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All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

An Interim Set of TNT Curves for LOX/LNG Explosions

Space launch companies are actively developing, or in some cases have already developed, new vehicles that use large quantities of liquid oxygen (LOX) and liquid natural gas (LNG) propellants. This propellant mixture currently lacks sound/verified explosive safety standards for use in explosive siting and flight safety analysis. In the U.S., multiple government and commercial organizations have conducted limited testing, or are preparing to conduct tests, intended to provide a sound technical basis for LOX/LNG explosive standards applicable to space launch vehicle ground and flight operations. The limited testing and analysis conducted so far indicates the potential to produce extremely energetic explosions due to the miscibility of LOX/LNG (methane), a unique feature relative to previously used propellant mixtures such as hydrogen/oxygen or kerosene/oxygen. MOX is a homogenous mixture of methane and oxygen that is possible because of methane’s 100% miscibility in LOX. Preliminary testing by N2L, Inc. described in this paper shows that it is possible to form MOX with a variety of mixing methods, and that MOX is a sensitive high explosive capable of producing overpressures greater than comparable masses of C-4. The limited amount and quality of large-scale LOX/LNG explosive test data, and the potential formation of high-explosive MOX, create significant unknowns in the determination of hazard areas (cleared of the public) for ground operations (such as a static fire test) and flights of launch vehicles with LOX/LNG propellant. This paper will review the methods used by NASA to develop and validate the LOX/LH2 blast model using large-scale explosive test programs such as Project PYRO, the Hydrogen-Oxygen Vertical Impact (HOVI), and the Large-Scale Hydrogen-Oxygen Explosion (LSHOE) tests. This paper will describe how the same process was used to prepare an interim LOX/LNG blast model that applies to various LV failure scenarios and conservatively accounts for potential MOX formation. This paper also summarizes past and future testing and modeling efforts funded by a consortium of NASA, the US Federal Aviation Administration (FAA), and the US Space Force (USSF). These test programs should be concluded within the next 3 years and are intended to provide empirical data for model verification and validation.

Liquid Oxygen↗
NASA NTRSDate not supplied

An Interim Set of TNT Curves for LOX/LNG Explosions

Space launch companies are actively developing, or in some cases have already developed, new vehicles that use large quantities of liquid oxygen (LOX) and liquid natural gas (LNG) propellants. This propellant mixture currently lacks sound/verified LOX/LNG explosive safety standards for use in explosive siting and flight safety analysis. In the U.S., multiple government and commercial organizations have conducted limited testing, or are preparing to conduct tests, intended to provide a sound technical basis for explosive standards applicable to space launch vehicle ground and flight operations. The limited testing and analysis conducted so far indicates the potential to produce extremely energetic explosions due to the miscibility of LOX/LNG (methane), a unique feature relative to previously used propellant mixtures such as hydrogen/oxygen or kerosene/oxygen. MOX is a homogenous mixture of methane and oxygen that is possible because of methane’s 100% miscibility in LOX. Preliminary testing by N2L, Inc. described in this paper shows that it is possible to form MOX with a variety of mixing methods, and that MOX is a sensitive high explosive capable of producing overpressures greater than comparable masses of C-4. The limited amount and quality of large-scale LOX/LNG explosive test data, and the potential formation of high-explosive MOX, create significant unknowns in the determination of hazard areas (cleared of the public) for ground operations (such as a static fire test) and flights of launch vehicles with LOX/LNG propellant. This paper will review the methods used by NASA to develop and validate the LOX/LH2 blast model using large-scale explosive test programs such as Project PYRO, the Hydrogen-Oxygen Vertical Impact (HOVI), and the Large-Scale Hydrogen-Oxygen Explosion (LSHOE) tests. This paper will describe how the same process was used to prepare an interim LOX/LNG blast model that applies to various LV failure scenarios and conservatively accounts for potential MOX formation. This paper also summarizes past and future testing and modeling efforts funded by a consortium of NASA, the US Federal Aviation Administration (FAA), and the US Space Force (USSF). These test programs should be concluded within the next 3 years and are intended to provide empirical data for model verification and validation.

Liquid Oxygen↗
NASA NTRS1983

High speed cryogenic self-acting, shaft seals for liquid rocket turbopumps

Three self acting lift pad liquid oxygen face seals and two self acting gaseous helium circumferential seals for high speed liquid oxygen turbopump were evaluated. The development of a technology for reliable, 10 hour life, multiple start seals for use in high speed liquid oxygen turbopumps is discussed.

Burcham, R. E.↗
NASA NTRS1987

Support of the eight-foot high temperature tunnel modification project

In order to meet the need for propulsion testing in the high supersonic range from Mach 4 to Mach 7, NASA has undertaken the modification of the Langley Eight Foot High Temperature Tunnel to add alternate Mach number capability and oxygen enrichment to allow the testing of operating engines at these Mach numbers and at true temperature tunnel. The transfer of liquid oxygen (LOX) from a storage vessel to a rocket engine generally requires the use of a pressurizing gas at high pressures. Although nitrogen is preferred, unfortunately, when gaseous nitrogen (GN2) is used as the pressurant to transfer liquid oxygen from a storage tank to the tunnel combustor, it contaminates the liquid oxygen and effects a loss of performance in the engine. The contamination of the LOX by the pressurizing GN2 is described, which may prove to be an important operational constraint. It is desirable to have reliable data concerning the penetration of GN2 into LOX during pressurization and the subsequent of self cleaning after blowdown.

Ngo, Kim Chi↗
NASA NTRS2009

Experimental Results of Integrated Refrigeration and Storage System Testing

Launch operations engineers at the Kennedy Space Center have identified an Integrated Refrigeration and Storage system as a promising technology to reduce launch costs and enable advanced cryogenic operations. This system uses a close cycle Brayton refrigerator to remove energy from the stored cryogenic propellant. This allows for the potential of a zero loss storage and transfer system, as well and control of the state of the propellant through densification or re-liquefaction. However, the behavior of the fluid in this type of system is different than typical cryogenic behavior, and there will be a learning curve associated with its use. A 400 liter research cryostat has been designed, fabricated and delivered to KSC to test the thermo fluid behavior of liquid oxygen as energy is removed from the cryogen by a simulated DC cycle cryocooler. Results of the initial testing phase focusing on heat exchanger characterization and zero loss storage operations using liquid oxygen are presented in this paper. Future plans for testing of oxygen densification tests and oxygen liquefaction tests will also be discussed. KEYWORDS: Liquid Oxygen, Refrigeration, Storage

Notardonato, W. U.↗
NASA NTRS2007

Cyanobacteria for Human Habitation beyond Earth

In light of the President s Moon/Mars initiative, lunar exploration has once again become a priority for NASA. In order to establish permanent bases on the Moon and proceed with human exploration of Mars, two key problems will be addressed: first, the production of O2 and second, the production of methane (CH4). While O2 is required for life support systems (LSS), both liquid O2 and CH4 are needed as an oxidizer and a propellant, respectively for the Lunar Surface Access Module (LSAM) and the Crew Exploration Vehicle (CEV). Unlike previous propulsion systems, the new CEV will use liquid oxygen (LO2) as an oxidizer and liquid methane (LCH4) as a propellant. Existing technology (e.g. hydrogen reduction) for the production of liquid oxygen from lunar regolith is very energy intensive and requires high temperature reactors. We propose an alternative approach using iron-tolerant cyanobacteria. We have found that iron-tolerant cyanobacteria (IT CB) are capable of etching iron-bearing minerals, which may lead to bonds breaking between Fe and O of common lunar mare basalt Fe-oxides including ilmenite, pseudobrookite, ferropseudobrookite, and armalcolite with the subsequent release of both Fe, Ti and oxygen as byproducts. We also propose to use CB biomass for CH4 production as carbon stock and a propellant. Both processes can be accomplished in an energy and cost effective manner because sunlight will be used as an energy source and allows the reactions at ambient temperatures between 10-60 C. Current evaluations include assessing the thermodynamics of such biogenic reactions using a variety of nutrients and atmospheric parameters, as well as assessing the rates and species variation effects of the driving reactions.

Brown, Igor↗
NASA NTRS2006

Iron-Tolerant Cyanobacteria for Human Habitation beyond Earth

In light of the President's Moon/Mars initiative, lunar exploration has once again become a priority for NASA. In order to establish permanent bases on the Moon and proceed with human exploration of Mars, two key problems will be addressed: first, the production of O2 and second, the production of methane (CH4). While O2 is required for life support systems (LSS), both liquid O2 and CH4 are needed as an oxidizer and a propellant, respectively for the Lunar Surface Access Module (LSAM) and the Crew Exploration Vehicle (CEV). Unlike previous propulsion systems, the new CEV will use liquid oxygen (LO2) as an oxidizer and liquid methane (LCH4) as a propellant. Existing technology (e.g. hydrogen reduction) for the production of liquid oxygen from lunar regolith is very energy intensive and requires high temperature reactors. We propose an alternative approach using iron-tolerant cyanobacteria. We have found that iron-tolerant cyanobacteria (IT CB) are capable of etching iron-bearing minerals, which may lead to bonds breaking between Fe and O of common lunar mare basalt Feoxides including ilmenite, pseudobrookite, ferropseudobrookite, and armalcolite with the subsequent release of both Fe, Ti and oxygen as by-products. We also propose to use CB biomass for CH4 production as carbon stock and a propellant. Both processes can be accomplished in an energy and cost effective manner because sunlight will be used as an energy source and allows the reactions at ambient temperatures between 10-60 C. Current evaluations include assessing the thermodynamics of such biogenic reactions using a variety of nutrients and atmospheric parameters, as well as assessing the rates and species variation effects of the driving reactions.

Brown, Igor↗
NASA NTRS1989

Design and test of an oxygen turbopump for a dual expander cycle rocket engine

A liquid oxygen (LOX) turbopump with an 860 R gaseous oxygen (GOX) turbine drive was designed for a 3750 lb thrust dual expander cycle rocket engine. This turbopump, which requires no interpropellant seals or system purges, features a 156 hp, single stage, full admission, impulse turbine; an axial flow inducer; a two-stage centrifugal pump with unshrouded impellers; long-life, LOX-lubricated, self-aligning, hydrostatic bearings; and a subcritical rotor design. It is constructed of Monel, a nickel-copper alloy, which has low ignition potential in oxygen. The pump was designed to deliver 34.7 gpm of 4655 psia liquid oxygen at a shaft speed of 75,000 rpm. The dual expander cycle rocket engine and the performance it requires of the LOX turbopump will be discussed as well as the design of the pump, turbine, bearings, and the turbopump rotordynamics. The test program and preliminary test results will also be presented.

Buckmann, P. S.↗
NASA NTRS2011

Catalytic Microtube Rocket Igniter

Devices that generate both high energy and high temperature are required to ignite reliably the propellant mixtures in combustion chambers like those present in rockets and other combustion systems. This catalytic microtube rocket igniter generates these conditions with a small, catalysis-based torch. While traditional spark plug systems can require anywhere from 50 W to multiple kW of power in different applications, this system has demonstrated ignition at less than 25 W. Reactants are fed to the igniter from the same tanks that feed the reactants to the rest of the rocket or combustion system. While this specific igniter was originally designed for liquid methane and liquid oxygen rockets, it can be easily operated with gaseous propellants or modified for hydrogen use in commercial combustion devices. For the present cryogenic propellant rocket case, the main propellant tanks liquid oxygen and liquid methane, respectively are regulated and split into different systems for the individual stages of the rocket and igniter. As the catalyst requires a gas phase for reaction, either the stored boil-off of the tanks can be used directly or one stream each of fuel and oxidizer can go through a heat exchanger/vaporizer that turns the liquid propellants into a gaseous form. For commercial applications, where the reactants are stored as gases, the system is simplified. The resulting gas-phase streams of fuel and oxidizer are then further divided for the individual components of the igniter. One stream each of the fuel and oxidizer is introduced to a mixing bottle/apparatus where they are mixed to a fuel-rich composition with an O/F mass-based mixture ratio of under 1.0. This premixed flow then feeds into the catalytic microtube device. The total flow is on the order of 0.01 g/s. The microtube device is composed of a pair of sub-millimeter diameter platinum tubes connected only at the outlet so that the two outlet flows are parallel to each other. The tubes are each approximately 10 cm long and are heated via direct electric resistive heating. This heating brings the gasses to their minimum required ignition temperature, which is lower than the auto-thermal ignition temperature, and causes the onset of both surface and gas phase ignition producing hot temperatures and a highly reacting flame. The combustion products from the catalytic tubes, which are below the melting point of platinum, are injected into the center of another combustion stage, called the primary augmenter. The reactants for this combustion stage come from the same source but the flows of non-premixed methane and oxygen gas are split off to a secondary mixing apparatus and can be mixed in a near-stoichiometric to highly lean mixture ratio. The primary augmenter is a component that has channels venting this mixed gas to impinge on each other in the center of the augmenter, perpendicular to the flow from the catalyst. The total crosssectional area of these channels is on a similar order as that of the catalyst. The augmenter has internal channels that act as a manifold to distribute equally the gas to the inward-venting channels. This stage creates a stable flame kernel as its flows, which are on the order of 0.01 g/s, are ignited by the combustion products of the catalyst. This stage is designed to produce combustion products in the flame kernel that exceed the autothermal ignition temperature of oxygen and methane.

Schneider, Steven J.↗
NASA NTRS1997

Performance Tests of a Liquid Hydrogen Propellant Densification Ground System for the X33/RLV

A concept for improving the performance of propulsion systems in expendable and single-stage-to-orbit (SSTO) launch vehicles much like the X33/RLV has been identified. The approach is to utilize densified cryogenic liquid hydrogen (LH2) and liquid oxygen (LOX) propellants to fuel the propulsion stage. The primary benefit for using this relatively high specific impulse densified propellant mixture is the subsequent reduction of the launch vehicle gross lift-off weight. Production of densified propellants however requires specialized equipment to actively subcool both the liquid oxygen and liquid hydrogen to temperatures below their normal boiling point. A propellant densification unit based on an external thermodynamic vent principle which operates at subatmospheric pressure and supercold temperatures provides a means for the LH2 and LOX densification process to occur. To demonstrate the production concept for the densification of the liquid hydrogen propellant, a system comprised of a multistage gaseous hydrogen compressor, LH2 recirculation pumps and a cryogenic LH2 heat exchanger was designed, built and tested at the NASA Lewis Research Center (LeRC). This paper presents the design configuration of the LH2 propellant densification production hardware, analytical details and results of performance testing conducted with the hydrogen densifier Ground Support Equipment (GSE).

Tomsik, Thomas M.↗
NASA NTRS2000

Injector for liquid fueled rocket engine

An injector for liquid fueled rocket engines wherein a generally flat core having a frustoconical dome attached to one side of the core to serve as a manifold for a first liquid, with the core having a generally circular configuration having an axis. The other side of the core has a plurality of concentric annular first slots and a plurality of annular concentric second slots alternating with the first slots, the second slots having a greater depth than said first slots. A bore extends through the core for inletting a second liquid into said core, the bore intersecting the second slots to feed the second liquid into the second slots. The core also has a plurality of first passageways leading from the manifold to the first annular slots for feeding the first liquid into said first slots. A faceplate brazed to said other side of the core is provided with apertures extending from the first and second slots through said face plate, these apertures being positioned to direct fuel and liquid oxygen into contact with each other in the combustion chamber. The first liquid may be liquid oxygen and the second liquid may be kerosene or liquid hydrogen.

Cornelius, Charles S.↗
NASA NTRS1977

A promising new cryogenic seal candidate

The Space Shuttle carries more than 500,000 kg of liquid oxygen (LO2) and 100,000 kg of liquid hydrogen (LH2) propellant. Seals for this main propellant system therefore must provide extremely high reliability under prolonged service at cryogenic temperatures. The chief threat to consistently reliable operation is not the cryogenic service itself, but the 325 K temperature variation encountered during each individual mission. Moreover, this reliability must be sustained during the entire 100-mission/ten-year projected life of the Space Shuttle. It was feared that conventional liquid oxygen seal materials (filled Teflon and Kel-F) could cold flow or compression set to such an extent during the high-temperature phases of the 100 successive missions that failures could result. Thus, in addition to LO2 compatibility and cryogenic serviceability, this program was directed toward seals that were not thermoplastic and had good compression set resistance

Merz, P. L.↗
NASA NTRS1990

Lunar habitat concept employing the Space Shuttle external tank

The Space Shuttle external tank, which consists of a liquid oxygen tank, an intertank structure, and a liquid hydrogen tank, is an expendable structure used for approximately 8.5 min during each launch. A concept for outfitting the liquid oxygen tank-intertank unit for a 12-person lunar habitat is described. The concept utilizes existing structures and openings for both man and equipment access without compromising the structural integrity of the tank. Living quarters, instrumentation, environmental control and life support, thermal control, and propulsion systems are installed at Space Station Freedom. The unmanned habitat is then transported to low lunar orbit and autonomously soft landed on the lunar surface. Design studies indicate that this concept is feasible by the year 2000 with concurrent development of a space transfer vehicle and manned cargo lander for crew changeover and resupply.

King, Charles B.↗
NASA NTRS1992

Orbital transfer vehicle oxygen turbopump technology. Volume 3: Hot oxygen testing

This report covers the work done in preparation for a liquid oxygen rocket engine turbopump test utilizing high pressure hot oxygen gas for the turbine drive. The turbopump (TPA) is designed to operate with 400 F oxygen turbine drive gas. The goal of this test program was to demonstrate the successful operation of the TPA under simulated engine conditions including the hot oxygen turbine drive. This testing follows a highly successful series of tests pumping liquid oxygen with gaseous nitrogen as the turbine drive gas. That testing included starting of the TPA with no assist to the hydrostatic bearing. The bearing start entailed a rubbing start until the pump generated enough pressure to support the bearing. The articulating, self-centering hydrostatic bearing exhibited no bearing load or stability problems. The TPA was refurbished for the hot gas drive tests and facility work was begun, but unfortunately funding cuts prohibited the actual testing.

Urke, Robert L.↗
NASA NTRS1988

Electromagnetic dampers for cryogenic applications

Cryogenic turbomachinery of the type used to pump high-pressure liquid hydrogen at -423 F and liquid oxygen at -297 F to the main engines of the Space Shuttle are subjected to lateral rotor vibrations from unbalance forces and transient loads. Conventional dampers which utilize viscous fluids such as lubricating oil cannot be used in turbopumps because the bearing components are filled with either liquid hydrogen or liquid oxygen, which have viscosity comparable to air and, therefore, are not effective in viscous dampers. Electromagentic dampers are currently being explored as a means of providing damping in cryogenic turbopumps because their damping effectiveness increases as temperature decreases and because they are compatible with the liquid hydrogen or liquid oxygen in the turbopumps.

Brown, Gerald V.↗
NASA NTRS2000

Propellant Densification Ground Testing Conducted for Launch Vehicles

The NASA Glenn Research Center at Lewis Field has taken the lead in the development of practical densified cryogenic propellants for launch vehicle applications. The technology of subcooling cryogenic propellants below their normal boiling point to produce a denser fluid is one of the key process technologies necessary to meet the challenge of single-stage-to-orbit and reusable launch vehicles. Densified propellants are critical to lowering launch costs because they enable more propellant to be packed into a given unit volume, thus improving the performance by reducing the overall size and weight of the launch vehicle. This two-pronged research and test program has evolved into (1) conducting tank loading tests using densified liquid hydrogen and (2) developing two large-scale propellant densification systems that will be performance tested next year at Glenn. The propellant-loading test program was undertaken at Glenn in coordination with Lockheed Martin Michoud Space Systems. In this testing, the liquid hydrogen recirculation and densification process was simulated, and the thermal stratification of the densified propellant was recorded throughout the tank. The test article was a flight-weight tank constructed from composite materials similar to those to be used on the X-33 launch vehicle. The tank geometry as designed by Lockheed Martin had two cylindrical lobes with a center septum. Liquid hydrogen flow rate, pressure data, and temperature data plotted over time were collected while the subscale tank was filled with 27 R (15 K) densified liquid hydrogen propellant. This testing has validated mathematical models and demonstrated the readiness of densified propellant technology for near-term use. It marks the first time that such a process has been carried out with a multiple-lobe, flight-similar tank. Glenn researchers have also been working on providing a process and critical test data for the continuous production of densified liquid hydrogen (LH2) and densified liquid oxygen (LO2). Each densification production process uses a high-efficiency, subatmospheric boiling bath heat exchanger to cool the working fluid. A near triple-point hydrogen boiling bath is used to condition and subcool hydrogen to 27 R (15 K), and a nitrogen boiling bath is used to cool the liquid oxygen to 120 R (66.7 K). Multistage centrifugal compressors operating at cryogenic inlet conditions maintain the heat exchanger bath vapor pressure below 1 atm. The LO2 propellant densification unit shown in the photograph has a 30 lb/sec capacity, whereas the LH2 unit was designed to process 8 lb/sec of propellant. Each densification unit will be transported to Glenn's South Forty area after all fabrication work is completed sometime late next year. There the LO2 and LH2 densifier performance tests will be conducted with another larger Lockheed Martin tank designated the Structural Test Article (STA). This liquid oxygen tank is a full-scale, flight-weight, prototype aluminum tank designed for the X-33. It has a capacity of 20,000 gallons of LO2. The tank loading and recirculation testing planned for next year with STA will provide the data necessary for full-scale development of propellant densification technology.

Tomsik, Thomas M.↗