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Using the results of previous experimental research by means of high-speed video recording, a mathematical model of ignition was developed for a typical gel fuel particle, based on an organic polymer thickener, in a high-temperature motionless air medium. The structure of such fuel is a rather dense polymer matrix with fine droplets of combustible liquid in the cells. The advantages of gel fuels over liquid rocket propellants dictate their prospects as an energy resource in aerospace industry. The mathematical model of the process under study was developed using the mathematical tools of continuum mechanics and chemical kinetics. It is represented by a system of time-dependent nonlinear partial differential equations with the corresponding initial and boundary conditions. The developed mathematical model describes the process in terms of a limiting mode: the rate of heat supply from the source to the fuel and combustible gas-vapor mixture is limited. This implies that the typical heating time is much longer than the time of chemical reactions. In this case, the factors limiting the process intensity are heat supply and diffusion. Exothermic reactions can be assumed to develop in equilibrium with due consideration of corresponding conditions in the gas medium around the fuel particle. Satisfactory verification results for the mathematical model and numerical algorithm made is possible to conclude that this approach can be used to reliably predict the ignition characteristics of such kind of gel fuels. Gel particles 0.25–2.00 mm in size were considered. They were heated in an air medium at 750–1473 K. The ignition delay times under such conditions ranged from 0.3 to 10.0 s.

A gas generating propellants are used as initiators of liquid rocket propellants turbopumps and have as desired characteristic a high-volume production of low-temperature gas. In this context, some formulations of composite propellant containing polyurethane (based on liquid hydroxyl-terminated polybutadiene), guanidine nitrate, ammonium perchlorate, and additives were evaluated and characterized in order to verify their potential as gas generator propellant, as well as to evaluate the influence of additives on mechanical properties. The formulations were prepared, analyzed, and tested for mechanical properties.

Liquid rocket propellants are subject to wide ranges of temperature and pressure when flowing through high performance liquid rocket engines. For example, as rocket propellant (RP) is used for regenerative cooling, near-wall fluid temperatures can exceed 700 K. The pressure can reach as high as 100 MPa when exiting the fuel pump. Therefore, accurate thermophysical property data, particularly density, are needed to accurately model the flow in rocket engines using computational fluid dynamics (CFD) and other engineering design tools. In this work, we report the density of two RP-2 samples at temperatures to 573 K and pressures to 100 MPa. A high-temperature, high-pressure (HTHP) variable-volume, windowed densimeter is used to determine the density. The experimental density data are satisfactorily correlated by the modified Tait equation that provides a means for interpolating the density within the experimental conditions investigated in this study. The HTHP perturbed-chain statistical associating fluid theory (PC-SAFT), Peng-Robinson (PR), and HTHP volume-translated PR equations of state are used to model RP-2 density over the entire temperature and pressure ranges.

Today's rocket chemical propulsion systems for in-orbit use rely on toxic liquid rocket propellants. The need for environmentally-friendly spacecraft and upper-stage high-performance engines can be seen. This paper covers the topic of optimisation of Geostationary Transfer Orbit apogee bipropellant rocket engines, using highly concentrated hydrogen peroxide as oxidiser and kerosene as fuel. Performance, structural and heat transfer analyses are described. In particular a detailed mass model for hydrogen peroxide/kerosene rocket propulsion systems is discussed. Special care is given to bipropellant rocket systems using staged combustion configurations where a catalyst bed is utilised. The optimisation process was conducted using Matlab software. The ultimate goal of this work was the development of a tool for hydrogen peroxide/hydrocarbon bipropellant rocket propulsion system optimisation in terms of given requirements and design constraints. The presented software may enable the development of advanced, environmentally-friendly satellites and highly-efficient architectures utilising storable green propellants, including small upper-stage propulsion systems.

The present research dealt with the concept of green hypergolic combination to replace the toxic hypergolic combinations. Hydrogen peroxide was selected as a green oxidizer. A novel recipe for the non-toxic hypergolic fuel (Stock 3) was suggested. Sodium borohydride was blended into the mixture of energetic hydrocarbon solvents as an ignition source for hypergolic ignition. The main ingredient of the mixture was diethylenetriamine. By mixing some amount of tetrahydrofuran with diethylenetriamine, the mixture became more flammable and volatile. The mixture of Stock 3 fuel remained stable for four months in the lab scale storability test. Through a simple drop test, the hypergolicity of the green hypergolic combination was verified. Comparing to the toxic hypergolic combination MMH/NTO as the reference, the theoretical performance of the green hypergolic combination would be achieved about 96.7% of the equilibrium specific impulse and about 105.7% of the density specific impulse. The applicability of the green hypergolic combination was successfully confirmed through the static hot-fire tests using 500 N scale hypergolic thruster.

A novel experimental method to evaluate the ignition characteristics of hypergolic propellants was investigated. Three hypergolic liquid fuels with different ignition characteristics were prepared, and their reaction rates with hydrogen peroxide were measured in a microreactor fabricated using a microelectromechanical systems lithography process. To prevent evaporation and boiling by the reaction heat, a low-concentration of hydrogen peroxide of 34.5 wt % was used as a reference oxidizer. The reaction rates of the hypergolic fuels were obtained from temperature measurements in the microreactor and correlated with the ignition delays that were separately measured by drop tests with high-test peroxide. The reaction rate measurements were fully compatible with the drop test results and proved to be useful as an alternative method for the selection of hypergolic propellants by ignition characteristics.

Metal hydrides are regarded as a series of promising hydrogen-supplying fuel for solid rocket propellants. Their effects on the energetic and combustion performances of propellants are closely related to their reaction mechanisms. Here we report a first attempt to determine the reaction mechanism of ZrH2, a high-density metal hydride, in the combustion of a double-base propellant to evaluate its potential as a fuel. ZrH2 is determined to possess good resistance to oxidation by nitrocellulose and nitroglycerine. Thus its combustion starts with dehydrogenation to generate H2 and metallic Zr. Subsequently, the newly formed Zr and H2 participate in the combustion and, especially, Zr melts and then combusts on the burning surface which favors the heat feedback to the propellant. This phenomenon is completely different from the combustion behavior of the traditional fuel Al, where the Al particles are ejected off the burning surface of the propellant to get into the luminous flame zone to burn. The findings in this work validate the potential of ZrH2 as a hydrogen-supplying fuel for double-base propellants.

A transition from hydrazine to less environmentally hazardous monopropellants with higher specific characteristics is considered to be perspective for thermocatalytic thrusters which are used for spacecraft attitude control and station keeping. Such propellants are commonly named as “green propellants”. Hydroxylammonium nitrate can be used as a basis for such a propellant. Hydroxylammonium nitrate-based propellant has a higher density, higher specific impulse and a lower freezing temperature compared to hydrazine. The high combustion temperature of the propellant (above 1800 0С) and strong oxidation properties of the combustion products impose specific requirements to the thruster design materials used. A laboratory model of the К100E thruster has been developed. The product operating capability with high dynamic characteristics has been demonstrated.

Two-dimensional transient laminar natural convection heat transfer in partially filled liquid propellant tanks, solving vorticity and energy equations

Gelled propellants for rocket propulsion present various advantages, such as storability with regard to liquid rocket propellants and a high specific impulse with regard to solid rocket propellants. However, one of the main limitations of gelled propellants in rocket propulsion is their low mixing performance resulting from their high viscosity. For these reasons, conventional injectors such as coaxial and pintle injectors are not suitable for spraying gelled propellants. To address this issue, we have herein devised new methods to increase the mixing performance. For the gelled propellant, liquid kerosene was compounded with 5 wt% of Thixatrol ST, a gelling agent, and shear viscosity data according to the shear rate were measured to confirm viscosity. Two types of injectors—flat type and deflector—were used to confirm the spray characteristics, such as spray angle, via the backlight image technique. The spray angle was formulated according to the total momentum ratio for comparison with the measured experimental data, and the result regarding the measured spray angle in the deflector type injector satisfies the expected spray angle. The dominant difference in spray characteristics between the flat type and deflector type injectors is in terms of a liquid jet stream, related to the breakup and atomization processes. In the case of the flat type injector, the liquid jet stream was detected even for a high total momentum ratio. However, in the case of the deflector type injector, the liquid jet stream disappeared with an increase in the total momentum ratio.

Hypergolic fuel has unique importance in missiles and satellite launch vehicles. Hydrazine and its derivatives (Monomethyl hydrazine, MMH and Unsymmetrical Dimethyl Hydrazine, UDMH) are the continuous choices of fuel in bipropellant though it is carcinogenic and toxic. Development of new, green and cost-effective fuel blends based on trust-worthy, ionic liquid and biofuel in different volume proportion have been reported. The blends exhibit reasonably good stability (chemical and thermal), low viscosity (η 1 g/cm3), low ignition delay and high performance in comparison with UDMH - a proven liquid fuel presently being used as promising fuels for liquid rocket application. The reported blends appear to be a technologically promising, eco-friendly and affordable fuel blend to replace currently used liquid toxic fuels.

Currently, toxic and volatile hydrazine derivatives are still the main fuel choices for liquid bipropellants, especially in some traditional rocket propulsion systems. Therefore, the search for safer hypergolic fuels as replacements for hydrazine derivatives has been one of the most challenging tasks. In this study, six imidazolylidene-borane compounds with zwitterionic structure have been synthesized and characterized, and their hypergolic reactivity has been studied. As expected, these compounds exhibited fast spontaneous combustion upon contact with white fuming nitric acid (WFNA). Among them, compound 5 showed excellent integrated properties including wide liquid operating range (-70-160 °C), superior loading density (0.99 g cm(-3) ), ultrafast ignition delay times with WFNA (15 ms), and high specific impulse (303.5 s), suggesting promising application potential as safer hypergolic fuels in liquid bipropellant formulations.

While shedding light on green space propulsion, the development of non-toxic hypergolic bipropellants has provided alternatives to the conventional, highly toxic propellants. Non-toxic energetic hydrocarbon fuels promoted by sodium borohydride were produced, and their hypergolicity with rocket grade hydrogen peroxide was confirmed. Ignition delays ranged from 4 ms to 17 ms, depending on the oxidizer concentration. A 500 N scale thruster utilizing a non-toxic hypergolic bipropellant combination was also developed. With an oxidizer-to-fuel ratio of 5.48, the injector was designed with pentad unlike-impinging jets, one fuel jet orifice and four oxidizer jet orifices for enhancing the mix and atomization of the hypergolic propellants. The operation test was performed, and burning time was established to be 3.5 s. For the demonstrator, the rising time was approximately 158 ms, and the pressure oscillation inside the combustion chamber was approximately ±7.69%. These results indicate the feasibility of this novel concept engine for future space missions.

With the increasing accessibility of commercial space flight, the environmental impacts of space launches will become increasingly significant in the coming years. Here, for the first time, a review is presented of the environmental impacts of space launches, specifically of emissions from commonly used solid and liquid rocket propellants. While there are a number of environmental impacts resulting from the launch of space vehicles, the depletion of stratospheric ozone is the most studied and most immediately concerning. Solid rocket motors are the subject of most of the environmental studies on rocket launches, while the now more commonly used liquid rocket propellants are underrepresented in the literature. The limited studies of emissions from rocket engines using liquid propellent reveal that while they do result in stratospheric ozone loss, solid rocket motors are responsible for orders of magnitude greater loss. The comparison of commonly used propellants highlights the environmental trade-offs that must be made when selecting a launch system. This review highlights the need for further study of the cumulative impacts that frequent space launches have on all areas of the environment, including global climate, ecosystem toxicity, and human toxicity, and with consideration given to all commonly used propellants, to ensure that the impacts are well characterised and well understood before the number of launches greatly increases.

A possibility of ballistic efficiency of space rockets with main rocket propulsion engines by means of developing a system of gasification of unburned residues in propellant tanks with the use of gas-generating compositions with self-sustained combustion is considered. The criterion for choosing gas-generating compositions is the increment of the characteristic velocity of the rocket stage acquired due to combustion of gasified propellant residues. The results of the present study show that the proposed method of gasification of liquid propellant residues increases the energy characteristics of the rocket.

Propellant grade hydrazine was prepared with 64 per cent yield and 95.5 per cent purity. Purity of the propellant grade hydrazine was determined using wet chemical, gas chromatographic (GC) and eudiometric methods. It was observed that the compositions containing blends of hydrazine-methyl alcohol-ammonium nitrate and hydrazine-methyl alcohol-ammonium perchlorate were not found to be frozen even after cooling to -65 °C for 30 minutes. Mono and bi-propellant thrusters were designed and developed to demonstrate the performance of prepared propellant grade hydrazine as a promising rocket fuel. Five static tests with 22 N thruster and one static test with 1 N thruster were performed successfully in mono-propellant mode. The hurdles of chamber pressure oscillations were overcome by compact packing of the catalyst. The desired decomposition and chamber pressure were achieved. One static test was performed successfully with 60 N bi-propellant thruster. The desired chamber pressure and thrust were achieved. The combustion was smooth and C* achieved was higher than that of UH-25, N2O4 combination. The performance of prepared propellant grade hydrazine shows it as a promising rocket fuels.Defence Science Journal, Vol. 65, No. 1, January 2015, pp.31-38, DOI:http://dx.doi.org/10.14429/dsj.65.7986

NASA Marshall Space Flight Center's Materials and Processes Department, with support from the Propulsion Systems Department, has renewed the development and maintenance of a hybrid test bed for exposing ablative thermal protection materials to an environment similar to that seen in solid rocket motors (SRM). The Solid Fuel Torch (SFT), operated during the Space Shuttle program, utilized gaseous oxygen for oxidizer and an aluminized hydroxyl-terminated polybutadiene (HTPB) fuel grain to expose a converging section of phenolic material to a 400 psi, 2-phase flow combustion environment. The configuration allows for up to a 2 foot long, 5 inch diameter fuel grain cartridge. Wanting to now test rubber insulation materials with a turn-back feature to mimic the geometry of an aft dome being impinged by alumina particles, the throat area has now been increased by several times to afford flow similarity. Combined with the desire to maintain a higher operating pressure, the oxidizer flow rate is being increased by a factor of 10. Out of these changes has arisen the need to characterize the fuel/oxidizer combination in a higher mass flux condition than has been previously tested at MSFC, and at which the literature has little to no reporting as well. For (especially) metalized fuels, hybrid references have pointed out possible dependence of fuel regression rate on a number of variables: mass flux, G - oxidizer only (G0), or - total mass flux (Gtot), Length, L, Pressure, P, and Diameter, D.

This paper presents experimental design and test results of the recently concluded 1-g inverted vertical outflow testing of two 325x2300 full scale liquid acquisition device (LAD) channels in liquid hydrogen (LH2). One of the channels had a perforated plate and internal cooling from a thermodynamic vent system (TVS) to enhance performance. The LADs were mounted in a tank to simulate 1-g outflow over a wide range of LH2 temperatures (20.3 - 24.2 K), pressures (100 - 350 kPa), and flow rates (0.010 - 0.055 kg/s). Results indicate that the breakdown point is dominated by liquid temperature, with a second order dependence on mass flow rate through the LAD. The best performance is always achieved in the coldest liquid states for both channels, consistent with bubble point theory. Higher flow rates cause the standard channel to break down relatively earlier than the TVS cooled channel. Both the internal TVS heat exchanger and subcooling the liquid in the propellant tank are shown to significantly improve LAD performance.

A review of the literature pertaining to hypergolic fuel systems, particularly using hydrazine or its derivatives and hydrogen peroxide, has been conducted. It has been shown that a large effort has been made towards minimizing the risks involved with the use of a toxic propellant such as the hydrazine. Substitution of hydrazines for nontoxic propellant formulations such as the use of high purity hydrogen peroxide with various types of fuels is one of the major areas of study for future hypergolic propellants. A series of criteria for future hypergolic propellants has been recommended, including low toxicity, wide temperature range applicability, short ignition delay, high specific impulse or density specific impulse, and storability at room temperature.

Since September 2001, NASA's In-Space Propulsion Technology (ISPT) program has been developing technologies for lowering the cost of planetary science missions. Recently completed is the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. Two other cost saving technologies nearing completion are the NEXT ion thruster and the Aerocapture technology project. Under development are several technologies for low-cost sample return missions. These include a low-cost Hall-effect thruster (HIVHAC) which will be completed in 2011, light-weight propellant tanks, and a Multi-Mission Earth Entry Vehicle (MMEEV). This paper will discuss the status of the technology development, the cost savings or performance benefits, and applicability of these in-space propulsion technologies to NASA's future Discovery, and New Frontiers missions, as well as their relevance for sample return missions.

Use of green propellants is a trend for future space propulsion. Hypergolic ionic liquid propellants, which are environmentally-benign while exhibiting energetic performances comparable to hydrazine, have shown great potential to meet the requirements of developing nontoxic high-performance propellant formulations for space propulsion applications. This Concept article presents a review of recent advances in the field of ionic liquid propellants.

Human exploration missions to Mars or other destinations in the solar system require large quantities of propellant to enable the transportation of required elements from Earth's sphere of influence to Mars. Current and proposed launch vehicles are incapable of launching all of the requisite mass on a single vehicle; hence, multiple launches and in-space aggregation are required to perform a Mars mission. This study examines the potential of reusable chemical propulsion stages based in cis-lunar space to meet the transportation objectives of the Evolvable Mars Campaign and identifies cis-lunar propellant supply requirements. These stages could be supplied with fuel and oxidizer delivered to cis-lunar space, either launched from Earth or other inner solar system sources such as the Moon or near Earth asteroids. The effects of uncertainty in the model parameters are evaluated through sensitivity analysis of key parameters including the liquid propellant combination, inert mass fraction of the vehicle, change in velocity margin, and change in payload masses. The outcomes of this research include a description of the transportation elements, the architecture that they enable, and an option for a campaign that meets the objectives of the Evolvable Mars Campaign. This provides a more complete understanding of the propellant requirements, as a function of time, that must be delivered to cis-lunar space. Over the selected sensitivity ranges for the current payload and schedule requirements of the 2016 point of departure of the Evolvable Mars Campaign destination systems, the resulting propellant delivery quantities are between 34 and 61 tonnes per year of hydrogen and oxygen propellant, or between 53 and 76 tonnes per year of methane and oxygen propellant, or between 74 and 92 tonnes per year of hypergolic propellant. These estimates can guide future propellant manufacture and/or delivery architectural analysis.

Over the past few decades, toxic and highly volatile hydrazine derivatives have been the main fuel choices for liquid bipropellants, especially in traditional hypergolic rocket engines. The search for new hypergolic fuels as replacements for hydrazine derivatives is of great interest to researchers. In this study, a series of N-alkylimidazole borane compounds has been synthesized and characterized. Interestingly, these compounds display promising applications as potential hypergolic fuels owing to their excellent physiochemical properties including low melting points, high thermal stability, low viscosities, and unique hypergolic reactivity. Compared with popular hypergolic ionic liquids, the cost-effective and scaling-up advantages of these materials highlight their promising potential as high-performance fuels in liquid bipropellant formulations.

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 (PMDs) 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. The present work reports on testing with liquid oxygen (LOX) at elevated pressures (and thus temperatures) (maximum pressure 1724 kPa and maximum temperature 122 K) as part of NASA’s continuing cryogenic LAD development program. These tests evaluate LAD performance for LOX stored in higher pressure vessels that may be used in propellant systems using pressure fed engines. Test data shows a significant drop in LAD bubble point values at higher liquid temperatures, consistent with lower liquid surface tension at those temperatures. Test data also indicates that there are no first order effects of helium solubility in LOX on LAD bubble point prediction. Test results here extend the range of data for LOX fluid conditions, and provide insight into factors affecting predicting LAD bubble point pressures.

Among the many factors that determine overall rocket performance, propellant density is important because it affects the size of the rocket. Thus, in order to decrease the size of a rocket, it may be desirable to increase the density of propellants. This study analyzes the concept of increasing the propellant density by employing a cooling source submerged in the liquid propellant. A simple, mathematical model was developed to predict the rate of densification and the propellant temperature profile. The mathematical model is generic and applicable to multiple propellants. The densification rate was determined experimentally by submerging a cooling source in liquid oxygen at constant, positive pressure, and measuring the time rate of change in temperature with respect to vertical position. The results from the mathematical model provided a reasonable fit when compared to experimental results.

Solid propellant gas generators play a role as a turbo-pump starter in liquid propellant propulsion systems by supplying pressurized gas to power turbines for engine start. Among the required combustion gas properties provided by solid propellant gas generators, the combustion gas temperature should not exceed a certain temperature which may damage the turbine blades. For such purposes, phase stabilized ammonium nitrate (AN)-based propellants have been widely used with a low combustion temperature. However, gas generator propellants with ammonium nitrate have historically exhibited incomplete combustion resulting in increased flame temperatures differing significantly from equilibrium values. In consideration of design requirements, an engineering model of solid propellant gas generator was manufactured using the combustion gas properties calculated by a chemical equilibrium code and then hot-fire tests were performed. Procedures for the correction of T 0 , k and M w of the combustion gas from the experimental results are introduced and the following effects on the design of the solid propellant gas generator are presented. From the experimental correction of the combustion gas properties, it is found that the amount of the propellant could be reduced while providing the same amount of available power to the turbines and consequently, the size of the gas generator could also be decreased.

The present study focuses on the combustion of liquid propellants used in engines for space propulsion applications (satellites and launch vehicles). The combustion of the following propellant combinations was studied: hydrogen/oxygen, hydrocarbon fuel/oxygen, hydrocarbon fuel/hydrogen peroxide, hydrazine/nitrogen tetroxide, Mono-methyl hydrazine/nitrogen tetroxide and Unsymmetrical dimethyl hydrazine/nitrogen tetroxide. The purpose of this paper is to determine the combustion flame temperature and the other thermochemical parameters (combustion products politropic parameter and molecular mass) as function of mixture ratio by mass of oxidizer to fuel (O/F). Furthermore, the vacuum specific impulse was also calculated at assumed pressure conditions of propulsion engines to show the effect of mixture ratio of the propellants on the performances. For the determination of the equilibrium composition at assumed temperature, nonlinear systems of equations were solved numerically using Lieberstein’s method. Some results were presented and compared with previous research in this area, the comparison shows good agreement between the results and the difference is less than 5%. As example, three cases of bi-propellant engines were studied using the programs of combustion and frozen fluid flow approximation: Space Shuttle Main Engine (SSME), Zenit second stage engine (RD-120) and Ariane-5 upper stage engine (Aestus).

This paper presents the hypergolic ignition test results of a potential environmentally friendly liquid propellant consisting of hydrogen peroxide oxidizer (with a concentration of 85%) and ethanolamine fuel for use in rocket engines. Open cup drop tests were conducted to study the effect of amount of metal salt catalyst in fuel and the initial temperatures of fuel and oxidizer on the ignition delay time. To test the hypergolic ignition of bipropellant formulation in a real rocket engine environment, a pressure-fed liquid propellant rocket engine (LPRE) was designed and developed. During the tests it was found that the amount of catalyst and the initial temperature of the fuel had a significant effect on the ignition delay of hypergolic bipropellant. However, the oxidizer temperature seemed to have almost no affect on the ignition delay. There was also significant difference between the ignition delay times from open cup tests and those from rocket engine static firings.

In the field of rocket propulsion system hydrogen peroxide has been used as mono-propellant and as the oxidizer of bi-propellants. At the beginning, hydrogen peroxide was used as mono-propellant for thrusters, but later it had been replaced by hydrazine, which has better specific impulse and storability. On the other hand, to drive turbo-pumps, hydrogen peroxide is still being utilized. As the oxidizer of bi-propellants it was used until 1970`s and from 1990`s hydrogen peroxide once again got back to developer`s interest, because one of the recent development purposes of rocket propulsion system is low-cost and ecologically-clean. Until now the storability of hydrogen peroxide has been remarkably improved. The combination of Kerosene/ also shows similar accelerating performance to Kerosene/ combination because of higher propellant density and higher O/F ratio, even though the propulsion performance is not as good as the combination of Kerosene/. Moreover, its combustion products are much cleaner than Kerosene/ combination.

An experimental study was conducted to evaluate the ignitibility of a family of potential gas-generator solid propellants under different CO 2 laser heat fluxes and chamber operating conditions. Six types of solid propellant were tested: one baseline and five variations. The ingredients of the baseline propellant were ammonium nitrate, guanidine nitrate, potassium nitrate, and polyvinyl alcohol. The other five propellant variations were obtained by adding different amounts of carbon black, ammonium perchlorate, RDX, and triamino guanidine nitrate to the baseline formulation. The propellants were tested in a high-pressure windowed strand burner, which was pressurized with either air or argon purge gases. Propellant ignition was accomplished by a 50-W CO 2 laser. The ignition delay times of the propellants were measured at pressures of 1 and 69 atm using high-speed video cameras. In this paper, the test results and comparison of the ignition and combustion characteristics of these propellants were presented.

The behavior of generalized diffusion coefficients and diffusion flows of individual substances in the multicomponent gas mixture is investigated. The combustion products of liquid rocket propellants being used are considered.

A solid rocket propellant based on glycidyl azide polymer (GAP) binder plasticized with nitrate esters and oxidized with a mixture of ammonium nitrate (AN) and triaminoguanidine nitrate (TAGN) was formulated and characterized. Non-lead ballistic modifiers were also included in order to obtain a propellant with non-acidic and non-toxic exhaust. This propellant was found to exhibit a burning rate approximately twice that of standard GAP/AN propellants. The exponent of the propellant is high compared to commonly used composite propellants but is still in the useable range at pressures below 13.8 MPa. This propellant may present a good compromise for applications requiring intermediate burn rate and impulse combined with low-smoke and non-toxic exhaust.

The goal of the project was to understand how changes in the rate of energy addition can be used to alter the combustion instability characteristics of liquid rocket engines. Fuels with increased energy, either due to higher heats of formation or energetic additives, presumably result in higher performance. This study seeks to understand how changes in combustion rate, due to fuel chemistry changes, might be used to develop high-performing, stable rocket engines. The overall objective of the project was to develop a fundamental understanding of how the spatial distribution of combustion and its temporal response to pressure oscillations dependson kinetic rates, flammability limits, and energy release density. The study combines basic drop combustion experiments, an in situ study using a spontaneously unstable model rocket combustor, and associated modeling of particles in combustion chamber gas flows.

NASA's exploration roadmap is focused on developing technologies and performing precursor missions to advance the state of the art for eventual human missions to Mars. One of the key components of this roadmap is various robotic missions to Near-Earth Objects, the Moon, and Mars to fill in some of the strategic knowledge gaps. The Resource Prospector (RP) project is one of these robotic precursor activities in the roadmap. RP is a multi-center and multi-institution project to investigate the polar regions of the Moon in search of volatiles. The mission is rated Class D and is approximately 10 days, assuming a five day direct Earth to Moon transfer. Because of the mission cost constraint, a trade study of the propulsion concepts was conducted with a focus on available low-cost hardware for reducing cost in development, while technical risk, system mass, and technology advancement requirements were also taken into consideration. The propulsion system for the lander is composed of a braking stage providing a high thrust to match the lander's velocity with the lunar surface and a lander stage performing the final lunar descent. For the braking stage, liquid oxygen (LOX) and liquid methane (LCH4) propulsion systems, derived from the Morpheus experimental lander, and storable bi-propellant systems, including the 4th stage Peacekeeper (PK) propulsion components and Space Shuttle orbital maneuvering engine (OME), and a solid motor were considered for the study. For the lander stage, the trade study included miniaturized Divert Attitude Control System (DACS) thrusters (Missile Defense Agency (MDA) heritage), their enhanced thruster versions, ISE-100 and ISE-5, and commercial-off-the-shelf (COTS) hardware. The lowest cost configuration of using the solid motor and the PK components while meeting the requirements was selected. The reference concept of the lander is shown in Figure 1. In the current reference configuration, the solid stage is the primary provider of delta-V. It will generate 15,000-lbf of thrust with a single burn of ~ 80's seconds. The lander stage is a bi-propellant, pressure-regulated, pulsing liquid propulsion system to perform all other functions.

Green (low toxicity) liquid rocket propellants have become attractive in recent years due to the features of the low cost and less environmental impact. However, the green propellants, such as kerosene/H2O2, usually have different operational conditions (i.e. relatively high O/F ratio) compared to conventional propellants because of their chemical properties. In this research, a new concept of the two-stage impinging type injector (O-F-F-O) is adopted for investigating the spray mixing at high O/F ratios between 3.75 and 6.25. The impinging distance, jet velocity and impinging angle for the two-stage impinging type injector are design parameters examined, where the impinging angle is more effective at spray atomization and droplet distribution. The PLIF technique is used to measure the droplet distribution so as to identify the spray characteristics. In order to simplify the development process of the injector, the predicted mixture ratio distribution from the individual fuel (F-F) and oxidizer (O-O) sprays by overlapping their averaged images is used to compare with the actual distribution from the two-stage impinging spray (O-F-F-O). At a constant total mass flow rate, results indicate that tendencies towards the variations of the average characteristic velocity (C) with increasing O/F ratios are similar for outcomes of the prediction and actual measurement. Also, there is obvious flow fields interaction between the fuel and oxidizer sprays and coordinating their relative intensities of sprays well can optimize the mixture ratio distribution of the two-stage impinging spray. Better mixing occurs when the fuel and oxidizer sprays have more similar and uniform distributions.

The Cassini spacecraft is in its final years. On September 15, 2017, Cassini will plunge deep into Saturn's atmosphere never to reemerge; thus concluding its second extended mission and 13 years in orbit around the ringed planet. As of October 2014, the spacecraft is four years in to its seven-year, second extended mission, the Cassini Solstice Mission (CSM). With three years left and only 2.5% of its loaded bipropellant and 37% of its loaded monopropellant remaining, the Cassini project actively manages the predicted end-of-mission propellant margins to maintain a high confidence in the spacecraft's ability to complete the CSM as designed. Accurate spacecraft navigation, rigorous remaining-propellant estimation, and frequent future propellant consumption prediction have resulted in efficient propellant use and a probability of sufficient propellant margin greater than 99%.

New developments in materials science in the areas of solution gelation processes and nanotechnology have led to the recent commercial production of aerogels. Concurrent with these advancements has been the development of new approaches to cryogenic thermal insulation systems. For example, thermal and physical characterizations of aerogel beads under cryogenic-vacuum conditions have been performed at the Cryogenics Test Laboratory of the NASA Kennedy Space Center. Aerogel-based insulation system demonstrations have also been conducted to improve performance for space launch applications. Subscale cryopumping experiments show the thermal insulating ability of these fully breathable nanoporous materials. For a properly executed thermal insulation system, these breathable aerogel systems are shown to not cryopump beyond the initial cooldown and thermal stabilization phase. New applications are being developed to augment the thermal protection systems of space launch vehicles, including the Space Shuttle External Tank. These applications include a cold-boundary temperature of 90 K with an ambient air environment in which both weather and flight aerodynamics are important considerations. Another application is a nitrogen-purged environment with a cold-boundary temperature of 20 K where both initial cooldown and launch ascent profiles must be considered. Experimental results and considerations for these flight system applications are discussed.

This paper reports results obtained from an experimental study of the combustion mechanism of aluminized propellants based on an energetic binder. The techniques used in this investigation include

For over thirty years the T-Burner has been the de facto standard for determining the unstable response of solid propellants. This paper reviews some of the different approaches that have been used, and briefly summarizes advantages and disadvantages of the different techniques. Data from a variety of sources have been reviewed and summarized showing the effect of various test conditions and propellant formulation variations. Much of the data taken in the 60's and early 70's were with so-called research propellants, with relatively low solids loading and usually monomodal AP as the oxidizer. Later data were obtained with propellants containing solids loading more typical and practical propellants. Several distinct differences were observed in the data for the different types of propellants. Data varying AP particle size distributions indicate that either additional fine AP or finer AP usually increases the propellant response. A significant reduction in the response can be achieved by reducing the amount of fine AP, while holding burning rate constant either by adding a catalyst or by varying the overall AP size distribution. A significant amount of data has been obtained for reduced smoke propellants containing stability additives. The data indicate that the additives provide significant particle damping, and most often, also reduce the propellant response. Data from double base propellants indicate a much broader response over a wide frequency range. This implies that double base propellants should be more unstable in tangential modes than composite propellants. This trend seems to be consistent with observed motor data. Measuring the response of metallized propellants has been a challenge due to the large damping due to metal oxide particles in the gas. Data from metallized propellants are not presented here, but will be summarized in a future paper.