the NRDS also offers many lessons learned on how best to setup, operate, emergency shutdown, and post-test examine NTP engines. The reactor and engine maintenance and disassembly facilities were used for assembly and inspection of radioactive engines after testing. Most reactor/ engines were run at test cell A or test cell C with open air exhaust. The Rover/NERVA program became aware of a new environmental regulation that would restrict the amount of radioactive particulates allowed to be release in open air and successfully demonstrated a scrubber concept with the NF-1. The ETS stand was the only one with a high altitude test chamber used for XE'. The ETS and other test cells showed the effects the engine's radiation had on the facility materials and instrumentation as well as side effects the ground test facility has back on the engine operation. The breakdown of Phoebus 1A at test cell C showed how the site was cleaned up and back to operation for five more engines before the program was cancelled.

Gerrish, Harold P.↗

NASA NTRS2006

Nitrile/Buna N Material Failure Assessment for an O-Ring used on the Gaseous Hydrogen Flow Control Valve (FCV) of the Space Shuttle Main Engine

After the rollout of Space Shuttle Discovery in April 2005 in preparation for return-to-flight, there was a failure of the Orbiter (OV-103) helium signature leak test in the gaseous hydrogen (GH2) system. Leakage was attributed to the Flow Control Valve (FCV) in Main Engine 3. The FCV determined to be the source of the leak for OV-103 is designated as LV-58. The nitrile/Buna N rubber O-ring seal was removed from LV-58, and failure analysis indicated radial cracks providing leak paths in one quadrant. Cracks were eventually found in 6 of 9 FCV O-rings among the three Shuttle Orbiters, though none were as severe as those for LV-58, OV-103. Testing by EM10 at MSFC on all 9 FCV O- rings included: laser dimensional, Shore A hardness and properties from a dynamic mechanical analyzer (DMA) and an Instron tensile machine. The following test data was obtained on the cracked quadrant of the LV-58, OV-103 O-ring: (1) the estimated compression set was only 9.5%, compared to none for the rest of the O-ring; (2) Shore A hardness for the O.D. was higher by almost 4 durometer points than for the rest of the O-ring; and (3) DMA data showed that the storage/elastic modulus E was almost 25% lower than for the rest of the O-ring. Of the 8 FCV O-rings tested on an Instron, 4 yielded tensile strengths that were below the MIL spec requirement of 1350 psi-a likely influence of rubber cracking. Comparisons were made between values of modulus determined by DNA (elastic) and Instron (Young s). Each nitrile/Buna N O-ring used in the FCV conforms to the MIL-P-25732C specification. A number of such O-rings taken from shelf storage at MSFC and Kennedy Space Center (KSC) were used to generate a reference curve of DMA glass transition temperature (Tg) vs. shelf storage time ranging from 8 to 26 years. A similar reference curve of TGA onset temperature (of rubber weight loss) vs. shelf storage time was also generated. The DMA and TGA data for the used FCV O-rings were compared to the reference curves. Correlations were also made between the DMA modulus (at 22 C) and Shore A hardness for all 9 of the FCV O-rings used among the three Shuttle Orbiters. The radial cracking in the FCV O-rings was determined to be due to ozone attack, as nitrile/Buna N rubber is susceptible to such attack. Nitrile/Buna N material under MIL-P25732C should be used in a hydraulic fluid environment to help protect it from cracking. However, the FCV O-rings were used in an air only environment. The FCV design has as much as a 9-mil gap that allows the O.D. of the O-ring to be directly exposed to ozone, pressurized air and some elevated temperatures, accelerating the weathering process that leads to O-ring cracking. Space Shuttle flights will likely not continue past 2010. Therefore, Shuttle management decided to continue using the nitrile/Buna N material for the FCVs, but have each O-ring replaced after 3 years to minimize any chances for crack initiation.

Wingard, Doug↗