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All sourcesNASA NTRSDOE OSTI
NASA NTRS2007

Optimized Heat Interception for Cryogen Tank Support

We consider means for using the cooling available in boil-off gas to intercept heat conducted through the support structure of a cryogen tank. A one-dimensional model of the structure coupled to a gas stream gives an analytical expression for heat leak in terms of flow rate for temperature independent-properties and laminar flow. A numerical model has been developed for heat transfer on a thin cylindrical tube with an attached vent line. The model is used to determine the vent path layout that will minimize heat flow into the cryogen tank. The results are useful for a number of applications, but the one of interest in this study is the minimization of the boil-off in large cryopropellant tanks in low Earth and low lunar orbit.

Canavan, Edgar R.↗
NASA NTRS2008

Active Costorage of Cryogenic Propellants for Exploration

Long-term storage of cryogenic propellants is a critical requirement for NASA's effort to return to the moon. Liquid hydrogen and liquid oxygen provide the highest specific impulse of any practical chemical propulsion system, and thus provides the greatest payload mass per unit of launch mass. Future manned missions will require vehicles with the flexibility to remain in orbit for months, necessitating long-term storage of these cryogenic liquids. For decades cryogenic scientific satellites have used cryogens to cool instruments. In many cases, the lifetime of the primary cryogen tank has been extended by intercepting much of the heat incident on the tank at an intermediate-temperature shield cooled either by a second cryogen tank or a mechanical cryocooler. For an LH2/LO2 propellant system, a combination of these ideas can be used, in which the shield around the LO2 tank is attached to, and at the same temperature as, the LO2 tank, but is actively cooled so as to remove all heat impinging on the tank and shield. This configuration eliminates liquid oxygen boil-off and cuts the liquid hydrogen boil-off to a small fraction of the unshielded rate. This paper studies the concept of active costorage as a means of long-term cryogenic propellant storage. The paper describes the design impact of an active costorage system for the Crew Exploration Vehicle (CEV). This paper also compares the spacecraft level impact of the active costorage concept with a passive storage option in relation to two different scales of spacecraft that will be used for the lunar exploration effort, the CEV and the Earth Departure Stage (EDS). Spacecraft level studies are performed to investigate the impact of scaling of the costorage technologies for the different components of the Lunar Architecture and for different mission durations.

Canavan, Edgar R.↗
NASA NTRS2010

Subcooling for Long Duration In-Space Cryogenic Propellant Storage

Cryogenic propellants such as hydrogen and oxygen are crucial for exploration of the solar system because of their superior specific impulse capability. Future missions may require vehicles to remain in space for months, necessitating long-term storage of these cryogens. A Thermodynamic Cryogen Subcooler (TCS) can ease the challenge of cryogenic fluid storage by removing energy from the cryogenic propellant through isobaric subcooling of the cryogen below its normal boiling point prior to launch. The isobaric subcooling of the cryogenic propellant will be performed by using a cold pressurant to maintain the tank pressure while the cryogen's temperature is simultaneously reduced using the TCS. The TCS hardware will be integrated into the launch infrastructure and there will be no significant addition to the launched dry mass. Heat leaks into all cryogenic propellant tanks, despite the use of the best insulation systems. However, the large heat capacity available in the subcooled cryogenic propellants allows the energy that leaks into the tank to be absorbed until the cryogen reaches its operational thermodynamic condition. During this period of heating of the subcooled cryogen there will be minimal loss of the propellant due to venting for pressure control. This simple technique can extend the operational life of a spacecraft or an orbital cryogenic depot for months with minimal mass penalty. In fact isobaric subcooling can more than double the in-space hold time of liquid hydrogen compared to normal boiling point hydrogen. A TCS for cryogenic propellants would thus provide an enhanced level of mission flexibility. Advances in the important components of the TCS will be discussed in this paper.

Mustafi, Shuvo↗
NASA NTRS2010

A Heat Transfer Investigation of Liquid and Two-Phase Methane

A heat transfer investigation was conducted for liquid and two-phase methane. The tests were conducted at the NASA Glenn Research Center Heated Tube Facility (HTF) using resistively heated tube sections to simulate conditions encountered in regeneratively cooled rocket engines. This testing is part of NASA s Propulsion and Cryogenics Advanced Development (PCAD) project. Nontoxic propellants, such as liquid oxygen/liquid methane (LO2/LCH4), offer potential benefits in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications. Regeneratively cooled thrust chambers are one solution for high performance, robust LO2/LCH4 engines, but cooling data on methane is limited. Several test runs were conducted using three different diameter Inconel 600 tubes, with nominal inner diameters of 0.0225-, 0.054-, and 0.075-in. The mass flow rate was varied from 0.005 to 0.07 lbm/sec. As the current focus of the PCAD project is on pressure fed engines for LO2/LCH4, the average test section outlet pressures were targeted to be 200 psia or 500 psia. The heat flux was incrementally increased for each test condition while the test section wall temperatures were monitored. A maximum average heat flux of 6.2 Btu/in.2 sec was achieved and, at times, the temperatures of the test sections reached in excess of 1800 R. The primary objective of the tests was to produce heat transfer correlations for methane in the liquid and two-phase regime. For two-phase flow testing, the critical heat flux values were determined where the fluid transitions from nucleate boiling to film boiling. A secondary goal of the testing was to measure system pressure drops in the two-phase regime.

VanNoord, Jonathan↗
NASA NTRS2012

Thermal Optimization and Assessment of a Long Duration Cryogenic Propellant Depot

A Cryogenic Propellant Depot (CPD) operating in Low Earth Orbit (LEO) could provide many near term benefits to NASA space exploration efforts. These benefits include elongation/extension of spacecraft missions and reduction of launch vehicle up-mass requirements. Some of the challenges include controlling cryogenic propellant evaporation and managing the high costs and long schedules associated with new spacecraft hardware development. This paper describes a conceptual CPD design that is thermally optimized to achieve extremely low propellant boil-off rates. The CPD design is based on existing launch vehicle architecture, and its thermal optimization is achieved using current passive thermal control technology. Results from an integrated thermal model are presented showing that this conceptual CPD design can achieve propellant boil-off rates well under 0.05% per day, even when subjected to the LEO thermal environment.

Honour, Ryan↗
NASA NTRS2003

System for Testing Thermal Insulation of Pipes

An apparatus and method have been developed for measuring the rates of leakage of heat into pipes carrying liquids, the purpose of the measurements being to quantify the thermal performance of the insulation system. The apparatus is designed primarily for testing pipes used to carry cryogenic liquids, but can also be used for measuring the thermal performance of other insulated pipes or piping systems. The basic measurement principle is straightforward: The outer surface of the pipe insulation is maintained at a fixed warmer temperature. The interior of the pipe is maintained in a narrow fixed lower-temperature range by means of a regular liquid (e.g., water) that is pumped through the pipe at a known flow rate or a cryogenic liquid (e.g., nitrogen) that is saturated at atmospheric pressure and replenished until steady-state conditions are achieved. In the case of water or another liquid pumped through, the inlet and outlet temperatures are measured and heat-leak power is calculated as the mass flow rate of the liquid multiplied by the specific heat of the liquid multiplied by the inlet-to-outlet temperature rise of the liquid. In the case of liquid nitrogen or another low-temperature boiling liquid, the heat-leak power is calculated as the rate of boil-off multiplied by the latent heat of vaporization of the liquid. Then the thermal-insulation performance of the pipe system can be calculated as a function of the measured heat-leak power, the inner and outer boundary temperatures, and the dimensions of the pipe. The apparatus can test as many as three pipes simultaneously. The pipes can have inner diameters up to .15 cm and outer diameters up to .20 cm. The lengths of the pipes may vary; typical lengths are of the order of 18 m. Two thermal guard boxes . one for each end of the pipe(s) under test . are used to make the inlet and outlet fluid connections to the pipe(s) (see figure). The connections include bellows that accommodate thermal expansion and contraction of the pipes. The guard boxes and pipe(s) are positioned so that the pipe(s) slope upward from the upstream to the downstream end at an angle of at least 2 . The upward slope allows vapor bubbles to accumulate at the downstream end. The thermal guard boxes keep the ends of the pipes at the lower interior temperature to prevent spurious lengthwise leakage of heat into the pipes. It is important to prevent this spurious heat leakage because, if it were allowed to occur, it could contribute a large error in the measured heat-leak power. The upstream thermal guard box includes a heat exchanger through which liquid flowing into the pipe(s) is subcooled to the saturation temperature corresponding to the ambient pressure. Conversely, this heat exchanger can also be used to warm the flowing liquid to a desired fixed temperature. The apparatus includes a temperature control device that is placed around each pipe under test. Each device is operated under thermostatic control to maintain the outer surface of the pipe insulation at the specified test temperature. All measurements are recorded on a portable data-acquisition system.

Fesmire, James E.↗
NASA NTRS2012

Supercritical CO2 Cleaning System for Planetary Protection and Contamination Control Applications

Current spacecraft-compatible cleaning protocols involve a vapor degreaser, liquid sonication, and alcohol wiping. These methods are not very effective in removing live and dead microbes from spacecraft piece parts of slightly complicated geometry, such as tubing and loosely fitted nuts and bolts. Contamination control practices are traditionally focused on cleaning and monitoring of particulate and oily residual. Vapor degreaser and outgassing bakeout have not been proven to be effective in removing some less volatile, hydrophilic biomolecules of significant relevance to life detection. A precision cleaning technology was developed using supercritical CO2 (SCC). SCC is used as both solvent and carrier for removing organic and particulate contaminants. Supercritical fluid, like SCC, is characterized by physical and thermal properties that are between those of the pure liquid and gas phases. The fluid density is a function of the temperature and pressure. Its solvating power can be adjusted by changing the pressure or temperature, or adding a secondary solvent such as alcohol or water. Unlike a regular organic solvent, SCC has higher diffusivities, lower viscosity, and lower surface tension. It readily penetrates porous and fibrous solids and can reach hard-to-reach surfaces of the parts with complex geometry. Importantly, the CO2 solvent does not leave any residue. The results using this new cleaning device demonstrated that both supercritical CO2 with 5% water as a co-solvent can achieve cleanliness levels of 0.01 mg/cm2 or less for contaminants of a wide range of hydrophobicities. Experiments under the same conditions using compressed Martian air mix, which consists of 95% CO2, produced similar cleaning effectiveness on the hydrophobic compounds. The main components of the SCC cleaning system are a high-pressure cleaning vessel, a boil-off vessel located downstream from the cleaning vessel, a syringe-type high-pressure pump, a heat exchanger, and a back pressure regulator (BPR). After soaking the parts to be cleaned in the clean vessel for a period, the CO2 with contaminants is flushed out of the cleaning vessel using fresh CO2 in a first-in-first-out (FIFO) method. The contaminants are either precipitating out in the boil-off container or being trapped in a filter subsystem. The parts to be cleaned are secured in a basket inside and can be rotated up to 1,400 rpm by a magnetic drive. The fluid flows within the vessel generate tangential forces on the parts surfaces, enhancing the cleaning effectiveness and shortening the soaking time. During the FIFO flushing, the pump subsystem pushes fresh CO2 into the cleaning vessel at a constant flow rate between 0.01 and 200 mL/min, while the BPR regulates the pressure in the cleaning vessel to within 0.1 bar by controlling the needle position in an outlet valve. The fresh CO2 gas flows through the heat exchanger at a given temperature before entering the cleaning vessel. A platinum resistance thermometer (PRT) reads the cleaning vessel interior temperature that can be controlled to within 0.1 K. As a result, cleaning vessel temperature remains constant during the FIFO flushing. There is no change in solvent power during FIFO flushing since both temperature and pressure inside the cleaning vessel remain unchanged, thus minimizing contaminants left behind. During decompression, both temperature and pressure are strictly controlled to prevent bubbles from generating in the cleaning vessel that could stir up the contaminants that sank to the bottom by gravity.

Lin, Ying↗
NASA NTRS2012

Thermal Optimization of an On-Orbit Long Duration Cryogenic Propellant Depot

A Cryogenic Propellant Depot (CPD) operating in Low Earth Orbit (LEO) could provide many near term benefits to NASA's space exploration efforts. These benefits include elongation/extension of spacecraft missions and requirement reduction of launch vehicle up-mass. Some of the challenges include controlling cryogenic propellant evaporation and managing the high costs and long schedules associated with the new development of spacecraft hardware. This paper describes a conceptual CPD design that is thermally optimized to achieve extremely low propellant boil-off rates. The CPD design is based on existing launch vehicle architecture, and its thermal optimization is achieved using current passive thermal control technology. Results from an integrated thermal model are presented showing that this conceptual CPD design can achieve propellant boil-off rates well under 0.05% per day, even when subjected to the LEO thermal environment.

Honour, Ryan↗
NASA NTRS2012

Formulation, Casting, and Evaluation of Paraffin-Based Solid Fuels Containing Energetic and Novel Additives for Hybrid Rockets

This investigation studied the inclusion of various additives to paraffin wax for use in a hybrid rocket motor. Some of the paraffin-based fuels were doped with various percentages of LiAlH4 (up to 10%). Addition of LiAlH4 at 10% was found to increase regression rates between 7 - 10% over baseline paraffin through tests in a gaseous oxygen hybrid rocket motor. Mass burn rates for paraffin grains with 10% LiAlH4 were also higher than those of the baseline paraffin. RDX was also cast into a paraffin sample via a novel casting process which involved dissolving RDX into dimethylformamide (DMF) solvent and then drawing a vacuum on the mixture of paraffin and RDX/DMF in order to evaporate out the DMF. It was found that although all DMF was removed, the process was not conducive to generating small RDX particles. The slow boiling generated an inhomogeneous mixture of paraffin and RDX. It is likely that superheating the DMF to cause rapid boiling would likely reduce RDX particle sizes. In addition to paraffin/LiAlH4 grains, multi-walled carbon nanotubes (MWNT) were cast in paraffin for testing in a hybrid rocket motor, and assorted samples containing a range of MWNT percentages in paraffin were imaged using SEM. The fuel samples showed good distribution of MWNT in the paraffin matrix, but the MWNT were often agglomerated, indicating that a change to the sonication and mixing processes were required to achieve better uniformity and debundled MWNT. Fuel grains with MWNT fuel grains had slightly lower regression rate, likely due to the increased thermal conductivity to the fuel subsurface, reducing the burning surface temperature.

Larson, Daniel B.↗
NASA NTRS2013

Large-Scale Liquid Hydrogen Tank Rapid Chill and Fill Testing for the Advanced Shuttle Upper Stage Concept

Cryogenic upper stages in the Space Shuttle program were prohibited primarily due to a safety risk of a 'return to launch site' abort. An upper stage concept addressed this concern by proposing that the stage be launched empty and filled using shuttle external tank residuals after the atmospheric pressure could no longer sustain an explosion. However, only about 5 minutes was allowed for tank fill. Liquid hydrogen testing was conducted within a near-ambient environment using the multipurpose hydrogen test bed 638.5 ft3 (18m3) cylindrical tank with a spray bar mounted longitudinally inside. Although the tank was filled within 5 minutes, chilldown of the tank structure was incomplete, and excessive tank pressures occurred upon vent valve closure. Elevated tank wall temperatures below the liquid level were clearly characteristic of film boiling. The test results have substantial implications for on-orbit cryogen transfer since the formation of a vapor film would be much less inhibited due to the reduced gravity. However, the heavy tank walls could become an asset in normal gravity testing for on-orbit transfer, i.e., if film boiling in a nonflight weight tank can be inhibited in normal gravity, then analytical modeling anchored with the data could be applied to reduced gravity environments with increased confidence.

Flachbart, R. H.↗
NASA NTRS2014

Simulation for Prediction of Entry Article Demise (SPEAD): an Analysis Tool for Spacecraft Safety Analysis and Ascent/Reentry Risk Assessment

For the purpose of performing safety analysis and risk assessment for a probable offnominal suborbital/orbital atmospheric reentry resulting in vehicle breakup, a synthesis of trajectory propagation coupled with thermal analysis and the evaluation of node failure is required to predict the sequence of events, the timeline, and the progressive demise of spacecraft components. To provide this capability, the Simulation for Prediction of Entry Article Demise (SPEAD) analysis tool was developed. This report discusses the capabilities, modeling, and validation of the SPEAD analysis tool. SPEAD is applicable for Earth or Mars, with the option for 3 or 6 degrees-of-freedom (DOF) trajectory propagation. The atmosphere and aerodynamics data are supplied in tables, for linear interpolation of up to 4 independent variables. The gravitation model can include up to 20 zonal harmonic coefficients. The modeling of a single motor is available and can be adapted to multiple motors. For thermal analysis, the aerodynamic radiative and free-molecular/continuum convective heating, black-body radiative cooling, conductive heat transfer between adjacent nodes, and node ablation are modeled. In a 6- DOF simulation, the local convective heating on a node is a function of Mach, angle-ofattack, and sideslip angle, and is dependent on 1) the location of the node in the spacecraft and its orientation to the flow modeled by an exposure factor, and 2) the geometries of the spacecraft and the node modeled by a heating factor and convective area. Node failure is evaluated using criteria based on melting temperature, reference heat load, g-load, or a combination of the above. The failure of a liquid propellant tank is evaluated based on burnout flux from nucleate boiling or excess internal pressure. Following a component failure, updates are made as needed to the spacecraft mass and aerodynamic properties, nodal exposure and heating factors, and nodal convective and conductive areas. This allows the trajectory to be propagated seamlessly in a single run, inclusive of the trajectories of components that have separated from the spacecraft. The node ablation simulates the decreasing mass and convective/reference areas, and variable heating factor. A built-in database provides the thermo-mechanical properties of For the purpose of performing safety analysis and risk assessment for a probable offnominal suborbital/orbital atmospheric reentry resulting in vehicle breakup, a synthesis of trajectory propagation coupled with thermal analysis and the evaluation of node failure is required to predict the sequence of events, the timeline, and the progressive demise of spacecraft components. To provide this capability, the Simulation for Prediction of Entry Article Demise (SPEAD) analysis tool was developed. This report discusses the capabilities, modeling, and validation of the SPEAD analysis tool. SPEAD is applicable for Earth or Mars, with the option for 3 or 6 degrees-of-freedom (DOF) trajectory propagation. The atmosphere and aerodynamics data are supplied in tables, for linear interpolation of up to 4 independent variables. The gravitation model can include up to 20 zonal harmonic coefficients. The modeling of a single motor is available and can be adapted to multiple motors. For thermal analysis, the aerodynamic radiative and free-molecular/continuum convective heating, black-body radiative cooling, conductive heat transfer between adjacent nodes, and node ablation are modeled. In a 6- DOF simulation, the local convective heating on a node is a function of Mach, angle-ofattack, and sideslip angle, and is dependent on 1) the location of the node in the spacecraft and its orientation to the flow modeled by an exposure factor, and 2) the geometries of the spacecraft and the node modeled by a heating factor and convective area. Node failure is evaluated using criteria based on melting temperature, reference heat load, g-load, or a combination of the above. The failure of a liquid propellant tank is evaluated based on burnout flux from nucleate boiling or excess internal pressure. Following a component failure, updates are made as needed to the spacecraft mass and aerodynamic properties, nodal exposure and heating factors, and nodal convective and conductive areas. This allows the trajectory to be propagated seamlessly in a single run, inclusive of the trajectories of components that have separated from the spacecraft. The node ablation simulates the decreasing mass and convective/reference areas, and variable heating factor. A built-in database provides the thermo-mechanical properties of

Ling, Lisa↗
NASA NTRS2014

Flow Visualization of Liquid Hydrogen Line Chilldown Tests

We present experimental measurements of wall and fluid temperature during chill-down tests of a warm cryogenic line with liquid hydrogen. Synchronized video and fluid temperature measurements are used to interpret stream temperature profiles versus time. When cold liquid hydrogen starts to flow into the warm line, a sequence of flow regimes, spanning from all-vapor at the outset to bubbly with continuum liquid at the end can be observed at a location far downstream of the cold inlet. In this paper we propose interpretations to the observed flow regimes and fluid temperature histories for two chilldown methods, viz. trickle (i.e. continuous) flow and pulse flow. Calculations of heat flux from the wall to the fluid versus wall temperature indicate the presence of the transition/nucleate boiling regimes only. The present tests, run at typical Reynolds numbers of approx O(10 (exp 5)), are in sharp contrast to similar tests conducted at lower Reynolds numbers where a well-defined film boiling region is observed.

Line chill down↗
NASA NTRS2014

Insulation Test Cryostat with Lift Mechanism

A multi-purpose, cylindrical thermal insulation test apparatus is used for testing insulation materials and systems of materials using a liquid boil-off calorimeter system for absolute measurement of the effective thermal conductivity (k-value) and heat flux of a specimen material at a fixed environmental condition (cold-side temperature, warm-side temperature, vacuum pressure level, and residual gas composition). The apparatus includes an inner vessel for receiving a liquid with a normal boiling point below ambient temperature, such as liquid nitrogen, enclosed within a vacuum chamber. A cold mass assembly, including the upper and lower guard chambers and a middle test vessel, is suspended from a lid of the vacuum canister. Each of the three chambers is filled and vented through a single feedthrough. All fluid and instrumentation feedthroughs are mounted and suspended from a top domed lid to allow easy removal of the cold mass. A lift mechanism allows manipulation of the cold mass assembly and insulation test article.

Fesmire, James E.↗
NASA NTRS2015

Proof-of-Concept Experiments on a Gallium-Based Ignitron for Pulsed Power Applications

Ignitrons are electrical switching devices that operate at switching times that are on the order of microseconds, can conduct high currents of thousands of amps, and are capable of holding off tens of thousands of volts between pulses. They consist of a liquid metal pool within an evacuated tube that serves both the cathode and the source of atoms and electrons for an arc discharge. Facing the liquid metal pool is an anode suspended above the cathode, with a smaller ignitor electrode tip located just above the surface of the cathode. The ignitron can be charged to significant voltages, with a potential difference of thousands of volts between anode and cathode. When an ignition pulse is delivered from the ignitor electrode to the cathode, a small amount of the liquid metal is vaporized and subsequently ionized, with the high voltage between the anode and cathode causing the gas to bridge the gap between the two electrodes. The electrons and ions move rapidly towards the anode and cathode, respectively, with the ions liberating still more atoms from the liquid metal cathode surface as a high-current plasma arc discharge is rapidly established. This arc continues in a self-sustaining fashion until the potential difference between the anode and cathode drops below some critical value. Ignitrons have been used in a variety of pulsed power applications, including the railroad industry, industrial chemical processing, and high-power arc welding. In addition, they might prove useful in terrestrial power grid applications, serving as high-current fault switches, quickly shunting dangerous high-current or high-voltage spikes safely to ground. The motivation for this work stemmed from the fact that high-power, high-reliability, pulsed power devices like the ignitron have been used for ground testing in-space pulsed electric thruster technologies, and the continued use of ignitrons could prove advantageous to the future development and testing of such thrusters. Previous ignitron designs have used mercury as the liquid metal cathode, owing to its presence as a liquid at room temperatures and a vapor pressure of 10 Pa (75 mtorr) at room temperature. While these are favorable properties, there are obvious environmental and personal safety concerns with the storage, handling, and use of mercury and its compounds. The purpose of the present work was to fabricate and test an ignitron that used as its cathode an alternate liquid metal that was safe to handle and store. To that end, an ignitron test article that used liquid gallium as the cathode material was developed and tested. Gallium is a metal that has a melting temperature of 29.76 C, which is slightly above room temperature, and a boiling point of over 2,300 C at atmospheric pressure. This property makes gallium the element with the largest relative difference between melting and boiling points. Gallium has a limited role in biology, and when ingested, it will be subsequently processed by the body and expelled rather than accumulating to toxic levels. The next section of this Technical Memorandum (TM) provides background information on the development of mercury-based ignitrons, which serves as the starting point for the development of the gallium-based variant. Afterwards, the experimental hardware and setup used in proof-of-concept testing of a basic gallium ignitron are presented. Experimental data, consisting of discharge voltage and current waveforms as well as high-speed imaging of the gallium arc discharge in the gallium ignitron test article, are presented to demonstrate the efficacy of the concept. Discussion of the data and suggestions on improvements for future iterations of the design are presented in the final two sections of this TM.

Ali, H. K.↗
NASA NTRS2012

Design of Accumulators and Liquid/Gas Charging of Single Phase Mechanically Pumped Fluid Loop Heat Rejection Systems

For single phase mechanically pumped fluid loops used for thermal control of spacecraft, a gas charged accumulator is typically used to modulate pressures within the loop. This is needed to accommodate changes in the working fluid volume due to changes in the operating temperatures as the spacecraft encounters varying thermal environments during its mission. Overall, the three key requirements on the accumulator to maintain an appropriate pressure range throughout the mission are: accommodation of the volume change of the fluid due to temperature changes, avoidance of pump cavitation and prevention of boiling in the liquid. The sizing and design of such an accumulator requires very careful and accurate accounting of temperature distribution within each element of the working fluid for the entire range of conditions expected, accurate knowledge of volume of each fluid element, assessment of corresponding pressures needed to avoid boiling in the liquid, as well as the pressures needed to avoid cavitation in the pump. The appropriate liquid and accumulator strokes required to accommodate the liquid volume change, as well as the appropriate gas volumes, require proper sizing to ensure that the correct pressure range is maintained during the mission. Additionally, a very careful assessment of the process for charging both the gas side and the liquid side of the accumulator is required to properly position the bellows and pressurize the system to a level commensurate with requirements. To achieve the accurate sizing of the accumulator and the charging of the system, sophisticated EXCEL based spreadsheets were developed to rapidly come up with an accumulator design and the corresponding charging parameters. These spreadsheets have proven to be computationally fast and accurate tools for this purpose. This paper will describe the entire process of designing and charging the system, using a case study of the Mars Science Laboratory (MSL) fluid loops, which is en route to Mars for an August 2012 landing.

HRS↗
NASA NTRS2015

Cryogenic Orbital Testbed (CRYOTE) Ground Test Article, Final Report

Liquid propulsion has been used since Robert Goddard started developing a liquid oxygen (LO2) and gasoline powered rocket and fired it in 1923 (Ref. 1). In the following decades engineers settled on the combination of liquid hydrogen (LH2) and LO2 as the most efficient propellant combination for in-space travel. Due to their low temperatures (LH2 at 20 K and LO2 at 90 K), they require special handling and procedures. General Dynamics began developing LO2 and LH2 upper stages in 1956 in the form of Centaur, these efforts were soon funded by the Department of Defense in conjunction with NASA (beginning in 1958) (Ref. 2). Meanwhile NASA also worked with McDonnell Douglas to develop the SIV-B stage for the Saturn V rocket. In the subsequent years, the engineers were able to push the Centaur to up to 9 hr of orbital lifetime and the SIV-B to up to 6 hr. Due to venting the resultant boil-off from the high heat loads through the foam insulation on the upper stages, both vehicles remained in a settled configuration throughout the flights, thus the two phases of propellant (liquid and vapor) were separated at a known location. The one exception to this were the Titan/Centaur missions, which thanks to the lower boil-off using three layers of multilayer insulation (MLI), were able to coast unsettled for up to 5.25 hr during direct geosynchronous orbit insertion missions. In the years since there has been a continuous effort to extend the life of these upper stages from hours to days or even months.

Joule-Thompson Effect↗
NASA NTRS2016

Insulation Test Cryostat with Lift Mechanism

A multi-purpose, cylindrical thermal insulation test apparatus is used for testing insulation materials and systems of materials using a liquid boil-off calorimeter system for absolute measurement of the effective thermal conductivity (k-value) and heat flux of a specimen material at a fixed environmental condition (cold-side temperature, warm-side temperature, vacuum pressure level, and residual gas composition). An inner vessel receives liquid with a normal boiling point below ambient temperature, such as liquid nitrogen, enclosed within a vacuum chamber. A cold mass assembly, including upper and lower guard chambers and middle test vessel, is suspended from a lid of the vacuum canister. Each of the three chambers is filled and vented through a single feedthrough. All fluid and instrumentation feedthroughs are mounted and suspended from a top domed lid allowing easy removal of the cold mass. A lift mechanism allows manipulation of the cold mass assembly and insulation test article.

Fesmire, James E.↗
NASA NTRS2017

Structural Heat Intercept, Insulation, and Vibration Evaluation Rig (SHIIVER)

Abstract: NASA is currently investigating methods to reduce the boil-off rate on large cryogenic upper stages. Two such methods to reduce the total heat load on existing upper stages are vapor cooling of the cryogenic tank support structure and integration of thick multilayer insulation systems to the upper stage of a launch vehicle. Previous efforts have flown a 2-layer MLI blanket and shown an improved thermal performance, and other efforts have ground-tested blankets up to 70 layers thick on tanks with diameters between 2 3 meters. However, thick multilayer insulation installation and testing in both thermal and structural modes has not been completed on a large scale tank. Similarly, multiple vapor cooled shields are common place on science payload helium dewars; however, minimal effort has gone into using boil-off to intercept heat on large structural surfaces associated with rocket stages. A majority of the vapor cooling effort focuses on metallic cylinders called skirts, which are the most common structural components for launch vehicles. In order to provide test data for comparison with analytical models, a 4 meter diameter test tank and assembly is currently being designed to include skirt structural systems with integral vapor cooling. In order to give a more representative result for actual applications, the technologies will be designed for an 8.4 m and scaled down to a 4 meter tank. In order to prove that the insulation system and vapor cooling attachment methods are structurally sound, acoustic testing will also be performed on the system. The test tank with insulation and vapor cooled shield installed will be tested thermally in the B2 test facility at NASAs Plum Brook Station both before and after being vibration tested at Plum Brooks Space Power Facility.

Johnson, Wesley↗