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At least 1,153 records · Page 64

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NASA NTRS2004

Thermodynamic Vent System Test in a Low Earth Orbit Simulation

A thermodynamic vent system for a cryogenic nitrogen tank was tested in a vacuum chamber simulating oxygen storage in low earth orbit. The nitrogen tank was surrounded by a cryo-shroud at -40 F. The tank was insulated with two layers of multi-layer insulation. Heat transfer into cryogenic tanks causes phase change and increases tank pressure which must be controlled. A thermodynamic vent system was used to control pressure as the location of vapor is unknown in low gravity and direct venting would be wasteful. The thermodynamic vent system consists of a Joule-Thomson valve and heat exchanger installed on the inlet side of the tank mixer-pump. The combination is used to extract thermal energy from the tank fluid, reducing temperature and ullage pressure. The system was sized so that the tank mixer-pump operated a small fraction of the time to limit motor heating. Initially the mixer used sub-cooled liquid to cool the liquid-vapor interface inducing condensation and pressure reduction. Later, the thermodynamic vent system was used. Pressure cycles were performed until steady-state operation was demonstrated. Three test runs were conducted at tank fills of 97, 80, and 63 percent. Each test was begun with a boil-off test to determine heat transfer into the tank. The lower tank fills had time averaged vent rates very close to steady-state boil-off rates showing the thermodynamic vent system was nearly as efficient as direct venting in normal gravity.

VanOverbeke, Thomas J.↗

NASA NTRS2004

High Power Density Motors

With the growing concerns of global warming, the need for pollution-free vehicles is ever increasing. Pollution-free flight is one of NASA's goals for the 21" Century. , One method of approaching that goal is hydrogen-fueled aircraft that use fuel cells or turbo- generators to develop electric power that can drive electric motors that turn the aircraft's propulsive fans or propellers. Hydrogen fuel would likely be carried as a liquid, stored in tanks at its boiling point of 20.5 K (-422.5 F). Conventional electric motors, however, are far too heavy (for a given horsepower) to use on aircraft. Fortunately the liquid hydrogen fuel can provide essentially free refrigeration that can be used to cool the windings of motors before the hydrogen is used for fuel. Either High Temperature Superconductors (HTS) or high purity metals such as copper or aluminum may be used in the motor windings. Superconductors have essentially zero electrical resistance to steady current. The electrical resistance of high purity aluminum or copper near liquid hydrogen temperature can be l/lOO* or less of the room temperature resistance. These conductors could provide higher motor efficiency than normal room-temperature motors achieve. But much more importantly, these conductors can carry ten to a hundred times more current than copper conductors do in normal motors operating at room temperature. This is a consequence of the low electrical resistance and of good heat transfer coefficients in boiling LH2. Thus the conductors can produce higher magnetic field strengths and consequently higher motor torque and power. Designs, analysis and actual cryogenic motor tests show that such cryogenic motors could produce three or more times as much power per unit weight as turbine engines can, whereas conventional motors produce only 1/5 as much power per weight as turbine engines. This summer work has been done with Litz wire to maximize the current density. The current is limited by the amount of heat it generates. By increasing the heat transfer out of the wire, the wires can carry a larger current and therefore produce more force. This was done by increasing the surface area of the wire to allow more coolant to flow over it. Litz wire was used because it can carry high frequency current. It also can be deformed into configurations that would increase the surface area. The best configuration was determined by heat transfer and force plots that were generated using Maxwell 2D. Future work will be done by testing and measuring the thrust force produced by the wires in a magnetic field.

Kascak, Daniel J.↗

NASA NTRS2002

Liquid Oxygen Propellant Densification Unit Ground Tested With a Large-Scale Flight-Weight Tank for the X-33 Reusable Launch Vehicle

Propellant densification has been identified as a critical technology in the development of single-stage-to-orbit reusable launch vehicles. Technology to create supercooled high-density liquid oxygen (LO2) and liquid hydrogen (LH2) is a key means to lowering launch vehicle costs. The densification of cryogenic propellants through subcooling allows 8 to 10 percent more propellant mass to be stored in a given unit volume, thereby improving the launch vehicle's overall performance. This allows for higher propellant mass fractions than would be possible with conventional normal boiling point cryogenic propellants, considering the normal boiling point of LO2 and LH2.

Tomsik, Thomas M.↗

NASA NTRS2006

Multiphase Flow Technology Impacts on Thermal Control Systems for Exploration

The Two-Phase Flow Facility (TPHIFFy) Project focused on bridging the critical knowledge gap by developing and demonstrating critical multiphase fluid products for advanced life support, thermal management and power conversion systems that are required to enable the Vision for Space Exploration. Safety and reliability of future systems will be enhanced by addressing critical microgravity fluid physics issues associated with flow boiling, condensation, phase separation, and system stability. The project included concept development, normal gravity testing, and reduced gravity aircraft flight campaigns, in preparation for the development of a space flight experiment implementation. Data will be utilized to develop predictive models that could be used for system design and operation. A single fluid, two-phase closed thermodynamic loop test bed was designed, assembled and tested. The major components in this test bed include: a boiler, a condenser, a phase separator and a circulating pump. The test loop was instrumented with flow meters, thermocouples, pressure transducers and both high speed and normal speed video cameras. A low boiling point surrogate fluid, FC-72, was selected based on scaling analyses using preliminary designs for operational systems. Preliminary results are presented which include flow regime transitions and some observations regarding system stability.

McQuillen, John↗

NASA NTRS2006

Chill Down Process of Hydrogen Transport Pipelines

A pseudo-steady model has been developed to predict the chilldown history of pipe wall temperature in the horizontal transport pipeline for cryogenic fluids. A new film boiling heat transfer model is developed by incorporating the stratified flow structure for cryogenic chilldown. A modified nucleate boiling heat transfer correlation for cryogenic chilldown process inside a horizontal pipe is proposed. The efficacy of the correlations is assessed by comparing the model predictions with measured values of wall temperature in several azimuthal positions in a well controlled experiment by Chung et al. (2004). The computed pipe wall temperature histories match well with the measured results. The present model captures important features of thermal interaction between the pipe wall and the cryogenic fluid, provides a simple and robust platform for predicting pipe wall chilldown history in long horizontal pipe at relatively low computational cost, and builds a foundation to incorporate the two-phase hydrodynamic interaction in the chilldown process.

Mei, Renwei↗

NASA NTRS2006

Cryogenic Tank Modeling for the Saturn AS-203 Experiment

A computational fluid dynamics (CFD) model is developed for the Saturn S-IVB liquid hydrogen (LH2) tank to simulate the 1966 AS-203 flight experiment. This significant experiment is the only known, adequately-instrumented, low-gravity, cryogenic self pressurization test that is well suited for CFD model validation. A 4000-cell, axisymmetric model predicts motion of the LH2 surface including boil-off and thermal stratification in the liquid and gas phases. The model is based on a modified version of the commercially available FLOW3D software. During the experiment, heat enters the LH2 tank through the tank forward dome, side wall, aft dome, and common bulkhead. In both model and test the liquid and gases thermally stratify in the low-gravity natural convection environment. LH2 boils at the free surface which in turn increases the pressure within the tank during the 5360 second experiment. The Saturn S-IVB tank model is shown to accurately simulate the self pressurization and thermal stratification in the 1966 AS-203 test. The average predicted pressurization rate is within 4% of the pressure rise rate suggested by test data. Ullage temperature results are also in good agreement with the test where the model predicts an ullage temperature rise rate within 6% of the measured data. The model is based on first principles only and includes no adjustments to bring the predictions closer to the test data. Although quantitative model validation is achieved or one specific case, a significant step is taken towards demonstrating general use of CFD for low-gravity cryogenic fluid modeling.

Grayson, Gary D.↗

NASA NTRS2006

Deep Space Network, Cryogenic HEMT LNAs

Exploration of the Solar System with automated spacecraft that are more than ten astronomical units (1 AU = 149,597,870.691 km) from earth requires very large antennae employing extremely sensitive receivers. A key figure of merit in the specification of the spacecraft-to-earth telecommunications link is the ratio of the antenna gain to operatio nal noise temperature (G/Top) of the system. The Deep Space Network (DSN) receivers are cryogenic, low-noise amplifiers (LNAs) which addres s the need to maintain Top as low as technology permits. Historicall y, the extra-ordinarily sensitive receive systems operated by the DSN have required ctyogenically cooled, ruby masers, operating at a physi cal temperature near the boiling point of helium, as the LNA. Althoug h masers continue to be used today, they are hand crafted at JPL and expensive to manufacture and maintain. Recent advances in the developm ent of indium phosphide (InP) based high electron mobility transistor s (HEMTs) combined with cryogenic cooling near the boiling point of h ydrogen have made this alternate technology comparable with and a fraction of the cost of maser technology. InP HEMT LNA modules are demons trating noise temperatures less than ten times the quantum noise limi t (10hf/k) from 1 to 100 GHz. To date, the lowest noise LNA modules developed for the DSN have demonstrated noise temperatures of 3.5 K and 8.5 K at 8.5 K at 32 GHz, respectively. Front-end receiver packages employing these modules have demonstrated operating system noise temperatures of 17 K at 8.4 GHz (on a 70m antenna at zenith) and 39 K at 3 2 GHz (on a 34m antenna at zenith). The development and demonstration of cryogenic, InP HEMT based front-end amplifiers for the DSN requir es accurate component and module characterization, and modeling from 1 to 100 GHz at physical temperatures down to 12 K. The characterizati on and modeling begins with the HEMT chip, proceeds to the multi-stag e HEMT LNA module, and culminates with the complete front-end cryogenic receiver package for the antenna. This presentation will provide a n overview of this development process. Examples will be shown for de vices, LNA modules, front-end receiver packages, antennae employing these packages and the improvements to the down-link capacity.

low-noise amplifiers (LNAs)↗

NASA NTRS1997

Liquid propellant densification

Super cooling the cryogenic liquid propellant in a vehicle propellant tank densities the propellant allowing the vehicle propellant tank to carry more fuel in the same volume tank while lowering the vapor pressure and thus the tank operating pressure. Lowering the tank operating pressure reduces the stress and therefore allows the walls of the tank to be thinner. Both the smaller tank volume and thinner tank wall results in an overall smaller and lighter vehicle with increased payload capability. The cryogenic propellant can be supercooled well below the normal boiling point temperature level by transporting the liquid propellant from the vehicle tanks to a ground based cooling unit which utilizes a combination of heat exchanger and compressor. The compressor lowers the coolant fluid bath pressure resulting in a low temperature boiling liquid which is subsequently used to cool the recirculating liquid. The cooled propellant is then returned to the vehicle propellant tank. In addition to reducing the vehicle size and weight the invention also allows location of the vent valve on the ground, elimination of on-board recirculation pumps or bleed systems, smaller and lighter engine pumps and valves, lighter and more stable ullage gas, and significant reduction in tank fill operation. All of these mentioned attributes provide lower vehicle weight and cost.

Lak, Tibor I.↗

NASA NTRS2007

Optimized Heat Interception for Cryogen Tank Support Structure

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, E. R.↗

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.↗