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NASA NTRS2017

Pressure-Fed LOX/LCH4 Reaction Control System for Spacecraft: Transient Modeling and Thermal Vacuum Hotfire Test Results

An integrated cryogenic liquid oxygen, liquid methane (LOX/LCH4) reaction control system (RCS) was tested at NASA Glenn Research Center's Plum Brook Station in the Spacecraft Propulsion Research Facility (B-2) under vacuum and thermal vacuum conditions. The RCS is a subsystem of the Integrated Cryogenic Propulsion Test Article (ICPTA), a pressure-fed LOX/LCH4 propulsion system composed of a single 2,800 lbf main engine, two 28 lbf RCS engines, and two 7 lbf RCS engines. Propellants are stored in four 48 inch diameter 5083 aluminum tanks that feed both the main engine and RCS engines in parallel. Helium stored cryogenically in a composite overwrapped pressure vessel (COPV) flows through a heat exchanger on the main engine before being used to pressurize the propellant tanks to a design operating pressure of 325 psi. The ICPTA is capable of simultaneous main engine and RCS operation. The RCS engines utilize a coil-on-plug (COP) ignition system designed for operation in a vacuum environment, eliminating corona discharge issues associated with a high voltage lead. There are two RCS pods on the ICPTA, with two engines on each pod. One of these two engines is a heritage flight engine from Project Morpheus. Its sea level nozzle was removed and replaced by an 85:1 nozzle machined using Inconel 718, resulting in a maximum thrust of 28 lbf under altitude conditions. The other engine is a scaled down version of the 28 lbf engine, designed to match the core and overall mixture ratios as well as other injector characteristics. This engine can produce a maximum thrust of 7 lbf with an 85:1 nozzle that was additively manufactured using Inconel 718. Both engines are film-cooled and capable of limited duration gas-gas and gas-liquid operation, as well as steady-state liquid-liquid operation. Each pod contains one of each version, such that two engines of the same thrust level can be fired as a couple on opposite pods. The RCS feed system is composed of symmetrical 3/8 inch lines that tap off of the main propellant manifold to send LOX and LCH4 outboard to the RCS pods. A Thermodynamic Vent System (TVS) is used to condition propellants at each pod by venting through an orifice and then routing the cold expansion products back through tubing that is welded along a large portion of the main RCS feed lines. Prior to final installation on the ICPTA, the RCS engines were tested in a small vacuum chamber at the Johnson Space Center (JSC) Energy Systems Test Area (ESTA) to verify functionality of the new COP ignition system and check out operation of the vacuum nozzles. After engine-level testing, the RCS engines were installed on the vehicle and a series of integrated hot-fire tests were performed at JSC consisting of various pulsing and steady-state firings as well as integrated main engine/RCS operation. The ICPTA was then integrated into the Plum Brook B-2 facility for vacuum and thermal/vacuum testing. Testing in the B-2 facility was composed of multiple thermal and pressure environments. The first set of tests were performed under ambient temperature and altitude pressure conditions. These tests consisted of a range of minimum impulse bit (MIB) pulsing sequences with low duty cycle, analogous to a coast phase in which the RCS is primarily used for station keeping. The primary goal of this sequence is to understand how propellant conditions were effected without an active TVS. In this scenario, consistent gas-gas operation is desirable since it results in a smaller MIB and more efficient propellant consumption. Multiple skin thermocouples are mounted on the feedlines, in addition to a submerged thermocouple on each commodity, in order to gather thermal data on the system. Higher duty cycle pulsing tests were then performed, analogous to an ascent or landing mission phase. The primary goal of this sequence was to examine how well the engines self-conditioned without active TVS when starting from a quiescent state. The TVS was then activated during some tests to demonstrate the capability to quickly condition the engines for higher pulsing demand scenarios. A thermocouple at the TVS outlet allows for the calculation of energy absorbed by the vented propellant. Lastly, tests with longer pulses and multiple engines firing either in sequence or simultaneously were run in order to gather transient system response data on waterhammer. Six total high-speed pressure transducers are installed on the RCS system, one sensor at the end of each propellant manifold line on the pods, and one at the tap-off location for each commodity. This will allow for the accurate characterization of waterhammer in the system under various propellant conditions and firing sequences. Other instrumentation for this test series includes nozzle throat thermocouples, chamber pressure measurement, heat soakback measurement, and tank wall plume impingement temperature measurement. The next set of tests were performed to demonstrate simultaneous main engine and RCS operation. Data from this test will be used to examine if there is any change to nominal operation of the RCS as a result of feed system interaction or other phenomenon. Some of these tests began under high vacuum conditions (target ambient pressure less than 1x10(exp -3) torr) and others began at altitude conditions. The last set of tests were performed with the B-2 cold wall active. Under these tests, many of the same low duty cycle MIB tests were repeated in order to characterize how propellant conditions changed with the lower heat leak. In this scenario the RCS manifold experiences much less heat leak, resulting in a change to how well the engines self-condition. As a result, an increase in maximum waterhammer pressures and a change in natural frequency of the system was expected due to higher density propellants. The lower heat leak should also result in a change to the MIB pulse profile, and data will be examined to understand how MIB repeatability is affected in the different operating environments. Parallel to the test efforts, a set of transient model development efforts were made to predict RCS performance. The primary effort was aimed at producing a SINDA/FLUINT model to predict propellant conditioning up to the engine inlet as a function of different environmental and operating parameters, with the goal of predicting chamber pressure, TVS performance, and propellant consumption over time. Preliminary results for this effort will be presented in comparison with test data. Additional modeling efforts were made using SINDA/FLUINT to predict waterhammer in the system since the software is capable of handling multiphase transient fluid dynamics. These results will be compared with the high-speed pressure transducer test data for validation purposes.

Atwell, Matthew J.↗
NASA NTRSDate not supplied

Energy Efficient Large-Scale Storage of Liquid Hydrogen

The world’s largest liquid hydrogen storage tanks were constructed in the mid-1960sat the NASA Kennedy Space Center. These two vacuum-jacketed, perlite powder insulated tanks, still in service today, have 3,200 m3 of useable capacity. In 2018, construction began on an additional storage tank at Launch Complex 39B. This new tank will give an additional storage capacity of 4,700 m3 for a total on-site storage capacity of roughly 8,000 m3. NASA’s new Space Launch System (SLS) heavy lift rocket for the Artemis program includes an LH2tank that makes up the bulk of the vehicle, holding 2,033 m3 of LH2 in its 8.4-m diameter by40-m height. The new storage tank includes two new energy-efficient technologies: a glass bubbles insulation system in lieu of perlite, and an Integrated Refrigeration and Storage (IRAS)heat exchanger for controlled storage capability. The evacuated glass bubbles insulation system is based on the prior two decades of research to prove the thermal performance benefits as well as the mechanical and vacuum integrity; and has been shown to reduce LH2 boiloff by 46%versus perlite in field demonstrations. The IRAS capability is centered on a heat exchanger system that is built within the inner vessel to reject heat from the bulk liquid through the future implementation of an external helium refrigerator. Controlled storage via IRAS, when fully implemented, will provide full control of the ullage pressure, zero boiloff, and even production of densified LH2 pending its adoption on future launch vehicles. The design basics are described along with main construction and testing processes involved. The key features of the new technology items and implications on simplified operations and long-term energy savings are addressed

cryogenics↗
NASA NTRS1966

Special treatment reduces helium permeation of glass in vacuum systems

Internal surfaces of the glass component of a vacuum system are exposed to cesium in gaseous form to reduce helium permeation. The cesium gas is derived from decomposition of cesium nitrate through heating. Several minutes of exposure of the internal surfaces of the glass vessel are sufficient to complete the treatment.

Bryant, P. J.↗
NASA NTRS1987

Space environmental considerations for a long-term cryogenic storage vessel

Information is given on the kind of protection that is needed against impact and perforation of a long-term cryogenic storage vessel in space by meteoroids and space debris. The long-term effects of the space environment on thermal control surfaces and coatings, and the question of whether the insulation and thermal control surfaces should be encased in a vacuum jacket shell are discussed.

Nakanishi, Shigeo↗
NASA NTRS2013

Lightweight Liquid Helium Dewar for High-Altitude Balloon Payloads

Astrophysical observations at millimeter wavelengths require large (2-to-5- meter diameter) telescopes carried to altitudes above 35 km by scientific research balloons. The scientific performance is greatly enhanced if the telescope is cooled to temperatures below 10 K with no emissive windows between the telescope and the sky. Standard liquid helium bucket dewars can contain a suitable telescope for telescope diameter less than two meters. However, the mass of a dewar large enough to hold a 3-to-5-meter diameter telescope would exceed the balloon lift capacity. The solution is to separate the functions of cryogen storage and in-flight thermal isolation, utilizing the unique physical conditions at balloon altitudes. Conventional dewars are launched cold: the vacuum walls necessary for thermal isolation must also withstand the pressure gradient at sea level and are correspondingly thick and heavy. The pressure at 40 km is less than 0.3% of sea level: a dewar designed for use only at 40 km can use ultra thin walls to achieve significant reductions in mass. This innovation concerns new construction and operational techniques to produce a lightweight liquid helium bucket dewar. The dewar is intended for use on high-altitude balloon payloads. The mass is low enough to allow a large (3-to-5-meter) diameter dewar to fly at altitudes above 35 km on conventional scientific research balloons without exceeding the lift capability of the balloon. The lightweight dewar has thin (250- micron) stainless steel walls. The walls are too thin to support the pressure gradient at sea level: the dewar launches warm with the vacuum space vented continuously during ascent to eliminate any pressure gradient across the walls. A commercial 500-liter storage dewar maintains a reservoir of liquid helium within a minimal (hence low mass) volume. Once a 40-km altitude is reached, the valve venting the vacuum space of the bucket dewar is closed to seal the vacuum space. A vacuum pump then evacuates the dewar vacuum space to provide the necessary thermal isolation. Liquid helium may then be transferred from the storage dewar into the bucket dewar to cool the telescope inside the bucket dewar. By splitting the functions of helium storage and in-flight thermal isolation, the parasitic mass associated with the dewar pressure vessel is eliminated to achieve factor-of-five or better reduction in mass. The lower mass allows flight on conventional scientific research balloons, even for telescopes 3 to 5 meters in diameter.

Kogut, Alan↗
NASA NTRS2006

Fabrication and Evaluation of Titanium and Zirconium based Wires for use during Extended, Deep Space, Missions

Novel materials and designs are necessary for transport vessels and propulsion systems to fulfill NASA's vision of easier access to space and the expansion of human exploration beyond low-earth orbit. Spacecraft components must necessarily be lighter and stronger than their predecessors and will likely be required to serve new purposes. Furthermore, they must be resilient to the thermal, vacuum, and radiation environment of space for extended periods of time and may need to perform in the near proximity of a nuclear fuel source. To this end research has been initiated to fabricate novel, composite, wires based on titanium and zirconium pearlitic alloys. It is expected that the fabricated wire will well endure in the space environment with application as tethers, sail components, fasteners, and a myriad of other (including earth-based) uses. A background on pearlitic wire, novel alloy development, microstructural characterization, and initial mechanical testing results will be presented and discussed.

Grugel, Richard N.↗
NASA NTRS1985

The development of a containment vessel and Dewar for the particle astrophysics magnet facility (ASTROMAG)

The ASTROMAG facility is the heart of a large charged particle detection and resolution system. ASTROMAG utilizes a superconducting magnet consisting of a large superconducting magnet coil with a stored magnetic energy of approximately 15 MJ. The active coil will have a mass of 1200 kg. This magnet will be cooled by a cryostat using a liquid helium Dewar for storage. The cryostat will have a series of gas-cooled shields with an external guard vacuum shield and an internal Dewar. The magnet and cryostat will be designed for shuttle or Delta launch and will be designed to withstand the internal pressure of expanded helium under full quench conditions when venting is prevented. The external guard vacuum shell is required to maintain a vacuum for Earth based testing and for cold launch of the cryostat and magnet. The magnet is designed to operate at 4.4 K with a peak field of 7.0 tesla. The superconducting material within the magnet is niobium titanium in a conductive matrix.

Source record↗
NASA NTRS1988

Use of pure nickel and LiOH for thermal energy storage

The solid to liquid phase transformation of LiOH has been proposed as an ideal candidate thermal energy storage media for a Rankine Cycle powered electrical generation unit envisioned in Space Station based solar dynamic systems. Due to the corrosive nature of molten hydroxides, long term containment of LiOH is of concern. Pure nickel is thought to be a suitably resistant material, and a program has been instituted to measure the effects of prolonged exposure of liquid and gaseous LiOH on the mechanical properties of pure nickel alloys. Results to date indicate that negligible weight and thickness changes occurred in Ni alloys exposed to LiOH for as long as 2500 hr at 775 K, and essentially no difference in 77-900 K tensile properties could be detected between LiOH exposed and vacuum annealed Ni specimens. Although there was little sign of outward damage, microstructural examination revealed that all hydroxide contaminated tensile test specimens had surface connected intergranular cracks along the gage lengths. Two other potential problems, which have strong implications with respect to a LiOH/Ni energy storage system, were also noted during the corrosion experiments. In particular stress corrosion cracking of weld joints in pressurized vessel and permeation of hydrogen through nickel were observed.

Whittenberger, J. D.↗
NASA NTRS1989

Unpressurized Container For Cryogenic Testing

Unpressurized cryostat makes mechanical testing of materials at low temperature more convenient. Maintains specimens at temperatures of -400 to -450 degree F without sealing them in gastight, vacuum-insulated container. Easy to insert and remove specimens and attach instrumentation wiring to them. Vents vapor continuously, so no danger of buildup of internal pressure from evaporating cryogenic liquid. Includes two concentric chambers with stainless-steel walls and fiber insulation. Specimen mounted in inner chamber, and such instruments as extensometers and thermocouples attached. Loose lid of polystyrene foam or other suitable material placed over vessel.

Walker, Susan B.↗
NASA NTRS2011

Hybrid Composite Cryogenic Tank Structure

A hybrid lightweight composite tank has been created using specially designed materials and manufacturing processes. The tank is produced by using a hybrid structure consisting of at least two reinforced composite material systems. The inner composite layer comprises a distinct fiber and resin matrix suitable for cryogenic use that is a braided-sleeve (and/or a filamentwound layer) aramid fiber preform that is placed on a removable mandrel (outfitted with metallic end fittings) and is infused (vacuum-assisted resin transfer molded) with a polyurethane resin matrix with a high ductility at low temperatures. This inner layer is allowed to cure and is encapsulated with a filamentwound outer composite layer of a distinct fiber resin system. Both inner and outer layer are in intimate contact, and can also be cured at the same time. The outer layer is a material that performs well for low temperature pressure vessels, and it can rely on the inner layer to act as a liner to contain the fluids. The outer layer can be a variety of materials, but the best embodiment may be the use of a continuous tow of carbon fiber (T-1000 carbon, or others), or other high-strength fibers combined with a high ductility epoxy resin matrix, or a polyurethane matrix, which performs well at low temperatures. After curing, the mandrel can be removed from the outer layer. While the hybrid structure is not limited to two particular materials, a preferred version of the tank has been demonstrated on an actual test tank article cycled at high pressures with liquid nitrogen and liquid hydrogen, and the best version is an inner layer of PBO (poly-pphenylenebenzobisoxazole) fibers with a polyurethane matrix and an outer layer of T-1000 carbon with a high elongation epoxy matrix suitable for cryogenic temperatures. A polyurethane matrix has also been used for the outer layer. The construction method is ideal because the fiber and resin of the inner layer has a high strain to failure at cryogenic temperatures, and will not crack or produce leaks. The outer layer serves as more of a high-performance structural unit for the inner layer, and can handle external environments.

DeLay, Thomas↗
NASA NTRS1977

Multistack nickel-hydrogen units

A typical 625 Inconel pressure vessel, equipped with four thermocouples, has voltage taps not only for the cell, but for each of the several packs inside the container. In a two-stack design, a weld ring is used to separate the packs which, joined together, form one solid group. In the four-stack design, stacks are separated mechanically with about half inch Plexiglas and are connected through the center. Electrolytes are forced out by vacuum. The two stack design is intended for 500 psi and was pressure tested up to 1,000. The problem of electrolyte shorts is discussed and possible solutions for its correction are given.

Smith, J.↗
NASA NTRS2006

Sensor Amplifier for the Venus Ground Ambient

Previous Venus Landers employed high temperature pressure vessels, with thermally protected electronics, to achieve successful missions, with a maximum surface lifetime of 127 minutes. Extending the operating range of electronic systems to the temperatures (480 C) and pressures (90 bar) of the Venus ground ambient would significantly increase the science return of future missions. Toward that end, the current work describes the innovative design of a sensor preamplifier, capable of working in the Venus ground ambient and designed using commercial components (thermionic vacuum tubes, wide band gap transistors, thick film resistors, advanced high temperature capacitors, and monometallic interfaces) To identify commercial components and electronic packaging materials that are capable of operation within the specified environment, a series of active devices, passive components, and packaging materials were screened for operability at 500C, assuming a 10x increase in the mission lifetime. In addition. component degradation as a function of time at 500(deg)C was evaluated. Based on the results of these preliminary evaluations, two amplifiers were developed.

Venus electronics↗
NASA NTRS2010

Insulation-Testing Cryostat With Lifting Mechanism

The figure depicts selected aspects of an apparatus for testing thermal-insulation materials for cryogenic systems at temperatures and under vacuum or atmospheric conditions representative of those encountered in use. This apparatus, called "Cryostat-100," is based on the established cryogen-boil-off calorimeter method, according to which the amount of heat that passes through an insulation specimen to a cryogenic fluid in a container, and thus the effective thermal conductance of the specimen, is taken to be proportional to the amount of the cryogenic fluid that boils off from the container. The design of Cryostat-100 is based partly on, and incorporates improvements over, the design of a similar prior apparatus called "Cryostat-1" described in "Improved Methods of Testing Cryogenic Insulation Materials" (KSC-12107 & KSC- 12108), NASA Tech Briefs, Vol. 24, No. 12 (December 2000), page 46. The design of Cryostat-100 also incorporates the best features of two other similar prior apparatuses called "Cryostat-2" (also described in the cited prior article) and "Cryostat- 4." Notable among the improvements in Cryostat-100 is the addition of a lifting mechanism that enables safe, rapid, reliable insertion and removal of insulation specimens and facilitates maintenance operations that involve lifting. As in Cryostat-1, the cold mass is a vertical stainless-steel cylindrical vessel subdivided into a larger measurement vessel with smaller thermal-guard vessels at both ends. During operation, all three vessels are kept filled with liquid nitrogen near saturation at ambient pressure (temperature .77.4 K). The cold mass of Cryostat-100 has a length of 1 m and diameter of 168 mm. Each specimen has a corresponding nominal length and inner diameter and a nominal thickness of 25.4 mm. Specimens that are shorter and have thicknesses between 0 and 50 mm are also acceptable. Bulk-fill, foam, clam-shell, multilayer insulation, and layered materials can be tested over a very wide range of thermal transmission: apparent thermal conductivity from 0.01 to 60 mW/m-K and heat flux from 0.1 to 500 W/sq m. A test in Cryostat-100 can be conducted at any desired gas pressure between ambient atmospheric pressure at one extreme and a vacuum with residual pressure <10(exp -5) torr (<1.33 10(exp -3) Pa) at the other extreme. The residual gas (and purge gas) is typically nitrogen, but can be any suitable purge gas (e.g., helium, argon, or carbon dioxide). Usually, the temperature on the warm boundary of the insulation specimen is maintained near the ambient value (approximately 293 K), while the boiling of liquid nitrogen at atmospheric pressure in the cold mass maintains the temperature on the cold boundary of the specimen at approximately 77 K.

Fesmire, James↗
NASA NTRS1990

Ongoing nickel-hydrogen energy storage device testing at George C. Marshall Space Flight Center

The primary objective of the testing is to characterize Ni-H2 cells for successful integration into the electrical power system (EPS) of the Hubble Space Telescope (HST). A broad spectrum of Ni-H2 design technology is encompassed by the testing configurations; tests include cells with dates of manufacture as early as 1976. The database includes cells of varied storage times, capacity, plate design, stack design, terminal configuration, pressure vessel thickness, separator material, potassium hydroxide (KOH) concentration, and thermal control. Currently, 196 Ni-H2 cells are being tested, grouped as follows: 12 RNH-35-3, 14 RNH-30-1, 22 HST cells (1 battery, flight spare lot), 132 HST cells (6 batteries, test modules 1 and 2, called TM1 and TM2), 12 HST cells (3 four-cell packs, TM1, TM2, flight spare module FSM), and 4 HST cells (engineering lot). In addition to the characterization and life testing, an extensive thermal vacuum and purge test was conducted in November 1989 and February 1990 using the HST FSM (3 batteries composed of 69 HST cells from the flight spare lot) to help verify thermal design. A report is presented of the progress, significant findings, and future objectives of the testing.

Lowery, John E.↗
NASA NTRS2012

Sapphire Viewports for a Venus Probe

A document discusses the creation of a viewport suitable for use on the surface of Venus. These viewports are rated for 500 C and 100 atm pressure with appropriate safety factors and reliability required for incorporation into a Venus Lander. Sapphire windows should easily withstand the chemical, pressure, and temperatures of the Venus surface. Novel fixture designs and seals appropriate to the environment are incorporated, as are materials compatible with exploration vessels. A test cell was fabricated, tested, and leak rate measured. The window features polish specification of the sides and corners, soft metal padding of the sapphire, and a metal C-ring seal. The system safety factor is greater than 2, and standard mechanical design theory was used to size the window, flange, and attachment bolts using available material property data. Maintenance involves simple cleaning of the window aperture surfaces. The only weakness of the system is its moderate rather than low leak rate for vacuum applications.

Bates, Stephen↗
NASA NTRS2001

Beagle 2: The Next Exobiology Mission to Mars

Beagle 2 is a 60 kg probe (with a 30 kg lander) developed in the United Kingdom for inclusion on the European Space Agency's 2003 Mars Express. Beagle 2 will deliver to the Martian surface a payload which consists of a high percentage of science instruments to landed spacecraft mass. Beagle 2 will be launched in June, 2003 with Mars Express on a Soyuz-Fregat rocket from the Baikonur Cosmodrome in Kazakhstan. Beagle 2 will land on Mars on December 26, 2003 in the Isidis Planitia basin (approximately 10 degrees N and 275 degrees W), a large sedimentary basin that overlies the boundary between ancient highlands and northern plains. Isidis Planitia, the third largest basin on Mars, which is possibly filled with sediment deposited at the bottom of long-standing lakes or seas, offers an ideal environment for preserving traces of life. Beagle 2 was developed to search for organic material and other volatiles on and below the surface of Mars in addition to the study of the inorganic chemistry and mineralogy. Beagle 2 will utilize a mechanical mole and grinder to obtain samples from below the surface, under rocks and inside rocks. A pair of stereo cameras will image the landing site along with a microscope for examination of surface and rock samples. Analyses will include both rock and soil samples at various wavelengths, X-ray spectrometer and Mossbauer spectrometer as well as a search for organics and other light element species (e.g. carbonates and water) and measurement of their isotopic compositions. Beagle 2 has as its focus the goal of establishing whether evidence for life existed in the past on Mars at the Isidis Planitia site or at least establishing if the conditions were ever suitable. Carbonates and organic components were first recognized as existing on Mars when they were found in the Martian meteorite Nakhla. Romanek et al showed the carbonates in ALH84001 were formed at low temperatures. McKay et al noted possible evidence of early life on Mars within the ALH84001 meteorite. Thomas-Keprta et al showed the magnetite biomarkers in ALH84001's carbonates are indistinguishable from those formed by magnetotactic bacteria found on Earth. Gibson et al showed there was significant evidence for liquid water and biogenic products present on Mars across a 3.9 billion year period. A mechanical arm (PAW) operates from the lander and is used for science operations along with sample acquisition. Instruments attached to the PAW include: stereo cameras, Mossbauer instrument, X-ray fluorescence instrument, microscope, environmental sensors, rock corer/grinder, a spoon, mirror, brushes, a mole attachment for acquisition of subsurface to depths of 1 to 2 meters and an illumination device. Each camera has 14 filters which have been optimized for mineralogy composition, dust and water vapor detection. The microscope's camera is designed for viewing the size and shape of dust particles, rock surfaces, microfossils, and characteristics of the samples prior to introduction into the gas analysis package (GAP). The camera has a resolution of 4 microns/pixel, features four color capability (red, green, blue and UV (ultraviolet) fluoresence), a depth of focus of 40 micrometers and translation stage of +3 millimeters. The heart of the Beagle 2's life detection package is the gas analysis package (GAP), which consists of a mass spectrometer with collectors at fixed masses for precise isotopic ratio measurements and voltage scanning for spectral analysis. Primary aim of the GAP is to search for the presence of bulk constituents, individual species, and isotopic fractionations for both extinct and extent life along with studying the low-temperature geochemistry of the hydrogen, carbon, nitrogen and oxygen components on Mars from both the surface and atmosphere. GAP is a magnetic sector mass spectrometer with the range of 1 to 140 amu which can be operated in both the static and dynamic modes. A triple Faraday collector array will be used for C, N and O ratios along with a double Faraday array for H/D. Pulse counting electron multiplier will be utilized for noble gases and selected organics. Anticipated detection limits are at the picomole level for operation in the static mode of operation and high precision isotopic measurements will be made in the dynamic mode. Sample processing and preparation system consists of reaction vessels along with references. Sample ovens capable of being heated are attached to the manifold for sample combustion. Surface, subsurface materials and interior rock specimens will be combusted in pure oxygen gas at various temperature intervals to release organic matter and volatiles. Combustion process will permit detection of all forms and all atoms of carbon present in the samples. A chemical processing system is capable of a variety of conversion reactions. Gases are manipulated either by cryogenic or chemical reactions and passed through the gas handling portion of the vacuum system. There are two modes of operation: quantitative analysis and precise isotopic measurements. Additional information is contained in the original extended abstract.

Gibson, Everett K., Jr.↗
NASA NTRS2016

Engineering Polymer Blends for Impact Damage Mitigation

Structures containing polymers such as DuPont's Surlyn® 8940, demonstrate puncture healing when impacted by a 9 millimeter projectile traveling from speeds near 300 meters per second (1,100 feet per second) to hypervelocity impacts in the micrometeoroid velocity range of 5 kilometers per second (16,000 feet per second). Surlyn® 8940 puncture heals over a temperature range of minus 30 degrees Centigrade to plus 70 degrees Centigrade and shows potential for use in pressurized vessels subject to impact damage. However, such polymers are difficult to process and limited in applicability due to their low thermal stability, poor chemical resistance and overall poor mechanical properties. In this work, several puncture healing engineered melt formulations were developed. Moldings of melt blend formulations were impacted with a 5.56 millimeter projectile with a nominal velocity of 945 meters per second (3,100 feet per second) at about 25 degrees Centigrade, 50 degrees Centigrade and 100 degrees Centigrade, depending upon the specific blend being investigated. Self-healing tendencies were determined using surface vacuum pressure tests and tensile tests after penetration using tensile dog-bone specimens (ASTM D 638-10). For the characterization of tensile properties both pristine and impacted specimens were tested to obtain tensile modulus, yield stress and tensile strength, where possible. Experimental results demonstrate a range of new puncture healing blends which mitigate damage in the ballistic velocity regime.

Gordon, Keith L.↗
NASA NTRS2017

Development of a Numerical Model of Hypervelocity Impact into a Pressurized Composite Overwrapped Pressure Vessel

As the outlook for space exploration becomes more ambitious and spacecraft travel deeper into space than ever before, it is increasingly important that propulsion systems perform reliably within the space environment. The increased reliability compels designers to increase design margin at the expense of system mass, which contrasts with the need to limit vehicle mass to maximize payload. Such are the factors that motivate the integration of high specific strength composite materials in the construction of pressure vessels commonly referred to as composite overwrapped pressure vessels (COPV). The COPV consists of a metallic liner for the inner shell of the COPV that is stiff, negates fluid permeation and serves as the anchor for composite laminates or filaments, but the liner itself cannot contain the stresses from the pressurant it contains. The compo-site-fiber reinforced polymer (CFRP) is wound around the liner using a combination of hoop (circumferential) and helical orientations. Careful consideration of wrap orientation allows the composite to evenly bear structural loading and creates the COPV's characteristic high strength to weight ratio. As the CFRP overwrap carries most of the stresses induced by pressurization, damage to the overwrap can affect mission duration, mission success and potentially cause loss-of-vehicle/loss-of-crew. For this reason, it is critical to establish a fundamental understanding of the mechanisms involved in the failure of a stressed composite such as that of the COPV. One of the greatest external threats to the integrity of a spacecraft's COPV is an impact from the meteoroid and orbital debris environments (MMOD). These impacts, even from submillimeter particles, generate extremely high stress states in the CFRP that can damage numerous fibers. As a result of this possibility, initial assumptions in survivability analysis for some human-rated NASA space-craft have assumed that any alteration of the vessel due to impact is considered a catastrophic failure. This assumption is conservative and made due to lack of knowledge on the level of allow-able damage to the composite overwrap that can be sustained and still allow successful completion of the mission. To quantify the allowable damage level to the composite overwrap involves assessing stress redistribution following damage as well as evaluating possible time-dependent mechanisms involved in the COPV response to an impact event. Limited published work in this subject has shown that COPV can withstand at least some level of damage due to high energy impacts. These observations have been confirmed and expanded upon in recent experimental research performed by NASA. This research has demonstrated that there is not only robustness in a COPV to compensate for CFRP damage, but has also identified two significant failure modes for pressurized COPV. The lowest threshold failure mode involves the perforation of the vessel, and the highest threshold failure mode is the catastrophic rupture. While both of these failure modes mean a loss of the COPV, system robustness affords some tolerance to the venting as opposed to the more catastrophic rupture. As a consequence, it is necessary to understand the conditions that result in the transition between these failure modes. The aforementioned experimental research has been performed in both the unpressurized and pressurized condition to identify the damage level that triggered the failure thresh-old. This COPV test program was sponsored by the NASA Engineering and Safety Center (NESC), and tests were performed at NASA White Sands Test Facility (WSTF). Planning and coordination were provided by NASA JSC Hypervelocity Impact Technology (HVIT) group, and the COPVs were provided by the ISS Program. Unpressurized testing has been conducted at the pressure of the vacuum test chamber, while, the pressurized testing has been conducted at 290 +/- 10 bar (4,200  100 psi) using nitrogen as the pressurizing gas, which corresponds to the design pressure for the target COPV. In this research, spherical aluminum projectiles with varying diameter has been chosen as the impactor. For the unpressurized COPV, the dependence of penetration up to the dependence of hole size in the liner has been obtained as a function of impact conditions. For the pressurized research, the dependence of penetration up to rupture has been obtained as a function of im-pact conditions. Two representative post-test photographs of the failed COPV's from a nor-mal impact into the COPV surface are shown in Fig. 1. These images display the dramatic difference between failure modes, venting (Fig. 1a) and rupture (Fig. 1b). For venting, liner perforation, severed composite fibers/tows and ply delamination are commonly observed damage characteristics of this COPV failure mode. In the case of rupture, the COPV typically experienced a separation of its domed regions and severe break-up of the cylindrical region. Fully understanding the transition from venting to rupture experimentally is costly and potentially unachievable for conditions that cannot be generated in the laboratory. These shortcomings have motivated the performance of three-dimensional numerical simulations to expand the existing experimental database. These simulations have been carried out with the nonlinear-structural-dynamics, analysis-tool, CTH. A typical pressure contour plot from an impact simulation of an entire COPV is shown in Fig. 2. To generate the COPV stress state without initiating a shock wave, the pressure in the simulated COPV is ramped up to the final pressure over a millisecond prior to impact of the projectile with nitrogen gas. Figure 2a shows the system in this initial condition. After one millisecond, a projectile is initiated into the simulation and impacts the COPV. Figure 2b shows the system after this impact. In the figure, the onset of venting is represented as the change in pressure (μbar), red to green, at the perforation site. Also seen in the figure is the eroded projectile that had passed into the COPV vessel with the generated shock wave in the pressurant propagating just ahead of the material. In this paper, pertinent experimental details and the development of the material constitutive models necessary for this work along with the efforts to validate their use are dis-cussed. The simulation results are presented and compared with the NASA experimental observations. While work is on-going from this effort, early observations pertinent to the failure threshold are presented.

Garcia, M. A.↗