Showing papers on "Propellant published in 2008"

The effect of evaporation on the wicking of liquids into a metallic weave

Abstract: Wicking of liquids into porous media is of great importance to many applications. One example are propellant management devices (PMD) used in spacecraft tanks. PMDs are designed to ensure gas free delivery of propellant during all acceleration conditions of the flight. This might be achieved by a metallic weave which is wetted by the propellant and thus prevents gas from entering below a critical bubble point pressure. In the case of cryogenic or volatile liquids the weave may dry out and refilling of the structure becomes an important issue. In this study we analyze the wicking of different liquids into a dry Dutch twilled weave (DTW 200 x 1400) by combining experimental and analytical approaches. Experiments were performed under isothermal and terrestrial conditions to investigate the role of evaporation for the capillary rise. The standard wicking model from Lucas and Washburn is enhanced to account for evaporation and gravity effects, too. By comparing the experimental results with the enhanced wicking model we find good qualitative agreement. It is also noted that evaporation may have a major impact on the wicking process.

Chemical Erosion of Carbon-Carbon/Graphite Nozzles in Solid-Propellant Rocket Motors

Abstract: A comprehensive theoretical/numerical framework is established and validated to study the chemical erosion of carbon-carbon/graphite nozzle materials in solid-rocket motors at practical operating conditions. The formulation takes into account detailed thermofluid dynamics for a multicomponent reacting flow, heterogeneous reactions at the nozzle surface, condensed-phase energy transport, and nozzle material properties. Many restrictive assumptions and approximations made in the previous models have been relaxed. Both metallized and nonmetallized AP/HTPB composite propellants are treated. The predicted nozzle surface recession rates compare well with three different sets of experimental data. The erosion rate follows the trend exhibited by the heat-flux distribution and is most severe in the throat region. H 2 O proved to be the most detrimental oxidizing species in dictating nozzle erosion, followed by much lesser contributions from OH and CO 2 , in that order. The erosion rate increases with increasing chamber pressure, mainly due to higher convective heat transfer and enhanced heterogeneous surface reactions. For nonmetallized propellants, the recession rate is dictated by heterogeneous chemical kinetics because the nozzle surface temperature is relatively low. For metallized propellants, the process is diffusion-controlled due to the high surface temperature. The erosion rate decreases with increasing aluminum content, a phenomenon resulting from reduced concentrations of oxidizing species H 2 O, OH, and CO 2 . The transition from the kinetics-controlled to diffusion-controlled mechanism occurs at a surface temperature of around 2800 K.

A Model of AP/HTPB Composite Propellant Combustion in Rocket-Motor Environments

Abstract: A comprehensive theoretical/numerical model for treating AP/HTPB composite-propellant combustion in a rocket-motor environment is presented. The formulation takes into account the conservation equations in both the gas and condensed phases, and accommodates finite-rate chemical kinetics and variable thermophysical properties. The processes in the two phases are coupled at the surface to determine the propellant burning behavior. An asymptotic analysis based on a large activation-energy approximation for the condensed-phase decomposition is applied to help resolve the combustion wave structure in the interfacial layer. A simplified global reaction is employed to characterize the final diffusion flame between the decomposition products of AP and the pyrolysis products of HTPB. Only laminar flows are considered here, to avoid complications arising from turbulence. A detailed parametric study is conducted on the gas-phase flame structures of AP/HTPB composite propellants. The dependence of burning rate, flame...

Constitutive modeling of solid propellant materials with evolving microstructural damage

Abstract: Solid propellants are composite materials with complex microstructure. In a generic form, the material consists of polymeric binder, crystal oxidizer (e.g., ammonium perchlorate), and fuel particles (e.g., aluminum). Severe stressing and extreme temperatures induce damage which is manifested in particle cracking, dewetting along particle/polymer interfaces, void nucleation and growth. Damage complicates the overall constitutive response of a solid propellant over and above the complexities associated with the differing constitutive properties of the particle and binder phases. Using rigorous homogenization theory for composite materials, we propose a general 3-D nonlinear macroscopic constitutive law that models microstructural damage evolution upon straining through continuous void formation and growth. The law addresses the viscous deformation rate within the framework of additive decomposition of the deformation rate and the concept of back stress is used to improve the model performance in stress relaxation. No restriction is placed on the magnitude of the strains. Experimental data from the standard relaxation and uniaxial tension tests are used to calibrate the model parameters in the case of a high elongation solid propellant. It is emphasized that the model parameters are descriptors of individual phase constitutive response and criticality conditions for particle decohesion which can systematically be determined through experiment. The model is used to predict the response of the material under more complex loading paths and to investigate the effect of crack tip damage on the mechanical behavior of a compact tension fracture specimen.

Metal-CO2 Propulsion for Mars Missions : Current Status and Opportunities

Abstract: *† Metal-CO2 propulsion is less known than other Mars in-situ resource utilization (ISRU) technologies This concept, based on using Martian carbon dioxide as an oxidizer in jet or rocket engines, offers the advantage of no chemical processing for CO2 and thus requires less power consumption than ISRU alternatives In this paper, we review various aspects of metal-CO2 propulsion, such as combustion of magnesium and aluminum in CO2, engine types and characteristics, production of liquid CO2 and metal fuel on Mars, and potential missions Lunar and terrestrial applications are also discussed I Introduction N-SITU Resource Utilization (ISRU) is recognized as an enabling technology for exploration of Mars, which can significantly reduce the mass, cost, and risk of robotic and human missions The critical element in future missions is the large mass of propellant for a Mars ascent vehicle, while power and propellant to accommodate a long stay and mobility on Mars are also important Transportation of propellant from Earth to Mars requires tremendous increase in the initial mass of hardware in low Earth orbit (and hence mission cost) as compared to prior no-return missions Fortunately, Mars possesses resources that can be used for propellant production The Martian atmosphere consisting of 95% CO2 is the obvious and most promising in-situ resource I

Test of B/Ti multilayer reactive igniters for a micro solid rocket array thruster☆

Abstract: In this study, reactive B/Ti multilayer igniters were investigated for the noncontact ignition of a micro solid rocket array thruster in vacuum. When current is supplied to the B/Ti multilayer igniter, the chemical reaction: 2B + Ti → TiB 2 + 1320 cal/g occurs, and sparkles are spread to a distance of several millimeters or more. The B/Ti multilayer igniters with three sizes were fabricated, and tested in six configurations of solid propellant. Although one rocket with ignition charge was ignited successfully, the noncontact ignition of the solid propellant was not achieved. However, the B/Ti multilayer igniters themselves generated small impulses of 10 −6 N s order, suggesting the possibility of self-propulsion.

Controlled organic rankine cycle system for recovery and conversion of thermal energy

Abstract: A system for controlled recovery of thermal energy and conversion to mechanical energy. The system collects thermal energy from a reciprocating engine, specifically from engine jacket fluid and/or engine exhaust and uses this thermal energy to generate a secondary power source by evaporating an organic propellant and using the gaseous propellant to drive an expander in production of mechanical energy. A monitoring module senses ambient and system conditions such as temperature, pressure, and flow of organic propellant at one or more locations; and a control module regulates system parameters based on monitored information to optimize secondary power output. A tertiary, or back-up power source may also be present. The system may be used to meet on-site power demands using primary, secondary, and tertiary power.

Optimal Design of Hybrid Rocket Motors for Microgravity Platform

Abstract: A nested direct/indirect method is used to find the optimal design for a microgravity platform which is based on a hybrid sounding rocket. The direct optimization of the parameters that affect the motor design is coupled with the indirect trajectory optimization to maximize a given mission performance index. A gas-pressure feed system is used, with three different propellant combinations. The feed system exploits a pressurizing gas, namely, helium, when hydrogen peroxide or liquid oxygen is used as an oxidizer. The simplest blowdown design is compared with a more complex pressurizing system, which has an additional gas tank that allows for a phase with constant propellant tank pressure. Only self-pressurization is considered with nitrous oxide; two different models are used to describe the behavior of the tank pressurization. The simplest model assumes liquid/vapor equilibrium. A two-phase model is also proposed: Saturated vapor and superheated liquid are considered and the liquid/vapor mass transfer evaluation is based on the liquid spinodal line. Results show that the different tank-pressurization models yield minimal differences of the optimal motor characteristics. Performance differs slightly due to the different mass of the residual oxidizer. The propellant comparison for the present case shows better performance for hydrogen peroxide/ polyethylene with respect to liquid oxygen/hydroxyl-terminated polybutadiene, while nitrous oxide/hydroxyl-terminated polybutadiene remains attractive for system simplicity and low costs.

Variability of The Rocket Propellants RP-1, RP-2, and TS-5: Application of a Composition- and Enthalpy-Explicit Distillation Curve Method†

Abstract: We have recently introduced several important improvements in the measurement of distillation curves for complex fluids This method is a significant improvement over current approaches, featuring a composition-explicit data channel for each distillate fraction (for both qualitative and quantitative analysis) and an assessment of the energy content of each distillate fraction, among other features The most significant modification is achieved with a new sampling approach that allows precise qualitative as well as quantitative analyses of each fraction, on the fly We have applied the new method to the measurement of a wide variety of fluids, including hydrocarbons, gasoline, jet fuel, diesel fuels (both petroleum-derived and biodiesel), and crude oils In this paper, we present the application of the technique to representative batches of the rocket propellants RP-1, RP-2, and TS-5 We not only present the distillation curves but also utilize the composition-explicit information to characterize distillat

Firearm having a new gas operating system

Abstract: M-16/AR-15 firearms modified with a new gas operating system are described. The modification of the firearm with the new gas operating system has a forward mounted gas and recoil spring system in which high pressure propellant gases from the cartridge expand in the barrel and operate the firearm.

Microstructure of Composite Propellants Using Simulated Packings and X-Ray Tomography

Abstract: C OMPOSITE solid propellants are heterogeneous media in which particles are embedded in a rubber binder. For industrial applications, such particles are commonly ammonium perchlorate (AP) and aluminum (Al), the sizes of which are typically below 100 m. As a general rule, simple mixture laws are often inappropriate for granular systems because they cannot capture the inherent microstructural features. This conclusion also stands for composite propellants, because various macroscale phenomena are strongly linked with the structure at microscale. They may be encountered in the field of detonation (e.g., hot-spot formation during deflagration– detonation transition), combustion (e.g., aluminum agglomeration and fine AP/binder premixed flames), mechanics (e.g., filler/binder adhesion), etc. This obviously motivates an improved knowledge of propellant microstructure and also the ability to “build” representative numericalmodels of structures as input data for detailedmultiphysics modeling (namely, computational fluid dynamics or finite element method codes). Those numerical microstructures are classically randompackings of spheres, which have been studied formany years because they serve as useful models for a variety of physical systems (granular, porous, anamorphous materials, etc.). Yet, random packings should faithfully reflect the features of the actual propellant spatial structure, not only global properties (as volume fraction, for instance), but also second-order spatial statistics. This can only be checked with some in situ experimental characterization of the structure. This work aims at comparing simulated random packings on a typical industrial AP/Al propellant together with experimental microstructure data to evaluate to which extent such simulated packing describe the actual microstructure. Experimental data are obtained through x-ray tomography of a propellant sample. This approach leads to genuine 3D statistical data in a nonintrusive way and allows a relative automatization of data sampling and processing. II. Random Packings

Low toxicity primer compositions for reduced energy ammunition

Abstract: A primer composition with reduced toxicity suited for reduced-energy ammunition comprises bismuth (III) oxide as the principal oxidizer and contains a portion of propellant composition mixed therein. This composition may also be used in a cartridge which is otherwise substantially free of any other propellant compound and preferably produces a residue which is substantially free of toxic substances.

Application of an Aerosol Shock Tube for the Kinetic Studies of n-Dodecane/Nano-Aluminum Slurries

Abstract: Aerosol shock tube methods were used to measure the ignition delay times of slurries of n-dodecane and nano-aluminum particles. The aerosol shock tube has the potential to allow the study of gel propellant chemical kinetics without interference from other physicalchemical processes. A Sono-Tek ultrasonic Spray nozzle was used to produce slurry aerosol droplets which were uniformly dispersed in the driven section of the shock tube. An array of laser extinction and absorption diagnostics was used to track species time histories, droplet breakup and evaporation, and measure ignition delay times. These experiments were conducted behind reflected shock waves for temperatures ranging from 1127 to 1249 K and pressures from 4.7 to 9.5 atm.

A Practical, Affordable Cryogenic Propellant Depot Based on ULA's Flight Experience

Abstract: Mankind is embarking on the next step in the journey of human exploration. We are returning to the moon and eventually moving to Mars and beyond. The current Exploration architecture seeks a balance between the need for a robust infrastructure on the lunar surface, and the performance limitations of Ares I and V. The ability to refuel or top-off propellant tanks from orbital propellant depots offers NASA the opportunity to cost effectively and reliably satisfy these opposing requirements. The ability to cache large orbital quantities of propellant is also an enabling capability for missions to Mars and beyond. This paper describes an option for a propellant depot that enables orbital refueling supporting Exploration, national security, science and other space endeavors. This proposed concept is launched using a single EELV medium class rocket and thus does not require any orbital assembly. The propellant depot provides cryogenic propellant storage that utilizes flight proven technologies augmented with technologies currently under development. The propellant depot system, propellant management, flight experience, and key technologies are also discussed. Options for refueling the propellant depot along with an overview of Exploration architecture impacts are also presented.

The performance of a piezoelectric-sensor-based SHM system under a combined cryogenic temperature and vibration environment

Abstract: A series of tests have been conducted to determine the survivability and functionality of a piezoelectric-sensor-based active structural health monitoring (SHM) SMART Tape system under the operating conditions of typical liquid rocket engines such as cryogenic temperature and vibration loads. The performance of different piezoelectric sensors and a low temperature adhesive under cryogenic temperature was first investigated. The active SHM system for liquid rocket engines was exposed to flight vibration and shock environments on a simulated large booster LOX-H2 engine propellant duct conditioned to cryogenic temperatures to evaluate the physical robustness of the built-in sensor network as well as operational survivability and functionality. Test results demonstrated that the developed SMART Tape system can withstand operational levels of vibration and shock energy on a representative rocket engine duct assembly, and is functional under the combined cryogenic temperature and vibration environment.

1,5‐Diamino‐4‐methyltetrazolium 5‐Nitrotetrazolate – Synthesis, Testing and Scale‐up

Abstract: 1,5-Diamino-4-methyltetrazolium 5-nitrotetrazolate (2b) was synthesized in high yield from 1,5-diamino-4-methyltetrazolium iodide (2a) and highly sensitive silver 5-nitrotetrazolate (AgNT). A safer synthesis, suitable for scale-up, is introduced involving reaction of the previously unreported 1-amino-5-imino-4-methyltetrazole free base (2) with ammonium 5-nitrotetrazolate. Both new compounds (2 and 2b) were fully characterized using vibrational (IR and Raman) and multinuclear NMR spectroscopy (1H, 13C, 14N, 15N), elemental analysis and single crystal X-ray diffraction. The hydrogen-bonding networks of both materials are described in terms of their graph-sets. Compound 2b is hydrolytically stable with a high melting point and concomitant decomposition at 160 °C. The sensitivity of the energetic salt 2b towards impact (>30 J) and friction (>360 N) was tested. The constant volume energy of combustion (ΔcU) of 2b was measured experimentally using bomb calorimetry. In addition, the detonation parameters (detonation pressure and velocity) of the nitrotetrazolate salt were calculated from the energy of formation, the crystal density and the molecular formula using the EXPLO5 computer code (P = 15.5·GPa, D = 6749 m s−1) and are similar to that of TNT and nitroguanidine making 2b of prospective interest in propellant charge formulations or, in combination with a suitable oxidizer, as a solid propellant.

A Comparison of Hydraulic and Propellant Fracture Propagation in a Shale Gas Reservoir

Abstract: This paper describes laboratory and field experiments that were conducted to compare hydraulic and propellant fracturing techniques in the Mancos Shale of Colorado. The Mancos is a Cretaceous shale approximately 2,000 feet thick (1) . Although mainly considered a source rock for other formations, the Mancos is also productive in certain areas. To better understand how stimulation technologies might be used to commercially develop the shale, four unique data sets were compared including results from both laboratory and field hydraulic and propellant fracturing. The results of this work are presented and include the following conclusions: the propellant fractures in the block test were constrained by relative strengths of the strata; the layering of the Mancos shale resulted in better height containment than expected in field test stimulations, although the mechanism of containment in the field tests is still not fully understood; and both propellant and hydraulic fracturing can provide stimulation benefits, but only when applied in appropriate situations.

Nitrous oxide fuel blend monopropellants

Abstract: Compositions and methods herein provide monopropellants comprising nitrous oxide mixed with organic fuels in particular proportions creating stable, storable, monopropellants which demonstrate high ISP performance. Due to physical properties of the nitrous molecule, fuel/nitrous blends demonstrate high degrees of miscibility as well as excellent chemical stability. While the monopropellants are particularly well suited for use as propulsion propellants, they also lend themselves well to power generation in demanding situations where some specific cycle creates useable work and for providing gas pressure and/or heat for inflating deployable materials.

Thrust Stand Micromass Balance for the Direct Measurement of Specific Impulse

Abstract: A technique has been developed to directly measure the specific impulse from pulsed thruster systems. The technique is especially useful for propulsion devices that use solid propellants, for which a direct measurement of the propellant mass flow is extremely difficult. A torsion balance is used with a horizontal axis of rotation. A thruster is placed on the balance such that the impulse of the thruster firing and the change in mass due to the expelled propellant act in the same direction. A combined impulse and steady-state force measurement (due to propellant mass loss) can then be decoupled to assess the ratio of the impulse to the weight of propellant expended, or the specific impulse. A model has been developed to show the utility of the technique for pulsed systems with a firing time less than the natural period of the balance. An experimental proof of principle study was also undertaken using the laser ablation of engineering-grade Buna, Viton, and Teflon propellants. Specific-impulse measurements on the order of 200 s have been demonstrated with this laser ablation thruster.

Transient Burning Rate Model for Solid Rocket Motor Internal Ballistic Simulations

Abstract: A general numerical model based on the Zeldovich-Novozhilov solid-phase energy conservation result for unsteady solid-propellant burning is presented in this paper. Unlike past models, the integrated temperature distribution in the solid phase is utilized directly for estimating instantaneous burning rate (rather than the thermal gradient at the burning surface). The burning model is general in the sense that the model may be incorporated for various propellant burning-rate mechanisms. Given the availability of pressure-related experimental data in the open literature, varying static pressure is the principal mechanism of interest in this study. The example predicted results presented in this paper are to a substantial extent consistent with the corresponding experimental firing response data.

Electrokinetic pumping of liquid propellants for small satellite microthruster applications

Abstract: For most orbital maneuvers, small satellites weighing less than 10 kg require propulsion systems capable of producing thrust in the micro-Newton to milli-Newton force range. At this scale, electrokinetic (EK) pumping offers a method to precisely meter liquid propellants under purely electrical control at pressures and flow rates well-suited for microthruster applications. After exploring a variety of materials and surface treatments for the electrokinetic pumping media, we have demonstrated EK pumping of anhydrous hydrazine using both packed-capillary and larger sintered-monolith pump designs. Hydrogen peroxide has proven difficult to directly pump electrokinetically, but we have shown the utility of delivering peroxide and other electrokinetically incompatible liquids indirectly using in-line reservoirs with fluidic isolation to separate the pump working fluid from the propellant. Directly- and indirectly-pumped propellants have been delivered to novel capillary microthrusters with integrated catalyst beds and plasma-formed micronozzle structures. Specific impulses up to 190 s have been shown for hydrazine in non-optimized capillary thrusters at a mass flow rate of 1.5 mg/s. Controlled thrust pulses with a maximum continuous thrust of 1.5 mN, minimum impulse bit of 7 μN-s, and average specific impulse of 110 s have been demonstrated with the electrokinetically pumped microthruster assembly.

Studies on Triple Base Gun Propellant Based on Two Energetic Azido Esters

Abstract: Triple base propellant is the workhorse propellant because it possesses several advantages like reduced flash, flame temperature, and erosion of the barrel as compared to double and single base propellant. Hence, efforts are going on worldwide to increase its performance by increasing its energy using energetic plasticizers and binders. In the present article, nonenergetic plasticizer dibutyl phthalate (DBP) is replaced by two energetic azido ester plasticizers, tris(azido acetoxy methyl) propane (TAAMP) and bis(azido acetoxy) bis(azido methyl) propane (BABAMP), in triple base composition and their different properties are studied. Experimental closed vessel (CV) results (loading density 0.2 g/cm3) clearly indicate that the triple base composition with 2% DBP has force constant 1018 J/g, which is increased to 1026 and 1030 J/g on replacement of DBP by 1 and 2% TAAMP, respectively. Mechanical properties of propellant compositions containing 1 and 2% TAAMP (compression strength 279 and 291 kgf/cm2, percenta...

Numerical investigations on dynamic process of muzzle flow

Abstract: The integrative process of a quiescent projectile accelerated by high-pressure gas to shoot out at a supersonic speed and beyond the range of a precursor flow field was simulated numerically. The calculation was based on ALE equations and a second-order precision Roe method that adopted chimera grids and a dynamic mesh. From the predicted results, the coupling and interaction among the precursor flow field, propellant gas flow field and high-speed projectile were discussed in detail. The shock-vortex interaction, shockwave reflection, shock-projectile interaction with shock diffraction, and shock focus were clearly demonstrated to explain the effect on the acceleration of the projectile.

Liquid Acquisition Device Testing with Sub-Cooled Liquid Oxygen

Abstract: When transferring propellant in space, it is most efficient to transfer single phase liquid from a propellant tank to an engine. In earth s gravity field or under acceleration, propellant transfer is fairly simple. However, in low gravity, withdrawing single-phase fluid becomes a challenge. A variety of propellant management devices (PMD) are used to ensure single-phase flow. One type of PMD, a liquid acquisition device (LAD) takes advantage of capillary flow and surface tension to acquire liquid. Previous experimental test programs conducted at NASA have collected LAD data for a number of cryogenic fluids, including: liquid nitrogen (LN2), liquid oxygen (LOX), liquid hydrogen (LH2), and liquid methane (LCH4). The present work reports on additional testing with sub-cooled LOX as part of NASA s continuing cryogenic LAD development program. Test results extend the range of LOX fluid conditions examined, and provide insight into factors affecting predicting LAD bubble point pressures.

Solid Rocket Combustion Instability—An American Historical Account

Abstract: T combustion processes are both important in rocket motors and difficult to understand. A wide range of transient processes are of interest, including ignition, quenching, transition from deflagration to detonation, and oscillatory combustion. These problems are often determined partly by factors outside the combustion zone of the propellant (e.g., an igniter, gas oscillations, etc.), but they all involve transient behavior of the combustion process, which is the central theme of this book. Much of the research on steady and transient combustion has been motivated by the problem of combustor instability in rocket motors that sometimes leads to destructive combustion-driven gas oscillations in the combustor. This chapter is a historical review of progress in the oscillatory combustor instability problem. Days of the Magicians

Real gas CFD simulation of supercritical H2-LOX combustion in the Mascotte single-injector combustor using a commercial CFD code

Abstract: Modern high performance rocket combustion engines operate at very high pressures up to 20 MPa. Propellants, typically hydrogen and oxygen, are injected at very low temperatures, below the critical temperature of oxygen (154.581 K), whereas the pressure has a supercritical value (p > pc = 50.43 bar) so that mixing and combustion occur at transcritical conditions. The material properties and equation of state in this region differ significantly from those at low pressures and high temperatures. The paper presents an implementation of real gas models into the commercial CFD-Code ANSYS CFX. This solver is validated with experimental data measured on the Mascotte test rig (V03) at ONERA which have been published in the 2 nd International Workshop on Rocket Combustion Modelling.

VASIMR Performance Measurements at Powers Exceeding 50 kW and Lunar Robotic Mission Applications

Abstract: Efficient plasma production and acceleration is observed in a new high power Variable Specific Impulse Magnetoplasma Rocket (VASIMR) experiment, the VX-100, using argon propellant. The Radio Frequency (RF) power exceeds 50 kW. A 100 % propellant utilization is achieved with ion fluxes up to 1.7×10/sec. We measure an ionization cost of 80±10 eV per ion-electron pair. Bulk argon ion flow velocities are measured up to 20 km/s. Thrust values based on plasma exhaust measurements exceed 1 N. A 200 kW superconducting device, the VX-200, and 150 m vacuum facility are described. We outline the operations concept for a solar-electric lunar cargo tug whose performance is extrapolated from the VX-100 experiment results. Due to the 5,000 second specific impulse of the VASIMR engine, the fraction of the initial mass in low Earth orbit (IMLEO) that arrives in low lunar orbit (LLO) is approximately double that of a chemical propulsion system that performs at a specific impulse of only 450 seconds. The effect of space photovoltaic power cost on the economic advantage of the solar-electric system is examined.