Propellant Losses Research Articles

Experimental measurement and molecular dynamics simulation analysis of thermal aging performance in composite solid propellants

As a complex mixed system, the aging mechanism of high-energy composite materials involves the aging of each component itself, interactions between components, and transformation of microstructural and thermodynamic properties. The study of a single component or performance parameter has great limitations for elaborating the cognitive complexity of aging processes, making it difficult to reveal the underlying principles of physicochemical property transformation. Thus, it can only be applied as an empirical summary. In view of this, this study comprehensively utilized numerical analysis and experimental detection to explore the evolution of micromolecular conformation and changes in macro aging performance from four aspects of composite systems, including binder aging, plasticizer migration, bond interface debonding, and mechanical property transformation. There was good consistency observed between molecular simulation and thermal aging experiments, indicating that there was basically no post curing process for HTPB propellants using TDI as the curing agent. In the early stage of aging, scission chain was the main process and, in the later stage of aging, oxidation crosslinking the main process. Oxidative crosslinking increased the mean square radius of gyration and glass transition temperature of the binder system, reduced the peak value of radial distribution function, and revealed the reason for the change of propellant hardness and CC and CO double bond content. However, chain degradation increased the number of polar groups, enhanced intermolecular polarity, reduced the solubility parameter, and increased the free volume fraction of the mixed system and the dioctyl sebacate diffusion coefficient. This resulted in an increase in the tangent value of the HTPB propellant loss angle from 0.43 to 0.44° and a decrease in the glass transition temperature from 203.3 to 202.0 K. The binding energy at the interfaces of AP/HTPB and Al/HTPB increased first and then decreased, which was consistent with SEM images of bonding interfaces. The joint effect of binding energy and mechanical properties of the binder system promoted the transformation of the stress-strain curve of aging propellant.

Liquid hydrogen tank chill and no-vent fill prediction using computational fluid dynamics

Cryogenic tank chill and fill is an important cryogenic fluid management technology that supports and enables many of NASA’s long-duration space missions. For no-vent fill, the receiver tank pressure remains below the supply tank pressure during the entire duration of the fill so that the tank does not require venting. This is especially advantageous for tank fill operations in low gravity where the position of the liquid is not always known and venting the tank may cause loss of propellant by venting liquid. In lieu of expensive tests conducted on-orbit, accurate computational models capable of predicting receiver tank pressure during cryogenic propellant tank fill may be used to reduce system and propellant mass as well as mission risk. However, these numerical models must be validated or anchored to test data. This study presents computational fluid dynamics (CFD) models with conjugate heat transfer that are used to predict a liquid hydrogen tank chill and no-vent fill ground test conducted at Lewis Research Center (now Glenn Research Center) in 1991. The specific test case chosen for model validation implemented an upward-facing jet near the bottom of a cylindrical 34 liter tank. CFD predictions are compared to experimental measurements of tank pressure, fill level, and wall temperature. The CFD results show reasonable agreement to the test data but overpredict the pressure collapse near the end of the fill despite the liquid jet penetrating the liquid-vapor interface.

Magnetic field optimization of hall thruster with large height-radius ratio for high specific impulse operation

Hall thrusters with large height-radius ratio not only have incalculable application values in reducing the volume and weight of thrusters, but also have the potential advantages of higher discharge performance and longer service life. However, the lower propellant density in the main ionization zone and the higher electron temperature in the channel aggravate the loss of propellant and current under high voltage, and significantly reduce the discharge efficiency under high specific impulse mode. To improve the discharge performance of Hall thrusters with large height-radius ratio under high voltage, an optimization scheme of internally loaded magnetic field was proposed in this work. The simulation results show that under the internally loaded magnetic field, both the ionization zone and the acceleration zone move toward the inside of the channel. Although the ion loss on the walls increases, the higher propellant density at the channel upstream greatly promotes the increase of ionization rate and significantly improves the propellant utilization efficiency. The second zone crossed by magnetic field lines in the channel can be established by the internally loaded magnetic field, which enhances the magnetic field intensity on the inner and outer walls, and reduces the electron temperature near the channel outlet significantly. So that the axial conduction of electrons is effectively restrained and the current utilization efficiency is greatly improved. With the introduction of internally loaded magnetic field, the total efficiency of HEP-1350PM can be increased by 7.2% at 400 V. Moreover, the performance optimization effect brought by the internally loaded magnetic field will be gradually amplified with the increase of discharge voltage, which makes the Hall thruster with large height-radius ratio expected to achieve high-efficiency discharge under higher specific impulse.

Modeling of pressure control in a cryogenic tank by a thermodynamic vent system considering flash evaporation

Thermodynamic vent system (TVS) has been recognized as a promising method for pressure control in cryogenic propellant tanks in microgravity environments. Prior simulations basically treated the liquid in the tanks quasi-static during the depressurization process via a venting operation. However, experimental evidence suggests that the liquid could be metastable and superheated at that moment. Therefore, improved models are needed for a more accurate simulation of the transient liquid-vapor phase change and heat transfer during the TVS depressurization. A new model that considers flash evaporation effect has been developed in the present work. The predicted pressure agrees with the experimental data better than those by conventional models. It is found that the flashing phenomenon plays a significant role in the depressurization, and behaves more intensely in the lower fill level cases. The flashing duration is mainly determined by the width of the pressure control band. The model has also been used to analyze the efficiency of the TVS in reducing propellant loss. It is found that the ineffectiveness of the heat exchanger and the unbalance of heating and cooling to the ullage in a full cycle are the main causes for the efficiency degradation of the TVS. Either increasing heat exchanging area, reducing orifice diameter of the throttling device or narrowing the pressure control band helps to improve the efficiency.

One-step green method to prepare progressive burning gun propellant through gradient denitration strategy

Gradiently denitrated gun propellant (GDGP) prepared by a “gradient denitration” strategy is obviously superior in progressive burning performance to the traditional deterred gun propellant. Currently, the preparation of GDGP employed a tedious two-step method involving organic solvents, which hinders the large-scale preparation of GDGP. In this paper, GDGP was successfully prepared via a novelty and environmentally friendly one-step method. The obtained samples were characterized by FT-IR, Raman, SEM and XPS. The results showed that the content of nitrate groups gradiently increased from the surface to the core in the surface layer of GDGP and the surface layer of GDGP exhibited a higher compaction than that of raw gun propellant, with a well-preserved nitrocellulose structure. The denitration process enabled the propellant surface with regressive energy density and good progressive burning performance, as confirmed by oxygen bomb and closed bomb test. At the same time, the effects of different solvents on the component loss of propellant were compared. The result showed that water caused the least component loss. Finally, the stability of GDGP was confirmed by methyl-violet test. This work not only provided environmentally friendly, simple and economic preparation of GDGP, but also confirmed the stability of GDGP prepared by this method.

Characterization of film-evaporating microcapillaries for water-based microthrusters

Compact micronewton thrusters can provide attitude control for small satellites to increase mission duration and enable constellation flying. Micronewton thrust control can also enhance missions that require precision pointing such as space telescopes, laser interferometers, and laser communication relays. Film-Evaporating MEMS Tunable Array (FEMTA) is a novel micropropulsion device that generates thrust by heating a micron-scale water capillary to induce controlled film-evaporation. Thrust stand tests under high vacuum have shown that FEMTA can produce controllable thrust of 150μN at 70s specific impulse using 0.65W of electrical power and ultra-pure deionized water as propellant. An undesired quiescent propellant loss rate is inherent to the current FEMTA design which limits its life span and reliability. To derive mitigations to this issue, the behavior of the fluid interface within the FEMTA micronozzle was characterized through direct experimentation. A test-bed FEMTA design was created which enabled direct observation of the liquid within the micronozzles and precise control over critical nozzle dimensions. These test-bed devices were used to measure contact angle, Laplace pressure, and total quiescent propellant loss rates for multiple nozzle configurations. Finally, a next generation FEMTA design was derived from the findings of these studies and its propulsive performance was measured under high vacuum on a micronewton thrust stand. Microfabrication was performed at Purdue’s Birck Nanotechnology Center and vacuum testing was completed at Purdue’s High Vacuum Lab.

Ground based experiment and numerical calculation on thermodynamic vent system in propellant tank for future cryogenic propulsion system

In planned planetary explorations, cryogenic liquids such as liquid hydrogen (LH2), liquid oxygen (LOX), and liquefied natural gases (LNG) are used as fuel and oxidants in the propulsion systems of spacecraft. Such explorations require long-term storage of those cryogens, as well as heat insulation technology to protect the heat from the outside and pressure control technology to suppress rises in pressure due to gas evaporation in the propellant tank. However, current vent systems that discharge the evaporated gas to the outside of the spacecraft are suboptimal because the propellant’s uncertain position in the tank when the spacecraft operates in microgravity environments causes a significant loss of propellant during venting. In response, we examined a method using a mixing jet in a thermodynamic vent system (TVS) that adjusts the tank pressure by cooling the inside of the tank and reducing boil-off gas.In this study, a verification experiment on the TVS performed by jet mixing using simulated liquid (LN2) was carried out. Subcooled mixing jets were supplied under vent-free conditions, and it was possible to reduce the temperature and pressure in the test tank. On the other hand, it was found that depending on the liquid level and supply flow rate of the mixing jet, the tip of the mixing jet could not reach the free surface, and reduction of tank pressure could not be realized.

LEU NTP Engine System Trades and Mission Options

The future of human exploration missions to Mars is dependent on solutions to the technology challenges being worked on by the National Aeronautics and Space Administration (NASA) and industry. One of the key architecture technologies involves propulsion that can transport the human crew from Earth orbit to other planets and back to Earth with the lowest risk to crew and the mission. Nuclear thermal propulsion (NTP) is a proven technology that provides the performance required to enable benefits in greater payload mass, shorter transit time, wider launch windows, and rapid mission aborts due to its high specific impulse and high thrust. Aerojet Rocketdyne (AR) has stayed engaged for several decades in working NTP engine systems and has worked with NASA recently to perform an extensive study on using low-enriched uranium NTP engine systems for a Mars campaign involving crewed missions from the 2030s through the 2050s. Aerojet Rocketdyne has used a consistent set of NASA ground rules and they are constantly updated as NASA adjusts its sights on obtaining a path to Mars, now via the Lunar Operations Platform-Gateway. Building on NASA’s work, AR has assessed NTP as the high-thrust propulsion option to transport the crew by looking at how it can provide more mission capability than chemical or other propulsion systems. The impacts of the NTP engine system on the Mars transfer vehicle configuration have been assessed via several trade studies since 2016, including thrust size, number of engine systems, liquid hydrogen stage size, reaction control system sizing, propellant losses, NASA Space Launch System (SLS) payload fairing size impact, and aggregation orbit. An AR study activity in 2018 included examining NTP stages derived from Mars crew mission elements to deliver extremely large cargo via multiple launches or directly off the NASA SLS. This paper provides an update on the results of the ongoing engine system and mission trade studies.

Basic Study on Thermodynamic Vent System in Propulsion System for Future Spacecraft

In planned planetary explorations, cryogenic liquids such as liquid hydrogen (LH2), liquid oxygen (LOX), and liquefied natural gases (LNG) are used as fuel and oxidants in the propulsion systems of spacecraft. Such explorations require long-term storage for those cryogens, as well as heat insulation technology to protect the heat from the outside and pressure control technology to suppress rises in pressure due to evaporation gas in the propellant tank. However, current vent systems that discharge the evaporated gas to the outside of the spacecraft are suboptimal, for the propellant’s uncertain position in the tank when spacecraft operate in microgravity environments causes the significant loss of propellant during venting. In response, we examined a method using jet mixing in a thermodynamic vent system (TVS) that adjusts the tank pressure by cooling the inside of the tank and reducing boil-off gas. To that end, thermal and fluid analyses were conducted to design a test tank system before the ground verification test of the TVS using liquid nitrogen (LN2) as the simulated cryogen. To evaluate our results, we compared experimental and numerical results regarding the formation of thermal stratification. The experimentally performed verification of the TVS function revealed that jet mixing can lower the liquid’s temperature in the tank after the experimental apparatus was modified to supply a stable subcooled mixing jet.

Ground operations demonstration unit for liquid hydrogen initial test results

NASA operations for handling cryogens in ground support equipment have not changed substantially in 50 years, despite major technology advances in the field of cryogenics. NASA loses approximately 50% of the hydrogen purchased because of a continuous heat leak into ground and flight vessels, transient chill down of warm cryogenic equipment, liquid bleeds, and vent losses. NASA Kennedy Space Center (KSC) needs to develop energy-efficient cryogenic ground systems to minimize propellant losses, simplify operations, and reduce cost associated with hydrogen usage. The GODU LH2 project has designed, assembled, and started testing of a prototype storage and distribution system for liquid hydrogen that represents an advanced end-to-end cryogenic propellant system for a ground launch complex. The project has multiple objectives including zero loss storage and transfer, liquefaction of gaseous hydrogen, and densification of liquid hydrogen. The system is unique because it uses an integrated refrigeration and storage system (IRAS) to control the state of the fluid. This paper will present and discuss the results of the initial phase of testing of the GODU LH2 system.

Development of a Ground Operations Demonstration Unit for Liquid Hydrogen at Kennedy Space Center

NASA operations for handling cryogens in ground support equipment have not changed substantially in 50 years, despite major technology advances in the field of cryogenics. NASA loses approximately 50% of the hydrogen purchased because of a continuous heat leak into ground and flight vessels, transient chill down of warm cryogenic equipment, liquid bleeds, and vent losses. NASA Kennedy Space Center (KSC) needs to develop energy-efficient cryogenic ground systems to minimize propellant losses, simplify operations, and reduce cost associated with hydrogen usage. The GODU LH2 project will design, assemble, and test a prototype storage and distribution system for liquid hydrogen that represents an advanced end-to-end cryogenic propellant system for a ground launch complex. The project has multiple objectives and will culminate with an operational demonstration of the loading of a simulated flight tank with densified propellants. The system will be unique because it uses an integrated refrigeration and storage system (IRAS) to control the state of the fluid. The integrated refrigerator is the critical feature enabling the testing of the following three functions: zero-loss storage and transfer, propellant densification/conditioning, and on-site liquefaction. This paper will discuss the test objectives, the design of the system, and the current status of the installation.

Propellant Losses Because of Particulate Emission in a Pulsed Plasma Thruster

Abstract : Propellant inefficiency material in particulate form is characterized in a laboratory pulsed plasma thruster (PPT) operating at 1 Hz with a 204 discharge energy (20 W). Exhaust deposits are collected and analyzed using a combination of a scanning electron microscope with energy dispersive x-ray analysis and microscopic imaging. Teflon(trademark) particulates are observed with characteristic dianietens ranging from over 100 micrometers down to less than 1 micrometer.

Retention of Freon in poly(n-butyl alpha-cyanoacrylate) sprayed films.

AbstractIt has been found that when n‐butyl α‐cyanoacrylate is sprayed on a blood surface from an aerosol containing Freon as the propellant, the polymeric cyanoacrylate film which forms contains 1.4% propellant under the spray conditions used. Studies of the rate of loss of propellant from the film indicate that ⅓ of the Freon initially present in the film still remains after 90 min of exposure of the film to air. Possible implications of these results are discussed.

Ground tests of orbital storability of liquid propellants.

The major source of propellant loss in long-duration space missions will be propellant boiloff caused by heat conduction into the propellant tank through structural and feed members that must penetrate the superinsulation to link the tank to the rocket engine, payload, etc. This paper describes tests of flight-type propellant tanks that were wrapped with multilayer superinsulation, filled with cryogenic (liquid N-2) or earth-storable (methanol or Freon-14) liquids, and tested in a large vacuum chamber with radiant heat inputs simulating a 260-naut-mile, 66.5° earth orbit. Spherical and cylindrical tanks of 2 to 8 ft in diameter were used. Their support systems simulated one in which the structural supports required during ascent or descent can be retracted, leaving only a wire suspension system during extended space flight. Results have been related to expected performance of actual propellants. These predictions suggest that boiloff rates will be less than 4%/yr for propellants other than LH> in spherical tanks of 10-ft diam (D.s) or larger; for H2, D8 > 15 ft may be needed to keep boiloff < 10%/yr.