Sample records for rocket propellant udmh

  1. Solid propellant rocket motor

    Dowler, W. L.; Shafer, J. I.; Behm, J. W.; Strand, L. D. (Inventor)


    The characteristics of a solid propellant rocket engine with a controlled rate of thrust buildup to a desired thrust level are discussed. The engine uses a regressive burning controlled flow solid propellant igniter and a progressive burning main solid propellant charge. The igniter is capable of operating in a vacuum and sustains the burning of the propellant below its normal combustion limit until the burning propellant surface and combustion chamber pressure have increased sufficiently to provide a stable chamber pressure.

  2. Pervaporation performance of PPO membranes in dehydration of highly hazardous mmh and udmh liquid propellants.

    Moulik, Siddhartha; Kumar, K Praveen; Bohra, Subha; Sridhar, Sundergopal


    Polyphenylene oxide (PPO) membranes synthesized from 2,6-dimethyl phenol monomer were subjected to pervaporation-based dehydration of the highly hazardous and hypergolic monomethyl hydrazine (MMH) and unsymmetrical dimethyl hydrazine (UDMH) liquid propellants. Membranes were characterized by TGA, DSC and SEM to study the effect of temperature besides morphologies of surface and cross-section of the films, respectively. Molecular dynamics (MD) simulation was used to study the diffusion behavior of solutions within the membrane. CFD method was employed to solve the governing mass transfer equations by considering the flux coupling. The modeling results were highlighted by the experimental data and were in good agreement. High separation factors (35-70) and reasonable water fluxes (0.1-0.2 kg/m(2)h) were observed for separation of the aqueous azeotropes of MMH (35 wt%) and UDMH (20 wt%) and their further enrichment to >90% purity. Effect of feed composition, membrane thickness and permeate pressure on separation performance of PPO membranes were investigated to determine optimum operating conditions.

  3. Low toxicity rocket propellants

    Wink, J.


    Hydrazine (N2H4) and its hypergolic mate nitrogen tetroxide (N2O4) are used on virtually all spacecraft and on a large number of launch vehicles. In recent years however, there has been an effort in identifying and developing alternatives to replace hydrazine as a rocket propellant.

  4. Low toxicity rocket propellants

    Wink, J.


    Hydrazine (N2H4) and its hypergolic mate nitrogen tetroxide (N2O4) are used on virtually all spacecraft and on a large number of launch vehicles. In recent years however, there has been an effort in identifying and developing alternatives to replace hydrazine as a rocket propellant.

  5. Alternate Propellant Thermal Rocket Project

    National Aeronautics and Space Administration — The Alternate Propellant Thermal Rocket (APTR) is a novel concept for propulsion of space exploration or orbit transfer vehicles. APTR propulsion is provided by...

  6. Environmentally compatible solid rocket propellants

    Jacox, James L.; Bradford, Daniel J.


    Hercules' clean propellant development research is exploring three major types of clean propellant: (1) chloride-free formulations (no chlorine containing ingredients), being developed on the Clean Propellant Development and Demonstration (CPDD) contract sponsored by Phillips Laboratory, Edwards Air Force Base, CA; (2) low HCl scavenged formulations (HCl-scavenger added to propellant oxidized with ammonium perchlorate (AP)); and (3) low HCl formulations oxidized with a combination of AN and AP (with or without an HCl scavenger) to provide a significant reduction (relative to current solid rocket boosters) in exhaust HCl. These propellants provide performance approaching that of current systems, with less than 2 percent HCl in the exhaust, a significant reduction (greater than or equal to 70 percent) in exhaust HCl levels. Excellent processing, safety, and mechanical properties were achieved using only readily available, low cost ingredients. Two formulations, a sodium nitrate (NaNO3) scavenged HTPB and a chloride-free hydroxy terminated polyether (HTPE) propellant, were characterized for ballistic, mechanical, and rheological properties. In addition, the hazards properties were demonstrated to provide two families of class 1.3, 'zero-card' propellants. Further characterization is planned which includes demonstration of ballistic tailorability in subscale (one to 70 pound) motors over the range of burn rates required for retrofit into current Hercules space booster designs (Titan 4 SRMU and Delta 2 GEM).

  7. Some typical solid propellant rocket motors

    Zandbergen, B.T.C.


    Typical Solid Propellant Rocket Motors (shortly referred to as Solid Rocket Motors; SRM's) are described with the purpose to form a database, which allows for comparative analysis and applications in practical SRM engineering.

  8. Some typical solid propellant rocket motors

    Zandbergen, B.T.C.


    Typical Solid Propellant Rocket Motors (shortly referred to as Solid Rocket Motors; SRM's) are described with the purpose to form a database, which allows for comparative analysis and applications in practical SRM engineering.

  9. Fuel-Cell Power Source Based on Onboard Rocket Propellants

    Ganapathi, Gani; Narayan, Sri


    The use of onboard rocket propellants (dense liquids at room temperature) in place of conventional cryogenic fuel-cell reactants (hydrogen and oxygen) eliminates the mass penalties associated with cryocooling and boil-off. The high energy content and density of the rocket propellants will also require no additional chemical processing. For a 30-day mission on the Moon that requires a continuous 100 watts of power, the reactant mass and volume would be reduced by 15 and 50 percent, respectively, even without accounting for boiloff losses. The savings increase further with increasing transit times. A high-temperature, solid oxide, electrolyte-based fuel-cell configuration, that can rapidly combine rocket propellants - both monopropellant system with hydrazine and bi-propellant systems such as monomethyl hydrazine/ unsymmetrical dimethyl hydrazine (MMH/UDMH) and nitrogen tetroxide (NTO) to produce electrical energy - overcomes the severe drawbacks of earlier attempts in 1963-1967 of using fuel reforming and aqueous media. The electrical energy available from such a fuel cell operating at 60-percent efficiency is estimated to be 1,500 Wh/kg of reactants. The proposed use of zirconia-based oxide electrolyte at 800-1,000 C will permit continuous operation, very high power densities, and substantially increased efficiency of conversion over any of the earlier attempts. The solid oxide fuel cell is also tolerant to a wide range of environmental temperatures. Such a system is built for easy refueling for exploration missions and for the ability to turn on after several years of transit. Specific examples of future missions are in-situ landers on Europa and Titan that will face extreme radiation and temperature environments, flyby missions to Saturn, and landed missions on the Moon with 14 day/night cycles.

  10. Solid propellants for rockets. Rocket suishin yaku

    Kubota, N. (Defense Agency, Tokyo (Japan). Technical Research and Development Inst.)


    Physical and chemical ProPerties and combustion characteristics of propellants differ according to the combination of oxidizers and fuel components. Composite smoke propellant, having crystalline ammonium perchlorate as an oxidizer and hydrocarbon Polymer as a fuel, has higher specific impulse and improved mechanical properties compared to smokeless double base propellant consisting of nitroglycerin and nirocellulose. Double base propellants with low specific impulse are combined with nitramines( RDX or HMX ) to make composite modified double based( CMDB ) propellants, as a result the smokeless property of double base propellant is preserved and the combustion efficiency is increased. With the combination of oxidizing agents and fuels, formation of various high functional propellants has been possible and energetic azide polymers have provided possibilities for fuels of propellants. 3 refs., 6 figs., 3 tabs.

  11. Specific Impulses Losses in Solid Propellant Rockets


    to use the collision function form proposed by Golovin to simplify this production term: 4C><=) <P- .: Accordingly: m hence, by integration: Now, we...November 21, 1940 in Paris, Seine. VFirst Thesis. "Contribution to the Study of Specific i Impulse Loss in Solid Propellant Rockets." Second Thesis

  12. 偏二甲肼污水的好氧生物降解及其动力学研究%The Aerobic Biodegradation of UDMH Wastewater and Its Kinetics Research

    廖琪丽; 王力; 彭清涛


    @@ Unsymmetrical dimethylhydrazine (UDMH) is a main liquid rocket propellant. There are plenty of toxic and stable contaminations in its wastewater. The harmful influence has been considered. Nowadays the disposal technology of wastewater containing UDMH has plenty of disadvantages, such as high energy consuming, low safety factor, numerous secondary contaminations, strong toxicity, and so on. It is necessary to find a new way for wastewater disposal. Biodegradation technology is low innocuity and cleanliness, has great degradation capacity, and can almost degrade all kinds of organic substances. In this paper the aerobic active sludge is used to treat wastewater containing UDMH which is one kind of biodegradation technology.

  13. Method of injecting fluid propellants into a rocket combustion chamber

    Schneider, Steven J. (Inventor)


    A rocket injector is provided with multiple sets of manifolds for supplying propellants to injector elements. Sensors transmit the temperatures of the propellants to a suitable controller which is operably connected to valves between these manifolds and propellant storage tanks. Additional valves are opened to furnish propellants to more of the manifolds when cryogenic propellant temperatures are sensed. Only a portion of the valves are opened to furnish propellants to some of the manifolds when lower temperatures are sensed.

  14. Integrated model of a composite propellant rocket

    Miccio, Francesco


    The combustion of composite solid propellants was investigated and an available numerical model was improved for taking into account the change of pressure, when the process occurs in a confined environment, as inside a rocket. The pressure increase upon ignition is correctly described by the improved model for both sandwich and dispersed particles propellants. In the latter case, self-induced fluctuations in the pressure and in all other computed variables occur, as consequence of the periodic rise and depletion of oxidizer particles from the binder matrix. The comparison with the results of the constant pressure model shows a different fluctuating profile of gas velocity, with a possible second order effect induced by the pressure fluctuations.

  15. Ammonium nitrate: a promising rocket propellant oxidizer

    Oommen; Jain


    Ammonium nitrate (AN) is extensively used in the area of fertilizers and explosives. It is present as the major component in most industrial explosives. Its use as an oxidizer in the area of propellants, however, is not as extensive as in explosive compositions or gas generators. With the growing demand for environmental friendly chlorine free propellants, many attempts have been made of late to investigate oxidizers producing innocuous combustion products. AN, unlike the widely used ammonium perchlorate, produces completely ecofriendly smokeless products. Besides, it is one of the cheapest and easily available compounds. However, its use in large rocket motors is restricted due to some of its adverse characteristics like hygroscopicity, near room temperature phase transformation involving a volume change, and low burning rate (BR) and energetics. The review is an attempt to consolidate the information available on the various issues pertaining to its use as a solid propellant oxidizer. Detailed discussions on the aspects relating to phase modifications, decomposition chemistry, and BR and energetics of AN-based propellants, are presented. To make the review more comprehensive brief descriptions of the history, manufacture, safety, physical and chemical properties and various other applications of the salt are also included. Copyright 1999 Elsevier Science B.V.

  16. Effect of Temperature on Mechanical Properties of Solid Rocket Propellants

    Himanshu Shekhar


    Full Text Available Mechanical properties of solid rocket propellants are dependent on temperature. Any change in temperature brings significant change in the tensile strength, percentage elongation, and elastic modulus of the propellant. Different classes of operational solid rocket propellants namely extruded double-base propellants, composite, extruded composite and nitrarte ester polyester propellants were evaluated at different temperatures in the operating range of the rockets and missiles preferably in the range of –50 oC to +55 oC. It was observed that for each class of propellant, as temperature reduces, propellant becomes hard. This is depicted by increase in elastic modulus and tensile strength of the material. However, trend of percentage elongation is not very uniform. Extruded double-base propellants show less percentage elongation (around 1 per cent at reduced temperature (–50 oC probably due to brittleness. So is the trend with case-bonded composite propellants. However, reverse trend is exhibited by cartridge-loaded composite propellants and nitrate ester polyester propellants. Such propellants show higher percentage elongation (6 per cent for CLCP and 35 per cent for NEPE at reduced temperature (–50 oC. This makes such propellants tough and more area under stress-strain curve at reduced temperature is observed.Defence Science Journal, 2011, 61(6, pp.529-533, DOI:

  17. Nuclear thermal rockets using indigenous extraterrestrial propellants

    Zubrin, Robert M.


    A preliminary examination of a concept for a Mars and outer solar system exploratory vehicle is presented. Propulsion is provided by utilizing a nuclear thermal reactor to heat a propellant volatile indigenous to the destination world to form a high thrust rocket exhaust. Candidate propellants, whose performance, materials compatibility, and ease of acquisition are examined and include carbon dioxide, water, methane, nitrogen, carbon monoxide, and argon. Ballistics and winged supersonic configurations are discussed. It is shown that the use of this method of propulsion potentially offers high payoff to a manned Mars mission. This is accomplished by sharply reducing the initial mission mass required in low earth orbit, and by providing Mars explorers with greatly enhanced mobility in traveling about the planet through the use of a vehicle that can refuel itself each time it lands. Thus, the nuclear landing craft is utilized in combination with a hydrogen-fueled nuclear-thermal interplanetary launch. By utilizing such a system in the outer solar system, a low level aerial reconnaissance of Titan combined with a multiple sample return from nearly every satellite of Saturn can be accomplished in a single launch of a Titan 4 or the Space Transportation System (STS). Similarly a multiple sample return from Callisto, Ganymede, and Europa can also be accomplished in one launch of a Titan 4 or the STS.

  18. Propellant Flow Actuated Piezoelectric Rocket Engine Igniter Project

    National Aeronautics and Space Administration — Spark ignition of a bi-propellant rocket engine is a classic, proven, and generally reliable process. However, timing can be critical, and the control logic,...

  19. Dynamical Model of Rocket Propellant Loading with Liquid Hydrogen

    National Aeronautics and Space Administration — A dynamical model describing the multi-stage process of rocket propellant loading has been developed. It accounts for both the nominal and faulty regimes of...

  20. Sensitivity of solid rocket propellants for card gap test

    Kimura, Eishu; Oyumi, Yoshio (Japan Defense Agency, Tokyo (Japan). Technical Research and Development Inst.)


    Card gap test, which is standardized in Japan Explosives Society, was modified in order to apply it to solid rocket propellants and carried out to evaluate sensitivities against shock stimuli. Solid propellants tested here were mainly azide polymer composite propellants, which contained ammonium nitrate (AN) as a main oxidizer. Double base propellant, composed nitroglycerin and nitrocellulose (NC), and ammonium perchlorate (AP)-based composite propellants. It is found that the sensitivity was dominated by the oxidizer characteristics. AP- and AN-based propellant had less sensitivity and HMX-based propellant showed higher sensitivity, and the adding of NC and TMETN contributed to worse sensitive for the card gap test. Good relationship was obtained between the card gap sensitivity and the oxygen balance of propellants tested here. (orig.)

  1. Burn Rate Modelling of Solid Rocket Propellants (Short Communication

    A.R. Kulkarni


    Full Text Available A generalised model of burning of a solid rocket propellant based on kinetics of propellant hasbeen developed. A complete set of variables has been formed after examining the existing models.Buckingham theorem provides the functional form of the model, such that the existing models are thesubcases of this generalised model. This proposed model has been validated by an experimental data.

  2. Solid propellant processing factor in rocket motor design


    The ways are described by which propellant processing is affected by choices made in designing rocket engines. Tradeoff studies, design proof or scaleup studies, and special design features are presented that are required to obtain high product quality, and optimum processing costs. Processing is considered to include the operational steps involved with the lining and preparation of the motor case for the grain; the procurement of propellant raw materials; and propellant mixing, casting or extrusion, curing, machining, and finishing. The design criteria, recommended practices, and propellant formulations are included.

  3. Refinement of Propellant Strand Burning Method to Suit Aluminised Composite Rocket Propellant


    ABSTRACT An epoxy coating was trialled as an inhibitor for composite rocket propellant strands burned in a Crawford–style bomb . The epoxy coating...into the Crawford bomb and pressurised with nitrogen to the specified pressure. Once pressure and temperature were stable at the required set- point...aggressive propellant. This method involved the coating of strands of propellant with a two-part epoxy polymer. Strands were dipped into a bath of epoxy

  4. Particle behavior in solid propellant rockets

    Netzer, D. W.; Diloreto, V. D.; Dubrov, E.


    The use of holography, high speed motion pictures, light scattering measurements, and post-fire particle collection/scanning electron microscopic examination to study the combustion of composite solid propellants is discussed. The relative advantages and disadvantages of the different experimental techniques for obtaining two-phase flow characteristics within the combustion environment of a solid propellant grain are evaluated. Combustion bomb studies using high speed motion pictures and post-fire residue analysis were completed for six low metal content propellants. Resolution capabilities and the relationships between post-fire residue and motion picture data are determined. Initial testing using a holocamera together with a 2D windowed motor is also described.

  5. Propellant removal from rocket motors containing double-base compositions

    Whinnery, L.; Griffiths, S.; Hruby, J.; Larson, R.; Lipkin, J.; Long, B.; Schoenfelder, C.


    The uncertain environmental consequences and regulations associated with using open burning/open detonation for the disposal of energetic materials are forcing both manufacturers and users to examine alternative disposal technologies. In general, these alternatives involve a material removal operation followed by processing steps that lead to reuse of valuable constituents and/or disposal of waste. While a number of post-removal processing options appear to be viable, the initial step of removing an energetic material, such as a solid rocket motor propellant, from its container remains a significant technological challenge. Large rocket motors containing highly energetic propellant, hazard class 1.1, are of particular concern because of their inherent handling hazards. We will describe the results of a study using thermal cycling to increase the surface area of inert propellant formulations. The propellant removal method studied employs thermal cycling to cryogenic temperatures (cryocycling). Using inert propellants and liquid nitrogen we have demonstrated that this process produces multiple cracks throughout the bulk of the grain. The properties of the actual and inert propellants are being measured, and a model is being developed to relate experiments on inert material to actual propellant. Possible methods to increase thermal gradients, crack propagation and initiation are also presented.

  6. Propellant development for the Advanced Solid Rocket Motor

    Landers, L. C.; Stanley, C. B.; Ricks, D. W.


    The properties of a propellant developed for the NASA Advanced Solid Rocket Motor (ASRM) are described in terms of its composition, performance, and compliance to NASA specifications. The class 1.3 HTPB/AP/A1 propellant employs an ester plasticizer and the content of ballistic solids is set at 88 percent. Ammonia evolution is prevented by the utilization of a neutral bonding agent which allows continuous mixing. The propellant also comprises a bimodal AP blend with one ground fraction, ground AP of at least 20 microns, and ferric oxide to control the burning rate. The propellant's characteristics are discussed in terms of tradeoffs in AP particle size and the types of Al powder, bonding agent, and HTPB polymer. The size and shape of the ballistic solids affect the processability, ballistic properties, and structural properties of the propellant. The revised baseline composition is based on maximizing the robustness of in-process viscosity, structural integrity, and burning-rate tailoring range.

  7. Ultrasonic investigation of mechanical properties of double base rocket propellants

    Schroeff, J.A. van der; Boer, R.S. de


    For a series of double base rocket propellants and for poly-methylmethacrylate (PMMA) the longitudinal and transverse sound wave velocities are measured at a frequency of 0.351 MHz in t h e temperature range of −40°C to +60°C. The relations between these acoustic properties and mechanical properties

  8. Metallic Hydrogen: A Game Changing Rocket Propellant

    Silvera, Isaac F.


    The objective of this research is to produce metallic hydrogen in the laboratory using an innovative approach, and to study its metastability properties. Current theoretical and experimental considerations expect that extremely high pressures of order 4-6 megabar are required to transform molecular hydrogen to the metallic phase. When metallic hydrogen is produced in the laboratory it will be extremely important to determine if it is metastable at modest temperatures, i.e. remains metallic when the pressure is released. Then it could be used as the most powerful chemical rocket fuel that exists and revolutionize rocketry, allowing single-stage rockets to enter orbit and chemically fueled rockets to explore our solar system.

  9. Dynamic mechanical analysis of double base rocket propellants

    Marcin Cegła


    Full Text Available The article presents dynamic mechanical analysis (DMA for solid rocket propellants testing. Principles of operation and measured values are briefly described. The authors refer to the previous research of PTFE material and literature data providing information about proper experimental conditions and influence of measurement frequency, load amplitude, and heating rate on the results of DMA tests. The experimental results of solid double-base rocket propellant testing obtained on the N Netzsch DMA 242 device are presented. Mechanical properties such as the dynamic storage modulus E´, the dynamic loss modulus E˝ and tan(δ were measured within temperature range from (–120°C to (+90°C at the heating rate of 1 K/min. The test sample was subjected to a dual cantilever multi-frequency test. Special attention was paid to determination of the glass transition temperature of the tested propellant in reference to the NATO standardization agreement 4540 as well as influence of the measurement frequency on the glass transition.[b]Keywords[/b]: Dynamic mechanical analysis, solid rocket propellants, glass transition temperature

  10. MEMS-Based Solid Propellant Rocket Array Thruster

    Tanaka, Shuji; Hosokawa, Ryuichiro; Tokudome, Shin-Ichiro; Hori, Keiichi; Saito, Hirobumi; Watanabe, Masashi; Esashi, Masayoshi

    The prototype of a solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft was completed and tested. The prototype has 10×10 φ0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 20×22 mm substrate. To realize such a dense array of micro-rockets, each ignition heater is powered from the backside of the thruster through an electrical feedthrough which passes along a propellant cylinder wall. Boron/potassium nitrate propellant (NAB) is used with/without lead rhodanide/potassium chlorate/nitrocellulose ignition aid (RK). Impulse thrust was measured by a pendulum method in air. Ignition required electric power of at least 3 4 W with RK and 4 6 W without RK. Measured impulse thrusts were from 2×10-5 Ns to 3×10-4 Ns after the calculation of compensation for air dumping.

  11. Scaling theory for liquid propellant rocket thrust chambers

    C. M. Sethna


    Full Text Available With the advent of the very large liquid propellant rocket, it has become necessary, if possible, to derive a rational scaling theory for combustion chamber design so as to enable relatively simple and economical initial tests to be carried out on small scaled models using scaled parameters of propellant mass flows, pressure etc., and from these to predict operating and design data for the full scale rocket. Owing to the complex and interdependent nature of the aerothermo-chemical processes in the chamber involving evaporation, diffusion and chemical reaction, the similarity criteria must necessarily extend over several, non-dimensional parameters, but it is still possible to evolve relatively simple rules for correlating the design and performance of the model and large scale motors-as shown by Penner-Tsien and Crocco. The paper concludes with a discussion of the accuracy and practical feasibility of such scaling rules.

  12. Characterization of rocket propellant combustion products

    Jenkins, R.A.; Nestor, C.W.; Thompson, C.V.; Gayle, T.M.; Ma, C.Y.; Tomkins, B.A.; Moody, R.L.


    The overall objective of the work described in this report is four-fold: to (a) develop a standardized and experimentally validated approach to the sampling and chemical and physical characterization of the exhaust products of scaled-down rocket launch motors fired under experimentally controlled conditions at the Army's Signature Characterization Facility (ASCF) at Redstone Arsenal in Huntsville, Alabama; (b) determine the composition of the exhaust produces; (c) assess the accuracy of a selected existing computer model for predicting the composition of major and minor chemical species; (d) recommended alternations to both the sampling and analysis strategy and the computer model in order to achieve greater congruence between chemical measurements and computer prediction. 34 refs., 2 figs., 35 tabs.

  13. A research on polyether glycol replaced APCP rocket propellant

    Lou, Tianyou; Bao, Chun Jia; Wang, Yiyang


    Ammonium perchlorate composite propellant (APCP) is a modern solid rocket propellant used in rocket vehicles. It differs from many traditional solid rocket propellants by the nature of how it is processed. APCP is cast into shape, as opposed to powder pressing it with black powder. This provides manufacturing regularity and repeatability, which are necessary requirements for use in the aerospace industry. For traditional APCP, ingredients normally used are ammonium peroxide, aluminum, Hydroxyl-terminated polybutadiene(HTPB), curing agency and other additives, the greatest disadvantage is that the fuel is too expensive. According to the price we collected in our country, a single kilogram of this fuel will cost 200 Yuan, which is about 35 dollars, for a fan who may use tons of the fuel in a single year, it definitely is a great deal of money. For this reason, we invented a new kind of APCP fuel. Changing adhesive agency from cross-linked htpb to cross linked polyether glycol gives a similar specific thrust, density and mechanical property while costs a lower price.

  14. Rocket Solid Propellant Alternative Based on Ammonium Dinitramide

    Grigore CICAN


    Full Text Available Due to the continuous run for a green environment the current article proposes a new type of solid propellant based on the fairly new synthesized oxidizer, ammonium dinitramide (ADN. Apart of having a higher specific impulse than the worldwide renowned oxidizer, ammonium perchlorate, ADN has the advantage, of leaving behind only nitrogen, oxygen and water after decomposing at high temperatures and therefore totally avoiding the formation of hydrogen chloride fumes. Based on the oxidizer to fuel ratios of the current formulations of the major rocket solid booster (e.g. Space Shuttle’s SRB, Ariane 5’s SRB which comprises mass variations of ammonium perchlorate oxidizer (70-75%, atomized aluminum powder (10-18% and polybutadiene binder (12-20% a new solid propellant was formulated. As previously stated, the new propellant formula and its variations use ADN as oxidizer and erythritol tetranitrate as fuel, keeping the same polybutadiene as binder.

  15. On the history of the development of solid-propellant rockets in the Soviet Union

    Pobedonostsev, Y. A.


    Pre-World War II Soviet solid-propellant rocket technology is reviewed. Research and development regarding solid composite preparations of pyroxyline TNT powder is described, as well as early work on rocket loading calculations, problems of flight stability, and aircraft rocket launching and ground rocket launching capabilities.

  16. Magnesium and Carbon Dioxide - A Rocket Propellant for Mars Missions

    Shafirovich, E. IA.; Shiriaev, A. A.; Goldshleger, U. I.


    A rocket engine for Mars missions is proposed that could utilize CO2 accumulated from the Martian atmosphere as an oxidizer. For use as possible fuel, various metals, their hydrides, and mixtures with hydrogen compounds are considered. Thermodynamic calculations show that beryllium fuels ensure the most impulse but poor inflammability of Be and high toxicity of its compounds put obstacles to their applications. Analysis of the engine performance for other metals together with the parameters of ignition and combustion show that magnesium seems to be the most promising fuel. Ballistic estimates imply that a hopper with the chemical rocket engine on Mg + CO2 propellant could be readily developed. This vehicle would be able to carry out 2-3 ballistic flights on Mars before the final ascent to orbit.

  17. SRM (Solid Rocket Motor) propellant and polymer materials structural modeling

    Moore, Carleton J.


    The following investigation reviews and evaluates the use of stress relaxation test data for the structural analysis of Solid Rocket Motor (SRM) propellants and other polymer materials used for liners, insulators, inhibitors, and seals. The stress relaxation data is examined and a new mathematical structural model is proposed. This model has potentially wide application to structural analysis of polymer materials and other materials generally characterized as being made of viscoelastic materials. A dynamic modulus is derived from the new model for stress relaxation modulus and is compared to the old viscoelastic model and experimental data.

  18. On the hydrodynamics of rocket propellant engine inducers and turbopumps

    d'Agostino, L.


    The lecture presents an overview of some recent results of the work carried out at Alta on the hydrodynamic design and rotordynamic fluid forces of cavitating turbopumps for liquid propellant feed systems of modern rocket engines. The reduced order models recently developed for preliminary geometric definition and noncavitating performance prediction of tapered-hub axial inducers and centrifugal turbopumps are illustrated. The experimental characterization of the rotordynamic forces acting on a whirling four-bladed, tapered-hub, variable-pitch high-head inducer, under different load and cavitation conditions is presented. Future perspectives of the work to be carried out at Alta in this area of research are briefly illustrated.

  19. Viscoelastic Modelling of Solid Rocket Propellants using Maxwell Fluid Model

    Himanshu Shekhar


    Full Text Available Maxwell fluid model consisting of a spring and a dashpot in series is applied for viscoelastic characterisation of solid rocket propellants. Suitable values of spring constant and damping coefficient wereemployed by least square variation of errors for generation of complete stress-strain curve in uniaxial tensile mode for case-bonded solid propellant formulations. Propellants from the same lot were tested at different strain rates. It was observed that change in spring constant, representing elastic part was very small with strain rate but damping constant varies significantly with variation in strain rate. For a typical propellant formulation, when strain rate was raised from 0.00037/s to 0.185/s, spring constant K changed from 5.5 MPato 7.9 MPa, but damping coefficient D was reduced from 1400 MPa-s to 4 MPa-s. For all strain rates, stress-strain curve was generated using Maxwell model and close matching with actual test curve was observed.This indicates validity of Maxwell fluid model for uniaxial tensile testing curves of case-bonded solid propellant formulations. It was established that at higher strain rate, damping coefficient becomes negligible as compared to spring constant. It was also observed that variation of spring constant is logarithmic with strain rate and that of damping coefficient follows power law. The correlation coefficients were introduced to ascertain spring constants and damping coefficients at any strain rate from that at a reference strain rate. Correlationfor spring constant needs a coefficient H, which is function of propellant formulation alone and not of test conditions and the equation developeds K2 = K1 + H ´ ln{(de2/dt/(de1/dt}. Similarly for damping coefficient D also another constant S is introduced and prediction formula is given by D2 = D1 ´ {(de2/dt/(de1/dt}S.Evaluating constants H and S at different strain rates validate this mathematical formulation for differentpropellant formulations

  20. Computational Thermochemistry of Jet Fuels and Rocket Propellants

    Crawford, T. Daniel


    The design of new high-energy density molecules as candidates for jet and rocket fuels is an important goal of modern chemical thermodynamics. The NASA Glenn Research Center is home to a database of thermodynamic data for over 2000 compounds related to this goal, in the form of least-squares fits of heat capacities, enthalpies, and entropies as functions of temperature over the range of 300 - 6000 K. The chemical equilibrium with applications (CEA) program written and maintained by researchers at NASA Glenn over the last fifty years, makes use of this database for modeling the performance of potential rocket propellants. During its long history, the NASA Glenn database has been developed based on experimental results and data published in the scientific literature such as the standard JANAF tables. The recent development of efficient computational techniques based on quantum chemical methods provides an alternative source of information for expansion of such databases. For example, it is now possible to model dissociation or combustion reactions of small molecules to high accuracy using techniques such as coupled cluster theory or density functional theory. Unfortunately, the current applicability of reliable computational models is limited to relatively small molecules containing only around a dozen (non-hydrogen) atoms. We propose to extend the applicability of coupled cluster theory- often referred to as the 'gold standard' of quantum chemical methods- to molecules containing 30-50 non-hydrogen atoms. The centerpiece of this work is the concept of local correlation, in which the description of the electron interactions- known as electron correlation effects- are reduced to only their most important localized components. Such an advance has the potential to greatly expand the current reach of computational thermochemistry and thus to have a significant impact on the theoretical study of jet and rocket propellants.

  1. 含碳偏二甲肼的点火和燃烧特性%Ignition and Combustion Characteristics of UDMH Containing Carbon

    张蒙正; 郝智超; 张玫; 李军


    添加碳粉是提高推进剂能量密度的途径之一.为了获得含碳偏二甲肼UDMH的点火特性和高效燃烧室设计方法,设计了撞击式喷注器液体火箭发动机燃烧室,通过实验手段,研究了含碳UDMH的燃烧特性、燃烧室特征长度对燃烧效率的影响以及冷却特性.结果表明,含碳UDMH燃烧室的起动、关机和脉冲性能与液体推进剂相当;在UDMH中添加碳元素将导致推进剂密度比冲增加,燃烧效率降低;增加燃烧室特征长度有利于提高含碳UDMH的燃烧效率;含碳UDMH适应于液膜冷却.%Addition of C to liquid propellant is one of the methods for improving energy density. To obtain the ignition characteristic of UDMH containing carbon and the design method for high-efficiency combustion chamber, liquid rocket engine chamber is designed for the like-on-like injector. The combustion characteristic of UDMH containing carbon, the effect of characteristic chamber length on combustion efficiency and the cooling characteristic are experimentally studied. The results show that effective film cooling can be obtained using UDMH containing carbon are equivalent to those of liquid propellant. It is also shown that the addition of C to UDMH can result in the increase of density specific impulse and the decrease of combustion efficiency, and that with the increase of characteristic length, combustion efficiency increases.

  2. Thermodynamic cycle analysis of solid propellant air-turbo-rocket

    CHEN Xiang; CHEN Yu-chun; TU Qiu-ye; ZHANG Hong; CAI Yuan-hu


    Solid propellant air-turbo-rocket (SPATR) is an air-breathing propulsion system. A numerical model of performance and characteristics analysis for SPATR was presented and the corresponding computer program was written according to the operation characteristics of SPATR. The influence on the SPATR performance at design point caused by the gas generator exit parameters and compressor pressure ratio had been computed and analyzed in detail. The off-design perform-ance of SPATR at sea level and high altitude had also been computed. The performance of thrust and specific impulse for SPATR with different solid propellant had been compared at off-design points, and the off-design performance comparison had been made between fuel-rich and oxygen-rich. The computation results indicated that SPATR operates within wide range of Maeh number (0 ~3) and altitude (0~12 km), and SPATR possesses high specific thrust (1 200 N/(kg/s)) and high specific impulse (7000 N/ (kg/s)) when fuel-air ratio of combustor equals fuel-air ratio.

  3. Lidar measurements of solid rocket propellant fire particle plumes.

    Brown, David M; Brown, Andrea M; Willitsford, Adam H; Dinello-Fass, Ryan; Airola, Marc B; Siegrist, Karen M; Thomas, Michael E; Chang, Yale


    This paper presents the first, to our knowledge, direct measurement of aerosol produced by an aluminized solid rocket propellant (SRP) fire on the ground. Such fires produce aluminum oxide particles small enough to loft high into the atmosphere and disperse over a wide area. These results can be applied to spacecraft launchpad accidents that expose spacecraft to such fires; during these fires, there is concern that some of the plutonium from the spacecraft power system will be carried with the aerosols. Accident-related lofting of this material would be the net result of many contributing processes that are currently being evaluated. To resolve the complexity of fire processes, a self-consistent model of the ground-level and upper-level parts of the plume was determined by merging ground-level optical measurements of the fire with lidar measurements of the aerosol plume at height during a series of SRP fire tests that simulated propellant fire accident scenarios. On the basis of the measurements and model results, the Johns Hopkins University Applied Physics Laboratory (JHU/APL) team was able to estimate the amount of aluminum oxide (alumina) lofted into the atmosphere above the fire. The quantification of this ratio is critical for a complete understanding of accident scenarios, because contaminants are transported through the plume. This paper provides an estimate for the mass of alumina lofted into the air.

  4. Fault Detection and Diagnosis Techniques for Liquid-Propellant Rocket Propellant Engines

    Wua, Jianjun; Tanb, Songlin


    Fault detection and diagnosis plays a pivotal role in the health-monitoring techniques for liquid- propellant rocket engines. This paper firstly gives a brief summary on the techniques of fault detection and diagnosis utilized in liquid-propellant rocket engines. Then, the applications of fault detection and diagnosis algorithms studied and developed to the Long March Main Engine System(LMME) are introduced. For fault detection, an analytical model-based detection algorithm, a time-series-analysis algorithm and a startup- transient detection algorithm based on nonlinear identification developed and evaluated through ground-test data of the LMME are given. For fault diagnosis, neural-network approaches, nonlinear-static-models based methods, and knowledge-based intelligent approaches are presented. Keywords: Fault detection; Fault diagnosis; Health monitoring; Neural networks; Fuzzy logic; Expert system; Long March main engines Contact author and full address: Dr. Jianjun Wu Department of Astronautical Engineering School of Aerospace and Material Engineering National University of Defense Technology Changsha, Hunan 410073 P.R.China Tel:86-731-4556611(O), 4573175(O), 2219923(H) Fax:86-731-4512301

  5. Unsteady Aerodynamic Investigation of the Propeller-Wing Interaction for a Rocket Launched Unmanned Air Vehicle

    G. Q. Zhang


    Full Text Available The aerodynamic characteristics of propeller-wing interaction for the rocket launched UAV have been investigated numerically by means of sliding mesh technology. The corresponding forces and moments have been collected for axial wing placements ranging from 0.056 to 0.5D and varied rotating speeds. The slipstream generated by the rotating propeller has little effects on the lift characteristics of the whole UAV. The drag can be seen to remain unchanged as the wing's location moves progressively closer to the propeller until 0.056D away from the propeller, where a nearly 20% increase occurred sharply. The propeller position has a negligible effect on the overall thrust and torque of the propeller. The efficiency affected by the installation angle of the propeller blade has also been analyzed. Based on the pressure cloud and streamlines, the vortices generated by propeller, propeller-wing interaction, and wing tip have also been captured and analyzed.

  6. Development of an advanced rocket propellant handler's suit

    Doerr, DonaldF.


    Most launch vehicles and satellites in the US inventory rely upon the use of hypergolic rocket propellants, many of which are toxic to humans. These fuels and oxidizers, such as hydrazine and nitrogen tetroxide have threshold limit values as low as 0.01 PPM. It is essential to provide space workers handling these agents whole body protection as they are universally hazardous not only to the respiratory system, but the skin as well. This paper describes a new method for powering a whole body protective garment to assure the safety of ground servicing crews. A new technology has been developed through the small business innovative research program at the Kennedy Space Center. Currently, liquid air is used in the environmental control unit (ECU) that powers the propellant handlers suit (PHE). However, liquid air exhibits problems with attitude dependence, oxygen enrichment, and difficulty with reliable quantity measurement. The new technology employs the storage of the supply air as a supercritical gas. This method of air storage overcomes all of three problems above while maintaining high density storage at relatively low vessel pressures (advantages of attitude independence and oxygen level stability. Thermal data revealed heat stress relief at least as good as liquid air supplied units. The application of supercritical air technology to this whole body protective ensemble marked an advancement in the state-of-the-art in personal protective equipment. Not only was long duration environmental control provided, but it was done without a high pressure vessel. The unit met human performance needs for attitude independence, oxygen stability, and relief of heat stress. This supercritical air (and oxygen) technology is suggested for microgravity applications in life support such as the Extravehicular Mobility Unit.

  7. Numerical Simulation of a Liquid Propellant Rocket Motor

    Nicolas M.C. Salvador; Marcelo M. Morales; Carlos E.S.S. Migueis; Demétrio Bastos-Netto


    This work presents a numerical simulation of the flow field in a liquid propellant rocket engine chamber and exit nozzle using techniques to allow the results to be taken as starting points for designing those propulsive systems.This was done using a Finite Volume method simulating the different flow regimes which usually take place in those systems. As the flow field has regions ranging from the low subsonic to the supersonic regimes, the numerical code used, initially developed for compressible flows only, was modified to work, proficiently in the whole velocity range. It is well known that codes have been developed in CFD, for either compressible or incompressible flows, the joint treatment of both together being complex even today, given the small number of references available in this area. Here an existing code for compressible flow was used and primitive variables,the pressure, the Cartesian components of the velocity and the temperature instead of the conserved variables were introduced in the Euler and Navier-Stokes equations. This was done to permit the treatment at any Mach number. Unstructured meshes with adaptive refinements were employed here. The convective terms were treated with upwind first and second order methods. The numerical stability was kept with artificial dissipation and in the spatial coverage one used a five stage Runge-Kutta scheme for the Fluid Mechanics and the VODE (Value of Ordinary Differential Equations) scheme along with the Chemkin II in the chemical reacting solution. During the development of this code simulating the flow in a rocket engine, comparison tests were made with several different types of internal and extemal flows, at different velocities, seeking to establish the confidence level of the techniques being used. These comparisons were done with existing theoretical results and with other codes already validated and well accepted by the CFD community.

  8. Rocket propellant reorientation and fluid management used in space commercialization

    Hung, R. J.; Lee, C. C.; Shyu, K. L.


    In a spacecraft design, the requirements of settled propellant are different for tank pressurization, engine restart, venting, or propellant transfer. The requirement to settle or to position liquid fuel over the outlet end of the spacecraft propellant tank prior main engine restart possess a microgravity fluid behavior problem. In this paper, the dynamical behavior of liquid propellant, fluid reorientation, and propellant resettling have been carried out.

  9. Method and device for protecting objects against Rocket Propelled Grenades (RPGs)

    Wentzel, C.M.


    A device and a method for protecting objects against rocket-propelled grenades having a hollow nose cone includes a netting of knotted and coated superstrong fibers disposed in front of the object, in such a manner that the nose cone of a rocket caught in the netting will penetrate one of the meshes

  10. Method and device for protecting objects against Rocket Propelled Grenades (RPGs)

    Wentzel, C.M.


    A device and a method for protecting objects against rocket-propelled grenades having a hollow nose cone includes a netting of knotted and coated superstrong fibers disposed in front of the object, in such a manner that the nose cone of a rocket caught in the netting will penetrate one of the meshes




    Full Text Available In this study, three dimensional modelling of extrusion forming of a double base solid rocket propellant is performed on Ansys® finite element simulation package. For the purpose of initial model construction and later comparisons with elastoviscoplastik model, the solid propellant is assumed to obey the elastic-plastic material response during the direct extrusion process. Taking into account the contact surface behavior with Coulomb friction and geometric and material nonlinearities, an incremental large large strain solution methodology has been adapted in the simulation. The hydrostatic pressure, stress, strain, and displacement values during extrusion of the solid rocket propellant are obtained from the simulation.

  12. Multi-Dimensional Combustion Instability Analysis of Solid Propellant Rocket Motors.


    RI D-R159 314 MULTI-DIMENSIONAL COMBUSTION INSTABLITY ANALYSIS OF 1/1 I SOLID PROPELLANT ROCK.. (U) ALABAMA UNIY IN HUNTSVILLE I DEPT OF MECHANICAL...STANDARDS MlICROCOPY RESOLUTION TEST CHART 0 0 0 03 V.%% f iSR.TR. 85-0567 NULTI-DIMNSIONAL COMBUSTION INSTABILITY ANALYSIS OF SOLID PROPELLANT ROCKET...analysis of solid propellant rocket motors. This research was motivated by the need for im- provement of the current practice in combustion instability

  13. Mathematical Modelling of In-Chamber Processes in Hydrocombined Propellant Solid Rocket Motors

    Nikolai A. Obukhov


    Full Text Available The special conditions of employment of commercial rockets in the sea environment has opened up new possibilities of improving motor performance. The interesting method suggests supplying water into the running motor. This paper reports the calculations and experiments carried out with solid propellant model setups. The results prove the validity of the proposed method and allow the refinement of calculation techniques for the prediction of solid rocket motor performance characteristics. The serviceability of the solid propellant charges working in combination with water is demonstrated. A mathematical model is proposed for the operation of a hydrocombined propellant motor with water and powdered additives applied to the combustion chamber."

  14. Combustion diagnosis for analysis of solid propellant rocket abort hazards: Role of spectroscopy

    Gill, W.; Cruz-Cabrera, A. A.; Donaldson, A. B.; Lim, J.; Sivathanu, Y.; Bystrom, E.; Haug, A.; Sharp, L.; Surmick, D. M.


    Solid rocket propellant plume temperatures have been measured using spectroscopic methods as part of an ongoing effort to specify the thermal-chemical-physical environment in and around a burning fragment of an exploded solid rocket at atmospheric pressures. Such specification is needed for launch safety studies where hazardous payloads become involved with large fragments of burning propellant. The propellant burns in an off-design condition producing a hot gas flame loaded with burning metal droplets. Each component of the flame (soot, droplets and gas) has a characteristic temperature, and it is only through the use of spectroscopy that their temperature can be independently identified.

  15. Estimation of Pressure Index and Temperature Sensitivity Coefficient of Solid Rocket Propellants by Static Evaluation

    Himanshu Shekhar


    Full Text Available Burning rate of a solid rocket propellant depends on pressure and temperature. Conventional strand burner and Crawford bomb test on propellant strands was conducted to assess these dependent parameters. However, behaviour of propellant in rocket motor is different from its behaviour in strand form. To overcome this anomaly, data from static evaluation of rocket motor was directly used for assessment of these burningrate controlling parameters. The conventional empirical power law (r=aoexp[p{T-To}]Pn was considered and a method was evolved for determination of pressure index (n and temperature sensitivity coefficient (p of burning rate for solid rocket propellants from static evaluation data. Effect of pressure index and temperature sensitivity coefficient on firing curve is also depicted. Propellant grain was fired in progressive mode to cover a very wide pressure range of 50 kg/cm2 to 250 kg/cm2 and propellant burning rate index was calculated to be 0.32 in the given pressure range. Propellant grain was fired at +35 °C and –20 °C temperatures and temperature sensitivity coefficient of burning rate was calculated to be 0.27 % per °C. Since both the values were evaluated from realised static evaluation curves, these are more realistic and accurate compared to data generated by conventional methods.Defence Science Journal, 2009, 59(6, pp.666-669, DOI:

  16. Characterization of booster-rocket propellants and their simulants

    Weirick, L.J.


    A series of shock-loading experiments on a composite and an energietic propellant and there simulants was conducted on a light-gas gun. The initial objectives were to obtain Hugoniot data, to investigate the pressure threshold at which a reaction occurs, and to measure spall threshold at various impact velocities. The Hugoniot data measured for the propellants fit the Hugoniot curves provided by the manufacturer of the propellants extremely well and the Hugoniot curves developed for the simulants matched those of the propellants. Threshold pressures to initiate reactions in the composite and energetic propellants were found to be 40 and 3 kbars, respectively. In spall tests, the composite propellant and its simulant exhibited spall strengths around 0.25 and 0.18 kbar, respectively. The energetic propellant and its simulant were somewhat stronger with spall strengths just above 0.33 and 0.22 kbar. 12 refs., 6 figs., 6 tabs.

  17. Design of Multi-Propellant Star Grains for Solid Propellant Rockets

    S. Krishnan


    Full Text Available A new approach to solve the geometry-problem of solid propellant star is presented. The basis of the approach is to take the web-thickness (a ballistic as well as a geometrical property as the characteristic length. The nondimensional characteristic parameters representing diameter, length, slenderness-ratio, and ignitor accommodation of the grain are all identified. Many particular cases of star configurations (from the configurations of single propellant to those of four different propellants can be analysed through the identified characteristic parameters. A better way of representing the single-propellant-star-performance in a design graph is explained. Two types of dual propellant grains are analysed in detail. The first type is characterised by its two distinct stages of burning (initially by single propellant burning and then by dual propellant burning; the second type has the dual propellant burning throughout. Suitability of the identified characteristic parameters to an optimisation study is demonstrated through examples.

  18. Verification of Model of Calculation of Intra-Chamber Parameters In Hybrid Solid-Propellant Rocket Engines

    Zhukov Ilya S.


    Full Text Available On the basis of obtained analytical estimate of characteristics of hybrid solid-propellant rocket engine verification of earlier developed physical and mathematical model of processes in a hybrid solid-propellant rocket engine for quasi-steady-state flow regime was performed. Comparative analysis of calculated and analytical data indicated satisfactory comparability of simulation results.

  19. Verification of Model of Calculation of Intra-Chamber Parameters In Hybrid Solid-Propellant Rocket Engines

    Zhukov Ilya S.; Borisov Boris V.; Bondarchuk Sergey S.; Zhukov Alexander S.


    On the basis of obtained analytical estimate of characteristics of hybrid solid-propellant rocket engine verification of earlier developed physical and mathematical model of processes in a hybrid solid-propellant rocket engine for quasi-steady-state flow regime was performed. Comparative analysis of calculated and analytical data indicated satisfactory comparability of simulation results.

  20. Development of a miniature solid propellant rocket motor for use in plume simulation studies

    Baran, W. J.


    A miniature solid propellant rocket motor has been developed to be used in a program to determine those parameters which must be duplicated in a cold gas flow to produce aerodynamic effects on an experimental model similar to those produced by hot, particle-laden exhaust plumes. Phenomena encountered during the testing of the miniature solid propellant motors included erosive propellant burning caused by high flow velocities parallel to the propellant surface, regressive propellant burning as a result of exposed propellant edges, the deposition of aluminum oxide on the nozzle surfaces sufficient to cause aerodynamic nozzle throat geometry changes, and thermal erosion of the nozzle throat at high chamber pressures. A series of tests was conducted to establish the stability of the rocket chamber pressure and the repeatibility of test conditions. Data are presented which define the tests selected to represent the final test matrix. Qualitative observations are also presented concerning the phenomena experienced based on the results of a large number or rocket tests not directly applicable to the final test matrix.

  1. Propellant Flow Actuated Piezoelectric Rocket Engine Igniter Project

    National Aeronautics and Space Administration — Under a Phase 1 effort, IES successfully developed and demonstrated a spark ignition concept where propellant flow drives a very simple fluid mechanical oscillator...

  2. Yuzhnoye's new liquid rocket engines as enablers for space exploration

    Degtyarev, Alexander; Kushnaryov, Alexander; Shulga, Vladimir; Ventskovsky, Oleg


    Advanced liquid rocket engines (LREs) are being created by Yuzhnoye Design Office of Ukraine based on the fifty-year experience of rocket engines' and propulsion systems' development. These LREs use both hypergolic (NTO+UDMH) and cryogenic (liquid oxygen+kerosene) propellants. First stage engines have a range of thrust from 40 to 250 t, while the upper stage (used in space) engines - from several kilograms to 50 t and a re-ignition feature. The engines are intended for both Ukraine"s independent access to space and international market.

  3. Laser Ignition Technology for Bi-Propellant Rocket Engine Applications

    Thomas, Matt; Bossard, John; Early, Jim; Trinh, Huu; Dennis, Jay; Turner, James (Technical Monitor)


    This viewgraph presentation gives an overview of laser ignition technology for bipropellant rocket engines applications. The objectives of this project include: (1) the selection test chambers and flows; (2) definition of the laser ignition setup; (3) pulse format optimization; (4) fiber optic coupled laser ignition system analysis; and (5) chamber integration issues definition. The testing concludes that rocket combustion chamber laser ignition is imminent. Support technologies (multiplexing, window durability/cleaning, and fiber optic durability) are feasible.

  4. Development of high temperature materials for solid propellant rocket nozzle applications

    Manning, C. R., Jr.; Lineback, L. D.


    Aspects of the development and characteristics of thermal shock resistant hafnia ceramic material for use in solid propellant rocket nozzles are presented. The investigation of thermal shock resistance factors for hafnia based composites, and the preparation and analysis of a model of elastic materials containing more than one crack are reported.

  5. Natural Rubber Based Fuel Rich Propellant for Ramjet Rocket

    Solomon Thomas; T. L. Varghese; Gupta, S. K.; T. S. Ram; V. N. Krishnamurthy


    Development of fuel rich propellants for air-breathing propulsion is one of the frontier areas of research. This carries less oxidiser than the normal propellants, uses ram air for complete combustion and is heavily metallised. Studies were conducted at VSSC for the development of hydroxyl-terminated natural rubber (HTNR)-based Mg and Mg-Al alloy-filled polyurethane systems. The HTNR prepolymer was modified to make it free of acid functionality so as to avoid reaction with the metals a...

  6. Natural Rubber Based Fuel Rich Propellant for Ramjet Rocket

    Solomon Thomas


    Full Text Available Development of fuel rich propellants for air-breathing propulsion is one of the frontier areas of research. This carries less oxidiser than the normal propellants, uses ram air for complete combustion and is heavily metallised. Studies were conducted at VSSC for the development of hydroxyl-terminated natural rubber (HTNR-based Mg and Mg-Al alloy-filled polyurethane systems. The HTNR prepolymer was modified to make it free of acid functionality so as to avoid reaction with the metals and was then reacted with calculated amount of propylene oxide in the presence of tertiary amine catalyst at room temperature. Various levels of Mg and Mg-Al alloy were used in different compositions and the effect on propellant characteristics was studied. TMP and butane diol were used for higher crosslinking, chain extension, etc and their effects were evaluated. Propellant compositions using 30-35 per cent by weight of Mg/Mg-Al alloy with the modified prepolymer have been developed. They exhibit good processibility and mechanical properties. The feasibility of room temperature curing was attempted. Motors weighing 2 kg with 100 mm OD and 200 mm length were successfully static-tested in the primary mode for performance evaluation.

  7. Radiographic Sensitivity of Flaw Detection in Solid Rocket Propellants

    G.G Puranik


    Full Text Available The sensitivity of flaw detection with x-ray radiographic methods is investigated here qualitatively in case of cast double base and cast composite propellants and for air pockets it is found to be 1.5 and 0.9 percent of the web respectively. General guidelines for the inspection of sustainer charges have also been laid down.

  8. Lead-Free Double-Base Propellant for the 2.75 Inch Rocket Motor

    Magill, B. T.; Nauflett, G. W.; Furrow, K. W.


    The current MK 66 2.75 inch Rocket Motor double-base propellant contains the lead-based ballistic modifier LC-12-15 to achieve the desired plateau and mesa burning rate characteristics. The use of lead compounds poses a concern for the environment and for personal safety due to the metal's toxic nature when introduced into the atmosphere by propellant manufacture, rocket motor firing, and disposal. Copper beta-resorcylate (copper 2,4-di-hydroxy-benzoate) was successfully used in propellant as a simple modifier in the mid 1970's. This and other compounds have also been mixed with lead salts to obtain more beneficial ballistic results. Synthesized complexes of lead and copper compounds soon replaced the mixtures. The complexes incorporate the lead, copper lack of organic liquids, which allows for easier propellant processing. About ten years ago, the Indian Head Division, Naval Surface Warfare Center (NSWC), initiated an effort to develop a lead-free propellant for use in missile systems. Several lead-free propellant candidate formulations were developed. About five years ago, NSWC, in conjunction with Alliant Techsystems, Radford Army Ammunition Plant, continued ballistic modifier investigations. A four component ballistic modifier system without lead for double-base propellants that provide adequate plateau and mesa burn rate characteristics was developed and patented. The ballistic modifier's system contains bismuth subsalicylate, 1.5 percent; copper salicylate, 1.0 percent, copper stannate, 0.77 percent; and carbon black, 0.1 percent. Action time and impulse data obtained through multiple static firings indicate that the new lead-free double-base propellant, while not a match for NOSIH-AA-2, will be a very suitable replacement in the 2.75 inch Rocket Motor. Accelerated aging of the double-base propellant containing the lead-free ballistic modifier showed that it had a much higher rate of stabilizer depletion than the AA-2. A comprehensive study showed that an

  9. Metallised Fuel rich Propellants for Solid Rocket Ramjet: A Review

    B. K. Athawale


    Full Text Available This paper reviews the research work carried out in the field of metallised fuel-rich propellants (FRP. Limitations and merits of various potential metals (Al, Mg, B, Be, or Zr as a component of FRP are discussed. The paper also includes a discussion on the combustion mechanism of metallised propellants, including problem areas and probable remedial measures. Zirconium and Ti appear to have potential to offer FRP with efficient combustion. Ideal performance is not achieved with current systems based on Al and B and further work is needed to develop FRP having all three desirable attributes, viz., ease of ignition, stable combustion and high specific impulse (I/sub sp/ in a single composition.

  10. Studies on Stress-Strain Curves of Aged Composite Solid Rocket Propellants

    Himanshu Shekhar


    Full Text Available Mechanical property evaluation of composite solid rocket propellants is used as a quick quality control tool for propellant development and production. However, stress-strain curves from uni-axial tensile testing can be utilised to assess the shelf-life of propellants also. Composite propellants (CP of two varieties cartridge-loaded (CLCP and case-bonded (CBCP are utilized in rocket and missile applications. Both classes of propellants were evaluated for mechanical properties namely tensile strength, modulus and percentage elongation using specimens conforming to ASTM D638 type IV at different ageing time. Both classes of propellants show almost identical variation in various mechanical properties with time. Tensile strength increases with time for both classes of propellants and percentage elongation reduces. Initial modulus is also found to decrease with time. Tensile strength is taken as degradation criteria and it is observed that CLCP has slower degradation rate than CBCP. This is because of two facts–(i higher initial tensile strength of CLCP (1.39 MPa compared to CBCP (0.665 MPa and (ii lower degradation rate of CLCP (0.0014 MPa/day with respect to CBCP (0.0025 MPa/day. For the studied composite propellants, a degradation criterion in the form of percentage change in tensile strength is evaluated and shelf life for different degradation criteria is tabulated for quick reference.Defence Science Journal, 2012, 62(2, pp.90-94, DOI:

  11. Regression rate behaviors of HTPB-based propellant combinations for hybrid rocket motor

    Sun, Xingliang; Tian, Hui; Li, Yuelong; Yu, Nanjia; Cai, Guobiao


    The purpose of this paper is to characterize the regression rate behavior of hybrid rocket motor propellant combinations, using hydrogen peroxide (HP), gaseous oxygen (GOX), nitrous oxide (N2O) as the oxidizer and hydroxyl-terminated poly-butadiene (HTPB) as the based fuel. In order to complete this research by experiment and simulation, a hybrid rocket motor test system and a numerical simulation model are established. Series of hybrid rocket motor firing tests are conducted burning different propellant combinations, and several of those are used as references for numerical simulations. The numerical simulation model is developed by combining the Navies-Stokes equations with the turbulence model, one-step global reaction model, and solid-gas coupling model. The distribution of regression rate along the axis is determined by applying simulation mode to predict the combustion process and heat transfer inside the hybrid rocket motor. The time-space averaged regression rate has a good agreement between the numerical value and experimental data. The results indicate that the N2O/HTPB and GOX/HTPB propellant combinations have a higher regression rate, since the enhancement effect of latter is significant due to its higher flame temperature. Furthermore, the containing of aluminum (Al) and/or ammonium perchlorate(AP) in the grain does enhance the regression rate, mainly due to the more energy released inside the chamber and heat feedback to the grain surface by the aluminum combustion.

  12. Regarding the evaluation of the solid rocket propellant response function to pressure coupling

    Ioan ION


    Full Text Available High frequency combustion instabilities imply a major risk for the solid rocket motor stableworking and they are directly linked to the propellant response to chamber pressure coupling. Thisarticle discusses a laboratory testing method for the measurement and evaluation of the pressurecoupled response for non-metalized propellants in a first stage. Experimental researches were donewith an adequate setup, built and improved in our lab, able to evaluate the propellant response byinterpreting the pressure oscillations damping in terms of propellant response. Our paper aims atdefining a linearized one-dimensional flow study model to analyze the disturbed operation of the solidpropellant rocket motors. Based on the applied model we can assert that the real part of propellantresponse is a function of the oscillations damping, acoustic energy in the motor chamber and variouslosses in the burning chamber. The imaginary part of propellant response mainly depends on thenormalized pulsation, on the burning chamber gas column and on the pressure oscillations frequency.Our research purpose was obviously to minimize the risk of the combustion instabilities effects on therocket motors working, by experimental investigations using jet modulating techniques and sustainedby an interesting study model based on the perturbation method.

  13. Attenuation studies of booster-rocket propellants and their simulants

    Weirick, L.J.


    A series of impact experiments on a composite propellant, an energetic propellant, and their simulants was recently completed using a light-gas gun. Previous experiments were done to obtain Hugoniot data, to investigate the pressure threshold at which a reaction occurs, and to measure spall damage at various impact velocities. The present studies measured the attenuation of shock waves in these materials, completing the shock characterization needed for material modeling. An initial impulse of 2.0 GPa magnitude and {approximately}0.6 {mu}s duration was imposed upon samples of various thicknesses. VISAR was used to measure the free-surface velocity at the back of the samples; these data were used to generate a curve of shock-wave attenuation versus sample thickness for each material. Results showed that all four materials attenuated the shock wave very similarly. Material thicknesses of 3.0, 7.62, 12.7, and 19.0 mm attenuated the shock wave {approximately}16%, 33%, 50%, and 66% respectively. 14 refs., 12 figs., 4 tabs.

  14. Burn-back Equations for High Volumetric Loading Single-grain Dual-thrust Rocket Propellant Configuration (Review Paper

    Himanshu Shekhar


    Full Text Available Dual-thrust mode is adopted in solid propellant rocket propulsion through tailoring of burning area, nozzle, rocket motor chamber, propellant type, multiple propellant blocks. In the present study, mathematical formulation has been evolved for generation of burning surface area with web burnt for a simple central blind hole in a solid cylindrical propellant geometry with proper partial inhibition on external and lateral surfaces. The burn-back equation has been validated by static firing and parametric study was conducted to understand effect of various control geometrical parameters. The system is utilised for high volumetric loading, single propellant, single composition, single-chamber, single nozzle dual-thrust mode of burning profiles in rocket application.Defence Science Journal, 2011, 61(2, pp.165-170, DOI:

  15. Thrust Vector Control of an Upper-Stage Rocket with Multiple Propellant Slosh Modes

    Jaime Rubio Hervas


    Full Text Available The thrust vector control problem for an upper-stage rocket with propellant slosh dynamics is considered. The control inputs are defined by the gimbal deflection angle of a main engine and a pitching moment about the center of mass of the spacecraft. The rocket acceleration due to the main engine thrust is assumed to be large enough so that surface tension forces do not significantly affect the propellant motion during main engine burns. A multi-mass-spring model of the sloshing fuel is introduced to represent the prominent sloshing modes. A nonlinear feedback controller is designed to control the translational velocity vector and the attitude of the spacecraft, while suppressing the sloshing modes. The effectiveness of the controller is illustrated through a simulation example.

  16. Mechanical characterization of composite solid rocket propellant based on hydroxy-terminated polybutadiene

    Gligorijević Nikola I.


    Full Text Available This paper presents the procedure of uniaxial mechanical characterization of composite solid rocket propellant based on hydroxy-terminated polybutadiene (HTPB, whose mechanical properties strongly depend on temperature, strain rate, natural aging and accumulated damage. A method of processing data is presented in order to determine time-temperature shift factor and master curves for tensile strength, ultimate strain and relaxation modulus, depending on reduced time. Functional dependences of these features represent an input for structural analysis of a rocket motor propellant grain. The effects of natural aging on the mechanical properties are also considered. [Projekat Ministarstva nauke Republike Srbije, br. TR 36050: Research and development of unmanned aircraft in support of traffic infrastructure monitoring

  17. Discrimination for ablative control mechanism in solid-propellant rocket nozzle


    The ablation in solid-propellant rocket nozzle is a coupling process resulted by chemistry, heat and mass transfer. Based on the heat and mass transfer theory, the aero-thermo-dynamic, and thermo-chemical kinetics, the thermal-chemical ablation model is established. Simulations are completed on the heat flow field and chemical ablation in the nozzle with different concentrations, frequency factors and activation energy of H2. The calculation results show that the concentration and the activation energy of H2 can provoke the transformation of control mechanism, whereas the influence brought by the frequency factor of H2 is feeble under a high-temperature and high-pressure combustion circumstance. The discrimination for ablative control mechanism is dependent on both concentration and activation energy of H2. This study will be useful in handling ablation and thermal protection problem in the design of solid-propellant rocket.

  18. State-space analysis of the dynamic characteristics of a variable thrust liquid propellant rocket engine

    Zhang, Yu-Lin

    This paper states the application of state-space method to the analysis of the dynamic characteristics of a variable thrust liquid propellant rocket engine and presents a set of state equations for describing the dynamic process of the engine. An efficient numerical method for solving these system equations is developed. The theoretical solutions agree well with the experimental data. The analysis leads to the following conclusion: the set coefficient of the pulse width, the working frequency of the solenoid valves and the deviation of the critical working points of these valves are important parameters for determining the dynamic response time and the control precision of this engine. The methods developed in this paper may be used effectively in the analysis of dynamic characteristics of variable thrust liquid propellant rocket engines.

  19. Analytical model for liquid rocket propellant feedline dynamics

    Holster, J. L.; Astleford, W. J.


    A generalized analytical model and computer program have been developed to predict the frequency response of arbitrary liquid propellant feedline designs. The analytical model is based on an extension of an existing distributed parameter representation of a viscous fluid transmission line with laminar flow which was modified to include the effects of a turbulent mean flow. The effects of dissolved ullage gases, wall elasticity, localized gas or vapor bubbles, bellows, forced changes in length due to structural excitation, complex side branches, and structural mounting stiffness are also included. Each line component is written as a four-terminal, pressure-flow relationship in matrix form in the Laplace domain; the transfer function relating the pressure response at the line terminal (inducer inlet) to the external excitation is obtained in the computer program by sequential matrix substitution.

  20. Low loss injector for liquid propellant rocket engines

    Vonpragenau, G. L. (Inventor)


    A low pressure loss injector element is disclosed for the main combustion chamber of a rocket engine which includes a lox post terminating in a cylindrical barrel. Received within the barrel is a lox plug which is threaded in the lox post and includes an interchangeable lox metering sieve which meters the lox into an annular lox passage. A second annular gas passage is coaxial with the annular lox passage. A cylindrical sleeve surrounds the annular gas passage and includes an interchangeable gas metering seive having metering orifices through which a hot gas passes into the annular passage. The jets which emerge from the annular lox passage and annular gas passage intersect in a recessed area away from the combustion area. Thus, mixing and combustion stability are enhanced.

  1. Navier-Stokes calculation of solid-propellant rocket motor internal flowfields

    Hsieh, Kwang-Chung; Yang, Vigor; Tseng, Jesse I. S.


    A comprehensive numerical analysis has been carried out to study the detailed physical and chemical processes involved in the combustion of homogeneous propellant in a rocket motor. The formulation is based on the time-dependent full Navier-Stokes equations, with special attention devoted to the chemical reactions in both gas and condensed phases. The turbulence closure is achieved using both the Baldwin-Lomax algebraic model and a modified k-epsilon two-equation scheme with a low Reynolds number and near-wall treatment. The effects of variable thermodynamic and transport properties are also included. The system of governing equations are solved using a multi-stage Runge-Kutta shceme with the source terms treated implicitly. Preliminary results clearly demonstrate the presence of various combustion regimes in the vicinity of propellant surface. The effects of propellant combustion on the motor internal flowfields are investigated in detail.

  2. Alternate propellants for the space shuttle solid rocket booster motors. [for reducing environmental impact of launches


    As part of the Shuttle Exhaust Effects Panel (SEEP) program for fiscal year 1973, a limited study was performed to determine the feasibility of minimizing the environmental impact associated with the operation of the solid rocket booster motors (SRBMs) in projected space shuttle launches. Eleven hypothetical and two existing limited-experience propellants were evaluated as possible alternates to a well-proven state-of-the-art reference propellant with respect to reducing emissions of primary concern: namely, hydrogen chloride (HCl) and aluminum oxide (Al2O3). The study showed that it would be possible to develop a new propellant to effect a considerable reduction of HCl or Al2O3 emissions. At the one extreme, a 23% reduction of HCl is possible along with a ll% reduction in Al2O3, whereas, at the other extreme, a 75% reduction of Al2O3 is possible, but with a resultant 5% increase in HCl.

  3. Low-Cost High-Performance Non-Toxic Self-Pressurizing Storable Liquid Bi-Propellant Pressure-Fed Rocket Engine Project

    National Aeronautics and Space Administration — Exquadrum proposes a high-performance liquid bi-propellant rocket engine that uses propellants that are non-toxic, self-pressurizing, and low cost. The proposed...

  4. Studies of Fission Fragment Rocket Engine Propelled Spacecraft

    Werka, Robert O.; Clark, Rodney; Sheldon, Rob; Percy, Thomas K.


    The NASA Office of Chief Technologist has funded from FY11 through FY14 successive studies of the physics, design, and spacecraft integration of a Fission Fragment Rocket Engine (FFRE) that directly converts the momentum of fission fragments continuously into spacecraft momentum at a theoretical specific impulse above one million seconds. While others have promised future propulsion advances if only you have the patience, the FFRE requires no waiting, no advances in physics and no advances in manufacturing processes. Such an engine unequivocally can create a new era of space exploration that can change spacecraft operation. The NIAC (NASA Institute for Advanced Concepts) Program Phase 1 study of FY11 first investigated how the revolutionary FFRE technology could be integrated into an advanced spacecraft. The FFRE combines existent technologies of low density fissioning dust trapped electrostatically and high field strength superconducting magnets for beam management. By organizing the nuclear core material to permit sufficient mean free path for escape of the fission fragments and by collimating the beam, this study showed the FFRE could convert nuclear power to thrust directly and efficiently at a delivered specific impulse of 527,000 seconds. The FY13 study showed that, without increasing the reactor power, adding a neutral gas to the fission fragment beam significantly increased the FFRE thrust through in a manner analogous to a jet engine afterburner. This frictional interaction of gas and beam resulted in an engine that continuously produced 1000 pound force of thrust at a delivered impulse of 32,000 seconds, thereby reducing the currently studied DRM 5 round trip mission to Mars from 3 years to 260 days. By decreasing the gas addition, this same engine can be tailored for much lower thrust at much higher impulse to match missions to more distant destinations. These studies created host spacecraft concepts configured for manned round trip journeys. While the

  5. New Frontiers AO: Advanced Materials Bi-propellant Rocket (AMBR) Engine Information Summary

    Liou, Larry C.


    The Advanced Material Bi-propellant Rocket (AMBR) engine is a high performance (I(sub sp)), higher thrust, radiation cooled, storable bi-propellant space engine of the same physical envelope as the High Performance Apogee Thruster (HiPAT(TradeMark)). To provide further information about the AMBR engine, this document provides details on performance, development, mission implementation, key spacecraft integration considerations, project participants and approach, contact information, system specifications, and a list of references. The In-Space Propulsion Technology (ISPT) project team at NASA Glenn Research Center (GRC) leads the technology development of the AMBR engine. Their NASA partners were Marshall Space Flight Center (MSFC) and Jet Propulsion Laboratory (JPL). Aerojet leads the industrial partners selected competitively for the technology development via the NASA Research Announcement (NRA) process.

  6. Solid propellant rocket motor internal ballistics performance variation analysis, phase 3

    Sforzini, R. H.; Foster, W. A., Jr.; Murph, J. E.; Adams, G. W., Jr.


    Results of research aimed at improving the predictability of off nominal internal ballistics performance of solid propellant rocket motors (SRMs) including thrust imbalance between two SRMs firing in parallel are reported. The potential effects of nozzle throat erosion on internal ballistic performance were studied and a propellant burning rate low postulated. The propellant burning rate model when coupled with the grain deformation model permits an excellent match between theoretical results and test data for the Titan IIIC, TU455.02, and the first Space Shuttle SRM (DM-1). Analysis of star grain deformation using an experimental model and a finite element model shows the star grain deformation effects for the Space Shuttle to be small in comparison to those of the circular perforated grain. An alternative technique was developed for predicting thrust imbalance without recourse to the Monte Carlo computer program. A scaling relationship used to relate theoretical results to test results may be applied to the alternative technique of predicting thrust imbalance or to the Monte Carlo evaluation. Extended investigation into the effect of strain rate on propellant burning rate leads to the conclusion that the thermoelastic effect is generally negligible for both steadily increasing pressure loads and oscillatory loads.

  7. Characterization of rocket propellant combustion products: Description of sampling and analysis methods for rocket exhaust characterization studies

    Jenkins, R.A.


    A systematic approach has been developed and experimentally validated for the sampling and chemical characterization of the rocket motor exhaust generated from the firing of scaled down test motors at the US Army's Signature Characterization Facility (ASCF) at Redstone Arsenal in Huntsville, Alabama. The overall strategy was to sample and analyze major exhaust constituents in near real time, while performing off-site analyses of samples collected for the determination of trace constituents of the particulate and vapor phases. Initial interference studies were performed using atmospheric pressure burns of 1 g quantities of propellants in small chambers at Oak Ridge National Laboratory. Carbon monoxide and carbon dioxide were determined using non-dispersive infrared instrumentation. Hydrogen cyanide, hydrogen chloride, and ammonia determinations were made using ion selective electrode technology. Oxides of nitrogen were determined using chemiluminescence instrumentation. Airborne particulate mass concentration was determined using infrared forward scattering measurements and a tapered element oscillating microbalance, as well as conventional gravimetry. Particulate phase metals were determined by collection on Teflon membrane filters, followed by inductively coupled plasma and atomic absorption analysis. Particulate phase polynuclear aromatic hydrocarbons (PAH) and nitro-PAH were collected using high volume sampling on a two stage filter. Target species were extracted, and quantified by gas chromatography/mass spectrometry (GC/MS). Vapor phase species were collected on multi-sorbent resin traps, and subjected to thermal desorption GC/MS for analysis. 11 refs., 1 fig., 1 tab.

  8. New discrimination method for ablative control mechanism in solid-propellant rocket nozzle


    A reasonable discrimination method for ablative control mechanism in solid-propellant rocket nozzle can improve the calculation accuracy of ablation rate. Based on the different rate constants for reactions of C with H2O and CO2,a new discrimination method for ablative control mechanism,which comprehensively considers the influence of nozzle surface temperature and gas component concentration,is presented. Using this new discrimination method,calculations were performed to simulate the nozzle throat insert ablation. The numerical results showed that the calculated ablation rate,which was more close to the measured values,was less than the value calculated by diffusion control mechanisms or by double control mechanisms. And H2O was proved to be the most detrimental oxidizing species in nozzle ablation.

  9. Thermo-mechanical concepts applied to modeling liquid propellant rocket engine stability

    Kassoy, David R.; Norris, Adam


    The response of a gas to transient, spatially distributed energy addition can be quantified mathematically using thermo-mechanical concepts available in the literature. The modeling demonstrates that the ratio of the energy addition time scale to the acoustic time scale of the affected volume, and the quantity of energy added to that volume during the former determine the whether the responses to heating can be described as occurring at nearly constant volume, fully compressible or nearly constant pressure. Each of these categories is characterized by significantly different mechanical responses. Application to idealized configurations of liquid propellant rocket engines provides an opportunity to identify physical conditions compatible with gasdynamic disturbances that are sources of engine instability. Air Force Office of Scientific Research.

  10. Characterization of typical platelet injector flow configurations. [liquid propellant rocket engines

    Hickox, C. E.


    A study to investigate the hydraulic atomization characteristics of several novel injector designs for use in liquid propellant rocket engines is presented. The injectors were manufactured from a series of thin stainless steel platelets through which orifices were very accurately formed by a photoetching process. These individual platelets were stacked together and the orifices aligned so as to produce flow passages of prescribed geometry. After alignment, the platelets were bonded into a single, 'platelet injector', unit by a diffusion bonding process. Because of the complex nature of the flow associated with platelet injectors, it was necessary to use experimental techniques, exclusively, throughout the study. Large scale models of the injectors were constructed from aluminum plates and the appropriate fluids were modeled using a glycerol-water solution. Stop-action photographs of test configurations, using spark-shadowgraph or stroboscopic back-lighting, are shown.

  11. Cold Flow Testing for Liquid Propellant Rocket Injector Scaling and Throttling

    Kenny, Jeremy R.; Moser, Marlow D.; Hulka, James; Jones, Gregg


    Scaling and throttling of combustion devices are important capabilities to demonstrate in development of liquid rocket engines for NASA's Space Exploration Mission. Scaling provides the ability to design new injectors and injection elements with predictable performance on the basis of test experience with existing injectors and elements, and could be a key aspect of future development programs. Throttling is the reduction of thrust with fixed designs and is a critical requirement in lunar and other planetary landing missions. A task in the Constellation University Institutes Program (CUIP) has been designed to evaluate spray characteristics when liquid propellant rocket engine injectors are scaled and throttled. The specific objectives of the present study are to characterize injection and primary atomization using cold flow simulations of the reacting sprays. These simulations can provide relevant information because the injection and primary atomization are believed to be the spray processes least affected by the propellant reaction. Cold flow studies also provide acceptable test conditions for a university environment. Three geometric scales - 1/4- scale, 1/2-scale, and full-scale - of two different injector element types - swirl coaxial and shear coaxial - will be designed, fabricated, and tested. A literature review is currently being conducted to revisit and compile the previous scaling documentation. Because it is simple to perform, throttling will also be examined in the present work by measuring primary atomization characteristics as the mass flow rate and pressure drop of the six injector element concepts are reduced, with corresponding changes in chamber backpressure. Simulants will include water and gaseous nitrogen, and an optically accessible chamber will be used for visual and laser-based diagnostics. The chamber will include curtain flow capability to repress recirculation, and additional gas injection to provide independent control of the

  12. Navier-Stokes analysis of solid propellant rocket motor internal flows

    Sabnis, J. S.; Gibeling, H. J.; Mcdonald, H.


    A multidimensional implicit Navier-Stokes analysis that uses numerical solution of the ensemble-averaged Navier-Stokes equations in a nonorthogonal, body-fitted, cylindrical coordinate system has been applied to the simulation of the steady mean flow in solid propellant rocket motor chambers. The calculation procedure incorporates a two-equation (k-epsilon) turbulence model and utilizes a consistently split, linearized block-implicit algorithm for numerical solution of the governing equations. The code was validated by comparing computed results with the experimental data obtained in cylindrical-port cold-flow tests. The agreement between the computed and experimentally measured mean axial velocities is excellent. The axial location of transition to turbulent flow predicted by the two-equation (k-epsilon) turbulence model used in the computations also agrees well with the experimental data. Computations performed to simulate the axisymmetric flowfield in the vicinity of the aft field joint in the Space Shuttle solid rocket motor using 14,725 grid points show the presence of a region of reversed axial flow near the downstream edge of the slot.

  13. Navier-Stokes analysis of solid propellant rocket motor internal flows

    Sabnis, J. S.; Gibeling, H. J.; Mcdonald, H.


    A multidimensional implicit Navier-Stokes analysis that uses numerical solution of the ensemble-averaged Navier-Stokes equations in a nonorthogonal, body-fitted, cylindrical coordinate system has been applied to the simulation of the steady mean flow in solid propellant rocket motor chambers. The calculation procedure incorporates a two-equation (k-epsilon) turbulence model and utilizes a consistently split, linearized block-implicit algorithm for numerical solution of the governing equations. The code was validated by comparing computed results with the experimental data obtained in cylindrical-port cold-flow tests. The agreement between the computed and experimentally measured mean axial velocities is excellent. The axial location of transition to turbulent flow predicted by the two-equation (k-epsilon) turbulence model used in the computations also agrees well with the experimental data. Computations performed to simulate the axisymmetric flowfield in the vicinity of the aft field joint in the Space Shuttle solid rocket motor using 14,725 grid points show the presence of a region of reversed axial flow near the downstream edge of the slot.

  14. Measurements of acoustic responses of gaseous propellant injectors. [for rocket combustion

    Janardan, B. A.; Daniel, B. R.; Zinn, B. T.


    Results are presented for an investigation intended to provide experimental data that can quantitatively describe the way in which various coaxial injector designs affect the stability of gaseous propellant rocket motors. The response factors of configurations that simulate the flow conditions in a gaseous-fuel injector element and a gaseous-oxidizer injector element are measured by using a modified impedance-tube technique and under cold-flow conditions simulating those observed in rocket motors with axial instability. The measured injector response factor data are presented and discussed. It is shown that there is reasonable agreement between the measured injector response factors and those predicted by the Feiler and Heidmann model (1967), and that the orifice length can be varied to shift the resonant frequency of the injector without any change in the magnitude of the response factor at resonance. A change in the injector open-area ratio is found to have a significant effect on the characteristics of the injector response factor.

  15. Metallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Heat Transfer and Combustion Measurements

    Palaszewski, Bryan; Zakany, James S.


    A series of rocket engine heat transfer experiments using metallized gelled liquid propellants was conducted. These experiments used a small 20- to 40-lb/f thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-percentage by weight loadings of aluminum particles. Gaseous oxygen was used as the oxidizer. Three different injectors were used during the testing: one for the baseline O(2)/RP-1 tests and two for the gelled and metallized gelled fuel firings. Heat transfer measurements were made with a rocket engine calorimeter chamber and nozzle with a total of 31 cooling channels. Each chamber used a water flow to carry heat away from the chamber and the attached thermocouples and flow meters allowed heat flux estimates at each of the 31 stations. The rocket engine Cstar efficiency for the RP-1 fuel was in the 65-69 percent range, while the gelled 0 percent by weight RP-1 and the 5-percent by weight RP-1 exhibited a Cstar efficiency range of 60 to 62% and 65 to 67%, respectively. The 55-percent by weight RP-1 fuel delivered a 42-47% Cstar efficiency. Comparisons of the heat flux and temperature profiles of the RP-1 and the metallized gelled RP-1/A1 fuels show that the peak nozzle heat fluxes with the metallized gelled O2/RP-1/A1 propellants are substantially higher than the baseline O2/RP-1: up to double the flux for the 55 percent by weight RP-1/A1 over the RP-1 fuel. Analyses showed that the heat transfer to the wall was significantly different for the RP-1/A1 at 55-percent by weight versus the RP-1 fuel. Also, a gellant and an aluminum combustion delay was inferred in the 0 percent and 5-percent by weight RP-1/A1 cases from the decrease in heat flux in the first part of the chamber. A large decrease in heat flux in the last half of the chamber was caused by fuel deposition in the chamber and nozzle. The engine combustion occurred well downstream of the injector face

  16. Developmental Studies on Metallised UDMH and Kerosene Gels

    T. L. Varghese; S. C. Gaindhar; John David; Josekutty Jose; R. M. Muthiah; S. S. Rao; K. N. Ninan; V. N. Krishnamurthy


    ...) and kerosene containing 30 per cent 15 micron Aluminium was studied. Metallised UDMH and kerosene gels were characterised with respect to pseudoplasticity, thixotropy, consistency and yield stress using Contrave's rheometer...

  17. Flight Investigation of the Performance of a Two-stage Solid-propellant Nike-deacon (DAN) Meteorological Sounding Rocket

    Heitkotter, Robert H


    A flight investigation of two Nike-Deacon (DAN) two-stage solid-propellant rocket vehicles indicated satisfactory performance may be expected from the DAN meteorological sounding rocket. Peak altitudes of 356,000 and 350,000 feet, respectively, were recorded for the two flight tests when both vehicles were launched from sea level at an elevation angle of 75 degrees. Performance calculations based on flight-test results show that altitudes between 358,000 feet and 487,000 feet may be attained with payloads varying between 60 pounds and 10 pounds.

  18. Solid propellants.

    Marsh, H. E., Jr.; Hutchison, J. J.


    The basic principles underlying propulsion by rocket motor are examined together with the configuration of a solid propellant motor. Solid propellants and their preparation are discussed, giving attention to homogeneous propellants, composite propellants, energetic considerations in choosing a solid propellant, the processing of composite propellants, and some examples of new developments. The performance of solid propellants is investigated, taking into account characteristics velocity, the specific impulse, and performance calculations. Aspects of propellant development considered include nonperformance requirements for solid propellants, the approach to development, propellant mechanical properties, and future trends.

  19. Bonding agensi za kompozitna raketna goriva / Bonding agents for composite rocket propellants

    Mirjana Petrić


    Full Text Available U radu je analiziran uticaj bonding agensd na mehaničke karakteristike i viskozitet kompozitnih raketnih goriva na bazi hidroksiterminiranog polibutadiena i toluendiizocijanata, odnosno izoforondiizocijanata. Komparirana je efikasnost bonding agensd (BA različite strukture trietilentetramina (TET, tris-(2-metil aziridinil fosfin oksida (MAPO i N,N-di (2-hidroksietil-4,4-dimetilhidantoina (DHE. Efikasnost delovanja bonding agensa utvrdenaje na osnovu parametara dobijenih iz testa jednoosnog zatezanja uzoraka (σm - zatezna čvrstoća pri maksimalnoj sili, εm relativno izduženje pri maksimalnoj sili i εp - relativno izduženje pri prekidu. Mehaničke karakteristike merene su u temperaturskom području od -50°C do 50°C. Analiziran je uticaj BA na brzinu promene viskoziteta goriva. / The paper analyzes the effects of bonding agents on mechanical properties and viscosity of composite rocket propellants based on hydroxiterminated polybutadiene and toluendiizocyanate, i.e. izophorondiizocyanate. The efficiency of bonding agents (BA of different structure has been compared including triethylentetramine (TET, tris-(2-methyl azyrinidile phosphine oxide (MAPO and N, N-di (2-hydroxiethyl-4,4-dimethylhydantoine (DHE. The BA efficiency has been determined on the basis of parameters obtained by uniaxial tensile tests (σm - tensile strength at maximum force, εm - relative allongation at maximum force and εp - relative allongation at fracture. The mechanical properties have been measured at the temperature range from -50°C to 50°C. The BA effects on propellant viscosity change rates have been analyzed.

  20. Rocket Propellant Ducts (Cryogenic Fuel Lines): First Cut Approximations and Design Guidance

    Brewer, William V.


    The design team has to set parameters before analysis can take place. Analysis is customarily a thorough and time consuming process which can take weeks or even months. Only when analysis is complete can the designer obtain feedback. If margins are negative, the process must be repeated to a greater or lesser degree until satisfactory results are achieved. Reduction of the number of iterations thru this loop would beneficially conserve time and resources. The task was to develop relatively simple, easy to use, guidelines and analytic tools that allow the designer to evaluate what effect various alternatives may have on performance as the design progresses. "Easy to use" is taken to mean closed form approximations and the use of graphic methods. "Simple" implies that 2-d and quasi 3-d approximations be exploited to whatever degree is useful before more resource intensive methods are applied. The objective is to avoid the grosser violation of performance margins at the outset. Initial efforts are focused on thermal expansion/contraction and rigid body kinematics as they relate to propellant duct displacements in the gimbal plane loop (GPL). The purpose of the loop is to place two flexible joints on the same two orthogonal intersecting axes as those of the rocket motor gimbals. This supposes the ducting will flex predictably with independent rotations corresponding to those of the motor gimbal actions. It can be shown that if GPL joint axes do not coincide with motor gimbal axes, displacement incompatibilities result in less predictable movement of the ducts.

  1. Experimental study of a valveless pulse detonation rocket engine using nontoxic hypergolic propellants

    Kan, Brandon K.

    A pulsed detonation rocket engine concept was explored through the use of hypergolic propellants in a fuel-centered pintle injector combustor. The combustor design yielded a simple open ended chamber with a pintle type injection element and pressure instrumentation. High-frequency pressure measurements from the first test series showed the presence of large pressure oscillations in excess of 2000 psia at frequencies between 400-600 hz during operation. High-speed video confirmed the high-frequency pulsed behavior and large amounts of after burning. Damaged hardware and instrumentation failure limited the amount of data gathered in the first test series, but the experiments met original test objectives of producing large over-pressures in an open chamber. A second test series proceeded by replacing hardware and instrumentation, and new data showed that pulsed events produced under expanded exhaust prior to pulsing, peak pressures around 8000 psi, and operating frequencies between 400-800 hz. Later hot-fires produced no pulsed behavior despite undamaged hardware. The research succeeded in producing pulsed combustion behavior using hypergolic fuels in a pintle injector setup and provided insights into design concepts that would assist future injector designs and experimental test setups.

  2. Feasibility of an advanced thrust termination assembly for a solid propellant rocket motor


    A total of 68 quench tests were conducted in a vented bomb assembly (VBA). Designed to simulate full-scale motor operating conditions, this laboratory apparatus uses a 2-inch-diameter, end-burning propellant charge and an insulated disc of consolidated hydrated aluminum sulfate along with the explosive charge necessary to disperse the salt and inject it onto the burning surface. The VBA was constructed to permit variation of motor design parameters of interest; i.e., weight of salt per unit burning surface area, weight of explosive per unit weight of salt, distance from salt surface to burning surface, incidence angle of salt injection, chamber pressure, and burn time. Completely satisfactory salt quenching, without re-ignition, occurred in only two VBA tests. These were accomplished with a quench charge ratio (QCR) of 0.023 lb salt per square inch of burning surface at dispersing charge ratios (DCR) of 13 and 28 lb of salt per lb of explosive. Candidate materials for insulating salt charges from the rocket combustion environment were evaluated in firings of 5-inch-diameter, uncured end-burner motors. A pressed, alumina ceramic fiber material was selected for further evaluation and use in the final demonstration motor.

  3. Propellant Vaporization as a Criterion for Rocket-Engine Design; Experimental Performance, Vaporization and Heat-Transfer Rates with Various Propellant Combinations

    Clark, Bruce J.; Hersch, Martin; Priem, Richard J.


    Experimental combustion efficiencies of eleven propellant combinations were determined as a function of chamber length. Efficiencies were measured in terms of characteristic exhaust velocities at three chamber lengths and in terms of gas velocities. The data were obtained in a nominal 200-pound-thrust rocket engine. Injector and engine configurations were kept essentially the same to allow comparison of the performance. The data, except for those on hydrazine and ammonia-fluorine, agreed with predicted results based on the assumption that vaporization of the propellants determines the rate of combustion. Decomposition in the liquid phase may be.responsible for the anomalous behavior of hydrazine. Over-all heat-transfer rates were also measured for each combination. These rates were close to the values predicted by standard heat-transfer calculations except for the combinations using ammonia.

  4. Residual Fuel Expulsion from a Simulated 50,000 Pound Thrust Liquid-Propellant Rocket Engine Having a Continuous Rocket-Type Igniter

    Messing, Wesley E.


    Tests have been conducted to determine the starting characteristics of a 50,000-pound-thrust rocket engine with the conditions of a quantity of fuel lying dormant in the simulated main thrust chamber. Ignition was provided by a smaller rocket firing rearwardly along the center line. Both alcohol-water and anhydrous ammonia were used as the residual fuel. The igniter successfully expelled the maximum amount of residual fuel (3 1/2 gal) in 2.9 seconds when the igniter.was equipped with a sonic discharge nozzle operating at propellant flow rates of 3 pounds per second. Lesser amounts of residual fuel required correspondingly lower expulsion times. When the igniter was equipped with a supersonic exhaust nozzle operating at a flow of 4 pounds per second, a slightly less effective expulsion rate was encountered.

  5. The application of near-infrared spectroscopy for the quality control analysis of rocket propellant fuel pre-mixes.

    Judge, Michael D


    The viability of near-infrared (NIR) spectroscopy as a technique for the quality control analysis of ingredient concentrations in a rocket propellant fuel liquid pre-mix was investigated. The pre-mix analyzed consisted of a polybutadiene pre-polymer, a plasticizer and two antioxidants. It was determined that NIR spectroscopy offered a fast and convenient method of verifying the percentage level of all four ingredients while requiring no sample preparation. The NIR methodology exhibited a high level of accuracy and precision. There was also a clear indication that the technique allowed monitoring of antioxidant depletion in the pre-mix on ageing.

  6. Numerical simulation of multi-phase combustion flow in solid rocket motors with metalized propellant%Nmerical simulation of multi-phase combustion flow in solid rocket motors with metalized propellant

    SHAFQAT Wahab; XIE Kan; LIU Yu


    Multi-phase flow field simulation has been performed on solid rocket motor and effect of multi-phases on the performance prediction of the solid rocket motor(SRM)is in- vestigation.During the combustion of aluminized propellant,the aluminum particles in the propellant melt and form liquid aluminum at the burning propellant surface.So the flow within the rocket motor is multi phase or two phase because it contains droplets and smoke particles of Al2O3.Flow simulations have been performed on a large scale motor,to observe the effect of the flowfield on the chamber and nozzle as well.Uniform particles diameters and Rosin-Rammler diameter distribution method that is based on the assumption that an expo- nential relationship exists between the droplet diameter,d and mass fraction of droplets with diameter greater than d have been used for the simulation of different distribution of Al2O3 droplets present in SRM.Particles sizes in the range of 1-1 00μm are used,as being the most common droplets.In this approach the complete range of particle sizes is divided into a set of discrete size ranges,each to be defined by single stream that is part of the group.Roe scheme-flux differencing splitting based on approximate Riemann problem has been used to simulate the effects of the multi-phase flowfeild.This is second order upwind scheme in which flux differencing splitting method is employed.To cater for the turbulence effect, Spalart-Allmaras model has been used.The results obtained show the great sensitivity of this diameters distribution and particles concentrations to the SRM flow dynamics,primarily at the motor chamber and nozzle exit.The results are shown with various sizes of the parti- cles concentrations and geometrical configurations including models for SRM and nozzle.The analysis also provides effect of multi-phase on performance prediction of solid rocket motor.

  7. Hybrid Rocket Technology

    Sankaran Venugopal; K K Rajesh; V Ramanujachari


    With their unique operational characteristics, hybrid rockets can potentially provide safer, lower-cost avenues for spacecraft and missiles than the current solid propellant and liquid propellant systems...


    Meshkov, N A; Valtseva, E A; Kharlamova, E N; Kulikova, A Z


    Since the late 1990s, the ongoing debate about the consequences of the rocket-space activities for the health of people residing near areas offall ofseparatingfrom parts of rockets. Some scientists (Kolyado IB et al., 2001, 2013; Shoikhet YN et al., 2005, 2008; Skrebtsova NV 2005, 2006, Sidorov PI et al., 2007) argue that the main cause of morbidity is the effect of unsymmetrical dimethyl hydrazine (UDMH). However, environmentalists find it only in areas offalling fragments of separated parts of carrier rockets. Presented in the article data were obtained as a result of perennial epidemiological and hygienic research. There was performed a hygienic assessment of the content of chemical substances in water soil andfood, nutritional status and health risk near areas of the district of falling 310 and 326. There were studied conditions of work and the health of military personnel at the sites of storage of propellant components. The relationship between revealed diseases and UDMH was not established, but there was their causality due to the influence of environmental factors characteristic of territories and living conditions. In the settlements near the area of falling district 310 the share of extremely anxious persons was shown to be 1.8 times higher than in controls, which is caused by cases of falling fragments stages of carrier rockets in the territory of settlements.

  9. The starting transient of solid propellant rocket motors with high internal gas velocities. Ph.D. Thesis

    Peretz, A.; Caveny, L. H.; Kuo, K. K.; Summerfield, M.


    A comprehensive analytical model which considers time and space development of the flow field in solid propellant rocket motors with high volumetric loading density is described. The gas dynamics in the motor chamber is governed by a set of hyperbolic partial differential equations, that are coupled with the ignition and flame spreading events, and with the axial variation of mass addition. The flame spreading rate is calculated by successive heating-to-ignition along the propellant surface. Experimental diagnostic studies have been performed with a rectangular window motor (50 cm grain length, 5 cm burning perimeter and 1 cm hydraulic port diameter), using a controllable head-end gaseous igniter. Tests were conducted with AP composite propellant at port-to-throat area ratios of 2.0, 1.5, 1.2, and 1.06, and head-end pressures from 35 to 70 atm. Calculated pressure transients and flame spreading rates are in very good agreement with those measured in the experimental system.

  10. Rocket

    K. Karmarkar


    Full Text Available The rockets of World War II represented, not the invention of a new weapon, but the modernization of a very old one. As early as 1232 A.D, the Chinese launched rockets against the Mongols. About a hundred years later the knowledge of ledge of rockets was quite widespread and they were used to set fire to buildings and to terrorize the enemy. But as cannon developed, rockets declined in warfare. However rockets were used occasionally as weapons till about 1530 A.D. About this time improvements in artillery-rifled gun barrel and mechanism to absorb recoil-established a standard of efficiency with which rockets could not compare until World War II brought pew conditions

  11. Propellant grain dynamics in aft attach ring of shuttle solid rocket booster

    Verderaime, V.


    An analytical technique for implementing simultaneously the temperature, dynamic strain, real modulus, and frequency properties of solid propellant in an unsymmetrical vibrating ring mode is presented. All dynamic parameters and sources are defined for a free vibrating ring-grain structure with initial displacement and related to a forced vibrating system to determine the change in real modulus. Propellant test data application is discussed. The technique was developed to determine the aft attach ring stiffness of the shuttle booster at lift-off.

  12. DURACON - Variable Emissivity Broadband Coatings for Liquid Propellant Rocket Nozzles Project

    National Aeronautics and Space Administration — The need exists for a fast drying, robust, low gloss, black, high emissivity coating that can be applied easily on aircraft rocket nozzles and nozzle extensions....

  13. Proposal for a Concept Assessment of a Fission Fragment Rocket Engine (FFRE) Propelled Spacecraft Project

    National Aeronautics and Space Administration — A new technology, the Fission Fragment Rocket Engine (FFRE), requires small amounts of readily available, energy dense, long lasting fuel, significant thrust at...

  14. Removing hydrochloric acid exhaust products from high performance solid rocket propellant using aluminum-lithium alloy

    Terry, Brandon C., E-mail: [School of Aeronautics and Astronautics, Purdue University, Zucrow Laboratories, 500 Allison Rd, West Lafayette, IN 47907 (United States); Sippel, Travis R. [Department of Mechanical Engineering, Iowa State University, 2025 Black Engineering, Ames, IA 50011 (United States); Pfeil, Mark A. [School of Aeronautics and Astronautics, Purdue University, Zucrow Laboratories, 500 Allison Rd, West Lafayette, IN 47907 (United States); Gunduz, I.Emre; Son, Steven F. [School of Mechanical Engineering, Purdue University, Zucrow Laboratories, 500 Allison Rd, West Lafayette, IN 47907 (United States)


    Highlights: • Al-Li alloy propellant has increased ideal specific impulse over neat aluminum. • Al-Li alloy propellant has a near complete reduction in HCl acid formation. • Reduction in HCl was verified with wet bomb experiments and DSC/TGA-MS/FTIR. - Abstract: Hydrochloric acid (HCl) pollution from perchlorate based propellants is well known for both launch site contamination, as well as the possible ozone layer depletion effects. Past efforts in developing environmentally cleaner solid propellants by scavenging the chlorine ion have focused on replacing a portion of the chorine-containing oxidant (i.e., ammonium perchlorate) with an alkali metal nitrate. The alkali metal (e.g., Li or Na) in the nitrate reacts with the chlorine ion to form an alkali metal chloride (i.e., a salt instead of HCl). While this technique can potentially reduce HCl formation, it also results in reduced ideal specific impulse (I{sub SP}). Here, we show using thermochemical calculations that using aluminum-lithium (Al-Li) alloy can reduce HCl formation by more than 95% (with lithium contents ≥15 mass%) and increase the ideal I{sub SP} by ∼7 s compared to neat aluminum (using 80/20 mass% Al-Li alloy). Two solid propellants were formulated using 80/20 Al-Li alloy or neat aluminum as fuel additives. The halide scavenging effect of Al-Li propellants was verified using wet bomb combustion experiments (75.5 ± 4.8% reduction in pH, ∝ [HCl], when compared to neat aluminum). Additionally, no measurable HCl evolution was detected using differential scanning calorimetry coupled with thermogravimetric analysis, mass spectrometry, and Fourier transform infrared absorption.

  15. Removing hydrochloric acid exhaust products from high performance solid rocket propellant using aluminum-lithium alloy.

    Terry, Brandon C; Sippel, Travis R; Pfeil, Mark A; Gunduz, I Emre; Son, Steven F


    Hydrochloric acid (HCl) pollution from perchlorate based propellants is well known for both launch site contamination, as well as the possible ozone layer depletion effects. Past efforts in developing environmentally cleaner solid propellants by scavenging the chlorine ion have focused on replacing a portion of the chorine-containing oxidant (i.e., ammonium perchlorate) with an alkali metal nitrate. The alkali metal (e.g., Li or Na) in the nitrate reacts with the chlorine ion to form an alkali metal chloride (i.e., a salt instead of HCl). While this technique can potentially reduce HCl formation, it also results in reduced ideal specific impulse (ISP). Here, we show using thermochemical calculations that using aluminum-lithium (Al-Li) alloy can reduce HCl formation by more than 95% (with lithium contents ≥15 mass%) and increase the ideal ISP by ∼7s compared to neat aluminum (using 80/20 mass% Al-Li alloy). Two solid propellants were formulated using 80/20 Al-Li alloy or neat aluminum as fuel additives. The halide scavenging effect of Al-Li propellants was verified using wet bomb combustion experiments (75.5±4.8% reduction in pH, ∝ [HCl], when compared to neat aluminum). Additionally, no measurable HCl evolution was detected using differential scanning calorimetry coupled with thermogravimetric analysis, mass spectrometry, and Fourier transform infrared absorption. Copyright © 2016 Elsevier B.V. All rights reserved.

  16. Viscoelastic properties of hydroxyl-terminated poly(butadiene based composite rocket propellants

    Brzić Saša J.


    Full Text Available In the present study, the viscoelastic response of three composite solid propellants based on hydroxyl-terminated poly(butadiene, ammonium perchlorate and aluminum has been investigated. The investigation was surveyed by dynamic mechanical analysis over a wide range of temperatures and frequencies. The mechanical properties of these materials are related to the macromolecular structure of the binder as well as to the content and nature of solid fillers. The storage modulus, loss modulus, loss factor and glass transition temperature for each propellant sample have been evaluated. The master curves of storage (log G' vs log ω and loss modulus (log G'' vs log ω were generated for each propellant. A comparison of logaT vs temperature curves for all propellants indicate conformance to Williams-Landel-Ferry equation. Choosing the glass transition as the reference temperature, WLF equation constants are determined. Fractional free volume at the glass transition temperature and thermal coefficient of free volume expansion values are in accordance with the consideration that Al is reinforcing filler.

  17. Experimental Study on the Combustion and Microexplosion of Freely Falling Gelled Unsymmetrical Dimethylhydrazine (UDMH Fuel Droplets

    Jianjun Wu


    Full Text Available The increasing demand for high energy density fuels and the concern for their safety have propelled research in the field of gelled propellants, where understanding the combustion of single gelled fuel droplets is the first stage to predict the spray combustion characteristics. The experiments utilized single-isolated freely falling gelled unsymmetrical dimethylhydrazine (UDMH droplets instead of the conventional suspended droplet approach, in order to eliminate the perturbation associated with the suspension mechanism. Morphological transformations of the gelled droplet involved in the combustion processes were monitored by employing a high-speed digital camera, while the effects of ambient pressure and oxygen concentration on burning rate constants were also studied. The experimental results show that four main phenomena (droplet deformation, bubble formation and growth, vapor jetting and luminous jetting flame with “horn” shape and three distinct phases were identified in the droplet combustion process; the high yield stress and polymer chain structure of polymer gellant are responsible for the appearance of bubbles with almost the same order of magnitude as the droplets. Increasing the ambient pressure can increase the burning rate, postpone the appearance of microexplosions, and weaken microexplosion intensity; while increasing the ambient oxygen concentration can promote the appearance of microexplosions, strengthen microexplosion intensity and increase the burning rate.

  18. Simple-1: Development stage of the data transmission system for a solid propellant mid-power rocket model

    Yarce, Andrés; Sebastián Rodríguez, Juan; Galvez, Julián; Gómez, Alejandro; García, Manuel J.


    This paper presents the development stage of a communication module for a solid propellant mid-power rocket model. The communication module was named. Simple-1 and this work considers its design, construction and testing. A rocket model Estes Ventris Series Pro II® was modified to introduce, on the top of the payload, several sensors in a CanSat form factor. The Printed Circuit Board (PCB) was designed and fabricated from Commercial Off The Shelf (COTS) components and assembled in a cylindrical rack structure similar to this small format satellite concept. The sensors data was processed using one Arduino Mini and transmitted using a radio module to a Software Defined Radio (SDR) HackRF based platform on the ground station. The Simple-1 was tested using a drone in successive releases, reaching altitudes from 200 to 300 meters. Different kind of data, in terms of altitude, position, atmospheric pressure and vehicle temperature were successfully measured, making possible the progress to a next stage of launching and analysis.

  19. Propellants

    Lt. Col. B. N Mitra


    Full Text Available There is little doubt that explosives had their origin in warfare. In the armed conflict between groups of individuals or of states, where each sought and still seeks to impose its will upon the other by force, it was inevitable that arms should grow and flourish. The sling, the bow and arrow, the sword and firearm typify evolution in warfare weapons. Bs a means of propelling missiles, the gun and gun powder were thought of. The history of explosives, therefore, may be said to begin with black powder.

  20. Pulsed-Laser, High Speed Photography of Rocket Propellant Surface Deflagration.


    quanti- tative statistical analysis would be of tremendous help to the modelers . This type of Information is available in high-speed movies of flows...fivJ,’t(,rriiilation to be completely consumed, movies sholld * mad- av-r extended time periods. Mu, !I of those cuantitative data could be organized...The emphasis in this program was to deliver to AFRPL a data base on the microscopic and transient combustion of propellants which can be used in models

  1. Assessment of Poisson's Ratio for Hydroxy-terminated Polybutadine-based Solid Rocket Propellants

    Himanshu Shekhar


    Full Text Available Poisson's ratio of hydroxy-terminated polybutadine (HTPB-based composite propellant is estimated from uni-axial tensile testing. Double dumbbell specimens as per ASTM D638 type IV standard were used and Poisson's ratio at break, obtained by change in volume of specimen, was calculated as approximately 0.25. It was also observed that Poisson's ratio is different along different lateral directions of the propellant specimen. Poisson's ratios in two orthogonal directions perpendicular to longitudinal axis were calculated as 0.17 and 0.30. As ASTM specimen has rectangular cross-section of approximate size 6 mm x 4 mm, the directional behaviour of Poisson's ratio may be attributed to initial dimensions. Prismatic propellant specimen with square cross-section and of 115 mm x 6 mm x 6 mm dimension do not show any variation wrt Young's modulus,tensile strength, and percentage elongation as compared to ASTM specimen. Directional behaviour of Poisson's ratio with almost similar numerical value was again observed, thus ruling out dependence of this behaviour on different initial dimensions of propellant cross-section. Further, Poisson's ratio varies linearly with strain even in linear portion of stress-strain curve in uni-axial tensile testing. The rate of reduction of Poisson's ratio with increase in strain is slower in linear region and it accelerates after dewetting due to formation of vacuoles. Variation of Poisson's ratio with strain has two different slopes in linear (slope = 0.3165 and nonlinear regions (slope = 0.61364. Numerical value of slope for variation of Poisson's ratio with strain almost doubles after dewetting. It must be noted that no change in volume does not necessarily indicate constant Poisson's ratioequal to 0.5. Composite propellants behave as compressible material in most of the regions and near-failure region or at higher strains; Poisson's ratio is not anywhere near to 0.5, instead it is near 0.25.Defence Science

  2. Flow visualization of a rocket injector spray using gelled propellant simulants

    Green, James M.; Rapp, Douglas C.; Roncace, James


    A study was conducted at NASA-Lewis to compare the atomization characteristics of gelled and nongelled propellant simulants. A gelled propellant simulant composed of water, sodium hydroxide, and an acrylic acid polymer resin (as the gelling agent) was used to simulate the viscosity of an aluminum/PR-1 metallized fuel gel. Water was used as a comparison fluid to isolate the rheological effects of the water-gel and to simulate nongelled RP-1. The water-gel was injected through the central orifice of a triplet injector element and the central post of a coaxial injector element. Nitrogen gas flowed through the outer orifices of the triplet injector element and through the annulus of the coaxial injector element and atomized the gelled and nongelled liquids. Photographs of the water-gel spray patterns at different operating conditions were compared with images obtained using water and nitrogen. A laser light was used for illumination of the sprays. The results of the testing showed that the water sprays produced a finer and more uniform atomization than the water-gel sprays. Rheological analysis of the water-gel showed poor atomization caused by high viscosity of water-gel delaying the transition to turbulence.

  3. Liquid propellant rocket engine combustion simulation with a time-accurate CFD method

    Chen, Y. S.; Shang, H. M.; Liaw, Paul; Hutt, J.


    Time-accurate computational fluid dynamics (CFD) algorithms are among the basic requirements as an engineering or research tool for realistic simulations of transient combustion phenomena, such as combustion instability, transient start-up, etc., inside the rocket engine combustion chamber. A time-accurate pressure based method is employed in the FDNS code for combustion model development. This is in connection with other program development activities such as spray combustion model development and efficient finite-rate chemistry solution method implementation. In the present study, a second-order time-accurate time-marching scheme is employed. For better spatial resolutions near discontinuities (e.g., shocks, contact discontinuities), a 3rd-order accurate TVD scheme for modeling the convection terms is implemented in the FDNS code. Necessary modification to the predictor/multi-corrector solution algorithm in order to maintain time-accurate wave propagation is also investigated. Benchmark 1-D and multidimensional test cases, which include the classical shock tube wave propagation problems, resonant pipe test case, unsteady flow development of a blast tube test case, and H2/O2 rocket engine chamber combustion start-up transient simulation, etc., are investigated to validate and demonstrate the accuracy and robustness of the present numerical scheme and solution algorithm.

  4. Nuclear magnetic resonance imaging of solid rocket propellants at 14.1 T.

    Maas, W E; Merwin, L H; Cory, D G


    Proton NMR images of solid propellant materials, consisting of a polybutadiene binder material filled with 82% solid particles, have been obtained at a magnetic field strength of 14.1 T and at a resolution of 8.5 x 8.5 micron. The images are the first of elastomeric materials obtained at a proton frequency of 600 MHz and have the highest spatial resolution yet reported. The images display a high contrast and are rich in information content. They reveal the distribution of individual filler particles in the polymer matrix as well as a thin polymer film of about 10-30 micron which is found to surround some of the larger filler particles.

  5. Influence of lead Inorganic Compounds on Combustion Rate of Double Base Rocket Propellants

    V. B. Pillai


    Full Text Available The influence of lead nitrate, red lead, lead chromate, lead floride and lead carbonate on the combustion behaviour of double base propellants in the pressure range-35-140kg/cm /sup 2/ was studied. While all these compounds increased burning rates in lower pressure range (35-60 kg/cm/sup 2/ and higher pressure range (120-140 kg/cm/sup 2/, only lead chromate and lead fluoride were effective in the intermediate pressure range of 60-105 kg/cm/sup 2/. None of these compounds were effective as platonizer, except lead fluoride, which lowered n value to 0.34 in the lower pressure range. Addition of carbon black along with lead compounds raised burning rates further and reduced n values significantly in the higher pressure regins. A probable mechanism on the role of lead compounds studied has been suggested based on burning rate and DTA results.

  6. Performance of a UTC FW-4S solid propellant rocket motor under the command effects of simulated altitude and rotational spin

    Merryman, H. L.; Smith, L. R.


    One United Technology Center FW-4S solid-propellant rocket motor was fired at an average simulated altitude of 103,000 ft while spinning about its axial centerline at 180 rpm. The objectives of the test program were to determine motor altitude ballistic performance including the measurement of the nonaxial thrust vector and to demonstrate structural integrity of the motor case and nozzle. These objectives are presented and discussed.

  7. Performance Evaluation and Experimental Studies on Metallised Gel Propellants

    T. L. Varghese


    Full Text Available Metallised gel propellants offer higher specific impulse and volumetric loading, reduced vaporisation loss, spillage and slosh problems and easy storage in comparison to the conventional liquid propellants. Theoretical performance analysis of gel propellant containing Al in unsymmetrical dimethyl hydrazine-dinitrogen tetroxide (UDMH-N/sub 2/O/sub 4} system shows peak Isp (vacuum condition of 316.7 s and 318.3 s at oxidiser/fuel (O/f ratios of 1.5 and 1.0, respectively for 30 per cent and 40 per cent UDMH-Al gel propellants, under standard conditions. The effect of other parameters like area ratio and chamber pressure on performance has been brought out in view of mission oriented applications. Aluminium has been found to be a better choice over magnesium in metallised gel propellants. Experimental studies on UDMH gellation using propellant grade (15 micrometerand pyrotechnic grade (1.5 micrometerAl in 500g batch level show that gellant(methyl cellulose concentration could be reduced by 50 percent using pyrotechnic grade Al. The pseudoplastic-thixotropic behaviour, flow rate through die holes, burst pressure tests and bulk density are studied. UDMH -25 to 30 per cent Al gels with both grades of Al are found to be stable, pseudoplastic (shear thinning and thixotropic (time-dependent shear thinning, but their flow pattern through die holes differ in nature.

  8. Velocity-coupled flow oscillations in a simulated solid-propellant rocket environment

    Yang, Vigor; Hsieh, Kwang-Chung; Tseng, Jesse I. S.


    A comprehensive numerical analysis has been carried out to study the unsteady flowfields in a simulated rocket-motor environment. The model is based on the time-dependent compressible Navier-Stokes equations with a two-equation turbulence closure scheme. Various important aspects of the coupling between acoustic oscillations and mean flowfields, including flow reversal, modification of transport properties, etc., are addressed. Results indicate that multi-dimensional effects play important roles in determining local flow structures and wave characteristics. In much of the domain, acoustic velocity nodal points are observed in the near-wall region. The classical one-dimensional theory fails to describe several important mechanisms associated with velocity-induced flow instabilities.

  9. Development of the platelet micro-orifice injector. [for liquid propellant rocket engines

    La Botz, R. J.


    For some time to come, liquid rocket engines will continue to provide the primary means of propulsion for space transportation. The injector represents a key to the optimization of engine and system performance. The present investigation is concerned with a unique injector design and fabrication process which has demonstrated performance capabilities beyond that achieved with more conventional approaches. This process, which is called the 'platelet process', makes it feasible to fabricate injectors with a pattern an order of magnitude finer than that obtainable by drilling. The fine pattern leads to an achievement of high combustion efficiencies. Platelet injectors have been identified as one of the significant technology advances contributing to the feasibility of advanced dual-fuel booster engines. Platelet injectors are employed in the Space Shuttle Orbit Maneuvering System (OMS) engines. Attention is given to injector design theory as it relates to pattern fineness, a description of platelet injectors, and test data obtained with three different platelet injectors.

  10. Rocket propulsion elements - An introduction to the engineering of rockets (6th revised and enlarged edition)

    Sutton, George P.

    The subject of rocket propulsion is treated with emphasis on the basic technology, performance, and design rationale. Attention is given to definitions and fundamentals, nozzle theory and thermodynamic relations, heat transfer, flight performance, chemical rocket propellant performance analysis, and liquid propellant rocket engine fundamentals. The discussion also covers solid propellant rocket fundamentals, hybrid propellant rockets, thrust vector control, selection of rocket propulsion systems, electric propulsion, and rocket testing.

  11. Uncertainty Quantification of Non-linear Oscillation Triggering in a Multi-injector Liquid-propellant Rocket Combustion Chamber

    Popov, Pavel; Sideris, Athanasios; Sirignano, William


    We examine the non-linear dynamics of the transverse modes of combustion-driven acoustic instability in a liquid-propellant rocket engine. Triggering can occur, whereby small perturbations from mean conditions decay, while larger disturbances grow to a limit-cycle of amplitude that may compare to the mean pressure. For a deterministic perturbation, the system is also deterministic, computed by coupled finite-volume solvers at low computational cost for a single realization. The randomness of the triggering disturbance is captured by treating the injector flow rates, local pressure disturbances, and sudden acceleration of the entire combustion chamber as random variables. The combustor chamber with its many sub-fields resulting from many injector ports may be viewed as a multi-scale complex system wherein the developing acoustic oscillation is the emergent structure. Numerical simulation of the resulting stochastic PDE system is performed using the polynomial chaos expansion method. The overall probability of unstable growth is assessed in different regions of the parameter space. We address, in particular, the seven-injector, rectangular Purdue University experimental combustion chamber. In addition to the novel geometry, new features include disturbances caused by engine acceleration and unsteady thruster nozzle flow.

  12. A detailed numerical simulation of a liquid-propellant rocket engine ground test experiment

    Lankford, D. W.; Simmons, M. A.; Heikkinen, B. D.


    A computational simulation of a Liquid Rocket Engine (LRE) ground test experiment was performed using two modeling approaches. The results of the models were compared with selected data to assess the validity of state-of-the-art computational tools for predicting the flowfield and radiative transfer in complex flow environments. The data used for comparison consisted of in-band station radiation measurements obtained in the near-field portion of the plume exhaust. The test article was a subscale LRE with an afterbody, resulting in a large base region. The flight conditions were such that afterburning regions were observed in the plume flowfield. A conventional standard modeling approach underpredicted the extent of afterburning and the associated radiation levels. These results were attributed to the absence of the base flow region which is not accounted for in this model. To assess the effects of the base region a Navier-Stokes model was applied. The results of this calculation indicate that the base recirculation effects are dominant features in the immediate expansion region and resulted in a much improved comparison. However, the downstream in-band station radiation data remained underpredicted by this model.

  13. Solid-propellant rocket motor internal ballistic performance variation analysis, phase 2

    Sforzini, R. H.; Foster, W. A., Jr.


    The Monte Carlo method was used to investigate thrust imbalance and its first time derivative throughtout the burning time of pairs of solid rocket motors firing in parallel. Results obtained compare favorably with Titan 3 C flight performance data. Statistical correlations of the thrust imbalance at various times with corresponding nominal trace slopes suggest several alternative methods of predicting thrust imbalance. The effect of circular-perforated grain deformation on internal ballistics is discussed, and a modified design analysis computer program which permits such an evaluation is presented. Comparisons with SRM firings indicate that grain deformation may account for a portion of the so-called scale factor on burning rate between large motors and strand burners or small ballistic test motors. Thermoelastic effects on burning rate are also investigated. Burning surface temperature is calculated by coupling the solid phase energy equation containing a strain rate term with a model of gas phase combustion zone using the Zeldovich-Novozhilov technique. Comparisons of solutions with and without the strain rate term indicate a small but possibly significant effect of the thermoelastic coupling.

  14. Asbestos Free Insulation Development for the Space Shuttle Solid Propellant Rocket Motor (RSRM)

    Allred, Larry D.; Eddy, Norman F.; McCool, A. A. (Technical Monitor)


    Asbestos has been used for many years as an ablation inhibitor in insulating materials. It has been a constituent of the AS/NBR insulation used to protect the steel case of the RSRM (Reusable Solid Rocket Motor) since its inception. This paper discusses the development of a potential replacement RSRM insulation design, several of the numerous design issues that were worked and processing problems that were resolved. The earlier design demonstration on FSM-5 (Flight Support Motor) of the selected 7% and 11% Kevlar(registered) filled EPDM (KF/EPDM) candidate materials was expanded. Full-scale process simulation articles were built and FSM-8 was manufactured using multiple Asbestos Free (AF) components and materials. Two major problems had to be overcome in developing the AF design. First, bondline corrosion, which occurred in the double-cured region of the aft dome, had to be eliminated. Second, KF/EPDM creates high levels of electrostatic energy (ESE), which does not readily dissipate from the insulation surface. An uncontrolled electrostatic discharge (ESD) of this surface energy during many phases of production could create serious safety hazards. Numerous processing changes were implemented and a conductive paint was developed to prevent exposed external insulation surfaces from generating ESE/ESD. Additionally, special internal instrumentation was incorporated into FSM-8 to record real-time internal motor environment data. These data included inhibitor insulation erosion rates and internal thermal environments. The FSM-8 static test was successfully conducted in February 2000 and much valuable data were obtained to characterize the AF insulation design.

  15. Design and evaluation of an oxidant-fuel-ratio-zoned rocket injector for high performance and ablative engine compatibility

    Winter, J. M.; Pavli, A. J.; Shinn, A. M., Jr.


    A method for temperature control of the combustion gases in the peripheral zone of a rocket combustor which would reduce ablative throat erosion, prevent melting of zirconia throat inserts, and maintain high combustion performance is discussed. Included are techniques for analyzing and predicting zoned injector performance, as well as the philosophy and method for accomplishing an optimum compromise between high performance and reduced effective gas temperature. The experimental work was done with a 1000-lbf rocket engine which used as propellants N2O4 and a blend of 50-percent N2H4 and 50-percent UDMH at 100-psia chamber pressure and an overall O/F of 2.0. The method selected to provide temperature control was to use 30 percent of the propellant to form a peripheral zone of combustion gases at an O/F of 1.31 and 2700 K. The remaining 70 percent of the propellant in the core was at an O/F of 2.45 to keep the overall O/F at 2.0.

  16. Enantioselective synthesis of aziridines using asymmetric transfer hydrogenation as a precursor for chiral derivatives used as bonding agent for rocket solid propellants

    Aparecida M. Kawamoto


    Full Text Available A rapid, expedient and enantioselective method for the synthesis of beta-hydroxy amines and monosubstituted aziridines in up to 99% e.e., via asymmetric transfer hydrogenation of a-amino ketones and cyclisation through treatment with tosyl chloride and base, is described. (1R,2R-N-(para-toluenesulfonyl-1,2-ethylenediamine with formic acid has been utilised as a ligand for the Ruthenium (II catalysed enantioselective transfer hydrogenation of the ketones.The chiral 2-methyl aziridine, which is a potentially more efficient bonding agent for Rocket Solid Propellant has been successfully achieved.

  17. Introduction to rocket science and engineering

    Taylor, Travis S


    What Are Rockets? The History of RocketsRockets of the Modern EraRocket Anatomy and NomenclatureWhy Are Rockets Needed? Missions and PayloadsTrajectoriesOrbitsOrbit Changes and ManeuversBallistic Missile TrajectoriesHow Do Rockets Work? ThrustSpecific ImpulseWeight Flow RateTsiolkovsky's Rocket EquationStagingRocket Dynamics, Guidance, and ControlHow Do Rocket Engines Work? The Basic Rocket EngineThermodynamic Expansion and the Rocket NozzleExit VelocityRocket Engine Area Ratio and LengthsRocket Engine Design ExampleAre All Rockets the Same? Solid Rocket EnginesLiquid Propellant Rocket Engines

  18. Technology Method Design of Assembly and Testing for Solid Propellant Rocket Engine of Aviation Seat%航空座椅固体火箭发动机装配及检测工艺技术设计



    本文对航空座椅某型固体火箭发动机部装、总装及检测、试验、包装技术难点等进行了工艺分析;介绍了固体火箭发动机装配全过程工艺流程、检测、试验方法及注意事项等,对于同类及新型火箭发动机的装配制造过程具有良好的借鉴、推广应用意义。%Aiming at the difficulty of solid propellant rocket engine of aviation seat to assembly, testing and packaging technology, the assembly, testing process and method for solid propellant rocket engine were introduced. It can be regarded as reference with application for solid propellant rocket engine assembly process.

  19. Soil factors of ecosystems' disturbance risk reduction under the impact of rocket fuel

    Krechetov, Pavel; Koroleva, Tatyana; Sharapova, Anna; Chernitsova, Olga


    Environmental impacts occur at all stages of space rocket launch. One of the most dangerous consequences of a missile launch is pollution by components of rocket fuels ((unsymmetrical dimethylhydrazine (UDMH)). The areas subjected to falls of the used stages of carrier rockets launched from the Baikonur cosmodrome occupy thousands of square kilometers of different natural landscapes: from dry steppes of Kazakhstan to the taiga of West Siberia and mountains of the Altai-Sayany region. The study aims at assessing the environmental risk of adverse effects of rocket fuel on the soil. Experimental studies have been performed on soil and rock samples with specified parameters of the material composition. The effect of organic matter, acid-base properties, particle size distribution, and mineralogy on the decrease in the concentration of UDMH in equilibrium solutions has been studied. It has been found that the soil factors are arranged in the following series according to the effect on UDMH mobility: acid-base properties > organic matter content >clay fraction mineralogy > particle size distribution. The estimation of the rate of self-purification of contaminated soil is carried out. Experimental study of the behavior of UDMH in soil allowed to define a model for calculating critical loads of UDMH in terrestrial ecosystems.

  20. Analysis of velocity-coupled response function data from the dual rotating valve. [combustion stability of solid rocket propellants

    Brown, R. S.; Waugh, R. C.


    The results of a re-evaluation of the propellant combustion data obtained using the dual valve approach for measuring velocity-coupling characteristics of solid propellants are presented. Data analysis and testing procedures are described. The velocity response is compared to pressure-coupled response data within the context of thermal wave response theory. This comparison shows important inconsistencies which cast doubt on inferring the velocity response from pressure-coupled response functions.

  1. Investigation of the role of aziridine bonding agents on the aging of the composite solid rocket propellant(CSRP)

    Amged A Ali; ZHANG Jian-wei; CAI Guo-biao


    The role of bonding agents on the aging characteristics is one of the important research topics.Aging program of the prepared propellant samples was conducting as follows:Five samples,two free of bonding agents,and three containing an aziridine based bonding agents (MAPO,HX-752,MAT4),four samples based on different bonding and curing agents all were aged at 70℃.The prepared bonding agent"MAT4"gave remarkable improvements and resulted in highly stable mechanical properties comparing with HX-752 or MAPO.The selected bonding agents family inhibited the rate of decomposition of the propellants during the aging periods and supported the propellant matrix against decomposition at the elevated temperatures.

  2. A Flight Investigation of the Damping in Roll and Rolling Effectiveness Including Aeroelastic Effects of Rocket Propelled Missile Models Having Cruciform, Triangular, Interdigitated Wings and Tails

    Hopko, R. N.


    The damping in roll and rolling effectiveness of two models of a missile having cruciform, triangular, interdigitated wings and tails have been determined through a Mach number range of 0.8 to 1.8 by utilizing rocket-propelled test vehicles. Results indicate that the damping in roll was relatively constant over the Mach umber range investigated. The rolling effectiveness was essentially constant at low supersonic speeds and increased with increasing mach numbers in excess of 1.4 over the Mach number range investigated. Aeroelastic effects increase the rolling-effectiveness parameters pb/2V divided by delta and decrease both the rolling-moment coefficient due to wing deflection and the damping-in-roll coefficient.

  3. Extended temperature range rocket injector

    Schneider, Steven J. (Inventor)


    A rocket injector is provided with multiple sets of manifolds for supplying propellants to injector elements. Sensors transmit the temperatures of the propellants to a suitable controller which is operably connnected to valves between these manifolds and propellant storage tanks. When cryogenic propellant temperatures are sensed, only a portion of the valves are opened to furnish propellants to some of the manifolds. When lower temperatures are sensed, additional valves are opened to furnish propellants to more of the manifolds.

  4. Characterization of rocket propellant combustion products. Chemical characterization and computer modeling of the exhaust products from four propellant formulations: Final report, September 23, 1987--April 1, 1990

    Jenkins, R.A.; Nestor, C.W.; Thompson, C.V.; Gayle, T.M.; Ma, C.Y.; Tomkins, B.A.; Moody, R.L.


    The overall objective of the work described in this report is four-fold: to (a) develop a standardized and experimentally validated approach to the sampling and chemical and physical characterization of the exhaust products of scaled-down rocket launch motors fired under experimentally controlled conditions at the Army`s Signature Characterization Facility (ASCF) at Redstone Arsenal in Huntsville, Alabama; (b) determine the composition of the exhaust produces; (c) assess the accuracy of a selected existing computer model for predicting the composition of major and minor chemical species; (d) recommended alternations to both the sampling and analysis strategy and the computer model in order to achieve greater congruence between chemical measurements and computer prediction. 34 refs., 2 figs., 35 tabs.

  5. Characterization and Fate of Gun and Rocket Propellant Residues on Testing and Training Ranges: Interim Report 1


    the gun barrel liners. Examples include wax , talc, and titanium dioxide. Grain size and shape The properties of the propellant are greatly...phthalate 2,4-dinitrotoluene nitroglycerin ethyl centralite triacetin 2,6-dinitrotoluene nitroguanidine akardite wax ethyl centralite diethylene...The extracts were filtered by passing each through a Millex-FH PTFE 0.45 syringe filter (Millipore Corp.). For reverse-phase high- performance liquid

  6. An Evaluation Of Rocket Parameters

    J. N. Beri


    Full Text Available The dependence of conventional parameters of internal ballistics of Solid Propellant Rockets using external burning cruciform charge, on the geometry of charge aad rocket motor is discussed and results applied in a special case.

  7. Numerical Analysis on the Thermal Safety of Solid Rocket Motor Propellant%固体发动机装药热安全性数值分析

    刘文一; 焦冀光


    Objective To investigate the safety of solid rocket motor(SRM)when it was cook-off. Methods Finite element model of solid SRM was established, and the temperature distribution and the explosion delay time of propellant in fast cook-off mode and slow cook- off mode were computed. Results Propellant reached its critical temperature (352 ℃) after 47 h slow cook-off, while it reached its critical temperature (355 ℃) after 697 s fast cook-off. Conclusion It was proven that the thermal diffusivity in fast cook-off mode was greater than that in slow cook-off mode, while the temperature gradient had an opposite trend. The reaction position of propellant was different in the two different working modes when it reached critical temperature, and the thermal storage capacity of propellant was dependent on its thickness.%目的:研究固体火箭发动机遭受火烤时的安全性。方法建立发动机有限元模型,计算推进剂在慢速烤燃和快速烤燃工况下的温度分布和爆炸延迟时间。结果推进剂慢烤47 h后达到临界温度,其值为352℃;快烤推进剂加热697 s后达到临界温度,临界温度为355℃。结论推进剂在快速烤燃模式下的热扩散速率大于慢速烤燃工况下,但是温度梯度则相反。两种工况下推进剂达到临界温度后开始反应的位置不同,推进剂厚度决定了其储热能力。

  8. Raman Spectroscopy for Instantaneous Multipoint, Multispecies Gas Concentration and Temperature Measurements in Rocket Engine Propellant Injector Flows

    Wehrmeyer, Joseph A.; Trinh, Huu Phuoc


    Propellant injector development at MSFC includes experimental analysis using optical techniques, such as Raman, fluorescence, or Mie scattering. For the application of spontaneous Raman scattering to hydrocarbon-fueled flows a technique needs to be developed to remove the interfering polycyclic aromatic hydrocarbon fluorescence from the relatively weak Raman signals. A current application of such a technique is to the analysis of the mixing and combustion performance of multijet, impinging-jet candidate fuel injectors for the baseline Mars ascent engine, which will burn methane and liquid oxygen produced in-situ on Mars to reduce the propellant mass transported to Mars for future manned Mars missions. The present technique takes advantage of the strongly polarized nature of Raman scattering. It is shown to be discernable from unpolarized fluorescence interference by subtracting one polarized image from another. Both of these polarized images are obtained from a single laser pulse by using a polarization-separating calcite rhomb mounted in the imaging spectrograph. A demonstration in a propane-air flame is presented.

  9. Solid propellant motor

    Shafer, J. I.; Marsh, H. E., Jr. (Inventor)


    A case bonded end burning solid propellant rocket motor is described. A propellant with sufficiently low modulus to avoid chamber buckling on cooling from cure and sufficiently high elongation to sustain the stresses induced without cracking is used. The propellant is zone cured within the motor case at high pressures equal to or approaching the pressure at which the motor will operate during combustion. A solid propellant motor with a burning time long enough that its spacecraft would be limited to a maximum acceleration of less than 1 g is provided by one version of the case bonded end burning solid propellant motor of the invention.

  10. 3D flash lidar performance in flight testing on the Morpheus autonomous, rocket-propelled lander to a lunar-like hazard field

    Roback, Vincent E.; Amzajerdian, Farzin; Bulyshev, Alexander E.; Brewster, Paul F.; Barnes, Bruce W.


    For the first time, a 3-D imaging Flash Lidar instrument has been used in flight to scan a lunar-like hazard field, build a 3-D Digital Elevation Map (DEM), identify a safe landing site, and, in concert with an experimental Guidance, Navigation, and Control system, help to guide the Morpheus autonomous, rocket-propelled, free-flying lander to that safe site on the hazard field. The flight tests served as the TRL 6 demo of the Autonomous Precision Landing and Hazard Detection and Avoidance Technology (ALHAT) system and included launch from NASA-Kennedy, a lunar-like descent trajectory from an altitude of 250m, and landing on a lunar-like hazard field of rocks, craters, hazardous slopes, and safe sites 400m down-range. The ALHAT project developed a system capable of enabling safe, precise crewed or robotic landings in challenging terrain on planetary bodies under any ambient lighting conditions. The Flash Lidar is a second generation, compact, real-time, air-cooled instrument. Based upon extensive on-ground characterization at flight ranges, the Flash Lidar was shown to be capable of imaging hazards from a slant range of 1 km with an 8 cm range precision and a range accuracy better than 35 cm, both at 1-σ. The Flash Lidar identified landing hazards as small as 30 cm from the maximum slant range which Morpheus could achieve (450 m); however, under certain wind conditions it was susceptible to scintillation arising from air heated by the rocket engine and to pre-triggering on a dust cloud created during launch and transported down-range by wind.

  11. 3-D Flash Lidar Performance in Flight Testing on the Morpheus Autonomous, Rocket-Propelled Lander to a Lunar-Like Hazard Field

    Roback, Vincent E.; Amzajerdian, Farzin; Bulyshev, Alexander E.; Brewster, Paul F.; Barnes, Bruce W.


    For the first time, a 3-D imaging Flash Lidar instrument has been used in flight to scan a lunar-like hazard field, build a 3-D Digital Elevation Map (DEM), identify a safe landing site, and, in concert with an experimental Guidance, Navigation, and Control (GN&C) system, help to guide the Morpheus autonomous, rocket-propelled, free-flying lander to that safe site on the hazard field. The flight tests served as the TRL 6 demo of the Autonomous Precision Landing and Hazard Detection and Avoidance Technology (ALHAT) system and included launch from NASA-Kennedy, a lunar-like descent trajectory from an altitude of 250m, and landing on a lunar-like hazard field of rocks, craters, hazardous slopes, and safe sites 400m down-range. The ALHAT project developed a system capable of enabling safe, precise crewed or robotic landings in challenging terrain on planetary bodies under any ambient lighting conditions. The Flash Lidar is a second generation, compact, real-time, air-cooled instrument. Based upon extensive on-ground characterization at flight ranges, the Flash Lidar was shown to be capable of imaging hazards from a slant range of 1 km with an 8 cm range precision and a range accuracy better than 35 cm, both at 1-delta. The Flash Lidar identified landing hazards as small as 30 cm from the maximum slant range which Morpheus could achieve (450 m); however, under certain wind conditions it was susceptible to scintillation arising from air heated by the rocket engine and to pre-triggering on a dust cloud created during launch and transported down-range by wind.

  12. FEM Modelling and Oscillation Analysis of Solid Propellant Rocket Motor%固体火箭发动机柔性喷管有限元建模及摆动分析

    王成林; 刘勇; 文立华


    The methods of FEM Modelling and Oscillation Analysis of Solid Propellant Rocket Motor were explores. The FEM model of flexible nozzle using the equivalent model of flexible joint based on the three-direction custom spring elements is build; modify the FEM model of the flexible joint according to the experiment data; and carry out the oscillation analysis of solid propellant rocket motor.%研究了固态火箭发动机柔性喷管有限元建模及摆动分析方法.利用基于自定义三向弹簧单元的柔性接头线性等效模型,建立了发动机柔性喷管有限元模型.根据试验数据对柔性接头模型进行修正,并对发动机柔性喷管进行了摆动分析.

  13. Energetic Properties of Rocket Propellants Evaluated through the Computational Determination of Heats of Formation of Nitrogen-Rich Compounds.

    Forquet, Valérian; Miró Sabaté, Carles; Chermette, Henry; Jacob, Guy; Labarthe, Émilie; Delalu, Henri; Darwich, Chaza


    The use of ab initio and DFT methods to calculate the enthalpies of formation of solid ionic compounds is described. The results obtained from the calculations are then compared with those from experimental measurements on nitrogen-rich salts of the 2,2-dimethyltriazanium cation (DMTZ) synthesized in our laboratory and on other nitrogen-rich ionic compounds. The importance of calculating accurate volumes and lattice enthalpies for the determination of heats of formation is also discussed. Furthermore, the crystal structure and hydrogen-bonding networks of the nitroformate salt of the DMTZ cation is described in detail. Lastly, the theoretical heats of formation were used to calculate the specific impulses (Isp ) of the salts of the DMTZ cation in view of a prospective application in propellant formulations.

  14. Photographic Study of Combustion in a Rocket Engine I : Variation in Combustion of Liquid Oxygen and Gasoline with Seven Methods of Propellant Injection

    Bellman, Donald R; Humphrey, Jack C


    Motion pictures at camera speeds up to 3000 frames per second were taken of the combustion of liquid oxygen and gasoline in a 100-pound-thrust rocket engine. The engine consisted of thin contour and injection plates clamped between two clear plastic sheets forming a two-dimensional engine with a view of the entire combustion chamber and nozzle. A photographic investigation was made of the effect of seven methods of propellant injection on the uniformity of combustion. From the photographs, it was found that the flame front extended almost to the faces of the injectors with most of the injection methods, all the injection systems resulted in a considerable nonuniformity of combustion, and luminosity rapidly decreased in the divergent part of the nozzle. Pressure vibration records indicated combustion vibrations that approximately corresponded to the resonant frequencies of the length and the thickness of the chamber. The combustion temperature divided by the molecular weight of the combustion gases as determined from the combustion photographs was about 50 to 70 percent of the theoretical value.

  15. The Thermal State Computational Research of the Low-Thrust Oxygen-Methane Gaseous-Propellant Rocket Engine in the Pulse Mode of Operation

    O. A. Vorozheeva


    Full Text Available Currently promising development direction of space propulsion engineering is to use, as spacecraft controls, low-thrust rocket engines (RDTM on clean fuels, such as oxygen-methane. Modern RDTM are characterized by a lack regenerative cooling and pulse mode of operation, during which there is accumulation of heat energy to lead to the high thermal stress of RDTM structural elements. To get an idea about the thermal state of its elements, which further will reduce the number of fire tests is therefore necessary in the development phase of a new product. Accordingly, the aim of this work is the mathematical modeling and computational study of the thermal state of gaseous oxygen-methane propellant RDMT operating in pulse mode.In this paper we consider a model RDTM working on gaseous propellants oxygen-methane in pulse mode.To calculate the temperature field of the chamber wall of model RDMT under consideration is used the mathematical model of non-stationary heat conduction in a two-dimensional axisymmetric formulation that takes into account both the axial heat leakages and the nonstationary processes occurring inside the chamber during pulse operation of RDMT.As a result of numerical study of the thermal state of model RDMT, are obtained the temperature fields during engine operation based on convective, conductive, and radiative mechanisms of heat transfer from the combustion products to the wall.It is shown that the elements of flanges of combustion chamber of model RDMT act as heat sinks structural elements. Temperatures in the wall of the combustion chamber during the engine mode of operation are considered relatively low.Raised temperatures can also occur in the mixing head in the feeding area of the oxidant into the combustion chamber.During engine operation in the area forming the critical section, there is an intensive heating of a wall, which can result in its melting, which in turn will increase the minimum nozzle throat area and hence

  16. Mars Integrated Propellant Production System Project

    National Aeronautics and Space Administration — The Integrated Mars In-Situ Propellant Production System (IMISPPS) is an end-to-end system that will produce rocket propellant on Mars from CO2 in the Martian...

  17. Mars Integrated Propellant Production System Project

    National Aeronautics and Space Administration — The Integrated Mars In-Situ Propellant Production System (IMISPPS) is an end-to-end system that will produce rocket propellant on Mars from CO2 in the Martian...

  18. Solid Propellant Grain Structural Integrity Analysis


    The structural properties of solid propellant rocket grains were studied to determine the propellant resistance to stresses. Grain geometry, thermal properties, mechanical properties, and failure modes are discussed along with design criteria and recommended practices.



  20. Lidar Sensor Performance in Closed-Loop Flight Testing of the Morpheus Rocket-Propelled Lander to a Lunar-Like Hazard Field

    Roback, V. Eric; Pierrottet, Diego F.; Amzajerdian, Farzin; Barnes, Bruce W.; Bulyshev, Alexander E.; Hines, Glenn D.; Petway, Larry B.; Brewster, Paul F.; Kempton, Kevin S.


    For the first time, a suite of three lidar sensors have been used in flight to scan a lunar-like hazard field, identify a safe landing site, and, in concert with an experimental Guidance, Navigation, and Control (GN&C) system, help to guide the Morpheus autonomous, rocket-propelled, free-flying lander to that safe site on the hazard field. The lidar sensors and GN&C system are part of the Autonomous Precision Landing and Hazard Detection and Avoidance Technology (ALHAT) project which has been seeking to develop a system capable of enabling safe, precise crewed or robotic landings in challenging terrain on planetary bodies under any ambient lighting conditions. The 3-D imaging Flash Lidar is a second generation, compact, real-time, aircooled instrument developed from a number of components from industry and NASA and is used as part of the ALHAT Hazard Detection System (HDS) to scan the hazard field and build a 3-D Digital Elevation Map (DEM) in near-real time for identifying safe sites. The Flash Lidar is capable of identifying a 30 cm hazard from a slant range of 1 km with its 8 cm range precision (1-s). The Flash Lidar is also used in Hazard Relative Navigation (HRN) to provide position updates down to a 250m slant range to the ALHAT navigation filter as it guides Morpheus to the safe site. The Navigation Doppler Lidar (NDL) system has been developed within NASA to provide velocity measurements with an accuracy of 0.2 cm/sec and range measurements with an accuracy of 17 cm both from a maximum range of 2,200 m to a minimum range of several meters above the ground. The NDLâ€"TM"s measurements are fed into the ALHAT navigation filter to provide lander guidance to the safe site. The Laser Altimeter (LA), also developed within NASA, provides range measurements with an accuracy of 5 cm from a maximum operational range of 30 km down to 1 m and, being a separate sensor from the Flash Lidar, can provide range along a separate vector. The LA measurements are also fed

  1. Liquid Rocket Engine Testing


    booster rocket engines • 6000-10000 psia capabilities – Can use gaseous nitrogen, helium, or hydrogen to pressurize propellant tanks 9Distribution A...Approved for Public Release; Distribution Unlimited. PA Clearance 16493 Simplified Test Stand Layout Oxidizer  TankFuel  Tank High  Pressure   Gas (GN2...requires large, complex facilities to deliver propellant at the proper pressure , temperature, and flow rates • The enormous energies involved

  2. Influence of Rocket Engine Characteristics on Shaft Sealing Technology Needs

    Keba, John E.


    This paper presents viewgraphs of The Influence of Rocket Engine Characteristics on Shaft Sealing Technology Needs. The topics include: 1) Rocket Turbomachinery Shaft Seals (Inter-Propellant-Seal (IPS) Systems, Lift-off Seal Systems, and Technology Development Needs); 2) Rocket Engine Characteristics (Engine cycles, propellants, missions, etc., Influence on shaft sealing requirements); and 3) Conclusions.

  3. Innovative Swirl Injector for LOX and Hydrocarbon Propellants Project

    National Aeronautics and Space Administration — Gases trapped in the propellant feed lines of space-based rocket engines due to cryogenic propellant boil-off or pressurant ingestion can result in poor combustion...

  4. Low acid producing solid propellants

    Bennett, Robert R.


    The potential environmental effects of the exhaust products of conventional rocket propellants have been assessed by various groups. Areas of concern have included stratospheric ozone, acid rain, toxicity, air quality and global warming. Some of the studies which have been performed on this subject have concluded that while the impacts of rocket use are extremely small, there are propellant development options which have the potential to reduce those impacts even further. This paper discusses the various solid propellant options which have been proposed as being more environmentally benign than current systems by reducing HCI emissions. These options include acid neutralized, acid scavenged, and nonchlorine propellants. An assessment of the acid reducing potential and the viability of each of these options is made, based on current information. Such an assessment is needed in order to judge whether the potential improvements justify the expenditures of developing the new propellant systems.

  5. Inter-Batch Variation and the Effect of Casting Vacuum on Ballistic and Mechanical Properties of a High Performing Cast Composite Rocket Propellant


    13 8 Scrape down - - - 9 Mix 30 0 45 10 Scrape down - - - 11 Add IPDI 30 1 4 12 Scrape down - - - 13 Mix 30 0 20 The dry clumps of material...http:// web Booster, HTPB, high performance composite propellant 19

  6. The mixing of solid propellant by an artificial muscle actuator

    岩崎, 祥大; 伴, 遼介; 吉浜, 舜; 中村, 太郎; 羽生, 宏人; Iwasaki, Akihiro; Ban, Ryosuke; Yoshihama, Shun; Nakamura, Taro; Habu, Hiroto


    This research aims to reduce the cost of the solid rocket motor production, mainly solid propellant. The production process of the solid rocket propellant are usually employed the multi-batch mixing. However, this study using a peristaltic pump as a mixer will lead to the continuous process. The pump system can mix the powder materials for propellant and we consider that it will make the slurry of the solid propellant efficiently by the mechanism of the fluid dynamics in the pump.

  7. Mars Rocket Propulsion System

    Zubrin, Robert; Harber, Dan; Nabors, Sammy


    A report discusses the methane and carbon monoxide/LOX (McLOx) rocket for ascent from Mars as well as other critical space propulsion tasks. The system offers a specific impulse over 370 s roughly 50 s higher than existing space-storable bio-propellants. Current Mars in-situ propellant production (ISPP) technologies produce impure methane and carbon monoxide in various combinations. While separation and purification of methane fuel is possible, it adds complexity to the propellant production process and discards an otherwise useful fuel product. The McLOx makes such complex and wasteful processes unnecessary by burning the methane/CO mixtures produced by the Mars ISPP systems without the need for further refinement. Despite the decrease in rocket-specific impulse caused by the CO admixture, the improvement offered by concomitant increased propellant density can provide a net improvement in stage performance. One advantage is the increase of the total amount of propellant produced, but with a decrease in mass and complexity of the required ISPP plant. Methane/CO fuel mixtures also may be produced by reprocessing the organic wastes of a Moon base or a space station, making McLOx engines key for a human Lunar initiative or the International Space Station (ISS) program. Because McLOx propellant components store at a common temperature, very lightweight and compact common bulkhead tanks can be employed, improving overall stage performance further.

  8. The mechanism of lead catalysis of double-base rocket propellant combustion: Super-rate burning and the plateau/mesa phenomena

    Graham, W.H.; Neidert, J.B. (Thiokol Corp., Huntsville, AL (United States))


    Although double-base propellants contain a fairly complex mixture of ingredients, the most unstable compounds present are the nitrate esters, nitrocellulose and nitrate ester plasticizers, which represent the majority of most formulations; usually the ballistic additives, a lead compound and often carbon, represent from 1% to 5% of the total propellant. Energetic additives such as RDX and HMX are often present, but these nitramines are less responsive to lead catalysis which is characterized by a region of super-rate burning at lower pressures followed by a plateau region at higher pressure as the catalytic effectiveness appears to decrease and often a mesa region as the burn rate actually decreases with pressure. To explain both the super-rate burning and the plateau/mesa phenomena, one should logically examine the chemistry of decomposition of the nitrate esters to explain both lead catalysis and the apparent loss of catalysis; fortunately, examination of the reported chemistry of nitrate esters and several more recent investigations, which used sophisticated experimental techniques to examine double-base propellant combustion, provide support for a comprehensive mechanism to explain both.

  9. Magnesium Based Rockets for Martian Exploration Project

    National Aeronautics and Space Administration — We propose to develop Mg rockets for Martian ascent vehicle applications. The propellant can be acquired in-situ from MgO in the Martian regolith (5.1% Mg by mass)...

  10. Nanoparticles for solid rocket propulsion

    Galfetti, L [Politecnico di Milano, SPLab, Milan (Italy); De Luca, L T [Politecnico di Milano, SPLab, Milan (Italy); Severini, F [Politecnico di Milano, SPLab, Milan (Italy); Meda, L [Polimeri Europa, Istituto G Donegani, Novara (Italy); Marra, G [Polimeri Europa, Istituto G Donegani, Novara (Italy); Marchetti, M [Universita di Roma ' La Sapienza' , Dipartimento di Ingegneria Aerospaziale ed Astronautica, Rome (Italy); Regi, M [Universita di Roma ' La Sapienza' , Dipartimento di Ingegneria Aerospaziale ed Astronautica, Rome (Italy); Bellucci, S [INFN, Laboratori Nazionali di Frascati, Frascati (Italy)


    The characterization of several differently sized aluminium powders, by BET (specific surface), EM (electron microscopy), XRD (x-ray diffraction), and XPS (x-ray photoelectron spectroscopy), was performed in order to evaluate their application in solid rocket propellant compositions. These aluminium powders were used in manufacturing several laboratory composite solid rocket propellants, based on ammonium perchlorate (AP) as oxidizer and hydroxil-terminated polybutadiene (HTPB) as binder. The reference formulation was an AP/HTPB/Al composition with 68/17/15% mass fractions respectively. The ballistic characterization of the propellants, in terms of steady burning rates, shows better performance for propellant compositions employing nano-aluminium when compared to micro-aluminium. Results obtained in the pressure range 1-70 bar show that by increasing the nano-Al mass fraction or decreasing the nano-Al size, larger steady burning rates are measured with essentially the same pressure sensitivity.

  11. Nanoparticles for solid rocket propulsion

    Galfetti, L.; DeLuca, L. T.; Severini, F.; Meda, L.; Marra, G.; Marchetti, M.; Regi, M.; Bellucci, S.


    The characterization of several differently sized aluminium powders, by BET (specific surface), EM (electron microscopy), XRD (x-ray diffraction), and XPS (x-ray photoelectron spectroscopy), was performed in order to evaluate their application in solid rocket propellant compositions. These aluminium powders were used in manufacturing several laboratory composite solid rocket propellants, based on ammonium perchlorate (AP) as oxidizer and hydroxil-terminated polybutadiene (HTPB) as binder. The reference formulation was an AP/HTPB/Al composition with 68/17/15% mass fractions respectively. The ballistic characterization of the propellants, in terms of steady burning rates, shows better performance for propellant compositions employing nano-aluminium when compared to micro-aluminium. Results obtained in the pressure range 1-70 bar show that by increasing the nano-Al mass fraction or decreasing the nano-Al size, larger steady burning rates are measured with essentially the same pressure sensitivity.

  12. Composite Solid Propellant Predictability and Quality Assurance

    Ramohalli, Kumar


    Reports are presented at the meeting at the University of Arizona on the study of predictable and reliable solid rocket motors. The following subject areas were covered: present state and trends in the research of solid propellants; the University of Arizona program in solid propellants, particularly in mixing (experimental and analytical results are presented).

  13. Materials characterization of propellants using ultrasonics

    Workman, Gary L.; Jones, David


    Propellant characteristics for solid rocket motors were not completely determined for its use as a processing variable in today's production facilities. A major effort to determine propellant characteristics obtainable through ultrasonic measurement techniques was performed in this task. The information obtained was then used to determine the uniformity of manufacturing methods and/or the ability to determine non-uniformity in processes.

  14. The scaling of the threshold conditions for solid propellant erosive burning

    Strand, L. D.; Nguyen, M. H.; Cohen, N. S.


    Rocket test firings were performed to measure the transition length threshold conditions while systematically varying various rocket motor parameters. These include the crossflow velocity, the chamber pressure, the propellant nonerosive burning rate, the propellant surface roughness, and the motor port diameter. The erosive burning trends with varying propellant burning rate, motor chamber pressure, and mass flow rate are consistent with published results.

  15. Measuring Combustion Advance in Solid Propellants

    Yang, L. C.


    Set of gauges on solid-propellant rocket motor with electrically insulating case measures advance of combustion front and local erosion rates of propellant and insulation. Data furnished by gauges aid in motor design, failure analysis, and performance prediction. Technique useful in determining propellant uniformity and electrical properties of exhaust plum. Gauges used both in flight and on ground. Foilgauge technique also useful in basic research on pulsed plasmas or combustion of solids.

  16. Hybrid Rocket Technology

    Sankaran Venugopal


    Full Text Available With their unique operational characteristics, hybrid rockets can potentially provide safer, lower-cost avenues for spacecraft and missiles than the current solid propellant and liquid propellant systems. Classical hybrids can be throttled for thrust tailoring, perform in-flight motor shutdown and restart. In classical hybrids, the fuel is stored in the form of a solid grain, requiring only half the feed system hardware of liquid bipropellant engines. The commonly used fuels are benign, nontoxic, and not hazardous to store and transport. Solid fuel grains are not highly susceptible to cracks, imperfections, and environmental temperature and are therefore safer to manufacture, store, transport, and use for launch. The status of development based on the experience of the last few decades indicating the maturity of the hybrid rocket technology is given in brief.Defence Science Journal, 2011, 61(3, pp.193-200, DOI:

  17. Thermodynamic cycle analysis on solid propellant air-turbo-rocket%固体推进剂吸气式涡轮火箭发动机的气动热力循环分析

    屠秋野; 丁朝霞; 陈玉春; 蔡元虎


    A numerical model for calculating performance of solid propellant air-turbo-rocket at design point was set up,and a relationship expression of fuel-air ratio of combustion based on compressor pressure ratio,tubine intet total temperature and turbine expansion ratio was put forward.And a relationship between the turbine expansion ratio and the bypass ratio was given.The effects of compressor pressure ratio, turbine inlet temperature,bypass ratio/turbine expansion ratio and flight Mach number on the specific thrust and specific impulse were analyzed quantitatively.%建立了固体推进剂吸气式涡轮火箭发动机的设计状态数值模型,提出了基于压气机增压比、涡轮前温度和涡轮落压比关系的燃烧室燃气与空气配比表达式,以及涡轮落压比和发动机涵道比的匹配关系.定量分析了压气机增压比、涡轮进口燃气总温、涵道比/涡轮落压比和飞行马赫数对固体推进剂吸气式涡轮火箭发动机的单位推力和比冲的影响.

  18. Adsorption and chemical reaction of gaseous mixtures of hydrogen chloride and water on aluminum oxide and application to solid-propellant rocket exhaust clouds

    Cofer, W. R., III; Pellett, G. L.


    Hydrogen chloride (HCl) and aluminum oxide (Al2O3) are major exhaust products of solid rocket motors (SRM). Samples of calcination-produced alumina were exposed to continuously flowing mixtures of gaseous HCl/H2O in nitrogen. Transient sorption rates, as well as maximum sorptive capacities, were found to be largely controlled by specific surface area for samples of alpha, theta, and gamma alumina. Sorption rates for small samples were characterized linearly with an empirical relationship that accounted for specific area and logarithmic time. Chemisorption occurred on all aluminas studied and appeared to form from the sorption of about a 2/5 HCl-to-H2O mole ratio. The chemisorbed phase was predominantly water soluble, yielding chloride/aluminum III ion mole ratios of about 3.3/1 suggestive of dissolved surface chlorides and/or oxychlorides. Isopiestic experiments in hydrochloric acid indicated that dissolution of alumina led to an increase in water-vapor pressure. Dissolution in aqueous SRM acid aerosol droplets, therefore, might be expected to promote evaporation.

  19. Performance Charts for Multistage Rocket Boosters

    MacKay, John S.; Weber, Richard J.


    Charts relating the stage propellant fractions are given for two-and three-stage rockets launching payloads into nominal low-altitude circular orbits about the earth. A simple method is described for extending these data to higher orbit or escape missions. Various combinations of stages using RP - liquid-oxygen and hydrogen - liquid-oxygen propellants are considered. However, the results can be generalized with little error to any other propellant combination.Charts relating the stage propellant fractions are given for two-and three-stage rockets launching payloads into nominal low-altitude circular orbits about the earth. A simple method is described for extending these data to higher orbit or escape missions. Various combinations of stages using RP - liquid-oxygen and hydrogen - liquid-oxygen propellants are considered. However, the results can be generalized with little error to any other propellant combination.

  20. Mobius propeller

    Leonid I. Gretchihin


    Full Text Available The article gives a detailed molecular-kinetic theory of the Mobius propeller functioning and shows the implementation of its computer modelling in different exploitation conditions. The mechanisation of the Mobius propeller can be carried out in such a way that, under certain conditions, it enables using this type of propellers as a heat pump. The developed theory of the Mobius propeller functioning has been experimentally verified in laboratory conditions.


    Lovinskaya, A V; Kolumbayeva, S Zh; Abilev, S K; Kolomiets, O L


    There was performed an assessment of genotoxic effects of rocket fuel component--unsymmetrical dimethylhydrazine (UDMH, heptyl)--on forming germ cells of male mice. Immunocytochemically there was studied the structure of meiotic nuclei at different times after the intraperitoneal administration of UDMH to male mice. There were revealed following types of disturbances of the structure of synaptonemal complexes (SCs) of meiotic chromosomes: single and multiple fragments of SCs associations of autosomes with a sex bivalent, atypical structure of the SCs with a frequency higher than the reference level. In addition, there were found the premature desinapsis of sex bivalents, the disorder offormation of the genital corpuscle and ring SCs. Established disorders in SCs of spermatocytes, analyzed at 38th day after the 10-days intoxication of animal by the component of rocket fuel, attest to the risk of permanent persistence of chromosomal abnormalities occurring in the pool of stem cells.

  2. Reducing Thrusts In Solid-Fuel Rockets

    Bement, Laurence J.


    Thrust-terminating system conceived to reduce thrust of solid-propellant rocket motor in controlled manner such that thrust loads not increased or decreased beyond predictable levels. Concept involves explosively cutting opposing venting pairs in case of rocket motor above nozzles to initiate venting of chamber and reduction of thrust. Vents sized and numbered to control amount and rate of reduction in thrust.

  3. Measuring Propellant Stress Relaxation Modulus Using Dynamic Mechanical Analyzer


    P. N., Singh, P. P., and Bhattacharya, B., “Determination of Activation Energy of Relaxation Events in Composite Solid Propellants by Dynamic...Article 3. DATES COVERED (From - To) 04 August 2016 – 29 March 2017 4. TITLE AND SUBTITLE Measuring Propellant Stress Relaxation Modulus Using Dynamic...ERC 14. ABSTRACT A method for determining the stress relaxation master curve of solid rocket propellants was developed. The propellant was tested in

  4. High-speed schlieren imaging of rocket exhaust plumes

    Coultas-McKenney, Caralyn; Winter, Kyle; Hargather, Michael


    Experiments are conducted to examine the exhaust of a variety of rocket engines. The rocket engines are mounted in a schlieren system to allow high-speed imaging of the engine exhaust during startup, steady state, and shutdown. A variety of rocket engines are explored including a research-scale liquid rocket engine, consumer/amateur solid rocket motors, and water bottle rockets. Comparisons of the exhaust characteristics, thrust and cost for this range of rockets is presented. The variety of nozzle designs, target functions, and propellant type provides unique variations in the schlieren imaging.

  5. Rocket Science at the Nanoscale.

    Li, Jinxing; Rozen, Isaac; Wang, Joseph


    Autonomous propulsion at the nanoscale represents one of the most challenging and demanding goals in nanotechnology. Over the past decade, numerous important advances in nanotechnology and material science have contributed to the creation of powerful self-propelled micro/nanomotors. In particular, micro- and nanoscale rockets (MNRs) offer impressive capabilities, including remarkable speeds, large cargo-towing forces, precise motion controls, and dynamic self-assembly, which have paved the way for designing multifunctional and intelligent nanoscale machines. These multipurpose nanoscale shuttles can propel and function in complex real-life media, actively transporting and releasing therapeutic payloads and remediation agents for diverse biomedical and environmental applications. This review discusses the challenges of designing efficient MNRs and presents an overview of their propulsion behavior, fabrication methods, potential rocket fuels, navigation strategies, practical applications, and the future prospects of rocket science and technology at the nanoscale.

  6. Low-thrust rocket trajectories

    Keaton, P.W.


    The development of low-thrust propulsion systems to complement chemical propulsion systems will greatly enhance the evolution of future space programs. Two advantages of low-thrust rockets are stressed: first, in a strong gravitational field, such as occurs near the Earth, freighter missions with low-thrust engines require one-tenth as much propellant as do chemical engines. Second, in a weak gravitational field, such as occurs in the region between Venus and Mars, low-thrust rockets are faster than chemical rockets with comparable propellant mass. The purpose here is to address the physics of low-thrust trajectories and to interpret the results with two simple models. Analytic analyses are used where possible - otherwise, the results of numerical calculations are presented in graphs. The author has attempted to make this a self-contained report. 57 refs., 10 figs.

  7. Low-thrust rocket trajectories

    Keaton, P.W.


    The development of low-thrust propulsion systems to complement chemical propulsion systems will greatly enhance the evolution of future space programs. Two advantages of low-thrust rockets are stressed: first, in a strong gravitational field, such as occurs near the Earth, freighter missions with low-thrust engines require one-tenth as much propellant as do chemical engines. Second, in a weak gravitational field, such as occurs in the region between Venus and Mars, low-thrust rockets are faster than chemical rockets with comparable propellant mass. The purpose here is to address the physics of low-thrust trajectories and to interpret the results with two simple models. Analytic analyses are used where possible - otherwise, the results of numerical calculations are presented in graphs. The author has attempted to make this a self-contained report.

  8. Rocket noise - A review

    McInerny, S. A.


    This paper reviews what is known about far-field rocket noise from the controlled studies of the late 1950s and 1960s and from launch data. The peak dimensionless frequency, the dependence of overall sound power on exhaust parameters, and the directivity of the overall sound power of rockets are compared to those of subsonic jets and turbo-jets. The location of the dominant sound source in the rocket exhaust plume and the mean flow velocity in this region are discussed and shown to provide a qualitative explanation for the low peak Strouhal number, fD(e)/V(e), and large angle of maximum directivity. Lastly, two empirical prediction methods are compared with data from launches of a Titan family vehicle (two, solid rocket motors of 5.7 x 10 to the 6th N thrust each) and the Saturn V (five, liquid oxygen/rocket propellant engines of 6.7 x 10 to the 6th N thrust, each). The agreement is favorable. In contrast, these methods appear to overpredict the far-field sound pressure levels generated by the Space Shuttle.

  9. Ignition Studies on Aluminised Propellant.

    K. A. Bhaskaran


    Full Text Available An experimental investigation on the ignition of metallised propellants (APIHTPB/AI has been carried out 10 determine the ignition delay, minimum ignition energy and corresponding heat flux,threshold heat flux for ignition and minimum ignition temperature, Ignition experiments were conductedusing a shock tube under convectiveheating conditions similar to those prevailingin a rocket motor. Heat flux at propellant location was measured by thin film heat flux gauge and also calculated from a ribbon thermocouple output under similar test conditions. The igntion delay was measured as the time lag between the arrival of hot gas at the propellant and the light emission due to actual ignition of the propellant. The experimental results indicate that the ignition delay characteristics are independent of pressure. The minimum energy for ignition obtained for the propellant is 1100J/m2 corresponding to the heat flux range of 80·120 WIcm2 for a gas velocity of 110 mls. The threshold heat flux required to ignite the propellant was 40 W/cm2 at a velocity of 110 mls. Heat flux corresponding to minimum ignition energy and the threshold heat flux increase with gas velocity. The threshold ignition temperature of the propellant was found to be 600 ± 20 K.

  10. Rocket Flight.

    Van Evera, Bill; Sterling, Donna R.


    Describes an activity for designing, building, and launching rockets that provides students with an intrinsically motivating and real-life application of what could have been classroom-only concepts. Includes rocket design guidelines and a sample grading rubric. (KHR)

  11. Pozicioniranje, orijentisanje i određivanje daljine do cilja na samohodnom višecevnom raketnom lansirnom sistemu korišćenjem GPS i elektronskih karata / Positioning, orientation and determination of the distance to target on a self-propelled multiple rocket launcher system using GPS and electronic maps

    Dragoljub J. Sekulović


    Full Text Available U radu je prikazano pozicioniranje i orijentisanje višecevnog raketnog sistema korišćenjem GPS prijemnika i određivanje rastojanja od lansera do cilja upotrebom geografskih koordinata. Koordinate koje se koriste za zadavanje pozicije cilja su geografske ili UTM, a dobijene su sa elektronske karte terena. / The GPS positioning and orientation of a self-propelled multiple rocket launcher is presented. The determination of the distance from the launcher to target using geographic coordinates is given as well The coordinates applied in determining the target position are either geographic or UTM coordinates while terrain electronic maps are obtained as a result.

  12. 基于动态云BP网络的液体火箭发动机故障诊断方法%Fault diagnosis method for liquid-propellant rocket engines based on the dynamic cloud-BP neural network

    刘垠杰; 黄强; 程玉强; 吴建军


    将云模型与BP(backpropagation)神经网络以串联方式有机结合,首先利用云变换方法进行网络的结构辨识和云模型的特征提取,同时通过在输入层引入单位延时环节描述发动机工作过程动态特性,研究提出了基于动态云BP网络的液体火箭发动机故障诊断方法.结合实际试车数据的验证结果表明,该方法能够准确识别发动机已有的3种故障模式,通过在试车数据中添加0期望、0.2标准差的随机噪声的方法来模拟环境噪声和测试过程中产生的随机噪声,根据持续性原则,方法仍能够正确进行故障检测与分类.方法单步运行时长为1.124x10-4,完全能够满足实时性要求.%A fault diagnosis method for liquid-propellant rocket engines was proposed based on the dynamic cloud-BP(back propagation) neural network in the way of the integration of cloud model and BP neural network.The Cloud transform method was used to identify the network configuration and to extract the cloud features.And a unit time-delay was also introduced into the input layer to describe the dynamic characteristics of the engine.Results with test data show that the method can isolate the existed 3 fault modes precisely.A 0 expectation,0.2 standard deviation noise was used to simulate the entironmental noise and stochastic noise,and the method can still detect and classify the fault accurately acount to lasting-rule.The method can run in real-time with the single processing time being 1.124×10-4 s.

  13. Advanced Solid Rocket Launcher and Its Evolution

    Morita, Yasuhiro; Imoto, Takayuki; Habu, Hiroto; Ohtsuka, Hirohito; Hori, Keiichi; Koreki, Takemasa; Fukuchi, Apollo; Uekusa, Yasuyuki; Akiba, Ryojiro

    The research on next generation solid propellant rockets is actively underway in various spectra. JAXA is developing the Advanced Solid Rocket (ASR) as a successor to the M-V launch vehicle, which was utilized over past ten years for space science programs including planetary missions. ASR is a result of the development of the next generation technology including a highly intelligent autonomous check-out system, which is connected to not only the solid rocket but also future transportation systems. It is expected to improve the efficiency of the launch system and double the cost performance. Far beyond this effort, the passion of the volunteers among the industry-government-academia cooperation has been united to establish the society of the freewheeling thinking “Next generation Solid Rocket Society (NSRS)”. It aims at a larger revolution than what the ASR provides so that the order of the cost performance is further improved. A study of the Low melting temperature Thermoplastic Propellant (LTP) is now at the experimental stage, which is expected to reform the manufacturing process of the solid rocket propellant and lead to a significant increase in cost performance. This paper indicates the direction of the big flow towards the next generation solid-propellant rockets: the concept of the intelligent ASR under development; and the innovation behind LTP.

  14. Liquid rocket engine injectors

    Gill, G. S.; Nurick, W. H.


    The injector in a liquid rocket engine atomizes and mixes the fuel with the oxidizer to produce efficient and stable combustion that will provide the required thrust without endangering hardware durability. Injectors usually take the form of a perforated disk at the head of the rocket engine combustion chamber, and have varied from a few inches to more than a yard in diameter. This monograph treats specifically bipropellant injectors, emphasis being placed on the liquid/liquid and liquid/gas injectors that have been developed for and used in flight-proven engines. The information provided has limited application to monopropellant injectors and gas/gas propellant systems. Critical problems that may arise during injector development and the approaches that lead to successful design are discussed.

  15. Processing solid propellants for recycling

    Whinnery, L.L.; Griffiths, S.K.; Handrock, J.L.; Lipkin, J.


    Rapid evolution in the structure of military forces worldwide is resulting in the retirement of numerous weapon systems. Many of these systems include rocket motors containing highly energetic propellants based on hazardous nitrocellulose/nitroglycerin (NC/NG) mixtures. Even as the surplus quantities of such material increases, however, current disposal methods -- principally open burning and open detonation (OB/OD) -- are coming under close scrutiny from environmental regulators. Environmentally conscious alternatives to disposal of propellant and explosives are thus receiving renewed interest. Recycle and reuse alternatives to OB/OD appear particularly attractive because some of the energetic materials in the inventories of surplus weapon systems represent potentially valuable resources to the commercial explosives and chemical industries. The ability to reclaim such resources is therefore likely to be a key requirement of any successful technology of the future in rocket motor demilitarization. This document consists of view graphs from the poster session.

  16. Biodegradation of rocket propellant waste, ammonium perchlorate

    Naqvi, S. M. Z.; Latif, A.


    The short term effects of ammonium perchlorate on selected organisms were studied. A long term experiment was also designed to assess the changes incurred by ammonium perchlorate on the nitrogen and chloride contents of soil within a period of 3 years. In addition, an attempt was made to produce methane gas from anaerobic fermentation of the aquatic weed, Alternanthera philoxeroides.

  17. Combustion Instabilities In Solid Propellant Rocket Motors


    34 AIAA Paper No. 98{3218. Avalon, G., Ugurtas, B. Grisch, F. and Bresson , F. (2000) \\Numerical Computations and Visualization Tests of the Flow Inside...Burners and Related Devices," Prog. in Comb. Energy and Science, Vol. 19, No. 4, pp. 313{364. Roberts , A.K. and Brownlee, W.G. (1971) \\Nonlinear

  18. Coaxial Propellant Injectors With Faceplate Annulus Control

    Horn, Mark D.; Miyata, Shinjiro; Farhangi, Shahram


    An improved design concept for coaxial propellant injectors for a rocket engine (or perhaps for a non-rocket combustion chamber) offers advantages of greater robustness, less complexity, fewer parts, lower cost, and less bulk, relative to prior injectors of equivalent functionality. This design concept is particularly well suited to small, tight-tolerance injectors, for which prior designs are not suitable because the practical implementation of those designs entails very high costs and difficulty in adhering to the tolerances.

  19. Rheology of composite solid propellants during motor casting

    Rogers, C. J.; Smith, P. L.; Klager, K.


    In a study conducted to evaluate flow parameters of uncured solid composite propellants during motor casting, two motors (1.8M-lb grain wt) were cast with a PBAN propellant exhibiting good flow characteristics in a 260-in. dia solid rocket motor. Attention is given to the effects of propellant compositional and processing variables on apparent viscosity as they pertain to rheological behavior and grain defect formation during casting. It is noted that optimized flow behavior is impaired with solid propellant loading. Non-Newtonian pseudoplastic flow is observed, which is dependent upon applied shear stress and the age of the uncured propellant.

  20. Biodegradation of Unsymmetrical Dimethylhydrazine in Solution and Soil by Bacteria Isolated from Activated Sludge

    Qili Liao


    Full Text Available The biodegradation effect and pathway of unsymmetrical dimethylhydrazine (UDMH, which is a major rocket propellant with highly toxic properties, with two strains isolated from the acclimated activated sludge were investigated in solution and in soil. The results demonstrated that Stenotrophomonas sp. M12 (M12 was able to degrade UDMH of 50 mg·L−1 as the sole carbon source in aqueous mineral salt medium (MSM, but could not degrade UDMH in soil. Comamonas sp. P4 (P4 barely degraded UDMH of 50 mg·L−1 as the sole carbon source in aqueous MSM, but the degrading capacity of P4 could be improved by the addition of an extra carbon source. Meanwhile, P4 was able to degrade UDMH of 100–600 mg·kg−1 in the soil. The degradation of UDMH in the soil was influenced by organic matter, autochthonous microorganisms, and metal ions. UDMH could inhibit metabolism of M12 and P4, and the inhibition influence was more severe in aqueous MSM than in soil. Oxygen content was important for M12 biodegrading UDMH, and co-metabolism helped P4 to self-detoxify and self-recover. The main intermediates of UDMH were identified by Gas Chromatography-Mass Spectrometer (GC/MS qualitative analysis, and the concentrations of UDMH and its important transformation products were determined in solution and soil. According to the determination results, the synchronous degradation theory was proposed, and the degradation pathway was discussed.

  1. Mars McLOX Rocket Propulsion System Project

    National Aeronautics and Space Administration — The Methane and Carbon Monoxide/LOX rocket (MCLOX) is a technology for accomplishing ascent from Mars. Current Mars in-situ propellant production (ISPP) technologies...

  2. Particle size reduction of propellants by cryocycling

    Whinnery, L.; Griffiths, S.; Lipkin, J. [and others


    Repeated exposure of a propellant to liquid nitrogen causes thermal stress gradients within the material resulting in cracking and particle size reduction. This process is termed cryocycling. The authors conducted a feasibility study, combining experiments on both inert and live propellants with three modeling approaches. These models provided optimized cycle times, predicted ultimate particle size, and allowed crack behavior to be explored. Process safety evaluations conducted separately indicated that cryocycling does not increase the sensitivity of the propellants examined. The results of this study suggest that cryocycling is a promising technology for the demilitarization of tactical rocket motors.

  3. Analysis of a Radioisotope Thermal Rocket Engine

    Machado-Rodriguez, Jonathan P.; Landis, Geoffrey A.


    The Triton Hopper is a concept for a vehicle to explore the surface of Neptunes moon Triton, which uses a radioisotope heated rocket engine and in-situ propellant acquisition. The initial Triton Hopper conceptual design stores pressurized Nitrogen in a spherical tank to be used as the propellant. The aim of the research was to investigate the benefits of storing propellant at ambient temperature and heating it through a thermal block during engine operation, as opposed to storing gas at a high temperature.

  4. Optimization of Construction of the rocket-assisted projectile

    Arkhipov Vladimir


    Full Text Available New scheme of the rocket motor of rocket-assisted projectile providing the increase in distance of flight due to controlled and optimal delay time of ignition of the solid-propellant charge of the SRM and increase in reliability of initiation of the SRM by means of the autonomous system of ignition excluding the influence of high pressure gases of the propellant charge in the gun barrel has been considered. Results of the analysis of effectiveness of using of the ignition delay device on motion characteristics of the rocket-assisted projectile has been presented.

  5. Propellant Grade Hydrazine in Mono/Bi-propellant Thrusters: Preparation and Performance Evaluation

    S. Krishnamachary


    Full Text Available 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:

  6. Simulation on Influence of Residual Rocket Motor Propellant on Warhead Damage Effect%火箭发动机残余装药对战斗部毁伤效果影响的模拟研究

    王宁; 赵孝彬; 王晨雪; 田军


    利用铅柱压缩量、铅壔扩孔值和冲击波超压与冲量等参数定量研究了推进剂残余装药对战斗部爆炸毁伤效果的影响,并探讨了影响战斗部装药/推进剂体系铅柱压缩量的因素.试验结果表明,四种推进剂的存在对战斗部装药/推进剂体系的铅柱压缩量均有不同程度的增益作用.HTPB四组元推进剂对战斗部装药/推进剂体系的铅柱压缩量和铅壔扩孔值的贡献量要高于HTPB三组元推进剂( HTPB-3),其铅壔扩孔值的增益量分别为81.4%和54%.比例距离范围为3.5~12.5 m·kg-1/3时,HTPB-3对1 kg战斗部装药PBXN/0.6 kg HTPB-3体系的超压TNT当量和冲量TNT当量分别贡献了18.7%和19.7%.%The influence of solid propellant on warhead damage effect was studied by these quantitative parameters including lead cylinder compression value,lead block volume,shock wave overpressure and impulse. Influencing factors that effect lead cylinder compression value of warhead charge/propellant system was also discussed. The results show that four kinds of propellants all could enhance lead cylinder compression value of warhead charge/propellant system in different degrees. Contribution to lead cylinder compression value of warhead charge/ propellant system by four groups HTPB propellant is more than that by three groups HTPB propellant ( HTPB-3). Increased percentage of lead block volume is 81.4% and 54% ,respectively. When scaled distance ranges from 3.5 m· kg-1/3 to 12.5 m ? Kg-1/3,HTPB-3 contributes to shock wave overpressure TNT equivalence and impulse TNT equivalence of 1 kg PBXN/0.6 kg HTPB-3 system by an increase of 18.7% and 19.7%,respectively.

  7. The development of space solid rocket motors in China

    Jianding, Huang; Dingyou, Ye


    China has undertaken to research and develop composite solid propellant rocket motors since 1958. At the request of the development of space technology, composite solid propellant rocket motor has developed from small to large, step by step. For the past thirty eight years, much progress has made, many technical obstacles, such as motor design, case materials and their processing technology, propellant formulations and manufacture, nozzles and thrust vector control, safe ignition, environment tests, nondestructive inspection and quality assurance, static firing test and measurement etc. have been solved. A serial of solid rocket motors have been offered for China's satellites launch. The systems of research, design, test and manufacture of solid rocket motors have been formed.

  8. Regenerative Cooling for Liquid Rocket Engines



    Heat transfer in the thrust chamber is of great importance in the design of liquid propellant rocket engines.Regenerative cooling is and advanced method which can ensure not only the proper running but also higher performance of a rocket engine.The theoretical model is complicated,it relates to fluid bynamics,heat transfer,combustion.etc…,In this paper,a regenerative cooling model is presented.Effects such as radiation,heat transfer to environment,variable thermal properties and coking are included in the model.This model can be applied to all kinds of liquid propellant rocket engines as well as similar constructions.The modularized computer code is completed in the work.

  9. Solid Propellant Test Article (SPTA) Test Firing


    The Marshall Space Flight Center (MSFC) engineers test fired a 26-foot long, 100,000-pound-thrust solid rocket motor for 30 seconds at the MSFC east test area, the first test firing of the Modified NASA Motor (M-NASA Motor). The M-NASA Motor was fired in a newly constructed stand. The motor is 48-inches in diameter and was loaded with two propellant cartridges weighing a total of approximately 12,000 pounds. The purpose of the test was to learn more about solid rocket motor insulation and nozzle materials and to provide young engineers additional hands-on expertise in solid rocket motor technology. The test is a part of NASA's Solid Propulsion Integrity Program, that is to provide NASA engineers with the techniques, engineering tools, and computer programs to be able to better design, build, and verify solid rocket motors.

  10. Tip-modified Propellers

    Andersen, Poul


    The paper deals with tip-modified propellers and the methods which, over a period of two decades, have been applied to develop such propellers. The development is driven by the urge to increase the efficiency of propellers and can be seen as analogous to fitting end plates and winglets to aircraft...... wings. The literature on four different designs is reviewed: the end-plate propeller; the two-sided, shifted end-plate propeller; the tip-fin propeller; and the bladelet propeller. The conclusion is that it is indeed possible to design tip-modified propellers that, relative to an optimum conventional...

  11. The Advanced Solid Rocket Motor

    Mitchell, Royce E.


    The Advanced Solid Rocket Motor will utilize improved design features and automated manufacturing methods to produce an inherently safer propulsive system for the Space Shuttle and future launch systems. This second-generation motor will also provide an additional 12,000 pounds of payload to orbit, enhancing the utility and efficiency of the Shuttle system. The new plant will feature strip-wound, asbestos-free insulation; propellant continuous mixing and casting; and extensive robotic systems. Following a series of static tests at the Stennis Space Center, MS flights are targeted to begin in early 1997.

  12. The Advanced Solid Rocket Motor

    Mitchell, Royce E.


    The Advanced Solid Rocket Motor will utilize improved design features and automated manufacturing methods to produce an inherently safer propulsive system for the Space Shuttle and future launch systems. This second-generation motor will also provide an additional 12,000 pounds of payload to orbit, enhancing the utility and efficiency of the Shuttle system. The new plant will feature strip-wound, asbestos-free insulation; propellant continuous mixing and casting; and extensive robotic systems. Following a series of static tests at the Stennis Space Center, MS flights are targeted to begin in early 1997.

  13. Dynamic characterization and analysis of space shuttle SRM solid propellant

    Hufferd, W. L.


    The dynamic response properties of the space shuttle solid rocket moter (TP-H1148) propellant were characterized and the expected limits of propellant variability were established. Dynamic shear modulus tests conducted on six production batches of TP-H1148 at various static and dynamic strain levels over the temperature range from 40 F to 90 F. A heat conduction analysis and dynamic response analysis of the space shuttle solid rocket motor (SRM) were also conducted. The dynamic test results show significant dependence on static and dynamic strain levels and considerable batch-to-batch and within-batch variability. However, the results of the SRM dynamic response analyses clearly demonstrate that the stiffness of the propellant has no consequential on the overall SRM dynamic response. Only the mass of the propellant needs to be considered in the dynamic analysis of the space shuttle SRM.

  14. Rocket Tablet,


    is a vast and desolate world, this is a strip of mir- aculous land! How many struggling dramas full of power and * grandeur were cheered, resisted and...rocket officers and men, a group enormous and powerful , marched into this land soaked with the fresh blood of our ancestors. This place is about to...and tough pestering said he wanted an American aircraft ob- tained on the battlefield to transport goods from Lanzhou, Xian, Beijing, Guangzhou and

  15. Thermal Decomposition Behavior and Non-isothermal Decomposition Reaction of Copper(Ⅱ) Salt of 4-Hydroxy-3,5-dinitropyridine Oxide and Its Application in Solid Rocket Propellant

    CHEN Pei陈沛; ZHAO Feng-Qi赵凤起; LUO Yang罗阳; HU Rong-Zu胡荣祖; GAO Sheng-Li高胜利; ZHENG Yu-Mei郑玉梅; DENG Min-Zhi邓敏智; GAO Yin高茵


    The thermal decomposition behavior and kinetic parameters of the exothermic decomposition reactions of the title compound in a temperature-programmed mode have been investigated by means of DSC, TG-DTG and lower rate thermolysis/FTIR. The possible reaction mechanism was proposed. The critical temperature of thermal explosion was calculated. The influence of the title compound on the combustion characteristic of composite modified double base propellant containing RDX has been explored with the strand burner. The results show that the kinetic model function in differential form, apparent activation energy Ea and pre-exponential factor A of the major exothermal explosion of the compound is 312.87 C. The kinetic equation of the major exothermic decomposition process of the title compound at 0.1 MPa could be expressed as:dα/dT=1016.42(1-α)e-2.502× 104/T. As an auxiliary catalyst, the title compound can help the main catalyst lead salt of 4-hydroxy-3,5-dinitropyridine oxide to enhance the burning rate and reduce the pressure exponent of RDX-CMDB propellant.

  16. Interactive Schematic Integration Within the Propellant System Modeling Environment

    Coote, David; Ryan, Harry; Burton, Kenneth; McKinney, Lee; Woodman, Don


    Task requirements for rocket propulsion test preparations of the test stand facilities drive the need to model the test facility propellant systems prior to constructing physical modifications. The Propellant System Modeling Environment (PSME) is an initiative designed to enable increased efficiency and expanded capabilities to a broader base of NASA engineers in the use of modeling and simulation (M&S) technologies for rocket propulsion test and launch mission requirements. PSME will enable a wider scope of users to utilize M&S of propulsion test and launch facilities for predictive and post-analysis functionality by offering a clean, easy-to-use, high-performance application environment.

  17. Bistable (latching) solenoid actuated propellant isolation valve

    Wichmann, H.; Deboi, H. H.


    The design, fabrication, assembly and test of a development configuration bistable (latching) solenoid actuated propellant isolation valve suitable for the control hydrazine and liquid fluorine to an 800 pound thrust rocket engine is described. The valve features a balanced poppet, utilizing metal bellows, a hard poppet/seat interface and a flexure support system for the internal moving components. This support system eliminates sliding surfaces, thereby rendering the valve free of self generated particles.

  18. Simulation methods of rocket fuel refrigerating with liquid nitrogen and intermediate heat carrier

    O. E. Denisov; A. V. Zolin; V. V. Chugunkov


    Temperature preparation of liquid propellant components (LPC) before fueling the tanks of rocket and space technology is the one of the operations performed by ground technological complexes on cosmodromes...

  19. Solid Rocket Booster-Illustration


    This illustration is a cutaway of the solid rocket booster (SRB) sections with callouts. The Shuttle's two SRB's are the largest solids ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds, augmenting the Shuttle's main propulsion system during liftoff. The major design drivers for the solid rocket motors (SRM's) were high thrust and reuse. The desired thrust was achieved by using state-of-the-art solid propellant and by using a long cylindrical motor with a specific core design that allows the propellant to burn in a carefully controlled marner. At burnout, the boosters separate from the external tank and drop by parachute to the ocean for recovery and subsequent refurbishment. The boosters are designed to survive water impact at almost 60 miles per hour, maintain flotation with minimal damage, and preclude corrosion of the hardware exposed to the harsh seawater environment. Under the project management of the Marshall Space Flight Center, the SRB's are assembled and refurbished by the United Space Boosters. The SRM's are provided by the Morton Thiokol Corporation.

  20. Ignition transient analysis of solid rocket motor

    Han, Samuel S.


    Measurement data on the performance of Space Shuttle Solid Rocket Motor show wide variations in the head-end pressure changes and the total thrust build-up during the ignition transient periods. To analyze the flow and thermal behavior in the tested solid rocket motors, a 1-dimensional, ideal gas flow model via the SIMPLE algorithm was developed. Numerical results showed that burning patterns in the star-shaped head-end segment of the propellant and the erosive burning rate are two important factors controlling the ignition transients. The objective of this study is to extend the model to include the effects of aluminum particle commonly used in solid propellants. To treat the effects of aluminum-oxide particles in the combustion gas, conservation of mass, momentum, and energy equations for the particles are added in the numerical formulation and integrated by an inter-phase-slip algorithm.

  1. Accelerated aging of AP/HTPB propellants and the influence of various environmental aging conditions

    Keizers, H.L.J.


    Preliminary resuits on accelerated aging of lab-scale produced AP/HTPB propellant and propellants from dissectioned rocket motors are discussed, including aging logic, storage conditions, test techniques and resuits on mechanical, ballistic and safety testing. The mam aging effect observed was harde

  2. Accelerated aging of AP/HTPB propellants and the influence of various environmental aging conditions

    Keizers, H.L.J.


    Preliminary resuits on accelerated aging of lab-scale produced AP/HTPB propellant and propellants from dissectioned rocket motors are discussed, including aging logic, storage conditions, test techniques and resuits on mechanical, ballistic and safety testing. The mam aging effect observed was

  3. Propellants: the feasibility of their manufacture in India

    R. K. Srivastava


    Full Text Available In the light of the recent Chinese aggression, the modernization of our defence forces has become a vital necessity. In this paper, one aspect of this problem has been considered in detail viz. the development of an indigenous rocket and missile force. While it is true that many factors e.g. rocket motors, propellants, guidance systems etc. are involved, it is also true that a start in one area will act as an impetus to developments in the other fields. Solid/liquid propellant and oxidizer systems have been considered, the properties of solid and liquid propellants evaluated and on the basis of such comparison, it has been concluded that effort concentrated on the development of liquid propellants will be well expended. Liquid propellant/oxidizer systems have been compared amongst themselves and it has been concluded that the hydrazine fuels oxidized by RFNA/WFNA/IRFNA would represent systems fulfilling the country's immediate military needs best. The availability of raw materials for the manufacture of hydrazine fuels (and also of some solid propellants has been considered and it is shown that the necessary raw materials are available in sufficient quantities to support an indigenous propellants industry.

  4. A review of research in low earth orbit propellant collection

    Singh, Lake A.; Walker, Mitchell L. R.


    This comprehensive review examines the efforts of previous researchers to develop concepts for propellant-collecting spacecraft, estimate the performance of these systems, and understand the physics involved. Rocket propulsion requires the spacecraft to expend two fundamental quantities: energy and propellant mass. A growing number of spacecraft collect the energy they need to execute propulsive maneuvers in-situ with solar panels. In contrast, every spacecraft using rocket propulsion has carried all of the propellant mass needed for the mission from the ground, which limits the range and mission capabilities. Numerous researchers have explored the concept of collecting propellant mass while in space. These concepts have varied in scale and complexity from chemical ramjets to fusion-driven interstellar vessels. Research into propellant-collecting concepts occurred in distinct eras. During the Cold War, concepts tended to be large, complex, and nuclear powered. After the Cold War, concepts transitioned to solar power sources and more effort has been devoted to detailed analysis of specific components of the propellant-collecting architecture. By detailing the major contributions and limitations of previous work, this review concisely presents the state-of-the-art and outlines five areas for continued research. These areas include air-compatible cathode technology, techniques to improve propellant utilization on atmospheric species, in-space compressor and liquefaction technology, improved hypersonic and hyperthermal free molecular flow inlet designs, and improved understanding of how design parameters affect system performance.

  5. Aircraft propeller control

    Day, Stanley G. (Inventor)


    In the invention, the speeds of both propellers in a counterrotating aircraft propeller pair are measured. Each speed is compared, using a feedback loop, with a demanded speed and, if actual speed does not equal demanded speed for either propeller, pitch of the proper propeller is changed in order to attain the demanded speed. A proportional/integral controller is used in the feedback loop. Further, phase of the propellers is measured and, if the phase does not equal a demanded phase, the speed of one propeller is changed, by changing pitch, until the proper phase is attained.

  6. Innovative Metallized Formulations for Solid Rocket Propulsion

    Luigi T DeLUCA; Luciano GALFETTI; Filippo MAGGI; Giovanni COLOMBO; Alice REINA; Stefano DOSSI; Daniele CONSONNI; Melissa BRAMBILLA


    Several metallized solid rocket propellants,AP/Metal/HTPB in the ratio 68/18/1 4,were experimentally analyzed at the Space Propulsion Laboratory of Politecnico di Milano.Effects of the metals (micrometric and nanometric Al,B,Mg,and a variety of dual metals) on the performance of the propellant were studied and contrasted to a conventional micrometric aluminum (30 μm average grain size) taken as reference.It is shown that the propellant microstructure plays a fundamental role in controlling the critical aggregation/agglomeration phenomena occurring below and near the burning surface.Two specific effects of microstructure in terms of steady burning rate and average agglomerate size are illustrated.

  7. Improved hybrid rocket fuel

    Dean, David L.


    McDonnell Douglas Aerospace, as part of its Independent R&D, has initiated development of a clean burning, high performance hybrid fuel for consideration as an alternative to the solid rocket thrust augmentation currently utilized by American space launch systems including Atlas, Delta, Pegasus, Space Shuttle, and Titan. It could also be used in single stage to orbit or as the only propulsion system in a new launch vehicle. Compared to solid propellants based on aluminum and ammonium perchlorate, this fuel is more environmentally benign in that it totally eliminates hydrogen chloride and aluminum oxide by products, producing only water, hydrogen, nitrogen, carbon oxides, and trace amounts of nitrogen oxides. Compared to other hybrid fuel formulations under development, this fuel is cheaper, denser, and faster burning. The specific impulse of this fuel is comparable to other hybrid fuels and is between that of solids and liquids. The fuel also requires less oxygen than similar hybrid fuels to produce maximum specific impulse, thus reducing oxygen delivery system requirements.

  8. Atmospheric scavenging of solid rocket exhaust effluents

    Fenton, D. L.; Purcell, R. Y.


    Solid propellant rocket exhaust was directly utilized to ascertain raindrop scavenging rates for hydrogen chloride. Two chambers were used to conduct the experiments; a large, rigid walled, spherical chamber stored the exhaust constituents, while the smaller chamber housing all the experiments was charged as required with rocket exhaust HCl. Surface uptake experiments demonstrated an HCl concentration dependence for distilled water. Sea water and brackish water HCl uptake was below the detection limit of the chlorine-ion analysis technique used. Plant life HCl uptake experiments were limited to corn and soybeans. Plant age effectively correlated the HCl uptake data. Metallic corrosion was not significant for single 20 minute exposures to the exhaust HCl under varying relative humidity. Characterization of the aluminum oxide particles substantiated the similarity between the constituents of the small scale rocket and the full size vehicles.

  9. Spark Ignition of Combustible Vapor in a Plastic Bottle as a Demonstration of Rocket Propulsion

    Mattox, J. R.


    I report an innovation that provides a compelling demonstration of rocket propulsion, appropriate for students of physics and other physical sciences. An electrical spark is initiated from a distance to cause the deflagration of a combustible vapor mixed with air in a lightweight plastic bottle that is consequently propelled as a rocket by the…

  10. Overview of the manufacturing sequence of the Advanced Solid Rocket Motor

    Chapman, John S.; Nix, Michael B.


    The manufacturing sequence of NASA's new Advanced Solid Rocket Motor, developed as a replacement of the Space Shuttle's existing Redesigned Solid Rocket Motor, is overviewed. Special attention is given to the case preparation, the propellant mix/cast, the nondestructuve evaluation, the motor finishing, and the refurbishment. The fabrication sequences of the case, the nozzle, and the igniter are described.

  11. 液体火箭发动机液膜冷却研究综述%Review of Research on Liquid Film Cooling for Liquid-propellant Rocket Engine

    周红玲; 杨成虎; 刘犇


    液膜冷却对降低燃烧室和喷注器头部温度有显著作用,而且通道结构比较简单,因此在载人航天液体推进系统用姿轨控发动机中得到了广泛应用。液膜冷却的传热过程主要包括对流传热和沸腾传热两种形式,传质过程主要包括液膜的蒸发和中心主气流对液膜的携带。对液膜冷却过程的研究工作进行了综述,讨论了液膜冷却的异常升温现象和发生机理。%Liquid film cooling is widely used in attitude and ahitude liquid rocket engine for manned spaceflight propulsion sys- tem. With a comparatively simple channel structure, it is an effective method to protect the combustor and injector from high temperature. The two main heat transfer patterns in liquid film cooling are convection transfer and boiling heat transfer, while the process of mass transfer mainly includes liquid film evaporating and the carrying of the film by main flow. Research on the heat and mass transfer during liquid film cooling is reviewed, and the phenomenon and mecha- nism of abnormal temperature rise of liquid film cooling are discussed.

  12. On Performance Evaluation of a New Liquid Propellant

    S. P. Panda


    Full Text Available A blend of 3-carene and cardanol in 70:30 weight proportion exhibits synergistic hypergolic ignition with red fuming nitric acid (RFNA as oxidizer. Attempts have been made to evaluate this new propellant by theoretical calculationof performance parameters and verification of the results by static firing of a 10 kg thrust rocket motor around 20 atmosphers of chamber pressure. At an oxidizer-to-fuel weight ratio (O/F of 3.34 (RFNA used had 21% N204 and 5% by weight of concentrated sulphuric acid as catalyst, the propellant produced a reasonably smooth pressure-time curve with an ignition delay of 35 milliseconds. The theoretical characteristic velocity value matched well with the experimental. No carbon residue was left in the rocket motor after firing. Specific impulse (theoretical of the propellant has been found to be 223.8 seconds at chamber pressure, 20 atmos and exist pressure, 1 atmos.

  13. Studies of solid propellant combustion with pulsed radiography

    Godai, T.; Tanemura, T.; Fujiwara, T.; Shimizu, M.


    Pulsed radiography was applied to observe solid propellant surface regression during rocket motor operation. Using a 150 KV flash X-ray system manufactured by the Field Emission Corporation and two kinds of film suppliers, images of the propellant surface of a 5 cm diameter end burning rocket motor were recorded on film. The repetition frame rate of 8 pulses per second and the pulse train length of 10 pulses are limited by the capability of the power supply and the heat build up within the X-ray tube, respectively. The experiment demonstrated the effectiveness of pulsed radiography for observing solid propellant surface regression. Measuring the position of burning surface images on film with a microdensitometer, quasi-instantaneous burning rate as a function of pressure and the variation of characteristic velocity with pressure and gas stay time were obtained. Other research items to which pulsed radiography can be applied are also suggested.

  14. Rocket Scientist for a Day: Investigating Alternatives for Chemical Propulsion

    Angelin, Marcus; Rahm, Martin; Gabrielsson, Erik; Gumaelius, Lena


    This laboratory experiment introduces rocket science from a chemistry perspective. The focus is set on chemical propulsion, including its environmental impact and future development. By combining lecture-based teaching with practical, theoretical, and computational exercises, the students get to evaluate different propellant alternatives. To…

  15. Study of Liquid Breakup Process in Solid Rocket Motor Nozzle


    34Chemical Erosion of Refractory - Metal Nozzle Inserts in Solid-Propellant Rocket Motors," J. Propulsion and Power, Vol. 25, no.1,, 2009. [4] E. Y. Wong...Paul A.;, "Gelcasting of Alumina," J. Am. Ceram . Soc. 74[3], pp. 612-618, 1991. [18] Blomquist , B. A.; Fink, J. K.; Leibowitz, L.;, "The

  16. Rocket Scientist for a Day: Investigating Alternatives for Chemical Propulsion

    Angelin, Marcus; Rahm, Martin; Gabrielsson, Erik; Gumaelius, Lena


    This laboratory experiment introduces rocket science from a chemistry perspective. The focus is set on chemical propulsion, including its environmental impact and future development. By combining lecture-based teaching with practical, theoretical, and computational exercises, the students get to evaluate different propellant alternatives. To…

  17. Development of small solid rocket boosters for the ILR-33 sounding rocket

    Nowakowski, Pawel; Okninski, Adam; Pakosz, Michal; Cieslinski, Dawid; Bartkowiak, Bartosz; Wolanski, Piotr


    This paper gives an overview of the development of a 6000 Newton-class solid rocket motor for suborbital applications. The design configuration and results of interior ballistics calculations are given. The initial use of the motor as the main propulsion system of the H1 experimental in-flight test platform, within the Polish Small Sounding Rocket Program, is presented. Comparisons of theoretical and experimental performance are shown. Both on-ground and in-flight tests are discussed. A novel composite-case manufacturing technology, which enabled to reach high propellant mass fractions, was validated and significant cost-reductions were achieved. This paper focuses on the process of adapting the design for use as the booster stage of the ILR-33 sounding rocket, under development at the Institute of Aviation in Warsaw, Poland. Parallel use of two of the flight-proven rocket motors along with the main stage is planned. The process of adapting the rocket motor for booster application consists of stage integration, aerothermodynamics and reliability analyses. The separation mechanism and environmental impact are also discussed within this paper. Detailed performance analysis with focus on propellant grain geometry is provided. The evolution of the design since the first flights of the H1 rocket is covered and modifications of the manufacturing process are described. Issues of simultaneous ignition of two motors and their non-identical performance are discussed. Further applications and potential for future development are outlined. The presented results are based on the initial work done by the Rocketry Group of the Warsaw University of Technology Students' Space Association. The continuation of the Polish Small Sounding Rocket Program on a larger scale at the Institute of Aviation proves the value of the outcomes of the initial educational project.

  18. Chemical reactions identified in the Titan 2, Titan 4, and Delta 2 propellant systems and their application to source modeling

    Prince, S. P.; Banning, D. W.; Wiseman, F. L.


    A series of tests involving the combustion of solid and liquid propellants used to fuel the Titan 2, Titan 4, and Delta 2 launch vehicles was performed. The purpose of these tests was to evaluate the nature and amounts of combustion gases from reacting these propellants in various proportions, and to apply the derived data to predicting toxic chemical emissions arising from a launch vehicle explosion. Propellants tested in this study included Aerozine-50 and nitrogen tetroxide (liquid propellants used in the Titan 2 and Titan 4 launch vehicles), PBAN solid propellant (used on the Titan 4 solid rocket motor), RP-1 and liquid oxygen (liquid propellants used to fuel the Delta 2 launch vehicle), and the Castor IVA solid rocket propellant used on the Delta 2 first stage engine. Tests were conducted in a 150-liter stainless steel combustion chamber in air at nominal pressure (0.8 atmospheres at Denver barometric conditions). Measurements of the chamber gas temperature and internal pressure were taken and gas samples were withdrawn and analyzed for expected combustion gases, unreacted propellants, organic vapors, and oxygen reacted from the air. A stainless steel witness plate was used to collect condensates which formed during the course of the propellant combustion tests. Results of this study suggest significantly different chemical fates for some of the rocket propellants than those predicted by chemical theory only. A description of the test parameters, results, and application to source predictions is presented.

  19. Rocket propulsion elements

    Sutton, George P


    The definitive text on rocket propulsion-now revised to reflect advancements in the field For sixty years, Sutton's Rocket Propulsion Elements has been regarded as the single most authoritative sourcebook on rocket propulsion technology. As with the previous edition, coauthored with Oscar Biblarz, the Eighth Edition of Rocket Propulsion Elements offers a thorough introduction to basic principles of rocket propulsion for guided missiles, space flight, or satellite flight. It describes the physical mechanisms and designs for various types of rockets' and provides an unders

  20. Development of Kabila rocket: A radioisotope heated thermionic plasma rocket engine

    Kalomba Mboyi


    Full Text Available A new type of plasma rocket engine, the Kabila rocket, using a radioisotope heated thermionic heating chamber instead of a conventional combustion chamber or catalyst bed is introduced and it achieves specific impulses similar to the ones of conventional solid and bipropellant rockets. Curium-244 is chosen as a radioisotope heat source and a thermal reductive layer is also used to obtain precise thermionic emissions. The self-sufficiency principle is applied by simultaneously heating up the emitting material with the radioisotope decay heat and by powering the different valves of the plasma rocket engine with the same radioisotope decay heat using a radioisotope thermoelectric generator. This rocket engine is then benchmarked against a 1 N hydrazine thruster configuration operated on one of the Pleiades-HR-1 constellation spacecraft. A maximal specific impulse and power saving of respectively 529 s and 32% are achieved with helium as propellant. Its advantages are its power saving capability, high specific impulses and simultaneous ease of storage and restart. It can however be extremely voluminous and potentially hazardous. The Kabila rocket is found to bring great benefits to the existing spacecraft and further research should optimize its geometric characteristics and investigate the physical principals of its operation.

  1. New Propellant Formulation Development


    initiators, JA2 19-Perf Hex propellants and Hagedorn-Plastic manufactured Nitrocellulose with Alkohol - German manufacturer. STAR-ATO goal is to develop... propellants in the U.S. Army’s small, medium and large caliber munitions are all nitrocellulose -based. As the Army drives continuous improvement in both...understanding the influence of nitrocellulose properties on propellant performance. Projectiles are getting heavier, ammunition is being exposed to

  2. Liquid fuel injection elements for rocket engines

    Cox, George B., Jr. (Inventor)


    Thrust chambers for liquid propellant rocket engines include three principal components. One of these components is an injector which contains a plurality of injection elements to meter the flow of propellants at a predetermined rate, and fuel to oxidizer mixture ratio, to introduce the mixture into the combustion chamber, and to cause them to be atomized within the combustion chamber so that even combustion takes place. Evolving from these injectors are tube injectors. These tube injectors have injection elements for injecting the oxidizer into the combustion chamber. The oxidizer and fuel must be metered at predetermined rates and mixture ratios in order to mix them within the combustion chamber so that combustion takes place smoothly and completely. Hence tube injectors are subject to improvement. An injection element for a liquid propellant rocket engine of the bipropellant type is provided which includes tangential fuel metering orifices, and a plurality of oxidizer tube injection elements whose injection tubes are also provided with tangential oxidizer entry slots and internal reed valves.

  3. Characterization of nal powders for rocket propulsion

    Merotto, L.; Galfetti, L.; Colombo, G.; DeLuca, L. T.


    Nanosized metal powders are known to significantly improve both solid and hybrid rocket performance, but have some drawbacks in terms of cost, safety, and possible influence on propellant mechanical properties. Performance enhancement through nanosized metal or metal hydride addition to solid fuels is currently under investigation also for hybrid propulsion. Therefore, a preburning characterization of the powders used in solid propellant or fuel manufacturing is useful to assess their effects on the ballistic properties and engine performance. An investigation concerning the comparative characterization of several aluminum powders having different particle size, age, and coating is presented. Surface area, morphology, chemical species concentration and characteristics, surface passivation layers, surface and subsurface chemical composition, ignition temperature and ignition delay are investigated. The aim of this characterization is to experimentally assess the effect of the nAl powder properties on ballistic characteristics of solid fuels and solidrocket composite-propellant performance, showing an increase in terms of Is caused by the decrease of two-phase losses in solid and a possible significant rf increase in hybrid rockets.

  4. 78 FR 41283 - Airworthiness Directives; Dowty Propellers Propellers


    ... Engineer, Boston Aircraft Certification Office, FAA, Engine and Propeller Directorate, 12 New England... Engineer, Boston Aircraft Certification Office, FAA, Engine and Propeller Directorate, 12 New England... Ganley, Acting Assistant Manager, Engine & Propeller Directorate, Aircraft Certification Service....

  5. Studies on an aerial propellant transfer space plane (APTSP)

    Jayan, N.; Biju Kumar, K.S.; Gupta, A.K.; Kashyap, A.K.; Venkatraman, K.; Mathews, J.; Makunda, H.S. [Indian Institute of Science, Bangalore (India). Dept. of Aerospace Engineering


    This paper presents a study of a fully reusable earth-to-orbit launch vehicle concept with horizontal take-off and landing, employing a turbojet engine for low speed, and a rocket for high-speed acceleration and space operations. This concept uses existing technology to the maximum possible extent, thereby reducing development time, cost and effort. It uses the experience in aerial filling of military aircraft for propellant filling at an altitude of 13 km at a flight speed of M=0.85. Aerial filling of propellant reduces the take-off weight significantly thereby minimizing the structural weight of the vehicle. The vehicle takes off horizontally and uses turbojet engines till the end of the propellant filling operation. The rocket engines provide thrust for the next phase till the injection of a satellite at LEO. A sensitivity analysis of the mission with respect to rocket engine specific impulse and overall vehicle structural factor is also presented in this paper. A conceptual design of space plane with a payload capability of 10 ton to LEO is carried out. The study shows that the realization of an aerial propellant transfer space plane is possible with limited development of new technology thus reducing the demands on the finances required for achieving the objectives. (author)

  6. Azido-based propellants

    Sayles, D.C.


    This patent describes an azido-based solid propellant composition having an improved burning rate comprising: a high energy plasticizer of tris-1,2,3(bis(1,2-difluoroamino)ethoxy)propane in an amount from about 24 to about 30 weight percent of the propellant composition; a curative and crosslinking agent of 4,5-epoxycyclohexylmethyl 4'5'-epoxycyclohexylcarboxylate in an amount from about 0.75 to about 1.5 weight percent of the propellant composition; a carboranyl burning rate catalyst of carboranyl-methyl propionate in an amount from about 2 to about 6 weight percent of the propellant composition; graphite linters of about 100 micrometers lengths in an amount from about 1 to about 3 weight percent of the propellant composition; aluminum powder in an amount from about 10 to about 12 weight percent of the propellant composition; aluminum flake in an amount from about 0.5 to about 2 weight percent of the propellant composition; ammonium perchlorate of about 0.9 micrometer diameter in an amount from about 46 to about 52 weight percent of the composition; a processing aid of lecithin in an amount from about 0.1 to about 0.2 weight percent of the propellant composition; and a binder of 2-azidoethyl acrylateacrylic acid copolymer in an amount from about 3 to about 8 weight percent of the propellant composition.

  7. Computational simulation of liquid rocket injector anomalies

    Przekwas, A. J.; Singhal, A. K.; Tam, L. T.; Davidian, K.


    A computer model has been developed to analyze the three-dimensional two-phase reactive flows in liquid fueled rocket combustors. The model is designed to study the influence of liquid propellant injection nonuniformities on the flow pattern, combustion and heat transfer within the combustor. The Eulerian-Lagrangian approach for simulating polidisperse spray flow, evaporation and combustion has been used. Full coupling between the phases is accounted for. A nonorthogonal, body fitted coordinate system along with a conservative control volume formulation is employed. The physical models built into the model include a kappa-epsilon turbulence model, a two-step chemical reaction, and the six-flux radiation model. Semiempirical models are used to describe all interphase coupling terms as well as chemical reaction rates. The purpose of this study was to demonstrate an analytical capability to predict the effects of reactant injection nonuniformities (injection anomalies) on combustion and heat transfer within the rocket combustion chamber. The results show promising application of the model to comprehensive modeling of liquid propellant rocket engines.

  8. Solar Thermal Rocket Propulsion

    Sercel, J. C.


    Paper analyzes potential of solar thermal rockets as means of propulsion for planetary spacecraft. Solar thermal rocket uses concentrated Sunlight to heat working fluid expelled through nozzle to produce thrust.

  9. Internal Flow Simulation of High-Performance Solid Rockets using a k-ωTurbulence Model



    @@ For technological reasons many high-performance solid rocket motors are made from segmented propellant grains with non-uniform port geometry. In this paper parametric studies have been carried out to examine the geometric dependence of transient flow features in solid rockets with non-uniform ports. Numerical computations have been carried out in an inert simulator of solid propellant rocket motor with the aid of a standard k-ω turbulence model. It was seen that the damping of the temperature fluctuation is faster in solid rocket with convergent port than with divergent port geometry. We inferred that the damping of the flow fluctuations using the port geometry is a meaningful objective for the suppression and control of the instability and/or pressure/thrust oscillations during the starting transient of solid rockets.

  10. History of solid propellants in the 20. century; Histoire des propergols solides au 20. siecle

    Jung, Ph.; Davenas, A.; McDonald, A.J.; Bret, P.; Moreau, J.P.; Boisson, J.; Kuentzmannn, P.; Maire, G.; Pontvianne, G.; Tranchant, J.; Evans, G.; Reydellet, D.; Vallet, G.; Eymard, M.; Pascal, Ph.; Kuentzmann, P.; Bonnevie, E.; Guery, J.F.; Lengelle, G.; Lhuillier, J.N.; Rat, R.; Keromnes, A.; Mathieu, D.; Simonetti, Ph.; Betin, P.; Thevenin, M.; Serra, J.J.; Delbac, P.; Lepeuple, G.; Miermont, H.; Guillot, J.; Vidal, M.; Citon, C.; Tauzia, J.M.; Chounet, G.; Cardin, J.; Longevialle, Y.; Uhrig, G.


    This colloquium has been jointly organized by the research center of history of sciences and techniques (CRHST) and the association of the friends of the gunpowder and pyrotechnical patrimony (A3P). It gathers historians of sciences and techniques and specialists of solid propellants and their applications who make a review of the approaches that have led to todays propellants efficiency and mastery. This books contains 2 introductive talks, 24 articles, a round table and some concluding remarks. The articles deal with: 1 - from the black powder rockets to the space shuttle: France, pioneer of solid propulsion, from Vaillant to Damblanc (1821-1938); the development of solid propellants in the 20. century; lessons learnt from the Challenger accident; 2 - the institutions: the laboratory of ballistics of Sevran-Livry (1945-1969); an historical overview of ONERA's researches on solid propellants; the cast propellants at the Direction of Explosives (1945-1955); 3 - the propellants: the manufacturing secrets of the extruded double base propellants; the development of cast double base propellants; the invention of composite propellants; 4 - space applications: the Diamant adventure; the solid propellant engines of Ariane 5, an endless story; P80, a new generation of solid propellant engines for space applications; 6 - physics and models: from ap{sup n} to 3-D simulations: the combustion of solid propellants in the 20. century; the mechanical behaviour of solid propellant loads (1960-70 years); composite propellants and static electricity (SE) or the occurrence of SE in the manufacturing and implementation of composite propellants; a priori calculation of the performances and synthesis of new energy materials for propellants; 6 - defense applications: French solid propellant rockets and missiles up to the 1960's; from PHI 1500 to PHI 1930 or the fabulous history of metallic and roving propulsion systems; the G2P, the propulsion system of the M4, the exploratory

  11. Rockets two classic papers

    Goddard, Robert


    Rockets, in the primitive form of fireworks, have existed since the Chinese invented them around the thirteenth century. But it was the work of American Robert Hutchings Goddard (1882-1945) and his development of liquid-fueled rockets that first produced a controlled rocket flight. Fascinated by rocketry since boyhood, Goddard designed, built, and launched the world's first liquid-fueled rocket in 1926. Ridiculed by the press for suggesting that rockets could be flown to the moon, he continued his experiments, supported partly by the Smithsonian Institution and defended by Charles Lindbergh. T

  12. Rocket Engine Innovations Advance Clean Energy


    During launch countdown, at approximately T-7 seconds, the Space Shuttle Main Engines (SSMEs) roar to life. When the controllers indicate normal operation, the solid rocket boosters ignite and the shuttle blasts off. Initially, the SSMEs throttle down to reduce stress during the period of maximum dynamic pressure, but soon after, they throttle up to propel the orbiter to 17,500 miles per hour. In just under 9 minutes, the three SSMEs burn over 1.6 million pounds of propellant, and temperatures inside the main combustion chamber reach 6,000 F. To cool the engines, liquid hydrogen circulates through miles of tubing at -423 F. From 1981to 2011, the Space Shuttle fleet carried crew and cargo into orbit to perform a myriad of unprecedented tasks. After 30 years and 135 missions, the feat of engineering known as the SSME boasted a 100-percent flight success rate.

  13. Metallic hydrogen: The most powerful rocket fuel yet to exist

    Silvera, Isaac F [Lyman Laboratory of Physics, Harvard University, Cambridge MA 02138 (United States); Cole, John W, E-mail: silvera@physics.harvard.ed [NASA MSFC, Huntsville, AL 35801 (United States)


    Wigner and Huntington first predicted that pressures of order 25 GPa were required for the transition of solid molecular hydrogen to the atomic metallic phase. Later it was predicted that metallic hydrogen might be a metastable material so that it remains metallic when pressure is released. Experimental pressures achieved on hydrogen have been more than an order of magnitude higher than the predicted transition pressure and yet it remains an insulator. We discuss the applications of metastable metallic hydrogen to rocketry. Metastable metallic hydrogen would be a very light-weight, low volume, powerful rocket propellant. One of the characteristics of a propellant is its specific impulse, I{sub sp}. Liquid (molecular) hydrogen-oxygen used in modern rockets has an Isp of {approx}460s; metallic hydrogen has a theoretical I{sub sp} of 1700s. Detailed analysis shows that such a fuel would allow single-stage rockets to enter into orbit or carry economical payloads to the moon. If pure metallic hydrogen is used as a propellant, the reaction chamber temperature is calculated to be greater than 6000 K, too high for currently known rocket engine materials. By diluting metallic hydrogen with liquid hydrogen or water, the reaction temperature can be reduced, yet there is still a significant performance improvement for the diluted mixture.

  14. Contact performance analysis on the joint of thrust vector control system in a certain solid propellant rocket motor%某固体火箭发动机推力向量控制系统接头接触性能分析

    刘文芝; 戴美魁; 韦广梅; 赵永忠


    以某固体火箭发动机推力向量控制系统摆动接头的单珠承载试验模型为计算模型,采用摩擦接触问题的Lagrange乘子法与弹塑性耦合的有限元理论,计算分析了摆动接头阳球试件在不同强化层厚度下的接触应力、变形及破坏机理;为降低系统摆动力矩,同时考虑大尺寸球面的表面强化和加工工艺,提出了满足接触性能的阳球表面强化层为1~1.4mm的厚度要求.通过与单珠承载试验对比分析,检验了有限元建模及算法的合理性;确定了小变形条件下摆动接头弹塑性摩擦接触应力限额为5 384.9 MPa.%In this paper, a single ball load-bearing experiment model of the swing joint of thrust vector control system in a solid propellant rocket motor was used as computation model. By using the Lagrangian multiplier method for friction contact question, and elasto-plastic couplings finite element method, the contact stress, deformation and destruct principle under different thickness of strengthening layer of the test swing joint convex sphere were computed and analyzed. To reduce system swing moment, and consider manufacturing techniques of large size intensify sphere, 1-1.4 mm thickness of strengthening layer satisfied for contact performance was presented. Rationality of FEM modeling and algorithm was validated by single ball load-bearing experiment. The elasto-plastic friction contact stress of the swing joint was limited to 5384.9 Mpa in small deformation condition.

  15. Nuclear rocket using indigenous Martian fuel NIMF

    Zubrin, Robert


    In the 1960's, Nuclear Thermal Rocket (NTR) engines were developed and ground tested capable of yielding isp of up to 900 s at thrusts up to 250 klb. Numerous trade studies have shown that such traditional hydrogen fueled NTR engines can reduce the inertial mass low earth orbit (IMLEO) of lunar missions by 35 percent and Mars missions by 50 to 65 percent. The same personnel and facilities used to revive the hydrogen NTR can also be used to develop NTR engines capable of using indigenous Martian volatiles as propellant. By putting this capacity of the NTR to work in a Mars descent/acent vehicle, the Nuclear rocket using Indigenous Martian Fuel (NIMF) can greatly reduce the IMLEO of a manned Mars mission, while giving the mission unlimited planetwide mobility.

  16. An Analysis of Rocket Propulsion Testing Costs

    Ramirez, Carmen; Rahman, Shamim


    The primary mission at NASA Stennis Space Center (SSC) is rocket propulsion testing. Such testing is commonly characterized as one of two types: production testing for certification and acceptance of engine hardware, and developmental testing for prototype evaluation or research and development (R&D) purposes. For programmatic reasons there is a continuing need to assess and evaluate the test costs for the various types of test campaigns that involve liquid rocket propellant test articles. Presently, in fact, there is a critical need to provide guidance on what represents a best value for testing and provide some key economic insights for decision-makers within NASA and the test customers outside the Agency. Hence, selected rocket propulsion test databases and references have been evaluated and analyzed with the intent to discover correlations of technical information and test costs that could help produce more reliable and accurate cost projections in the future. The process of searching, collecting, and validating propulsion test cost information presented some unique obstacles which then led to a set of recommendations for improvement in order to facilitate future cost information gathering and analysis. In summary, this historical account and evaluation of rocket propulsion test cost information will enhance understanding of the various kinds of project cost information; identify certain trends of interest to the aerospace testing community.

  17. Scaled Rocket Testing in Hypersonic Flow

    Dufrene, Aaron; MacLean, Matthew; Carr, Zakary; Parker, Ron; Holden, Michael; Mehta, Manish


    NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was strongly based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Detailed base heating results are outside of the scope of the current work, rather test methodology and techniques are presented along with broader applicability toward scaled rocket testing in supersonic and hypersonic flow.

  18. Solid Hydrogen Formed for Atomic Propellants

    Palaszewski, Bryan A.


    Several experiments on the formation of solid hydrogen particles in liquid helium were recently conducted at the NASA Glenn Research Center at Lewis Field. The solid hydrogen experiments are the first step toward seeing these particles and determining their shape and size. The particles will ultimately store atoms of boron, carbon, or hydrogen, forming an atomic propellant. Atomic propellants will allow rocket vehicles to carry payloads many times heavier than possible with existing rockets or allow them to be much smaller and lighter. Solid hydrogen particles are preferred for storing atoms. Hydrogen is generally an excellent fuel with a low molecular weight. Very low temperature hydrogen particles (T < 4 K) can prevent the atoms from recombining, making it possible for their lifetime to be controlled. Also, particles that are less than 1 mm in diameter are preferred because they can flow easily into a pipe when suspended in liquid helium. The particles and atoms must remain at this low temperature until the fuel is introduced into the engine combustion (or recombination) chamber. Experiments were, therefore, planned to look at the particles and observe their formation and any changes while in liquid helium.

  19. The use of electrical discharge for ignition and control of combustion of solid propellants

    Tachibana, Takeshi; Kobayashi, Tsuruo; Matsuda, Takashi; Kimura, Itsuro


    As the first step of the study of the combustion control of solid propellants by electrical discharges, the effects of an arc discharge, which flows along the burning surface, on the burning rate and on the increase of enthalpy of the combustion product were investigated. For specially devised composite propellants, which are composed of Al and Teflon powders, it was shown that the combination can be controlled by an arc discharge; the combustion continues when the arc discharge is applied and is interrupted when the arc discharge breaks. In the present investigation, it was also shown that an arc discharge coupled with a high-frequency electrical discharge has potential as an effective ignition method for solid propellants. For the application of this type of combustion control to an ignitor for a solid propellant rocket motor or to a control rocket motor, this method lacks flexibility in the configuration scale and needs relatively high electric power at the present stage.

  20. Application of Closed Vessel Technique for the Evaluation of Burning Rates of Propellants at Low Pressures

    D. Vittal


    Full Text Available Closed vessel technique has been well established for the evaluation of burning characteristics of gun, mortar and small arms propellants at high pressures of about 750 kg/cm/sup 2/ - 3000 kg/cm/sup 2/ propellants in the pressure range up to about 200 kg/cm/sup 2/ (19.6 MPa. One of the modern trends in armaments technology is development of short range, high efficiency rockets and rocket assisted projectiles where the chamber pressure are in the range of 100 kg/cm/sup 2/ - 800 kg/cm/sup 2/ (9.8 MPa-78.5 MPa. An extension of the closed vessel technique is now presented for the measurement of rates of burning of propellants in this pressure range and a few experimental results on some conventional propellants are given.

  1. Rocket injector head

    Green, C. W., Jr. (Inventor)


    A high number of liquid oxygen and gaseous hydrogen orifices per unit area are provided in an injector head designed to give intimate mixing and more thorough combustion. The injector head comprises a main body portion, a cooperating plate member as a flow chamber for one propellant, a cooperating manifold portion for the second propellant, and an annular end plate for enclosing an annular propellant groove formed around the outer edge of the body. All the openings for one propellant are located at the same angle with respect to a radial plane to permit a short combustion chamber.

  2. Mechanical Slosh Models for Rocket-Propelled Spacecraft

    Jang, Jiann-Woei; Alaniz, Abram; Yang, Lee; Powers. Joseph; Hall, Charles


    Several analytical mechanical slosh models for a cylindrical tank with flat bottom are reviewed. Even though spacecrafts use cylinder shaped tanks, most of those tanks usually have elliptical domes. To extend the application of the analytical models for a cylindrical tank with elliptical domes, the modified slosh parameter models are proposed in this report by mapping an elliptical dome cylindrical tank to a flat top/bottom cylindrical tank while maintaining the equivalent liquid volume. For the low Bond number case, the low-g slosh models were also studied. Those low-g models can be used for Bond number > 10. The current low-g slosh models were also modified to extend their applications for the case that liquid height is smaller than the tank radius. All modified slosh models are implemented in MATLAB m-functions and are collected in the developed MST (Mechanical Slosh Toolbox).

  3. Modeling of Nonlinear Combustion Instability in Solid Propellant Rocket Motors


    results of two solutions that tesi the ability of the LW+H+ACM to trect such problems are discussed. In both cases, the initial Pressure distur$.;nce...piezoelectric transducers. Thermal shielding of the pressure transducers was accomplished using vulcanized rubber (RTV). Based on shock i,,be calibration tests...potted with vulcanized rubber (RTV) which is utilized to provide thermal shielding 24 45. 16.0 � AFT CLOSURE WI0.45 250. for the pressure transducers

  4. Metallic Hydrogen - Potentially a High Energy Rocket Propellant

    Cole, John; Silvera, Ike


    Pure metallic hydrogen is predicted to have a specific impulse (Isp) of 1700 seconds, but the reaction temperature is too high for current engine materials. Diluting metallic hydrogen with liquid hydrogen can reduce the reaction temperature to levels compatible with current material limits and still provide an Isp greater than 900 s. Metallic hydrogen has not yet been produced on earth, but experimental techniques exist that may change this situation. This paper will provide a brief description of metallic hydrogen and the status of experiments that may soon produce detectable quantities of this material in the lab. Also provided are some characteristics for diluted metallic hydrogen engines and launch vehicles.

  5. A Study of Flame Physics and Solid Propellant Rocket Physics


    including suggestions for reducing the burden, to the Department of Defense, Executive Services and Communications Directorate 10704-0188...was re- placed with a 3-step kinetics model which gives improved agreement with experiental data for burning rates of a variety of heterogeneous

  6. Solid Propellant Flame Spectroscopy


    400 jm to reach the maximum flame temperature, a distance that can be reduced by replacing the HTPB binder with a polyester or CMDB binder. The...the dark zone for propellants similar to HIX2 is 2-2.5 mm at 1.8 MPa (18 atm, 265 psia) (Ref. 22,187). In contrast, the dark zone for HMX CMDB ...propellants eliminates the dark zone is not surprising, since TMETN is a nitrate ester as was the double-base matrix of Kubota’s HMX CMDB propellant. A

  7. Propeller tone bursts

    Succi, G. P.; Munro, D. H.; Ingard, K. U.


    Intense high frequency (25-38 kHz) tone bursts have been observed in acoustic tests of a scale model of a general aviation propeller. The amplitude of the tone burst is approximately equal to the amplitude of the propeller noise signature. The conditions necessary for the production of these tone bursts are described. The experiments indicate that the origin of these bursts is a periodic flow oscillation on the suction surface of the propeller blade tips which may be due to the interaction between an oscillating shock wave and a laminar boundary layer.

  8. High temperature propellant development

    Anderson, F. A.


    It is reported that the neccessary technology has been developed and demonstrated for the manufacture of heat-sterilizable solid propellants which meet specific ballistic goals. It is shown that: (1) phosphate doping of ammonium perchlorate significantly enhances the thermal stability of the substance; (2) grinding the ammonium perchlorate to reduce particle size further increases thermal stability; and (3) unsaturated polymers such as the polybutadienes can be successfully used in a heat-sterilizable propellant system. Among the topics considered by the study are oxidizers, dopants, binders, and the thermal cycling of 70 lb and 600 lb propellant grains.

  9. Propeller TAP flap

    Thomsen, Jørn Bo; Bille, Camilla; Wamberg, Peter


    The aim of this study was to examine if a propeller thoracodorsal artery perforator (TAP) flap can be used for breast reconstruction. Fifteen women were reconstructed using a propeller TAP flap, an implant, and an ADM. Preoperative colour Doppler ultrasonography was used for patient selection...... major complications needing additional surgery. One flap was lost due to a vascular problem. Breast reconstruction can be performed by a propeller TAP flap without cutting the descending branch of the thoracodorsal vessels. However, the authors would recommend that a small cuff of muscle is left around...

  10. Advanced Computer Science on Internal Ballistics of Solid Rocket Motors

    Shimada, Toru; Kato, Kazushige; Sekino, Nobuhiro; Tsuboi, Nobuyuki; Seike, Yoshio; Fukunaga, Mihoko; Daimon, Yu; Hasegawa, Hiroshi; Asakawa, Hiroya

    In this paper, described is the development of a numerical simulation system, what we call “Advanced Computer Science on SRM Internal Ballistics (ACSSIB)”, for the purpose of improvement of performance and reliability of solid rocket motors (SRM). The ACSSIB system is consisting of a casting simulation code of solid propellant slurry, correlation database of local burning-rate of cured propellant in terms of local slurry flow characteristics, and a numerical code for the internal ballistics of SRM, as well as relevant hardware. This paper describes mainly the objectives, the contents of this R&D, and the output of the fiscal year of 2008.

  11. Performance optimization of marine propellers

    Chang-Sup Lee


    In this paper, a design method for increasing performance of the marine propellers including the WCT propeller is suggested. It is described to maximize the performance of the propeller by adjusting expanded areas of the propeller blade. Results show that efficiency can be increased up to over 2% through the suggested design method.

  12. Current status of free radicals and electronically excited metastable species as high energy propellants

    Rosen, G.


    A survey is presented of free radicals and electronically excited metastable species as high energy propellants for rocket engines. Nascent or atomic forms of diatomic gases are considered free radicals as well as the highly reactive diatomic triatomic molecules that posess unpaired electrons. Manufacturing and storage problems are described, and a review of current experimental work related to the manufacture of atomic hydrogen propellants is presented.

  13. Heat transfer analysis of fuel assemblies in a heterogeneous gas core nuclear rocket

    Watanabe, Yoichi; Appelbaum, Jacob; Diaz, Nils; Maya, Isaac


    Heat transfer problems of a heterogeneous gaseous core nuclear rocket were studied. The reactor core consists of 1.5-m long hexagonal fuel assemblies filled with pressurized uranium tetrafluoride (UF4) gas. The fuel gas temperature ranges from 3500 to 7000 K at a nominal operating condition of 40 atm. Each fuel assembly has seven coolant tubes, through which hydrogen propellant flows. The propellant temperature is not constrained by the fuel temperature but by the maximum temperature of the graphite coolant tube. For a core achieving a fission power density of 1000 MW/cu m, the propellant core exit temperature can be as high as 3200 K. The physical size of a 1250 MW gaseous core nuclear rocket is comparable with that of a NERVA-type solid core nuclear rocket. The engine can deliver a specific impulse of 1020 seconds and a thrust of 330 kN.

  14. Analyses of Cryogenic Propellant Tank Pressurization based upon Ground Experiments

    Ludwig, Carina; Dreyer, Michael


    The pressurization system of cryogenic propellant rockets requires on-board pressurant gas. The objective of this study was to analyze the influence of the pressurant gas temperature on the required pressurant gas mass in terms of lowering the launcher mass. First, ground experiments were performed in order to investigate the pressurization process with regard to the influence of the pressurant gas inlet temperature. Second, a system study for the cryogenic upper stage of a sma...

  15. Development of a solid propellant viscoelastic dynamic model

    Hufferd, W. L.; Fitzgerald, J. E.


    The results of a one year study to develop a dynamic response model for the Space Shuttle Solid Rocket Motor (SRM) propellant are presented. An extensive literature survey was conducted, from which it was concluded that the only significant variables affecting the dynamic response of the SRM propellant are temperature and frequency. Based on this study, and experimental data on propellants related to the SRM propellant, a dynamic constitutive model was developed in the form of a simple power law with temperature incorporated in the form of a modified power law. A computer program was generated which performs a least-squares curve-fit of laboratory data to determine the model parameters and it calculates dynamic moduli at any desired temperature and frequency. Additional studies investigated dynamic scaling laws and the extent of coupling between the SRM propellant and motor cases. It was found, in agreement with other investigations, that the propellant provides all of the mass and damping characteristics whereas the case provides all of the stiffness.

  16. Grooved Fuel Rings for Nuclear Thermal Rocket Engines

    Emrich, William


    An alternative design concept for nuclear thermal rocket engines for interplanetary spacecraft calls for the use of grooved-ring fuel elements. Beyond spacecraft rocket engines, this concept also has potential for the design of terrestrial and spacecraft nuclear electric-power plants. The grooved ring fuel design attempts to retain the best features of the particle bed fuel element while eliminating most of its design deficiencies. In the grooved ring design, the hydrogen propellant enters the fuel element in a manner similar to that of the Particle Bed Reactor (PBR) fuel element.

  17. Magnetic resonance imaging (MRI) of solid rocket components

    Wallner, A.S. [Missouri Western State College, St. Joseph, MO (United States); Nissan, R.A.; Merwin, L.H. [Naval Air Warfare Center, China Lake, CA (United States)] [and others


    The evaluation of solid rocket components has become an area of great interest. Studying these materials with MRI offers a great advantage to observe knit lines, regions of inhomogeneity, voids, defects, plasticizer rich/poor areas and solids distribution because of the nondestructive nature of the technique. Aspects of sample preparation, spectroscopic relaxation studies, and MRI as a method of studying these systems will be discussed. Initial images show the ability to image propellant, liner, and explosive materials with an in-plane resolution of 70 {mu}m/pixel. These initial images show that MRI can be developed as a viable nondestructive evaluation method of solid rocket components.

  18. The flight of uncontrolled rockets

    Gantmakher, F R; Dryden, H L


    International Series of Monographs on Aeronautics and Astronautics, Division VII, Volume 5: The Flight of Uncontrolled Rockets focuses on external ballistics of uncontrolled rockets. The book first discusses the equations of motion of rockets. The rocket as a system of changing composition; application of solidification principle to rockets; rotational motion of rockets; and equations of motion of the center of mass of rockets are described. The text looks at the calculation of trajectory of rockets and the fundamentals of rocket dispersion. The selection further focuses on the dispersion of f

  19. Water rocket - Electrolysis propulsion and fuel cell power

    Carter, P H; Dittman, M D; Kare, J T; Militsky, F; Myers, B; Weisberg, A H


    Water Rocket is the collective name for an integrated set of technologies that offer new options for spacecraft propulsion, power, energy storage, and structure. Low pressure water stored on the spacecraft is electrolyzed to generate, separate, and pressurize gaseous hydrogen and oxygen. These gases, stored in lightweight pressure tanks, can be burned to generate thrust or recombined to produce electric power. As a rocket propulsion system, Water Rocket provides the highest feasible chemical specific impulse (-400 seconds). Even higher specific impulse propulsion can be achieved by combining Water Rocket with other advanced propulsion technologies, such as arcjet or electric thrusters. With innovative pressure tank technology, Water Rocket's specific energy [Wh/kg] can exceed that of the best foreseeable batteries by an order of magnitude, and the tanks can often serve as vehicle structural elements. For pulsed power applications, Water Rocket propellants can be used to drive very high power density generators, such as MHD devices or detonation-driven pulse generators. A space vehicle using Water Rocket propulsion can be totally inert and non-hazardous during assembly and launch. These features are particularly important for the timely development and flight qualification of new classes of spacecraft, such as microsats, nanosats, and refuelable spacecraft.

  20. Hydrodynamics of Ship Propellers

    Breslin, John P.; Andersen, Poul

    This book deals with flows over propellers operating behind ships, and the hydrodynamic forces and moments which the propeller generates on the shaft and on the ship hull.The first part of the text is devoted to fundamentals of the flow about hydrofoil sections (with and without cavitation......) and about wings. It then treats propellers in uniform flow, first via advanced actuator disc modelling, and then using lifting-line theory. Pragmatic guidance is given for design and evaluation of performance, including the use of computer modelling.The second part covers the development of unsteady forces...... arising from operation in non-uniform hull wakes. First, by a number of simplifications, various aspects of the problem are dealt with separately until the full problem of a non-cavitating, wide-bladed propeller in a wake is treated by a new and completely developed theory. Next, the complicated problem...

  1. Investigation of Advanced Propellants to Enable Single Stage to Orbit Launch Vehicles


    Jeigh Shelley Department of Mechanical Engineering Raymond Moszée Air Force Research Laboratory ABSTRACT INVESTIGATION OF ADVANCED...Engineering Raymond Moszée Air Force Research Laboratory For the University Graduate Committee: Dean, Division of Graduate Studies...September 2005. 33. Bayer , M.J., “Comparative Assessment of Rocket-Propelled Single-Stage-to- Orbit Concepts,” Journal of Spacecraft and Rockets, Vol. 40

  2. Experimental Evaluation of a Subscale Gaseous Hydrogen/gaseous Oxygen Coaxial Rocket Injector

    Smith, Timothy D.; Klem, Mark D.; Breisacher, Kevin J.; Farhangi, Shahram; Sutton, Robert


    The next generation reusable launch vehicle may utilize a Full-Flow Stage Combustion (FFSC) rocket engine cycle. One of the key technologies required is the development of an injector that uses gaseous oxygen and gaseous hydrogen as propellants. Gas-gas propellant injection provides an engine with increased stability margin over a range of throttle set points. This paper summarizes an injector design and testing effort that evaluated a coaxial rocket injector for use with gaseous oxygen and gaseous hydrogen propellants. A total of 19 hot-fire tests were conducted up to a chamber pressure of 1030 psia, over a range of 3.3 to 6.7 for injector element mixture ratio. Post-test condition of the hardware was also used to assess injector face cooling. Results show that high combustion performance levels could be achieved with gas-gas propellants and there were no problems with excessive face heating for the conditions tested.

  3. Characterization of aluminum/RP-1 gel propellant properties

    Rapp, Douglas C.; Zurawski, Robert L.


    Research efforts are being conducted by the NASA Lewis Research Center to formulate and characterize the properties of Al/RP-1 and RP-1 gelled propellants for rocket propulsion systems. Twenty four different compositions of gelled fuels were formualted with 5 and 16 micron, atomized aluminum powder in RP-1. The total solids concentration in the propellant varied from 5 to 60 wt percent. Tests were conducted to evaluate the stability and rheological characteristics of the fuels. Physical separation of the solids occurred in fuels with less than 50 wt percent solids concentration. The rheological characteristics of the Al/RP-1 fuels varied with solids concentration. Both thixotropic and rheopectic gel behavior were observed. The unmetallized RP-1 gels, which were formulated by a different technique than the Al/RP-1 gels, were highly viscoelastic. A history of research efforts which were conducted to formulate and characterize the properties of metallized propellants for various applications is also given.

  4. Laboratory test methods for combustion stability properties of solid propellants

    Strand, L. D.; Brown, R. S.


    An overview is presented of experimental methods for determining the combustion-stability properties of solid propellants. The methods are generally based on either the temporal response to an initial disturbance or on external methods for generating the required oscillations. The size distribution of condensed-phase combustion products are characterized by means of the experimental approaches. The 'T-burner' approach is shown to assist in the derivation of pressure-coupled driving contributions and particle damping in solid-propellant rocket motors. Other techniques examined include the rotating-valve apparatus, the impedance tube, the modulated throat-acoustic damping burner, and the magnetic flowmeter. The paper shows that experimental methods do not exist for measuring the interactions between acoustic velocity oscillations and burning propellant.

  5. Composite propellant tank study for very low cost space transportation

    Moser, D. J.; Keith, E. L.


    A study of life-cycle cost is conducted to determine acceptable options for composite propellant tanks at low cost and weight and for use at moderate pressures. The review examines all cost issues relevant to the production, mass, applications, and reliability of the tanks for pressure-fed rockets. Specific attention is given to the manufacturing and life-cycle issues relevant to the use of composite materials in this application since composites are effective materials for liquid propellant tanks. Specific costs and parametric considerations are given for several tank candidates with 62,303-lb capacities. The mass sensitivity of the fourth stage for the concept vehicle is shown to be high, and the use of a 325-psi fourth-stage tank is shown to yield the minimum cost/lb for the stage. Wound S-glass/epoxy composites can be employed as cost-effective replacements for steel in the design of liquid-propellant tanks.

  6. Ballistic anomalies in solid rocket motors due to migration effects

    Pröbster, M.; Schmucker, R. H.

    Double base and composite propellants are generally used for rocket motors, whereby double base propellants basically consist of nitrocellulose plasticized with an explosive plasticizer, mostly nitroglycerine, and in some cases with an additional inert plasticizer and ballistic additives. Composite propellants consist of an oxidizer like ammonium perchlorate and of aluminum, binder and plasticizer and often contain liquid or solid burning rate catalysts. A common feature of both propellants is that they contain smaller or larger amounts of chemically unbonded liquid species which tend to migrate. If these propellants loose part of the plasticizer by migration into the insulation layer, not only will there be a change in mechanical propellant properties but also the bond between propellant and insulation may degrade. However, depending on the severity of these effects, the change in the ballistic properties of the propellant grain caused by plasticizer migration may be of even more importance. In the past, most emphasis was placed on the behaviour of end-burning configurations. However, more recent theoretical and experimental studies revealed that not only for end-burning grain configurations but also for internal burning configurations there is a common effect which is responsible for ballistic anomalies: migration of liquid species from the propellant into the insulation. By using a plasticizer equilibrated insulation in an internal burning configuration the liquid species migration and thus the previously observed ballistic anomalies are avoided. Using this approach for end-burning configurations provides similar positive results. The various factors affecting plasticizer migration are studied and discussed, and several methods to prevent liquid species migration are described as well as methods to obtain plasticizer resistant insulations.

  7. The Guggenheim Aeronautics Laboratory at Caltech and the creation of the modern rocket motor (1936-1946): How the dynamics of rocket theory became reality

    Zibit, Benjamin Seth

    This thesis explores and unfolds the story of discovery in rocketry at The California Institute of Technology---specifically at Caltech's Guggenheim Aeronautics Laboratory---in the 1930s and 1940s. Caltech was home to a small group of engineering students and experimenters who, beginning in the winter of 1935--1936, formed a study and research team destined to change the face of rocket science in the United States. The group, known as the Guggenheim Aeronautics Laboratory (GALCIT, for short) Rocket Research Group, invented a new type of solid-rocket propellant, made distinct and influential discoveries in the theory of rocket combustion and design, founded the Jet Propulsion Laboratory, and incorporated the first American industrial concern devoted entirely to rocket motor production: The Aerojet Corporation. The theoretical work of team members, Frank Malina, Hsueh-shen Tsien, Homer J. Stewart, and Mark Mills, is examined in this thesis in detail. The author scrutinizes Frank Malina's doctoral thesis (both its assumptions and its mathematics), and finds that, although Malina's key assertions, his formulae, hold, his work is shown to make key assumptions about rocket dynamics which only stand the test of validity if certain approximations, rather than exact measurements, are accepted. Malina studied the important connection between motor-nozzle design and thrust; in his Ph.D. thesis, he developed mathematical statements which more precisely defined the design/thrust relation. One of Malina's colleagues on the Rocket Research Team, John Whiteside Parsons, created a new type of solid propellant in the winter of 1941--1942. This propellant, known as a composite propellant (because it simply was a relatively inert amalgam of propellant and oxidizer in non-powder form), became the forerunner of all modern solid propellants, and has become one of the seminal discoveries in the field of Twentieth Century rocketry. The latter chapters of this dissertation discuss the

  8. Computational simulation of liquid fuel rocket injectors

    Landrum, D. Brian


    A major component of any liquid propellant rocket is the propellant injection system. Issues of interest include the degree of liquid vaporization and its impact on the combustion process, the pressure and temperature fields in the combustion chamber, and the cooling of the injector face and chamber walls. The Finite Difference Navier-Stokes (FDNS) code is a primary computational tool used in the MSFC Computational Fluid Dynamics Branch. The branch has dedicated a significant amount of resources to development of this code for prediction of both liquid and solid fuel rocket performance. The FDNS code is currently being upgraded to include the capability to model liquid/gas multi-phase flows for fuel injection simulation. An important aspect of this effort is benchmarking the code capabilities to predict existing experimental injection data. The objective of this MSFC/ASEE Summer Faculty Fellowship term was to evaluate the capabilities of the modified FDNS code to predict flow fields with liquid injection. Comparisons were made between code predictions and existing experimental data. A significant portion of the effort included a search for appropriate validation data. Also, code simulation deficiencies were identified.

  9. Solid Rocket Launch Vehicle Explosion Environments

    Richardson, E. H.; Blackwood, J. M.; Hays, M. J.; Skinner, T.


    Empirical explosion data from full scale solid rocket launch vehicle accidents and tests were collected from all available literature from the 1950s to the present. In general data included peak blast overpressure, blast impulse, fragment size, fragment speed, and fragment dispersion. Most propellants were 1.1 explosives but a few were 1.3. Oftentimes the data from a single accident was disjointed and/or missing key aspects. Despite this fact, once the data as a whole was digitized, categorized, and plotted clear trends appeared. Particular emphasis was placed on tests or accidents that would be applicable to scenarios from which a crew might need to escape. Therefore, such tests where a large quantity of high explosive was used to initiate the solid rocket explosion were differentiated. Also, high speed ground impacts or tests used to simulate such were also culled. It was found that the explosions from all accidents and applicable tests could be described using only the pressurized gas energy stored in the chamber at the time of failure. Additionally, fragmentation trends were produced. Only one accident mentioned the elusive "small" propellant fragments, but upon further analysis it was found that these were most likely produced as secondary fragments when larger primary fragments impacted the ground. Finally, a brief discussion of how this data is used in a new launch vehicle explosion model for improving crew/payload survival is presented.

  10. Standard Molded Composite Rocket Pyrogen Igniter - A progress report

    Lucy, M. H.


    The pyrogen igniter has the function to furnish a controlled, high temperature, high pressure gas to ignite solid propellant surfaces in a rocket motor. Present pyrogens consist of numerous inert components. The Standard Molded Pyrogen Igniter (SMPI) consists of three basic parts, a cap with several integrally molded features, an ignition pellet retainer plate, and a tube with additional integrally molded features. A description is presented of an investigation which indicates that the SMPI concept is a viable approach to the design and manufacture of pyrogen igniters for solid propellant rocket motors. For some applications, combining the structural and thermal properties of molded composites can result in the manufacture of lighter assemblies at considerable cost reduction. It is demonstrated that high strength, thin walled tubes with high length to diameter ratios can be fabricated from reinforced plastic molding compound using the displacement compression process.

  11. Study of Cavitation/Vaporization in Liquid Rocket Thruster Injectors


    Caveny, L. H., and Summerfield, M., Aluminized Solid Propellants Burning in a Rocket Motor Flowfield, AIAA Journal, Vol. 16, No. 7, 1978, pp. 736-739. [2...the swirl chamber, and the pulsator and manifold are made of 304 stainless steel . Figure 1: Nomenclature defined for swirl injector. Table 1...Wayne, NJ, 2009 [17] MATLAB, Matrix Laboratory, Software Package, R2009a, The MathWorks, Natick, MA, 2009. [18] Coleman, H.W. and Steele , W.G

  12. The Ion Rocket


    discharge velocity w and the speci- fic impulse lap respectively cannot be increased. At this limit condition the thermal rocket oecouos "choked up...structural quality is 900 t, 3) In the case of an atomic-driven thermal rocket ’,;lth specific Ipipulse ISjy«8C0 sec and thrust to weight ratio « 1, the

  13. Model Rockets and Microchips.

    Fitzsimmons, Charles P.


    Points out the instructional applications and program possibilities of a unit on model rocketry. Describes the ways that microcomputers can assist in model rocket design and in problem calculations. Provides a descriptive listing of model rocket software for the Apple II microcomputer. (ML)

  14. Advanced Tactical Booster Technologies: Applications for Long-Range Rocket Systems


    System HIMARS [3] which can employ the MGM- 140 Army Tactical Missile System (ATacMS) solid propellant missile [4] to achieve the required range...launcher. 15. SUBJECT TERMS solid rocket; optimisation; artillery 16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF ABSTRACT 18. NUMBER OF PAGES 19a...the volumetrically constrained environment of a land-based launcher. Keywords— solid rocket; optimisation; artillery I. INTRODUCTION The Australian

  15. Feasibility Study and Demonstration of an Aluminum and Ice Solid Propellant

    Timothee L. Pourpoint


    Full Text Available Aluminum-water reactions have been proposed and studied for several decades for underwater propulsion systems and applications requiring hydrogen generation. Aluminum and water have also been proposed as a frozen propellant, and there have been proposals for other refrigerated propellants that could be mixed, frozen in situ, and used as solid propellants. However, little work has been done to determine the feasibility of these concepts. With the recent availability of nanoscale aluminum, a simple binary formulation with water is now feasible. Nanosized aluminum has a lower ignition temperature than micron-sized aluminum particles, partly due to its high surface area, and burning times are much faster than micron aluminum. Frozen nanoscale aluminum and water mixtures are stable, as well as insensitive to electrostatic discharge, impact, and shock. Here we report a study of the feasibility of an nAl-ice propellant in small-scale rocket experiments. The focus here is not to develop an optimized propellant; however improved formulations are possible. Several static motor experiments have been conducted, including using a flight-weight casing. The flight weight casing was used in the first sounding rocket test of an aluminum-ice propellant, establishing a proof of concept for simple propellant mixtures making use of nanoscale particles.

  16. Parallelization of Rocket Engine System Software (Press)

    Cezzar, Ruknet


    The main goal is to assess parallelization requirements for the Rocket Engine Numeric Simulator (RENS) project which, aside from gathering information on liquid-propelled rocket engines and setting forth requirements, involve a large FORTRAN based package at NASA Lewis Research Center and TDK software developed by SUBR/UWF. The ultimate aim is to develop, test, integrate, and suitably deploy a family of software packages on various aspects and facets of rocket engines using liquid-propellants. At present, all project efforts by the funding agency, NASA Lewis Research Center, and the HBCU participants are disseminated over the internet using world wide web home pages. Considering obviously expensive methods of actual field trails, the benefits of software simulators are potentially enormous. When realized, these benefits will be analogous to those provided by numerous CAD/CAM packages and flight-training simulators. According to the overall task assignments, Hampton University's role is to collect all available software, place them in a common format, assess and evaluate, define interfaces, and provide integration. Most importantly, the HU's mission is to see to it that the real-time performance is assured. This involves source code translations, porting, and distribution. The porting will be done in two phases: First, place all software on Cray XMP platform using FORTRAN. After testing and evaluation on the Cray X-MP, the code will be translated to C + + and ported to the parallel nCUBE platform. At present, we are evaluating another option of distributed processing over local area networks using Sun NFS, Ethernet, TCP/IP. Considering the heterogeneous nature of the present software (e.g., first started as an expert system using LISP machines) which now involve FORTRAN code, the effort is expected to be quite challenging.

  17. Propellers in yaw

    Ribner, Herbert S


    It was realized as early as 1909 that a propeller in yaw develops a side force like that of a fin. In 1917, R. G. Harris expressed this force in terms of the torque coefficient for the unyawed propeller. Of several attempts to express the side force directly in terms of the shape of the blades, however, none has been completely satisfactory. An analysis that incorporates induction effects not adequately covered in previous work and that gives good agreement with experiment over a wide range of operating conditions is presented. The present analysis shows that the fin analogy may be extended to the form of the side-force expression and that the effective fin area may be taken as the projected side area of the propeller.

  18. Self-propelled droplets

    Seemann, Ralf; Fleury, Jean-Baptiste; Maass, Corinna C.


    Self-propelled droplets are a special kind of self-propelled matter that are easily fabricated by standard microfluidic tools and locomote for a certain time without external sources of energy. The typical driving mechanism is a Marangoni flow due to gradients in the interfacial energy on the droplet interface. In this article we review the hydrodynamic prerequisites for self-sustained locomotion and present two examples to realize those conditions for emulsion droplets, i.e. droplets stabilized by a surfactant layer in a surrounding immiscible liquid. One possibility to achieve self-propelled motion relies on chemical reactions affecting the surface active properties of the surfactant molecules. The other relies on micellar solubilization of the droplet phase into the surrounding liquid phase. Remarkable cruising ranges can be achieved in both cases and the relative insensitivity to their own `exhausts' allows to additionally study collective phenomena.

  19. NASA's Advanced solid rocket motor

    Mitchell, Royce E.

    The Advanced Solid Rocket Motor (ASRM) will not only bring increased safety, reliability and performance for the Space Shuttle Booster, it will enhance overall Shuttle safety by effectively eliminating 174 failure points in the Space Shuttle Main Engine throttling system and by reducing the exposure time to aborts due to main engine loss or shutdown. In some missions, the vulnerability time to Return-to-Launch Site aborts is halved. The ASRM uses case joints which will close or remain static under the effects of motor ignition and pressurization. The case itself is constructed of the weldable steel alloy HP 9-4-0.30, having very high strength and with superior fracture toughness and stress corrosion resistance. The internal insulation is strip-wound and is free of asbestos. The nozzle employs light weight ablative parts and is some 5,000 pounds lighter than the Shuttle motor used to date. The payload performance of the ASRM-powered Shuttle is 12,000 pounds higher than that provided by the present motor. This is of particular benefit for payloads delivered to higher inclinations and/or altitudes. The ASRM facility uses state-of-the-art manufacturing techniques, including continuous propellant mixing and direct casting.

  20. NASA's Advanced solid rocket motor

    Mitchell, Royce E.


    The Advanced Solid Rocket Motor (ASRM) will not only bring increased safety, reliability and performance for the Space Shuttle Booster, it will enhance overall Shuttle safety by effectively eliminating 174 failure points in the Space Shuttle Main Engine throttling system and by reducing the exposure time to aborts due to main engine loss or shutdown. In some missions, the vulnerability time to Return-to-Launch Site aborts is halved. The ASRM uses case joints which will close or remain static under the effects of motor ignition and pressurization. The case itself is constructed of the weldable steel alloy HP 9-4-0.30, having very high strength and with superior fracture toughness and stress corrosion resistance. The internal insulation is strip-wound and is free of asbestos. The nozzle employs light weight ablative parts and is some 5,000 pounds lighter than the Shuttle motor used to date. The payload performance of the ASRM-powered Shuttle is 12,000 pounds higher than that provided by the present motor. This is of particular benefit for payloads delivered to higher inclinations and/or altitudes. The ASRM facility uses state-of-the-art manufacturing techniques, including continuous propellant mixing and direct casting.

  1. 78 FR 45052 - Airworthiness Directives; Hartzell Propeller, Inc. Propellers


    ... Manager, Engine & Propeller Directorate, Aircraft Certification Service. BILLING CODE 4910-13-P ...-07-AD; Amendment 39-17520; AD 2013-15-04] RIN 2120-AA64 Airworthiness Directives; Hartzell Propeller, Inc. Propellers AGENCY: Federal Aviation Administration (FAA), DOT. ACTION: Final rule. SUMMARY:...

  2. The Alfred Nobel rocket camera. An early aerial photography attempt

    Ingemar Skoog, A.


    Alfred Nobel (1833-1896), mainly known for his invention of dynamite and the creation of the Nobel Prices, was an engineer and inventor active in many fields of science and engineering, e.g. chemistry, medicine, mechanics, metallurgy, optics, armoury and rocketry. Amongst his inventions in rocketry was the smokeless solid propellant ballistite (i.e. cordite) patented for the first time in 1887. As a very wealthy person he actively supported many Swedish inventors in their work. One of them was W.T. Unge, who was devoted to the development of rockets and their applications. Nobel and Unge had several rocket patents together and also jointly worked on various rocket applications. In mid-1896 Nobel applied for patents in England and France for "An Improved Mode of Obtaining Photographic Maps and Earth or Ground Measurements" using a photographic camera carried by a "…balloon, rocket or missile…". During the remaining of 1896 the mechanical design of the camera mechanism was pursued and cameras manufactured. In April 1897 (after the death of Alfred Nobel) the first aerial photos were taken by these cameras. These photos might be the first documented aerial photos taken by a rocket borne camera. Cameras and photos from 1897 have been preserved. Nobel did not only develop the rocket borne camera but also proposed methods on how to use the photographs taken for ground measurements and preparing maps.

  3. Technology of foamed propellants

    Boehnlein-Mauss, Jutta; Kroeber, Hartmut [Fraunhofer Institut fuer Chemische Technologie ICT, Pfinztal (Germany)


    Foamed propellants are based on crystalline explosives bonded in energetic reaction polymers. Due to their porous structures they are distinguished by high burning rates. Energy content and material characteristics can be varied by using different energetic fillers, energetic polymers and porous structures. Foamed charges can be produced easily by the reaction injection moulding process. For the manufacturing of foamed propellants a semi-continuous remote controlled production plant in pilot scale was set up and a modified reaction injection moulding process was applied. (Abstract Copyright [2009], Wiley Periodicals, Inc.)

  4. Radiation/convection coupling in rocket motors and plumes

    Farmer, R. C.; Saladino, A. J.


    The three commonly used propellant systems - H2/O2, RP-1/O2, and solid propellants - primarily radiate as molecular emitters, non-scattering small particles, and scattering larger particles, respectively. Present technology has accepted the uncoupling of the radiation analysis from that of the flowfield. This approximation becomes increasingly inaccurate as one considers plumes, interior rocket chambers, and nuclear rocket propulsion devices. This study will develop a hierarchy of methods which will address radiation/convection coupling in all of the aforementioned propulsion systems. The nature of the radiation/convection coupled problem is that the divergence of the radiative heat flux must be included in the energy equation and that the local, volume-averaged intensity of the radiation must be determined by a solution of the radiative transfer equation (RTE). The intensity is approximated by solving the RTE along several lines of sight (LOS) for each point in the flowfield. Such a procedure is extremely costly; therefore, further approximations are needed. Modified differential approximations are being developed for this purpose. It is not obvious which order of approximations are required for a given rocket motor analysis. Therefore, LOS calculations have been made for typical rocket motor operating conditions in order to select the type approximations required. The results of these radiation calculations, and the interpretation of these intensity predictions are presented herein.

  5. 'RCHX-1-STORM' first Slovenian meteorological rocket program

    Kerstein, Aleksander; Matko, Drago; Trauner, Amalija; Britovšek, Zvone


    Astronautic and Rocket Society Celje (ARSC) formed a special working team for research and development of a small meteorological hail suppression rocket in the 70th. The hail suppression system was established in former Yugoslavia in the late 60th as an attempt to protect important agricultural regions from one of the summer's most vicious storm. In this time Slovenia was a part of Yugoslavia as one of the federal republic with relative high developed agricultural region production. The Rocket program 'RCHX-STORM' was a second attempt, for Slovenia indigenously developed in the production of meteorological hail suppression rocket. ARSC has designed a family of small sounding rocket that were based on highly promising hybrid propellant propulsion. Hybrid propulsion was selected for this family because it was offering low cost, save production and operation and simple logistics. Conventional sounding rockets use solid propellant motor for their propulsion. The introduction of hybrid motors has enabled a considerable decrease in overall cost. The transportation handling and storage procedures were greatly simplified due to the fact that a hybrid motor was not considered as explosive matter. A hybrid motor may also be designed to stand a severe environment without resorting to conditioning arrangements. The program started in the late 70th when the team ARSC was integrated in the Research and Development Institute in Celje (RDIC). The development program aimed to produce three types of meteorological rockets with diameters 76, 120 and 160 mm. Development of the RCHX-76 engine and rocket vehicle including flight certification has been undertaken by a joint team comprising of the ARCS, RDIC and the company Cestno podjetje Celje (CPC), Road building company Celje. Many new techniques and methods were used in this program such as computer simulation of external and internal ballistics, composite materials for rocket construction, intensive static testing of models and

  6. Rheokinetic Analysis of Hydroxy Terminated Polybutadiene Based Solid Propellant Slurry

    Abhay K Mahanta


    Full Text Available The cure kinetics of propellant slurry based on hydroxy-terminated polybutadiene (HTPB and toluene diisocyanate (TDI polyurethane reaction has been studied by viscosity build up method. The viscosity (ɳ–time (t plots conform to the exponential function ɳ = aebt, where a & b are empirical constants. The rate constants (k for viscosity build up at various shear rate (rpm, evaluated from the slope of dɳ/dt versus ɳ plots at different temperatures, were found to vary from 0.0032 to 0.0052 min-1. It was observed that the increasing shear rate did not have significant effect on the reaction rate constants for viscosity build up of the propellant slurry. The activation energy (Eɳ, calculated from the Arrhenius plots, was found to be 13.17±1.78 kJ mole-1, whereas the activation enthalpy (∆Hɳ* and entropy (∆Sɳ* of the propellant slurry, calculated from Eyring relationship, were found to be 10.48±1.78 kJ mole-1 and –258.51± 5.38 J mole-1K-1, respectively. The reaction quenching temperature of the propellant slurry was found to be -9 ° C, based upon the experimental data. This opens up an avenue for a “freeze-and-store”, then “warm-up and cast”, mode of manufacturing of very large solid rocket propellant grains.

  7. Another Look at Rocket Thrust

    Hester, Brooke; Burris, Jennifer


    Rocket propulsion is often introduced as an example of Newton's third law. The rocket exerts a force on the exhaust gas being ejected; the gas exerts an equal and opposite force--the thrust--on the rocket. Equivalently, in the absence of a net external force, the total momentum of the system, rocket plus ejected gas, remains constant. The law of…

  8. Another Look at Rocket Thrust

    Hester, Brooke; Burris, Jennifer


    Rocket propulsion is often introduced as an example of Newton's third law. The rocket exerts a force on the exhaust gas being ejected; the gas exerts an equal and opposite force--the thrust--on the rocket. Equivalently, in the absence of a net external force, the total momentum of the system, rocket plus ejected gas, remains constant. The law of…

  9. Spark Ignition of Combustible Vapor in a Plastic Bottle as a Demonstration of Rocket Propulsion

    Mattox, J. R.


    I report an innovation that provides a compelling demonstration of rocket propulsion, appropriate for students of physics and other physical sciences. An electrical spark is initiated from a distance to cause the deflagration of a combustible vapor mixed with air in a lightweight plastic bottle that is consequently propelled as a rocket by the release of combustion products, i.e., a "whoosh rocket." My recommendation is that the standard fuel for pedagogical whoosh demonstrations be isopropanol, and the recommended vessel is the 3.8-L high-density polyethylene (HDPE) bottle.

  10. Bibliography of Books and Published Reports on Gas Turbines, Jet Propulsion, and Rocket Power Plants


    Note No. 19󈧟 (1949). R. Tennant and P. Kahn, Super -sonic thrust. Aeroplane 77, 74 (1949). 1950. 0. E. Balje, A contribution to the design o-" radial...Meteor’s stable-mate. Flight 53, 429 (1948). Napier Naiad. Flight 54, 180 (1948). Provoked Attacker. Flight 53, 247 (1948). Saunders-Roe SR/45... Super -altitude research rocket revealed by the Navy. Aviation 46, 40 (June 1947). J. H. Wyld, The liquid-propellant rocket motor. J. Am. Rocket Soc

  11. Solid Propellant Test Article (SPTA) Test Stand


    This photograph shows the Solid Propellant Test Article (SPTA) test stand with the Modified Nasa Motor (M-NASA) test article at the Marshall Space Flight Center (MSFC). The SPTA test stand, 12-feet wide by 12-feet long by 24-feet high, was built in 1989 to provide comparative performance data on nozzle and case insulation material and to verify thermostructural analysis models. A modified NASA 48-inch solid motor (M-NASA motor) with a 12-foot blast tube and 10-inch throat makes up the SPTA. The M-NASA motor is being used to evaluate solid rocket motor internal non-asbestos insulation materials, nozzle designs, materials, and new inspection techniques. New internal motor case instrumentation techniques are also being evaluated.

  12. Rocket University at KSC

    Sullivan, Steven J.


    "Rocket University" is an exciting new initiative at Kennedy Space Center led by NASA's Engineering and Technology Directorate. This hands-on experience has been established to develop, refine & maintain targeted flight engineering skills to enable the Agency and KSC strategic goals. Through "RocketU", KSC is developing a nimble, rapid flight engineering life cycle systems knowledge base. Ongoing activities in RocketU develop and test new technologies and potential customer systems through small scale vehicles, build and maintain flight experience through balloon and small-scale rocket missions, and enable a revolving fresh perspective of engineers with hands on expertise back into the large scale NASA programs, providing a more experienced multi-disciplined set of systems engineers. This overview will define the Program, highlight aspects of the training curriculum, and identify recent accomplishments and activities.

  13. Disposal of Liquid Propellants


    SYNTHESIS OF LIQUID PROPELLANT Hydroxylammonium nitrate (HAN), prepared via the electrolysis of nitric acid, is commercially available as a high-purity...stack gases, and brine solution from the wet scrubber (82). 5 Applicability/Limitation Most types of solid, liquid, and gaseous organic wastes or

  14. Study of solid rocket motor for space shuttle booster, Volume 3: Program acquisition planning


    The program planning acquisition functions for the development of the solid propellant rocket engine for the space shuttle booster is presented. The subjects discussed are: (1) program management, (2) contracts administration, (3) systems engineering, (4) configuration management, and (5) maintenance engineering. The plans for manufacturing, testing, and operations support are included.

  15. High Frequency Combustion Instabilities of LOx/CH4 Spray Flames in Rocket Engine Combustion Chambers

    Sliphorst, M.


    Ever since the early stages of space transportation in the 1940’s, and the related liquid propellant rocket engine development, combustion instability has been a major issue. High frequency combustion instability (HFCI) is the interaction between combustion and the acoustic field in the combustion c

  16. High Frequency Combustion Instabilities of LOx/CH4 Spray Flames in Rocket Engine Combustion Chambers

    Sliphorst, M.


    Ever since the early stages of space transportation in the 1940’s, and the related liquid propellant rocket engine development, combustion instability has been a major issue. High frequency combustion instability (HFCI) is the interaction between combustion and the acoustic field in the combustion c

  17. Space shuttle with common fuel tank for liquid rocket booster and main engines (supertanker space shuttle)

    Thorpe, Douglas G.


    An operation and schedule enhancement is shown that replaces the four-body cluster (Space Shuttle Orbiter (SSO), external tank, and two solid rocket boosters) with a simpler two-body cluster (SSO and liquid rocket booster/external tank). At staging velocity, the booster unit (liquid-fueled booster engines and vehicle support structure) is jettisoned while the remaining SSO and supertank continues on to orbit. The simpler two-bodied cluster reduces the processing and stack time until SSO mate from 57 days (for the solid rocket booster) to 20 days (for the liquid rocket booster). The areas in which liquid booster systems are superior to solid rocket boosters are discussed. Alternative and future generation vehicles are reviewed to reveal greater performance and operations enhancements with more modifications to the current methods of propulsion design philosophy, e.g., combined cycle engines, and concentric propellant tanks.

  18. Improved method of measuring pressure coupled response for composite solid propellants

    Su, Wanxing; Wang, Ningfei; Li, Junwei; Zhao, Yandong; Yan, Mi


    Pressure coupled response is one of the main causes of combustion instability in the solid rocket motor. It is also a characteristic parameter for predicting the stability. The pressure coupled response function is usually measured by different methods to evaluate the performance of new propellant. Based on T-burner and "burning surface doubled and secondary attenuation", an improved method for measuring the pressure coupled response of composite propellant is introduced in this article. A computational fluid dynamics (CFD) study has also been conducted to validate the method and to understand the pressure oscillation phenomenon in T-burner. Three rounds of tests were carried out on the same batch of aluminized AP/HTPB composite solid propellant. The experimental results show that the sample propellant had a high response function under the conditions of high pressure (~11.5 MPa) and low frequency (~140 Hz). The numerically predicted oscillation frequency and amplitude are consistent with the experimental results. One practical solid rocket motor using this sample propellant was found to experience pressure oscillation at the end of burning. This confirms that the sample propellant is prone to combustion instability. Finally, acoustic pressure distribution and phase difference in T-burner were analyzed. Both the experimental and numerical results are found to be associated with similar acoustic pressure distribution. And the phase difference analysis showed that the pressure oscillations at the head end of the T-burner are 180° out of phase from those in the aft end of the T-burner.

  19. Quantitative computer representation of propellant processing

    Hicks, M. D.; Nikravesh, P. E.


    With the technology currently available for the manufacture of propellants, it is possible to control the variance of the total specific impulse obtained from the rocket boosters to within approximately five percent. Though at first inspection this may appear to be a reasonable amount of control, when it is considered that any uncertainty in the total kinetic energy delivered to the spacecraft translates into a design with less total usable payload, even this degree of uncertainty becomes unacceptable. There is strong motivation to control the variance in the specific impulse of the shuttle's solid boosters. Any small gains in the predictability and reliability of the booster would lead to a very substantial payoff in earth-to-orbit payload. The purpose of this study is to examine one aspect of the manufacture of solid propellants, namely, the mixing process. The traditional approach of computational fluid mechanics is notoriously complex and time consuming. Certain simplifications are made, yet certain fundamental aspects of the mixing process are investigated as a whole. It is possible to consider a mixing process in a mathematical sense as an operator, F, which maps a domain back upon itself. An operator which demonstrates good mixing should be able to spread any subset of the domain completely and evenly throughout the whole domain by successive applications of the mixing operator, F. Two and three dimensional models are developed and graphical visualization two and three dimensional mixing processes are presented.

  20. Simulation Research of Vaporization and Pressure Variation in a Cryogenic Propellant Tank at the Launch Site

    Chen, Liang; Liang, Guo-zhu


    In order to improve depiction of pressure variation and investigate the interrelation among the physical processes in propellant tanks, a 2D axial symmetry Volume-of-Fluid (VOF) CFD model is established to simulate a large-sized liquid propellant tank when the rocket is preparing for launch with propellant loaded at the launch site. The numerical model is considered with propellant free convection, heat transfer between the tank and the external environment, thermal exchange between propellant and inner tank wall surfaces, gas compressibility, and phase change modeled under the assumption of thermodynamic equilibrium. Vaporization rate of the vented LH2 tank and prediction of pressure change in the tank pressurized with GHe are obtained through simulation. We analysis the distributions of phase, temperature, and velocity vectors to reveal interactions among the propellant's own convection motion, heat transfer and phase change. The results show that the vaporization rate is mainly affected by heat leaks though the tank wall when the tank is vented, but it does not completely accord with the trend of the leakage because of convection motion and temperature nonuniformity of the liquid propellant in the tank. We also find that the main factors on pressure variation in the pressurized tank are the heat transfer on the tank wall surface bonding the ullage and propellant vaporization which has comparatively less influence.

  1. Shuttle Redesigned Solid Rocket Motor aluminum oxide investigations

    Blomshield, Fred S.; Kraeutle, Karl J.; Stalnaker, Richard A.


    During the launch of STS-54, a 15 psi pressure blip was observed in the ballistic pressure trace of one of the two Space Shuttle Redesigned Solid Rocket Motors (RSRM). One possible scenario for the observed pressure increase deals with aluminum oxide slag formation in the RSRM. The purpose of this investigation was to examine changes which may have occurred in the aluminum oxide formation in shuttle solid propellant due to changes in the ammonium perchlorate. Aluminum oxide formation from three propellants, all having the same formulation, but containing ammonium perchlorate from different manufacturers, will be compared. Three methods have been used to look for possible differences among the propellants. The first method was to examine window bomb movies of the propellants burning at 100, 300 and 600 psia. The motor operating pressure during the pressure blip was around 600 psia. The second method used small samples of propellant which were fired in a combustion bomb which quenched the burning aluminum particles soon after they left the propellant surface. The bomb was fired in both argon and Nitrogen atmospheres at various pressures. Products from this device were examined by optical microscopy. The third method used larger propellant samples fired into a particle collection device which allowed the aluminum to react and combust more completely. This device was pressurized with Nitrogen to motor operating pressures. The collected products were subdivided into size fractions by screening and sedimentation and analyzed optically with an optical microscope. the results from all three methods indicate very small changes in the size distribution of combustion products.

  2. Additive Manufacturing a Liquid Hydrogen Rocket Engine

    Jones, Carl P.; Robertson, Elizabeth H.; Koelbl, Mary Beth; Singer, Chris


    Space Propulsion is a 5 day event being held from 2nd May to the 6th May 2016 at the Rome Marriott Park Hotel in Rome, Italy. This event showcases products like Propulsion sub-systems and components, Production and manufacturing issues, Liquid, Solid, Hybrid and Air-breathing Propulsion Systems for Launcher and Upper Stages, Overview of current programmes, AIV issues and tools, Flight testing and experience, Technology building blocks for Future Space Transportation Propulsion Systems : Launchers, Exploration platforms & Space Tourism, Green Propulsion for Space Transportation, New propellants, Rocket propulsion & global environment, Cost related aspects of Space Transportation propulsion, Modelling, Pressure-Thrust oscillations issues, Impact of new requirements and regulations on design etc. in the Automotive, Manufacturing, Fabrication, Repair & Maintenance industries.

  3. Origin of how steam rockets can reduce space transport cost by orders of magnitude

    Zuppero, Anthony; Larson, Thomas K.; Schnitzler, Bruce G.; Werner, James E.; Rice, John W.; Hill, Thomas J.; Richins, William D.; Parlier, Lynn


    A brief sketch shows the origin of why and how thermal rocket propulsion has the unique potential to dramatically reduce the cost of space transportation for most inner solar system missions of interest. Orders of magnitude reduction in cost are apparently possible when compared to all processes requiring electrolysis for the production of rocket fuels or propellants and to all electric propulsion systems. An order of magnitude advantage can be attributed to rocket propellant tank factors associated with storing water propellant, compared to cryogenic liquids. An order of magnitude can also be attributed to the simplicity of the extraction and processing of ice on the lunar surface, into an easily stored, non-cryogenic rocket propellant (water). A nuclear heated thermal rocket can deliver thousands of times its mass to Low Earth Orbit from the Lunar surface, providing the equivalent to orders of magnitude drop in launch cost for mass in Earth orbit. Mass includes water ice. These cost reductions depend (exponentially) on the mission delta-v requirements being less than about 6 km/s, or about 3 times the specific velocity of steam rockets (2 km/s, from Isp 200 sec). Such missions include: from the lunar surface to Low Lunar Orbit, (LLO), from LLO to lunar escape, from Low Earth Orbit (LEO) to Geosynchronous Orbit (GEO), from LEO to Earth Escape, from LEO to Mars Transfer Orbit, from LLO to GEO, missions returning payloads from about 10% of the periodic comets using propulsive capture to orbits around Earth itself, and fast, 100 day missions from Lunar Escape to Mars. All the assertions depend entirely and completely on the existence of abundant, nearly pure ice at the permanently dark North and South Poles of the Moon.

  4. Evaluation of Vortex Chamber Concepts for Liquid Rocket Engine Applications

    Trinh, Huu Phuoc; Knuth, Williams; Michaels, Scott; Turner, James E. (Technical Monitor)


    Rocket-based combined-cycle engines (RBBC) being considered at NASA for future generation launch vehicles feature clusters of small rocket thrusters as part of the engine components. Depending on specific RBBC concepts, these thrusters may be operated at various operating conditions including power level and/or propellant mixture ratio variations. To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for the subject cycle engine application. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concepts not only offer the system simplicity but they also would enhance the combustion performance. The test results showed that the chamber performance was markedly high even at a low chamber length-to- diameter ratio (L/D). This incentive can be translated to a convenience in the thrust chamber packaging.

  5. Soil-geochemical factors of rocket fuel migration in the landscape

    Krechetov, P. P.; Kasimov, N. S.; Koroleva, T. V.


    The effect of different soil-geochemical factors on the migration of asymmetric dimethylhydrazine (UDMH) in the landscape has been studied. Experimental studies have been performed on soil and rock samples with specified parameters of the material composition. The effect of organic matter, acid-base properties, particle size distribution, and mineralogy on the decrease in the concentration of UDMH in equilibrium solutions has been studied. It has been found that the soil-geochemical factors are arranged in the following series according to the effect on UDMH mobility: acid-base properties > organic matter content > clay fraction mineralogy > particle size distribution.

  6. Rocket Flight Path

    Jamie Waters


    Full Text Available This project uses Newton’s Second Law of Motion, Euler’s method, basic physics, and basic calculus to model the flight path of a rocket. From this, one can find the height and velocity at any point from launch to the maximum altitude, or apogee. This can then be compared to the actual values to see if the method of estimation is a plausible. The rocket used for this project is modeled after Bullistic-1 which was launched by the Society of Aeronautics and Rocketry at the University of South Florida.

  7. LANTR-based Mars missions: Go to phobos for propellant?

    Stancati, Michael L.; Jacobs, Mark K.; Rauwolf, Gerald A.


    Two of the high-leverage propulsion technologies that have been proposed for human Mars missions-the Nuclear Thermal Rocket (NTR) engine and In Situ Propellant Production (ISPP)-show even greater potential when combined. Many previous studies have demonstrated the efficacy of manufacturing return propellant in situ to reduce the delivered mass requirement for the Earth launch and outbound transportation elements for any round trip mission. For human Mars exploration, this advantage may well be enabling, given current launch vehicle capability projections and reasonable expectations for a constrained program budget. NASA has proposed that the same LOX-Augmented NTR (LANTR) engine concept designed for use on lunar stages could also be used for Mars vehicle configurations, and that the tanks could be filled with propellants from Phobos for the return trip. This approach preserves the strategy of using a few common design elements for both lunar and Mars missions, while also making a significant mass performance improvement for the Mars return stage. We characterize the likely impact on performance of ``steady-state'' Earth-Mars transportation, as compared to Mars-only ISPP alternatives, and offer a preview of potential cost savings (work still in progress) for steady-state operation with Phobos propellants.

  8. Modeling and Fault Simulation of Propellant Filling System

    Jiang, Yunchun; Liu, Weidong; Hou, Xiaobo


    Propellant filling system is one of the key ground plants in launching site of rocket that use liquid propellant. There is an urgent demand for ensuring and improving its reliability and safety, and there is no doubt that Failure Mode Effect Analysis (FMEA) is a good approach to meet it. Driven by the request to get more fault information for FMEA, and because of the high expense of propellant filling, in this paper, the working process of the propellant filling system in fault condition was studied by simulating based on AMESim. Firstly, based on analyzing its structure and function, the filling system was modular decomposed, and the mathematic models of every module were given, based on which the whole filling system was modeled in AMESim. Secondly, a general method of fault injecting into dynamic system was proposed, and as an example, two typical faults - leakage and blockage - were injected into the model of filling system, based on which one can get two fault models in AMESim. After that, fault simulation was processed and the dynamic characteristics of several key parameters were analyzed under fault conditions. The results show that the model can simulate effectively the two faults, and can be used to provide guidance for the filling system maintain and amelioration.

  9. A feasibility study on using inkjet technology, micropumps, and MEMs as fuel injectors for bipropellant rocket engines

    Glynne-Jones, Peter; Coletti, Michele; White, Neil M.; Gabriel, Stephen; Bramanti, Cristina


    Control over drop size distributions, injection rates, and geometrical distribution of fuel and oxidizer sprays in bi-propellant rocket engines has the potential to produce more efficient, more stable, less polluting rocket engines. This control also offers the potential of an engine that can be throttled, working efficiently over a wide range of output thrusts. Inkjet printing technologies, MEMS fuel atomizers, and piezoelectric injectors similar in concept to those used in diesel engines ar...

  10. Rockets in World War I


    World War I enlisted rockets once again for military purposes. French pilots rigged rockets to the wing struts of their airplanes and aimed them at enemy observation balloons filled with highly inflammable hydrogen.

  11. Infrared signature modelling of a rocket jet plume - comparison with flight measurements

    Rialland, V.; Guy, A.; Gueyffier, D.; Perez, P.; Roblin, A.; Smithson, T.


    The infrared signature modelling of rocket plumes is a challenging problem involving rocket geometry, propellant composition, combustion modelling, trajectory calculations, fluid mechanics, atmosphere modelling, calculation of gas and particles radiative properties and of radiative transfer through the atmosphere. This paper presents ONERA simulation tools chained together to achieve infrared signature prediction, and the comparison of the estimated and measured signatures of an in-flight rocket plume. We consider the case of a solid rocket motor with aluminized propellant, the Black Brant sounding rocket. The calculation case reproduces the conditions of an experimental rocket launch, performed at White Sands in 1997, for which we obtained high quality infrared signature data sets from DRDC Valcartier. The jet plume is calculated using an in-house CFD software called CEDRE. The plume infrared signature is then computed on the spectral interval 1900-5000 cm-1 with a step of 5 cm-1. The models and their hypotheses are presented and discussed. Then the resulting plume properties, radiance and spectra are detailed. Finally, the estimated infrared signature is compared with the spectral imaging measurements. The discrepancies are analyzed and discussed.

  12. The engineering of a nuclear thermal landing and ascent vehicle utilizing indigenous Martian propellant

    Zubrin, Robert M.


    The following paper reports on a design study of a novel space transportation concept known as a 'NIMF' (Nuclear rocket using Indigenous Martian Fuel). The NIMF is a ballistic vehicle which obtains its propellant out of the Martian air by compression and liquefaction of atmospheric CO2. This propellant is subsequently used to generate rocket thrust at a specific impulse of 264 s by being heated to high temperature (2800 K) gas in the NIMFs' nuclear thermal rocket engines. The vehicle is designed to provide surface to orbit and surface to surface transportation, as well as housing, for a crew of three astronauts. It is capable of refueling itself for a flight to its maximum orbit in less than 50 days. The ballistic NIMF has a mass of 44.7 tonnes and, with the assumed 2800 K propellant temperature, is capable of attaining highly energetic (250 km by 34,000 km elliptical) orbits. This allows it to rendezvous with interplanetary transfer vehicles which are only very loosely bound into orbit around Mars. If a propellant temperature of 2000 K is assumed, then low Mars orbit can be attained; while if 3100 K is assumed, then the ballistic NIMF is capable of injecting itself onto a minimum energy transfer orbit to Earth in a direct ascent from the Martian surface.

  13. Hydrocarbon Rocket Technology Impact Forecasting

    Stuber, Eric; Prasadh, Nishant; Edwards, Stephen; Mavris, Dimitri N.


    Ever since the Apollo program ended, the development of launch propulsion systems in the US has fallen drastically, with only two new booster engine developments, the SSME and the RS-68, occurring in the past few decades.1 In recent years, however, there has been an increased interest in pursuing more effective launch propulsion technologies in the U.S., exemplified by the NASA Office of the Chief Technologist s inclusion of Launch Propulsion Systems as the first technological area in the Space Technology Roadmaps2. One area of particular interest to both government agencies and commercial entities has been the development of hydrocarbon engines; NASA and the Air Force Research Lab3 have expressed interest in the use of hydrocarbon fuels for their respective SLS Booster and Reusable Booster System concepts, and two major commercially-developed launch vehicles SpaceX s Falcon 9 and Orbital Sciences Antares feature engines that use RP-1 kerosene fuel. Compared to engines powered by liquid hydrogen, hydrocarbon-fueled engines have a greater propellant density (usually resulting in a lighter overall engine), produce greater propulsive force, possess easier fuel handling and loading, and for reusable vehicle concepts can provide a shorter turnaround time between launches. These benefits suggest that a hydrocarbon-fueled launch vehicle would allow for a cheap and frequent means of access to space.1 However, the time and money required for the development of a new engine still presents a major challenge. Long and costly design, development, testing and evaluation (DDT&E) programs underscore the importance of identifying critical technologies and prioritizing investment efforts. Trade studies must be performed on engine concepts examining the affordability, operability, and reliability of each concept, and quantifying the impacts of proposed technologies. These studies can be performed through use of the Technology Impact Forecasting (TIF) method. The Technology Impact

  14. Baking Soda and Vinegar Rockets

    Claycomb, James R.; Zachary, Christopher; Tran, Quoc


    Rocket experiments demonstrating conservation of momentum will never fail to generate enthusiasm in undergraduate physics laboratories. In this paper, we describe tests on rockets from two vendors that combine baking soda and vinegar for propulsion. The experiment compared two analytical approximations for the maximum rocket height to the…

  15. Baking Soda and Vinegar Rockets

    Claycomb, James R.; Zachary, Christopher; Tran, Quoc


    Rocket experiments demonstrating conservation of momentum will never fail to generate enthusiasm in undergraduate physics laboratories. In this paper, we describe tests on rockets from two vendors that combine baking soda and vinegar for propulsion. The experiment compared two analytical approximations for the maximum rocket height to the…

  16. Introduction to Rocket Propulsion


    Von Braun; 1966. 4. Introduction to Ordnance Technology; IHSP 76-129; 1976. 5. Physics; D. Halliday and R. Resnick ; 1963. 6. Physics Tells Why: Luke Sky- walker in Star Wars when he said "Don’t get cocky." We never plan for EVERYTHING, though we like to think we do. As we’ve said, rocket

  17. This "Is" Rocket Science!

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela


    Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical…

  18. The Relativistic Rocket

    Antippa, Adel F.


    We solve the problem of the relativistic rocket by making use of the relation between Lorentzian and Galilean velocities, as well as the laws of superposition of successive collinear Lorentz boosts in the limit of infinitesimal boosts. The solution is conceptually simple, and technically straightforward, and provides an example of a powerful…

  19. Rocketing to the Skies


    ONE sunny morning,we startedfor Yanqi Lake,Huairou District,Beijing,to try“rocket bungy”,so farthe only facility for this sport inChina.On the way there,wequestioned our courage and heartendurance. Entering the gate we saw,towering over a banner saying,

  20. Performance evaluation of commercial copper chromites as burning rate catalyst for solid propellants

    Milton Faria Diniz


    Full Text Available Copper chromites are well known as burning rate catalysts for the combustion of composite solid propellants, used as a source of energy for rocket propulsion. The propellant burning rate depends upon the catalyst characteristics such as chemical composition and specific surface area. In this work, copper chromite samples from different suppliers were characterized by chemical analysis, FT-IR spectroscopy and by surface area measurement (BET. The samples were then evaluated as burning rate catalyst in a typical composite propellant formulation based on HTPB binder, ammonium perchlorate and aluminum. The obtained surface area values are very close to those informed by the catalyst suppliers. The propellant processing as well as its mechanical properties were not substantially affected by the type of catalyst. Some copper chromite catalysts caused an increase in the propellant burning rate in comparison to the iron oxide catalyst. The results show that in addition to the surface area, other parameters like chemical composition, crystalline structure and the presence of impurities might be affecting the catalyst performance. All evaluated copper chromite samples may be used as burning rate catalyst in composite solid propellant formulations, with slight advantages for the SX14, Cu-0202P and Cu-1800P samples, which led to the highest burning rate propellants.

  1. Computational survey of representative energetic materials as propellants for microthruster applications

    Fuchs, Brian; Stec, Daniel, III


    Microthrusters are critical for the development of terrestrial micromissiles and nano air vehicles for reconnaissance, surveillance, and sensor emplacement. With the maturation of MEMS manufacturing technology, the physical components of the thrusters can be readily fabricated. The thruster type that is the most straightforward is chemical combustion of a propellant that is ignited by a heating element giving a single shot thrust. Arrays of MEMS manufactured thrusters can be ganged to give multiple firings. The basic model for such a system is a solid rocket motor. The desired elements for the propellant of a chemical thruster are high specific impulse (I sp), high temperature and pressure, and low molecular weight combustion gases. Since the combustion chamber of a microthruster is extremely small, the propellant material must be able to ignite, sustain and complete its burn inside the chamber. The propellant can be either a solid or a liquid. There are a large number of energetic materials available as candidates for a propellant for microthrusters. There has been no systematic evaluation of the available energetic materials as propellant candidates for microthrusters. This report summarizes computations done on a series of energetic materials to address their suitabilities as microthruster propellants.

  2. Solid rocket motor fire tests: Phases 1 and 2

    Chang, Yale; Hunter, Lawrence W.; Han, David K.; Thomas, Michael E.; Cain, Russell P.; Lennon, Andrew M.


    JHU/APL conducted a series of open-air burns of small blocks (3 to 10 kg) of solid rocket motor (SRM) propellant at the Thiokol Elkton MD facility to elucidate the thermal environment under burning propellant. The propellant was TP-H-3340A for the STAR 48 motor, with a weight ratio of 71/18/11 for the ammonium perchlorate, aluminum, and HTPB binder. Combustion inhibitor applied on the blocks allowed burning on the bottom and/or sides only. Burns were conducted on sand and concrete to simulate near-launch pad surfaces, and on graphite to simulate a low-recession surface. Unique test fixturing allowed propellant self-levitation while constraining lateral motion. Optics instrumentation consisted of a longwave infrared imaging pyrometer, a midwave spectroradiometer, and a UV/visible spectroradiometer. In-situ instrumentation consisted of rod calorimeters, Gardon gauges, elevated thermocouples, flush thermocouples, a two-color pyrometer, and Knudsen cells. Witness materials consisted of yttria, ceria, alumina, tungsten, iridium, and platinum/rhodium. Objectives of the tests were to determine propellant burn characteristics such as burn rate and self-levitation, to determine heat fluxes and temperatures, and to carry out materials analyses. A summary of qualitative results: alumina coated almost all surfaces, the concrete spalled, sand moisture content matters, the propellant self-levitated, the test fixtures worked as designed, and bottom-burning propellant does not self-extinguish. A summary of quantitative results: burn rate averaged 1.15 mm/s, thermocouples peaked at 2070 C, pyrometer readings matched MWIR data at about 2400 C, the volume-averaged plume temperatures were 2300-2400 C with peaks of 2400-2600 C, and the heat fluxes peaked at 125 W/cm2. These results are higher than other researchers' measurements of top-burning propellant in chimneys, and will be used, along with Phase 3 test results, to analyze hardware response to these environments, including General

  3. Producing propellants from water in lunar soil using solar lasers

    de Morais Mendonca Teles, Antonio

    The exploration of the Solar System is directly related to the efficiency of engines designed to explore it, and consequently, to the propulsion techniques, materials and propellants for those engines. With the present day propulsion techniques it is necessary great quantities of propellants to impulse a manned spacecraft to Mars and beyond in the Solar System, which makes these operations financially very expensive because of the costs involved in launching it from planet Earth, due to its high gravity field strength. To solve this problem, it is needed a planetary place with smaller gravity field strength, near to the Earth and with great quantities of substances at the surface necessary for the in-situ production of propellants for spacecrafts. The only place available is Earth's natural satellite the Moon. So, here in this paper, I propose the creation of a Lunar Propellant Manufacturer. It is a robot-spacecraft which can be launched from Earth using an Energia Rocket, and to land on the Moon in an area (principally near to the north pole where it was discovered water molecules ice recently) with great quantities of oxygen and hydrogen (propellants) in the silicate soil, previously observed and mapped by spacecrafts in lunar orbit, for the extraction of those molecules from the soil and the in-situ production of the necessary propellants. The Lunar Propellant Manufacturer (LPM) spacecraft consists of: 1) a landing system with four legs (extendable) and rovers -when the spacecraft touches down, the legs retract in order that two apparatuses, analogue to tractor's wheeled belts parallel sided and below the spacecraft, can touch firmly the ground -it will be necessary for the displacement of the spacecraft to new areas with richer propellants content, when the early place has already exhausted in propellants; 2) a digging machine -a long, resistant extendable arm with an excavator hand, in the outer part of the spacecraft -it will extend itself to the ground

  4. Boiler and Pressure Balls Monopropellant Thermal Rocket Engine

    Greene, William D. (Inventor)


    The proposed technology is a rocket engine cycle utilizing as the propulsive fluid a low molecular weight, cryogenic fluid, typically liquid hydrogen, pressure driven, heated, and expelled through a nozzle to generate high velocity and high specific impulse discharge gas. The proposed technology feeds the propellant through the engine cycle without the use of a separate pressurization fluid and without the use of turbomachinery. Advantages of the proposed technology are found in those elements of state-of-the-art systems that it avoids. It does not require a separate pressurization fluid or a thick-walled primary propellant tank as is typically required for a classical pressure-fed system. Further, it does not require the acceptance of intrinsic reliability risks associated with the use of turbomachinery

  5. Scale Effects on Solid Rocket Combustion Instability Behaviour

    David R. Greatrix


    Full Text Available The ability to understand and predict the expected internal behaviour of a given solid-propellant rocket motor under transient conditions is important. Research towards predicting and quantifying undesirable transient axial combustion instability symptoms necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. A numerical model incorporating pertinent elements, such as a representative transient, frequency-dependent combustion response to pressure wave activity above the burning propellant surface, is applied to the investigation of scale effects (motor size, i.e., grain length and internal port diameter on influencing instability-related behaviour in a cylindrical-grain motor. The results of this investigation reveal that the motor’s size has a significant influence on transient pressure wave magnitude and structure, and on the appearance and magnitude of an associated base pressure rise.

  6. Modelling of composite propellant properties

    Keizers, H.L.J.; Hordijk, A.C.; Vliet, L.D. van; Bouquet, F.


    State-of-the-art composite propellants are based on solid particles (AP, Aluminium) in a polymeric HTPB based binder system. The usability of a propellant for a particular application is dependent on a large number of properties. These different properties sometimes result in contradictory requireme

  7. Thermohydraulic Design Analysis Modeling for Korea Advanced NUclear Thermal Engine Rocket for Space Application

    Nam, Seung Hyun; Choi, Jae Young; Venneria, Paolo F.; Jeong, Yong Hoon; Chang, Soon Heung [KAIST, Daejeon (Korea, Republic of)


    Space exploration is a realistic and profitable goal for long-term humanity survival, although the harsh space environment imposes lots of severe challenges to space pioneers. To date, almost all space programs have relied upon Chemical Rockets (CRs) rating superior thrust level to transit from the Earth's surface to its orbit. However, CRs inherently have insurmountable barrier to carry out deep space missions beyond Earth's orbit due to its low propellant efficiency, and ensuing enormous propellant requirement and launch costs. Meanwhile, nuclear rockets typically offer at least two times the propellant efficiency of a CR and thus notably reduce the propellant demand. Particularly, a Nuclear Thermal Rocket (NTR) is a leading candidate for near-term manned missions to Mars and beyond because it satisfies a relatively high thrust as well as a high efficiency. The superior efficiency of NTRs is due to both high energy density of nuclear fuel and the low molecular weight propellant of Hydrogen (H{sub 2}) over the chemical reaction by-products. A NTR uses thermal energy released from a nuclear fission reactor to heat the H{sub 2} propellant and then exhausted the highly heated propellant through a propelling nozzle to produce thrust. A propellant efficiency parameter of rocket engines is specific impulse (I{sub s}p) which represents the ratio of the thrust over the propellant consumption rate. If the average exhaust H{sub 2} temperature of a NTR is around 3,000 K, the I{sub s}p can be achieved as high as 1,000 s as compared with only 450 - 500 s of the best CRs. For this reason, NTRs are favored for various space applications such as orbital tugs, lunar transports, and manned missions to Mars and beyond. The best known NTR development effort was conducted from 1955 to1974 under the ROVER and NERVA programs in the USA. These programs had successfully designed and tested many different reactors and engines. After these projects, the researches on NERVA derived

  8. Cryogenic Propellant Storage and Transfer Project

    National Aeronautics and Space Administration — The Cryogenic Propellant Storage and Transfer project will demonstrate the capability to safely and efficiently store, transfer and measure cryogenic propellants,...

  9. Recent Experimental Efforts on High-Pressure Supercritical Injection for Liquid Rockets and Their Implications

    Bruce Chehroudi


    Full Text Available Pressure and temperature of the liquid rocket thrust chambers into which propellants are injected have been in an ascending trajectory to gain higher specific impulse. It is quite possible then that the thermodynamic condition into which liquid propellants are injected reaches or surpasses the critical point of one or more of the injected fluids. For example, in cryogenic hydrogen/oxygen liquid rocket engines, such as Space Shuttle Main Engine (SSME or Vulcain (Ariane 5, the injected liquid oxygen finds itself in a supercritical condition. Very little detailed information was available on the behavior of liquid jets under such a harsh environment nearly two decades ago. The author had the opportunity to be intimately involved in the evolutionary understanding of injection processes at the Air Force Research Laboratory (AFRL, spanning sub- to supercritical conditions during this period. The information included here attempts to present a coherent summary of experimental achievements pertinent to liquid rockets, focusing only on the injection of nonreacting cryogenic liquids into a high-pressure environment surpassing the critical point of at least one of the propellants. Moreover, some implications of the results acquired under such an environment are offered in the context of the liquid rocket combustion instability problem.

  10. Attitude Dynamics of a Spinning Rocket with Internal Fluid Whirling Motion

    Marius Ionut MARMUREANU


    Full Text Available This paper evaluates the impact that helical motion of fluid products of combustion within the combustion chamber of a rocket can have on the attitude dynamics of rocket systems. By developing the study presented by Sookgaew (2004, we determined the configuration of the Coriolis moment components, which catch the impact of the combustion product’s whirling motion, for the radial and centripetal propellant burn pattern specific to S-5M and S-5K solid rocket motors. We continue the investigation of the effects of internal whirling motion of fluid products of combustion on the attitude behavior of variable mass systems of the rocket type by examining the spin motion and transverse attitude motion of such systems. The results obtained show that internal fluid whirling motion can cause appreciable deviations in spin rate predictions, and also affects the frequencies of the transverse angular velocity components.

  11. Rocket injector anomalies study. Volume 2: Results of parametric studies

    Przekwas, A. J.; Singhal, A. K.; Tam, L. T.


    The employment of a existing computer program to simulate three dimensional two phase gas spray flows in liquid propellant rocket engines. This was accomplished by modification of an existing three dimensional computer program (REFLAN3D) with Euler/Lagrange approach for simulating two phase spray flow, evaporation and combustion. The modified code is referred to as REFLAN3D-SPRAY. Computational studies of the model rocket engine combustion chamber are presented. The parametric studies of the two phase flow and combustion shows qualitatively correct response for variations in geometrical and physical parameters. The injection nonuniformity test with blocked central fuel injector holes shows significant changes in the central flame core and minor influence on the wall heat transfer fluxes.

  12. Characteristics of response factors of coaxial gaseous rocket injectors

    Janardan, B. A.; Daniel, B. R.; Zinn, B. T.


    The results of an experimental investigation undertaken to determine the frequency dependence of the response factors of various gaseous propellant rocket injectors subject to axial instabilities are presented. The injector response factors were determined, using the modified impedance-tube technique, under cold-flow conditions simulating those observed in unstable rocket motors. The tested injectors included a gaseous-fuel injector element, a gaseous-oxidizer injector element and a coaxial injector with both fuel and oxidizer elements. Emphasis was given to the determination of the dependence of the injector response factor upon the open-area ratio of the injector, the length of the injector orifice, and the pressure drop across the injector orifices. The measured data are shown to be in reasonable agreement with the corresponding injector response factor data predicted by the Feiler and Heidmann model.

  13. Reusable, flyback liquid rocket booster for the Space Shuttle

    Benton, Mark G.


    This paper outlines a preliminary design for an unmanned, reusable, flyback liquid rocket booster (LRB) as an evolutionary follow-on to the Shuttle solid rocket booster (SRB). Previous Shuttle liquid-propellant booster concepts are reviewed in order to gain insight into these designs. The operating costs, environmental impacts, and abort options of the SRB are discussed. The LRB flight profile and advantages of LRB use are discussed. The preliminary design for the LRB is outlined in detail using calculations and drawings. This design maximizes the use of existing hardware and proven technology to minimize cost and development time. The LRB design is presented as a more capable, more environmentally acceptable, and safer Shuttle booster.

  14. Testing of electroformed deposited iridium/powder metallurgy rhenium rockets

    Reed, Brian D.; Dickerson, Robert


    High-temperature, oxidation-resistant chamber materials offer the thermal margin for high performance and extended lifetimes for radiation-cooled rockets. Rhenium (Re) coated with iridium (Ir) allow hours of operation at 2200 C on Earth-storable propellants. One process for manufacturing Ir/Re rocket chambers is the fabrication of Re substrates by powder metallurgy (PM) and the application of Ir coatings by using electroformed deposition (ED). ED Ir coatings, however, have been found to be porous and poorly adherent. The integrity of ED Ir coatings could be improved by densification after the electroforming process. This report summarizes the testing of two 22-N, ED Ir/PM Re rocket chambers that were subjected to post-deposition treatments in an effort to densify the Ir coating. One chamber was vacuum annealed, while the other chamber was subjected to hot isostatic pressure (HIP). The chambers were tested on gaseous oxygen/gaseous hydrogen propellants, at mixture ratios that simulated the oxidizing environments of Earth-storable propellants. ne annealed ED Ir/PM Re chamber was tested for a total of 24 firings and 4.58 hr at a mixture ratio of 4.2. After only 9 firings, the annealed ED Ir coating began to blister and spall upstream of the throat. The blistering and spalling were similar to what had been experienced with unannealed, as-deposited ED Ir coatings. The HIP ED Ir/PM Re chamber was tested for a total of 91 firings and 11.45 hr at mixture ratios of 3.2 and 4.2. The HIP ED Ir coating remained adherent to the Re substrate throughout testing; there were no visible signs of coating degradation. Metallography revealed, however, thinning of the HIP Ir coating and occasional pores in the Re layer upstream of the throat. Pinholes in the Ir coating may have provided a path for oxidation of the Re substrate at these locations. The HIP ED Ir coating proved to be more effective than vacuum annealed and as-deposited ED Ir. Further densification is still required to

  15. Gas velocity and temperature near a liquid rocket injector face

    Boylan, D. M.; Ohara, J.


    The gas flow near the injector of a liquid propellant rocket was investigated by rapidly inserting butt-welded platinum-platinum rhodium thermocouples through the injector into the chamber. The transient responses of the thermocouples were analyzed to determine average gas temperatures and velocities. A method of fitting exponential curves to repeated measurements of the transient temperature at several positions near the injector face produced consistent results. Preliminary tests yielded gas flow directions and gas compositions at the injector face. Average gas temperatures were found to be between 3100 (1700) and 3500 F (1950 C) and the average gas velocities between 550 (170) and 840 feet/second (260 m/sec).

  16. Lunar soft landing with minimum-mass propulsion system using H2O2/kerosene bipropellant rocket system

    Moon, Yongjun; Kwon, Sejin


    Minimum-mass propulsion specifications using a H2O2/kerosene bipropellant rocket system for a small lunar lander were derived. A multivariable optimization was conducted with propulsion specifications and propellant consumptions obtained by solving optimal control problems for a lunar soft landing. In this paper, the optimal specifications, trajectory, and mass budget are presented.

  17. Design of Propellers for Motorsoarers

    Larrabee, E. E.


    A method was developed for the design of propellers of minimum induced loss matched to an arbitrary operating point characterized by disc loading (thrust or power), air density, shaft speed, flight speed, and number of blades. A consistent procedure is outlined to predict the performance of these propellers under off design conditions, or to predict the performance of propellers of general geometry. The examples discussed include a man powered airplane, a hang glider with a 7.5 kW (10 hp) 8,000 rpm engine, and an airplane-like motorsoarer.

  18. Magnetic propeller in symbiotic stars

    Panferov, Alexander; Mikolajewski, Maciej


    Rapidly spinning magnetic white dwarfs in symbiotic stars may pass through the propeller stage. It is believed that a magnetic propeller acts in two such stars CH Cyg and MWC 560. We review a diversity of manifestations of the propeller there. In these systems in a quiescent state the accretion onto a white dwarf from the strong enough wind of a companion star is suppressed by the magnetic field, and the hot component luminosity is low. Since the gas stored in the envelope eventually settles ...

  19. Liquid Propellants for Advanced Gun Ammunitions

    K. P. Rao


    Full Text Available With constant improvements, the conventional solid propellants for guns have almost reached their limit in performance. Liquid gun propellants are promising new comers capable of surpassing these performance limits and have numerous advantages over solid propellants. A method has been worked out to predict the internal ballistics of a liquid propellant gun and illustrated in a typical application.

  20. 14 CFR 35.22 - Feathering propellers.


    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Feathering propellers. 35.22 Section 35.22 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: PROPELLERS Design and Construction § 35.22 Feathering propellers. (a) Feathering propellers...

  1. 14 CFR 23.905 - Propellers.


    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propellers. 23.905 Section 23.905 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS... Propellers. (a) Each propeller must have a type certificate. (b) Engine power and propeller shaft...

  2. Draft environmental impact statement: Space Shuttle Advanced Solid Rocket Motor Program


    The proposed action is design, development, testing, and evaluation of Advanced Solid Rocket Motors (ASRM) to replace the motors currently used to launch the Space Shuttle. The proposed action includes design, construction, and operation of new government-owned, contractor-operated facilities for manufacturing and testing the ASRM's. The proposed action also includes transport of propellant-filled rocket motor segments from the manufacturing facility to the testing and launch sites and the return of used and/or refurbished segments to the manufacturing site.

  3. Nuclear Thermal Rocket Propulsion Systems


    NUCLEAR THERMAL ROCKET PROPULSION SYSTEMS, IAA WHITE PAPER PARIS, FRANCE, MARCH 2005 Lt Col Timothy J. Lawrence U.S. Air Force Academy...YYYY) 18-03-2005 2. REPORT TYPE White Paper 3. DATES COVERED (From - To) 18 Mar 2005 4. TITLE AND SUBTITLE NUCLEAR THERMAL ROCKET PROPULSION...reduce radiation exposure, is to have a high energy system like a nuclear thermal rocket that can get the payload to the destination in the fastest

  4. Computational Fluid Dynamics Simulation of Combustion Instability in Solid Rocket Motor : Implementation of Pressure Coupled Response Function

    S. Saha


    Full Text Available Combustion instability in solid propellant rocket motor is numerically simulated by implementing propellant response function with quasi steady homogeneous one dimensional formulation. The convolution integral of propellant response with pressure history is implemented through a user defined function in commercial computational fluid dynamics software. The methodology is validated against literature reported motor test and other simulation results. Computed amplitude of pressure fluctuations compare closely with the literarture data. The growth rate of pressure oscillations of a cylindrical grain solid rocket motor is determined for different response functions at the fundamental longitudinal frequency. It is observed that for response function more than a critical value, the motor exhibits exponential growth rate of pressure oscillations.

  5. Studies on Decomposition and Combustion Mechanism of Solid Fuel Rich Propellants


    energetic plasticizers in place of conventional inert binders and inert plasticizers like organic phthalates and acetates . FRP formulations based on poly...9), 1994, pp. 269-278. 16) Singh, H; Shekhar, H. Science and Technology of rocket propellants, Printed by Print well, Tow -Chowk, Darbhanga, (India...diazido 3,3 nitroazo pentane) -- -80 554 -- Diethylene glycol bis- azido acetate 1 -100 -392 -- Trimethylol nitro methane 1.4 -72 -230

  6. Formulation & Storage Studies on Hydrazine-Based Gelled Propellants.

    Mohan Verma


    Full Text Available The current trend in the area of high energy storable liquid rocket propellant research is to develop gelled/metallized systems and to explore the feasibility of their application in rocket motors. The idea stems from the fact that the conversion of a conventional liquid propellant to a gelled state and its subsequent metallization has the potential to significantly enhance the performance and density specific impulse. However, it is mandatory that the gelation of the liquid propellant be brought about with a minimum concentration of gellant and the gel formed should show a shear-thinning behaviour and good storage life. Keeping this in view, an experimental program to prepare thixotropic gels of hydrazine and its methyl substitutedderivatives like monomethyl hydrazineand unsymmetrical dimethyl hydrazinewas conducted under ambient conditions. The gellants investigated included particulate, chemical and synthetic materials. The gelation of the liquid fuels ..ould be brought about with gellant concentration as low as I wt per cent in some cases. Metallized gels using up to 40 wt per cent of AlIMg metal powders could be formulated. These heterogeneous systems showed a reduction in critical gellant concentration with degree of metallization. Further, less gelation time is observed if the gellant concentration is kept fixed and metal content is increased. The storage studies on gelled systems conducted over a period of three months showed good stability. The metallized systems, however, showed the settling of metal powders in meagre concentration which does not pose a major problem as these gels could be re-homogenized with stirring.

  7. Rocket Assembly and Checkout Facility

    Federal Laboratory Consortium — FUNCTION: Integrates, tests, and calibrates scientific instruments flown on sounding rocket payloads. The scientific instruments are assembled on an optical bench;...

  8. Combustion and Performance Analyses of Coaxial Element Injectors with Liquid Oxygen/Liquid Methane Propellants

    Hulka, J. R.; Jones, G. W.


    Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for in-space vehicles. This propellant combination has not been previously used in a flight-qualified engine system, so limited test data and analysis results are available at this stage of early development. NASA has funded several hardware-oriented activities with oxygen and methane propellants over the past several years with the Propulsion and Cryogenic Advanced Development (PCAD) project, under the Exploration Technology Development Program. As part of this effort, the NASA Marshall Space Flight Center has conducted combustion, performance, and combustion stability analyses of several of the configurations. This paper summarizes the analyses of combustion and performance as a follow-up to a paper published in the 2008 JANNAF/LPS meeting. Combustion stability analyses are presented in a separate paper. The current paper includes test and analysis results of coaxial element injectors using liquid oxygen and liquid methane or gaseous methane propellants. Several thrust chamber configurations have been modeled, including thrust chambers with multi-element swirl coax element injectors tested at the NASA MSFC, and a uni-element chamber with shear and swirl coax injectors tested at The Pennsylvania State University. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interaction Design and Analysis (ROCCID), and the Coaxial Injector Combustion Model (CICM). Significant effort was applied to show how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied or developed in the future. This paper describes the test hardware configurations, presents the results of all the analyses, and compares the results from the two analytical methods

  9. A two-phase restricted equilibrium model for combustion of metalized solid propellants

    Sabnis, J. S.; Dejong, F. J.; Gibeling, H. J.


    An Eulerian-Lagrangian two-phase approach was adopted to model the multi-phase reacting internal flow in a solid rocket with a metalized propellant. An Eulerian description was used to analyze the motion of the continuous phase which includes the gas as well as the small (micron-sized) particulates, while a Lagrangian description is used for the analysis of the discrete phase which consists of the larger particulates in the motor chamber. The particulates consist of Al and Al2O3 such that the particulate composition is 100 percent Al at injection from the propellant surface with Al2O3 fraction increasing due to combustion along the particle trajectory. An empirical model is used to compute the combustion rate for agglomerates while the continuous phase chemistry is treated using chemical equilibrium. The computer code was used to simulate the reacting flow in a solid rocket motor with an AP/HTPB/Al propellant. The computed results show the existence of an extended combustion zone in the chamber rather than a thin reaction region. The presence of the extended combustion zone results in the chamber flow field and chemical being far from isothermal (as would be predicted by a surface combustion assumption). The temperature in the chamber increases from about 2600 K at the propellant surface to about 3350 K in the core. Similarly the chemical composition and the density of the propellant gas also show spatially non-uniform distribution in the chamber. The analysis developed under the present effort provides a more sophisticated tool for solid rocket internal flow predictions than is presently available, and can be useful in studying apparent anomalies and improving the simple correlations currently in use. The code can be used in the analysis of combustion efficiency, thermal load in the internal insulation, plume radiation, etc.

  10. Regenerative Gas Dryer for In-Situ Propellant Production

    Paz, Aaron


    Rocket propellant can be produced anywhere that water is found by splitting it into hydrogen and oxygen, potentially saving several tons of mass per mission and enabling the long term presence of humans in space beyond LEO. When water is split into hydrogen and oxygen, the gaseous products can be very humid (several thousand ppm). Propellant-grade gases need to be extremely dry before being converted into cryogenic liquids (less than 26 ppm water for grade B Oxygen). The primary objective of this project is to design, build and test a regenerative gas drying system that can take humid gas from a water electrolysis system and provide dry gas (less than 26ppm water) to the inlet of a liquefaction system for long durations. State of the art work in this area attempted to use vacuum as a means to regenerate desiccant, but it was observed that water would migrate to the dry zone without a sweep gas present to direct the desorbed vapor. Further work attempted to use CO2 as a sweep gas, but this resulted in a corrosive carbonic acid. In order for in-situ propellant production to work, we need a way to continuously dry humid gas that addresses these issues.

  11. The University of Arizona program in solid propellants

    Ramohalli, Kumar


    The University of Arizona program is aimed at introducing scientific rigor to the predictability and quality assurance of composite solid propellants. Two separate approaches are followed: to use the modern analytical techniques to experimentally study carefully controlled propellant batches to discern trends in mixing, casting, and cure; and to examine a vast bank of data, that has fairly detailed information on the ingredients, processing, and rocket firing results. The experimental and analytical work is described briefly. The principle findings were that: (1) pre- (dry) blending of the coarse and fine ammonium perchlorate can significantly improve the uniformity of mixing; (2) the Fourier transformed IR spectra of the uncured and cured polymer have valuable data on the state of the fuel; (3) there are considerable non-uniformities in the propellant slurry composition near the solid surfaces (blades, walls) compared to the bulk slurry; and (4) in situ measurements of slurry viscosity continuously during mixing can give a good indication of the state of the slurry. Several important observations in the study of the data bank are discussed.

  12. Propeller aircraft interior noise model

    Pope, L. D.; Wilby, E. G.; Wilby, J. F.


    An analytical model was developed to predict the interior noise of propeller-driven aircraft. The fuselage model is that of a cylinder with a structurally-integral floor. The cabin sidewall is stiffened by stringers and ring frames, and the floor by longitudinal beams. The cabin interior is covered with a sidewall treatments consisting of layers of porous material and an impervious trim septum. Representation of the propeller pressure field is utilized as input data in the form of the propeller noise signature at a series of locations on a grid over the fuselage structure. Results obtained from the analytical model are compared with test data measured by NASA in a scale model cylindrical fuselage excited by a model propeller.

  13. Rocket + Science = Dialogue

    Morris,Bruce; Sullivan, Greg; Burkey, Martin


    It's a cliche that rocket engineers and space scientists don t see eye-to-eye. That goes double for rocket engineers working on human spaceflight and scientists working on space telescopes and planetary probes. They work fundamentally different problems but often feel that they are competing for the same pot of money. Put the two groups together for a weekend, and the results could be unscientific or perhaps combustible. Fortunately, that wasn't the case when NASA put heavy lift launch vehicle designers together with astronomers and planetary scientists for two weekend workshops in 2008. The goal was to bring the top people from both groups together to see how the mass and volume capabilities of NASA's Ares V heavy lift launch vehicle could benefit the science community. Ares V is part of NASA's Constellation Program for resuming human exploration beyond low Earth orbit, starting with missions to the Moon. In the current mission scenario, Ares V launches a lunar lander into Earth orbit. A smaller Ares I rocket launches the Orion crew vehicle with up to four astronauts. Orion docks with the lander, attached to the Ares V Earth departure stage. The stage fires its engine to send the mated spacecraft to the Moon. Standing 360 feet high and weighing 7.4 million pounds, NASA's new heavy lifter will be bigger than the 1960s-era Saturn V. It can launch almost 60 percent more payload to translunar insertion together with the Ares I and 35 percent more mass to low Earth orbit than the Saturn V. This super-sized capability is, in short, designed to send more people to more places to do more things than the six Apollo missions.

  14. Contra rotative propeller performance estimation

    Coca Casanueva, Vladimir


    Due to the continuous increase in the fuel price, the propeller engine solution (the most efficient in fuel saving terms) becomes very attractive to airlines and thus, to aircraft manufacturers. However, airlines aren’t ready to fly an aircraft at lower cruise Mach number than the traditional Mach 0,84, which jeopardizes the fuel efficiency of propellers. At this stage is where the contra-rotative concept appears, which let us to increase the cruise speed while reducing fuel consumption...


    Hartono Hartono


    Full Text Available Increase of the price of world oil pushs liner to conduct action is economic from every operational unit in it’sship armada. One other most dominant in usage of fuel is Main engine is working to turn around propeller asship actuator. On that account ship owner wants design propeller which is economic for it’s ship to canreduce usage of fuel of 20% when sailing.

  16. Combustion characteristics of SMX and SMX based propellants

    Reese, David A.

    This work investigates the combustion of the new solid nitrate ester 2,3-hydroxymethyl-2,3-dinitro-1,4-butanediol tetranitrate (SMX, C6H 8N6O16). SMX was synthesized for the first time in 2008. It has a melting point of 85 °C and oxygen balance of 0% to CO 2, allowing it to be used as an energetic additive or oxidizer in solid propellants. In addition to its neat combustion characteristics, this work also explores the use of SMX as a potential replacement for nitroglycerin (NG) in double base gun propellants and as a replacement for ammonium perchlorate in composite rocket propellants. The physical properties, sensitivity characteristics, and combustion behaviors of neat SMX were investigated. Its combustion is stable at pressures of up to at least 27.5 MPa (n = 0.81). The observed flame structure is nearly identical to that of other double base propellant ingredients, with a primary flame attached at the surface, a thick isothermal dark zone, and a luminous secondary flame wherein final recombination reactions occur. As a result, the burning rate and primary flame structure can be modeled using existing one-dimensional steady state techniques. A zero gas-phase activation energy approximation results in a good fit between modeled and observed behavior. Additionally, SMX was considered as a replacement for nitroglycerin in a double base propellant. Thermochemical calculations indicate improved performance when compared with the common double base propellant JA2 at SMX loadings above 40 wt-%. Also, since SMX is a room temperature solid, migration may be avoided. Like other nitrate esters, SMX is susceptible to decomposition over long-term storage due to the presence of excess acid in the crystals; the addition of stabilizers (e.g., derivatives of urea) during synthesis should be sufficient to prevent this. the addition of Both unplasticized and plasticized propellants were formulated. Thermal analysis of unplasticized propellant showed a distinct melt

  17. Single stage to orbit mass budgets derived from propellant density and specific impulse

    Whitehead, J.C.


    The trade between specific impulse (Isp) and density is examined in view of Single Stage To Orbit (SSTO) requirements. Mass allocations for vehicle hardware are derived from these two properties, for several propellant combinations and a dual-fuel case. This comparative analysis, based on flight-proven hardware, indicates that the higher density of several alternative propellants compensates for reduced Isp, when compared with cryogenic oxygen and hydrogen. Approximately half the orbiting mass of a rocket- propelled SSTO vehicle must be allocated to propulsion hardware and residuals. Using hydrogen as the only fuel requires a slightly greater fraction of orbiting mass for propulsion, because hydrogen engines and tanks are heavier than those for denser fuels. The advantage of burning both a dense fuel and hydrogen in succession depends strongly on tripropellant engine weight. The implications of the calculations for SSTO vehicle design are discussed, especially with regard to the necessity to minimize non-tankage structure.

  18. Aircraft Propeller Hub Repair

    Muth, Thomas R. [Oak Ridge National Lab. (ORNL), Oak Ridge, TN (United States); Peter, William H. [Oak Ridge National Lab. (ORNL), Oak Ridge, TN (United States)


    The team performed a literature review, conducted residual stress measurements, performed failure analysis, and demonstrated a solid state additive manufacturing repair technique on samples removed from a scrapped propeller hub. The team evaluated multiple options for hub repair that included existing metal buildup technologies that the Federal Aviation Administration (FAA) has already embraced, such as cold spray, high velocity oxy-fuel deposition (HVOF), and plasma spray. In addition the team helped Piedmont Propulsion Systems, LLC (PPS) evaluate three potential solutions that could be deployed at different stages in the life cycle of aluminum alloy hubs, in addition to the conventional spray coating method for repair. For new hubs, a machining practice to prevent fretting with the steel drive shaft was recommended. For hubs that were refurbished with some material remaining above the minimal material condition (MMC), a silver interface applied by an electromagnetic pulse additive manufacturing method was recommended. For hubs that were at or below the MMC, a solid state additive manufacturing technique using ultrasonic welding (UW) of thin layers of 7075 aluminum to the hub interface was recommended. A cladding demonstration using the UW technique achieved mechanical bonding of the layers showing promise as a viable repair method.

  19. Raman studies for stockpile reliability of missiles by detecting degradation of propellant stabilizers

    Farley, Carlton; Kassu, Aschalew; Mills, Jonathan; Bibb, Jonathan; Curley, Michael; Ruffin, Paul; Sharma, Anup; Rice, Jeremy; McDonald, Brian


    The objective of this study is to demonstrate a sensitive Raman technique for sensing degradation of propellant stabilizers like MNA and 2-NDPA that are commonly used in some missiles. The functionality of missiles and rockets are often evaluated by being fired or decomposed at routine time-intervals after prolonged storage. However, these destructive testing techniques for determining long-term rocket motor aging and shelf-life are extremely costly. If successful, the Raman technique could be utilized to determine the health of propellant stabilizers without dismantling the missiles as is commonly done at present. Raman technique is to measure concentrations of propellant stabilizers between 0.1-2% in glycerin. Two different lasers at 785 nm and 532 nm are used for developing this technique. A secondary objective is to develop a theoretical model that predicts temperature as a function of time and position inside the cylindrical storage container of MNA or 2-NDPA stabilizer. This model can help in understanding the thermal degradation of propellant stabilizers.

  20. Effect of Additives on Liner Properties of Case-bonded Composite Propellants

    S. B. Navale


    Full Text Available A thin layer of liner is applied to ensure a good bond between the insulator and the propellant in case-bonded rocket motors. It also acts as a protective shield for the insulatorby providing a limited fire protection effect. Liner compositions should preferably be based on the same binder system used in the propellant formulations. As the liner has to hold the propellant and the insulator without debond under all the environmental conditions, it plays a key role in predicted performance of a rocket motor. Hence, studies were carried out to improve the liner properties using various hydroxyl compounds, such as butanediol, cardanol, trimethylol propane, pyrogallol, etc as additives. Butanediol and phloroglucinol combination gave the best results in terms of mechanical properties and interface properties for the liner compositions. The effect of filler content on the liner properties was also studied. The results showed that higher filler content does not affect interface properties. Considering the fire retardancy effect and reinforcement of antimony trioxide (S£203, the formulation containing higher Sb2O3 was selected. The studies on pot life/castable life of liner showed that propellant could be cast up to 6 days after liner coating, without adversely affecting the bonding and the bond strength.

  1. High Frequency Combustion Instabilities of LOx/CH4 Spray Flames in Rocket Engine Combustion Chambers

    Sliphorst, M.


    Ever since the early stages of space transportation in the 1940’s, and the related liquid propellant rocket engine development, combustion instability has been a major issue. High frequency combustion instability (HFCI) is the interaction between combustion and the acoustic field in the combustion chamber. It destroys the thermal boundary layer wall increasing heat transfer and could lead to compromised performance, and ultimately to destruction of the engine and mission loss. The main object...

  2. The open-cycle gas-core nuclear rocket engine - Some engineering considerations.

    Taylor, M. F.; Whitmarsh, C. L., Jr.; Sirocky, P. J., Jr.; Iwanczyk, L. C.


    A preliminary design study of a conceptual 6000-MW open-cycle gas-core nuclear rocket engine system was made. The engine has a thrust of 44,200 lb and a specific impulse of 4400 sec. The nuclear fuel is uranium-235 and the propellant is hydrogen. Critical fuel mass was calculated for several reactor configurations. Major components of the reactor (reflector, pressure vessel) and the waste heat rejection system were considered conceptually and were sized.

  3. A History of Collapse Factor Modeling and Empirical Data for Cryogenic Propellant Tanks

    deQuay, Laurence; Hodge, B. Keith


    One of the major technical problems associated with cryogenic liquid propellant systems used to supply rocket engines and their subassemblies and components is the phenomenon of propellant tank pressurant and ullage gas collapse. This collapse is mainly caused by heat transfer from ullage gas to tank walls and interfacing propellant, which are both at temperatures well below those of this gas. Mass transfer between ullage gas and cryogenic propellant can also occur and have minor to significant secondary effects that can increase or decrease ullage gas collapse. Pressurant gas is supplied into cryogenic propellant tanks in order to initially pressurize these tanks and then maintain required pressures as propellant is expelled from these tanks. The net effect of pressurant and ullage gas collapse is increased total mass and mass flow rate requirements of pressurant gases. For flight vehicles this leads to significant and undesirable weight penalties. For rocket engine component and subassembly ground test facilities this results in significantly increased facility hardware, construction, and operational costs. "Collapse Factor" is a parameter used to quantify the pressurant and ullage gas collapse. Accurate prediction of collapse factors, through analytical methods and modeling tools, and collection and evaluation of collapse factor data has evolved over the years since the start of space exploration programs in the 1950 s. Through the years, numerous documents have been published to preserve results of studies associated with the collapse factor phenomenon. This paper presents a summary and selected details of prior literature that document the aforementioned studies. Additionally other literature that present studies and results of heat and mass transfer processes, related to or providing important insights or analytical methods for the studies of collapse factor, are presented.

  4. What fuel for a rocket?

    Miranda, E N


    Elementary concepts from general physics and thermodynamics have been used to analyze rocket propulsion. Making some reasonable assumptions, an expression for the exit velocity of the gases is found. From that expression one can conclude what are the desired properties for a rocket fuel.

  5. Rocket launchers as passive controllers

    Cochran, J. E., Jr.; Gunnels, R. T.; McCutchen, R. K., Jr.


    A concept is advanced for using the motion of launchers of a free-flight launcher/rocket system which is caused by random imperfections of the rockets launched from it to reduce the total error caused by the imperfections. This concept is called 'passive launcher control' because no feedback is generated by an active energy source after an error is sensed; only the feedback inherent in the launcher/rocket interaction is used. Relatively simple launcher models with two degrees of freedom, pitch and yaw, were used in conjunction with a more detailed, variable-mass model in a digital simulation code to obtain rocket trajectories with and without thrust misalignment and dynamic imbalance. Angular deviations of rocket velocities and linear deviations of the positions of rocket centers of mass at burnout were computed for cases in which the launcher was allowed to move ('flexible' launcher) and was constrained so that it did not rotate ('rigid' launcher) and ratios of flexible to rigid deviations were determined. Curves of these error ratios versus launcher frequency are presented. These show that a launcher which has a transverse moment of inertia about its pivot point of the same magnitude as that of the centroidal transverse moments of inertia of the rockets launched from it can be tuned to passively reduce the errors caused by rocket imperfections.

  6. Aspects of Propeller Developements for a Submarine

    Andersen, Poul; kappel, Jens Julius; Spangenberg, Eugen


    Design and development of propellers for submarines are in some ways different from propellers for surface vessels. The most important demand is low acoustic signature that has priority over propeller efficiency, and the submarine propeller must be optimized with respect to acoustics rather than...... efficiency. Moreover the operating conditions of a submarine propeller are quite different. These aspects are discussed as well as the weighing of the various propeller parameters against the design objectives. The noise generated by the propeller can be characterized as thrust noise due to the inhomogeneous...... wake field of the submarine, trailing-edge noise and noise caused by turbulence in the inflow. The items discussed are demonstrated in a case study where a propeller of the Kappel type was developed. Three stages of the development are presented, including a design of an 8-bladed propeller where...

  7. Assessment of tbe Performance of Ablative Insulators Under Realistic Solid Rocket Motor Operating Conditions (a Doctoral Dissertation)

    Martin, Heath Thomas


    Ablative insulators are used in the interior surfaces of solid rocket motors to prevent the mechanical structure of the rocket from failing due to intense heating by the high-temperature solid-propellant combustion products. The complexity of the ablation process underscores the need for ablative material response data procured from a realistic solid rocket motor environment, where all of the potential contributions to material degradation are present and in their appropriate proportions. For this purpose, the present study examines ablative material behavior in a laboratory-scale solid rocket motor. The test apparatus includes a planar, two-dimensional flow channel in which flat ablative material samples are installed downstream of an aluminized solid propellant grain and imaged via real-time X-ray radiography. In this way, the in-situ transient thermal response of an ablator to all of the thermal, chemical, and mechanical erosion mechanisms present in a solid rocket environment can be observed and recorded. The ablative material is instrumented with multiple micro-thermocouples, so that in-depth temperature histories are known. Both total heat flux and thermal radiation flux gauges have been designed, fabricated, and tested to characterize the thermal environment to which the ablative material samples are exposed. These tests not only allow different ablative materials to be compared in a realistic solid rocket motor environment but also improve the understanding of the mechanisms that influence the erosion behavior of a given ablative material.

  8. Development Testing of 1-Newton ADN-Based Rocket Engines

    Anflo, K.; Gronland, T.-A.; Bergman, G.; Nedar, R.; Thormählen, P.


    With the objective to reduce operational hazards and improve specific and density impulse as compared with hydrazine, the Research and Development (R&D) of a new monopropellant for space applications based on AmmoniumDiNitramide (ADN), was first proposed in 1997. This pioneering work has been described in previous papers1,2,3,4 . From the discussion above, it is clear that cost savings as well as risk reduction are the main drivers to develop a new generation of reduced hazard propellants. However, this alone is not enough to convince a spacecraft builder to choose a new technology. Cost, risk and schedule reduction are good incentives, but a spacecraft supplier will ask for evidence that this new propulsion system meets a number of requirements within the following areas: This paper describes the ongoing effort to develop a storable liquid monopropellant blend, based on AND, and its specific rocket engines. After building and testing more than 20 experimental rocket engines, the first Engineering Model (EM-1) has now accumulated more than 1 hour of firing-time. The results from test firings have validated the design. Specific impulse, combustion stability, blow-down capability and short pulse capability are amongst the requirements that have been demonstrated. The LMP-103x propellant candidate has been stored for more than 1 year and initial material compatibility screening and testing has started. 1. Performance &life 2. Impact on spacecraft design &operation 3. Flight heritage Hereafter, the essential requirements for some of these areas are outlined. These issues are discussed in detail in a previous paper1 . The use of "Commercial Of The Shelf" (COTS) propulsion system components as much as possible is essential to minimize the overall cost, risk and schedule. This leads to the conclusion that the Technology Readiness Level (TRL) 5 has been reached for the thruster and propellant. Furthermore, that the concept of ADN-based propulsion is feasible.

  9. Computing Q-D Relationships for Storage of Rocket Fuels

    Jester, Keith


    The Quantity Distance Measurement Tool is a GIS BASEP computer program that aids safety engineers by calculating quantity-distance (Q-D) relationships for vessels that contain explosive chemicals used in testing rocket engines. (Q-D relationships are standard relationships between specified quantities of specified explosive materials and minimum distances by which they must be separated from persons, objects, and other explosives to obtain specified types and degrees of protection.) The program uses customized geographic-information-system (GIS) software and calculates Q-D relationships in accordance with NASA's Safety Standard For Explosives, Propellants, and Pyrotechnics. Displays generated by the program enable the identification of hazards, showing the relationships of propellant-storage-vessel safety buffers to inhabited facilities and public roads. Current Q-D information is calculated and maintained in graphical form for all vessels that contain propellants or other chemicals, the explosiveness of which is expressed in TNT equivalents [amounts of trinitrotoluene (TNT) having equivalent explosive effects]. The program is useful in the acquisition, siting, construction, and/or modification of storage vessels and other facilities in the development of an improved test-facility safety program.

  10. Nuclear Rocket Engine Reactor

    Lanin, Anatoly


    The development of a nuclear rocket engine reactor (NRER ) is presented in this book. The working capacity of an active zone NRER under mechanical and thermal load, intensive neutron fluxes, high energy generation (up to 30 MBT/l) in a working medium (hydrogen) at temperatures up to 3100 K is displayed. Design principles and bearing capacity of reactors area discussed on the basis of simulation experiments and test data of a prototype reactor. Property data of dense constructional, porous thermal insulating and fuel materials like carbide and uranium carbide compounds in the temperatures interval 300 - 3000 K are presented. Technological aspects of strength and thermal strength resistance of materials are considered. The design procedure of possible emergency processes in the NRER is developed and risks for their origination are evaluated. Prospects of the NRER development for pilotless space devices and piloted interplanetary ships are viewed.

  11. Erosive burning of solid propellants

    King, Merrill K.


    Presented here is a review of the experimental and modeling work concerning erosive burning of solid propellants (augmentation of burning rate by flow of product gases across a burning surface). A brief introduction describes the motor design problems caused by this phenomenon, particularly for low port/throat area ratio motors and nozzleless motors. Various experimental techniques for measuring crossflow sensitivity of solid propellant burning rates are described, with the conclusion that accurate simulation of the flow, including upstream flow development, in actual motors is important since the degree of erosive burning depends not only on local mean crossflow velocity and propellant nature, but also upon this upstream development. In the modeling area, a brief review of simplified models and correlating equations is presented, followed by a description of more complex numerical analysis models. Both composite and double-base propellant models are reviewed. A second generation composite model is shown to give good agreement with data obtained in a series of tests in which composite propellant composition and heterogeneity (particle size distribution) were systematically varied. Finally, the use of numerical models for the development of erosive burning correlations is described, and a brief discussion of scaling is presented.

  12. Application of Chaboche Model in Rocket Thrust Chamber Analysis

    Asraff, Ahmedul Kabir; Suresh Babu, Sheela; Babu, Aneena; Eapen, Reeba


    Liquid Propellant Rocket Engines are commonly used in space technology. Thrust chamber is one of the most important subsystems of a rocket engine. The thrust chamber generates propulsive thrust force for flight of the rocket by ejection of combustion products at supersonic speeds. Often double walled construction is employed for these chambers. The thrust chamber investigated here has its hot inner wall fabricated out of a high thermal conductive material like copper alloy and outer wall made of stainless steel. Inner wall is subjected to high thermal and pressure loads during operation of engine due to which it will be in the plastic regime. Main reasons for the failure of such chambers are fatigue in the plastic range (called as low cycle fatigue since the number of cycles to failure will be low in plastic range), creep and thermal ratcheting. Elasto plastic material models are required to simulate the above effects through a cyclic stress analysis. This paper gives the details of cyclic stress analysis carried out for the thrust chamber using different plasticity model combinations available in ANSYS (Version 15) FE code. The best model among the above is applied in the cyclic stress analysis of two dimensional (plane strain and axisymmetric) and three dimensional finite element models of thrust chamber. Cyclic life of the chamber is calculated from stress-strain graph obtained from above analyses.

  13. Cycle Trades for Nuclear Thermal Rocket Propulsion Systems

    White, C.; Guidos, M.; Greene, W.


    Nuclear fission has been used as a reliable source for utility power in the United States for decades. Even in the 1940's, long before the United States had a viable space program, the theoretical benefits of nuclear power as applied to space travel were being explored. These benefits include long-life operation and high performance, particularly in the form of vehicle power density, enabling longer-lasting space missions. The configurations for nuclear rocket systems and chemical rocket systems are similar except that a nuclear rocket utilizes a fission reactor as its heat source. This thermal energy can be utilized directly to heat propellants that are then accelerated through a nozzle to generate thrust or it can be used as part of an electricity generation system. The former approach is Nuclear Thermal Propulsion (NTP) and the latter is Nuclear Electric Propulsion (NEP), which is then used to power thruster technologies such as ion thrusters. This paper will explore a number of indirect-NTP engine cycle configurations using assumed performance constraints and requirements, discuss the advantages and disadvantages of each cycle configuration, and present preliminary performance and size results. This paper is intended to lay the groundwork for future efforts in the development of a practical NTP system or a combined NTP/NEP hybrid system.

  14. MHD thrust vectoring of a rocket engine

    Labaune, Julien; Packan, Denis; Tholin, Fabien; Chemartin, Laurent; Stillace, Thierry; Masson, Frederic


    In this work, the possibility to use MagnetoHydroDynamics (MHD) to vectorize the thrust of a solid propellant rocket engine exhaust is investigated. Using a magnetic field for vectoring offers a mass gain and a reusability advantage compared to standard gimbaled, elastomer-joint systems. Analytical and numerical models were used to evaluate the flow deviation with a 1 Tesla magnetic field inside the nozzle. The fluid flow in the resistive MHD approximation is calculated using the KRONOS code from ONERA, coupling the hypersonic CFD platform CEDRE and the electrical code SATURNE from EDF. A critical parameter of these simulations is the electrical conductivity, which was evaluated using a set of equilibrium calculations with 25 species. Two models were used: local thermodynamic equilibrium and frozen flow. In both cases, chlorine captures a large fraction of free electrons, limiting the electrical conductivity to a value inadequate for thrust vectoring applications. However, when using chlorine-free propergols with 1% in mass of alkali, an MHD thrust vectoring of several degrees was obtained.

  15. Measurements of reactive gaseous rocket injector admittances

    Janardan, B. A.; Daniel, B. R.; Bell, W. A.; Zinn, B. T.


    The paper describes the results of an experimental study of the quantitative determination of the capabilities of the combustion processes associated with coaxial gaseous propellant rocket injectors to drive combustor pressure oscillations. The data, obtained by employing the modified impedance tube technique with compressed air as the oxidizer and acetylene gas as the fuel, describe the frequency dependence of the admittance of the combined injector-combustion process. The measured data are compared with the predictions of the Feiler and Heidmann analytical model utilizing different values for the characteristic combustion time tau sub b. The values of tau sub b which result in a best fit between the measured and predicted data are indicated for different equivalence ratios. It is shown that for the coaxial injector investigated in this study the tau sub b varies between 0.7 and 1.2 msec for equivalence ratios in the range of 0.57 to 1.31. In addition, the experimental data indicate that the tested injector system could drive combustion instabilities over a frequency range that is in qualitative agreement with the predictions of the Feiler and Heidmann model.

  16. Particle Size Distributions Measured in the Stratospheric Plumes of Three Rockets During the ACCENT Missions

    Wiedinmyer, C.; Brock, C. A.; Reeves, J. M.; Ross, M. N.; Schmid, O.; Toohey, D.; Wilson, J. C.


    The global impact of particles emitted by rocket engines on stratospheric ozone is not well understood, mainly due to the lack of comprehensive in situ measurements of the size distributions of these emitted particles. During the Atmospheric Chemistry of Combustion Emissions Near the Tropopause (ACCENT) missions in 1999, the NASA WB-57F aircraft carried the University of Denver N-MASS and FCAS instruments into the stratospheric plumes from three rockets. Size distributions of particles with diameters from 4 to approximately 2000 nm were calculated from the instrument measurements using numerical inversion techniques. The data have been averaged over 30-second intervals. The particle size distributions observed in all of the rocket plumes included a dominant mode near 60 nm diameter, probably composed of alumina particles. A smaller mode at approximately 25 nm, possibly composed of soot particles, was seen in only the plumes of rockets that used liquid oxygen and kerosene as a propellant. Aircraft exhaust emitted by the WB-57F was also sampled; the size distributions within these plumes are consistent with prior measurements in aircraft plumes. The size distributions for all rocket intercepts have been fitted to bimodal, lognormal distributions to provide input for global models of the stratosphere. Our data suggest that previous estimates of the solid rocket motor alumina size distributions may underestimate the alumina surface area emission index, and so underestimate the particle surface area available for heterogeneous chlorine activation reactions in the global stratosphere.

  17. Numerical and experimental analysis of heat transfer in injector plate of hydrogen peroxide hybrid rocket motor

    Cai, Guobiao; Li, Chengen; Tian, Hui


    This paper is aimed to analyze heat transfer in injector plate of hydrogen peroxide hybrid rocket motor by two-dimensional axisymmetric numerical simulations and full-scale firing tests. Long-time working, which is an advantage of hybrid rocket motor over conventional solid rocket motor, puts forward new challenges for thermal protection. Thermal environments of full-scale hybrid rocket motors designed for long-time firing tests are studied through steady-state coupled numerical simulations of flow field and heat transfer in chamber head. The motor adopts 98% hydrogen peroxide (98HP) oxidizer and hydroxyl-terminated poly-butadiene (HTPB) based fuel as the propellants. Simulation results reveal that flowing liquid 98HP in head oxidizer chamber could cool the injector plate of the motor. The cooling of 98HP is similar to the regenerative cooling in liquid rocket engines. However, the temperature of the 98HP in periphery portion of the head oxidizer chamber is higher than its boiling point. In order to prevent the liquid 98HP from unexpected decomposition, a thermal protection method for chamber head utilizing silica-phenolics annular insulating board is proposed. The simulation results show that the annular insulating board could effectively decrease the temperature of the 98HP in head oxidizer chamber. Besides, the thermal protection method for long-time working hydrogen peroxide hybrid rocket motor is verified through full-scale firing tests. The ablation of the insulating board in oxygen-rich environment is also analyzed.

  18. Chaotic dynamics of propeller singing

    YU Dapeng; ZHAO Deyou; WANG Yu


    The system of propeller singing is proved for the first time to have the character of chaotic dynamics through the study of the signal time series. The estimation of the topolog- ical dimension, the confirmation of the number of independent variable and the description of the character of attractor trajectory in reconstructed phase space are implemented during the analysis of the system. The result indicates that the system of propeller singing can be recon- structed by the optional delay time tD = 1, the minimal embedding dimension dE = 8, and the reconstructed topological parameter with the fractional correlation dimension D2 = 5.1579 and the positive maximum Lyapunov exponent λtD=0.0771. The results provide a new approach to the further study of the propeller singing phenomenon.

  19. Cavitation simulation on marine propellers

    Shin, Keun Woo

    is reproduced in the simulation. The overall results suggest the possibility of the cavitation model in the RANS solver to be used for practical applications in propeller design process as a complementary tool to the cavitation tunnel test and the other numerical methods. The outstanding issue for cloudy...... developed in the last decade. They show the potential for the simulation of propeller cavitation with robustness, but they are still to be more proved for practical applications. In the present work, hydrodynamic and numerical characteristics of several cavitation models developed for a viscous flow solver...... are investigated, and one of the cavitation models is verified for the cavitation simulation on marine propellers. Three cavitation models with a vapor transport equation and a cavitation model with a barotropic state law are implemented in the in-house RANS solver, EllipSys. The numerical results for cavitating...

  20. Combustion chemistry of solid propellants

    Baer, A. D.; Ryan, N. W.


    Several studies are described of the chemistry of solid propellant combustion which employed a fast-scanning optical spectrometer. Expanded abstracts are presented for four of the studies which were previously reported. One study of the ignition of composite propellants yielded data which suggested early ammonium perchlorate decomposition and reaction. The results of a study of the spatial distribution of molecular species in flames from uncatalyzed and copper or lead catalyzed double-based propellants support previously published conclusions concerning the site of action of these metal catalysts. A study of the ammonium-perchlorate-polymeric-fuel-binder reaction in thin films, made by use of infrared absorption spectrometry, yielded a characterization of a rapid condensed-phase reaction which is likely important during the ignition transient and the burning process.

  1. Nuclear-thermal rocket thrust transient effects on minimum-fuel lunar trajectories

    Rivas, Matthew L.


    A technically viable option for low-cost minimum-fuel Lunar transfers with short trip times is the use of nuclear thermal rockets. However, little work has been done on the effects the associated thrust transients have on these optimal trajectories. The nominal thrust level of an engine is not immediately reached when the rocket is turned ``on.'' Similarly, when the engine is turned ``off'', the thrust and specific impulse levels decrease over a period of time which is directly related to both the flow effecs of the engine and cooling requirements. This paper presents an analysis of these effects on a typical optimal Lunar transfer. Several different models simulating the transient effects are used. They range from simple ``mass dumps'' to account for the extra required propellant to curve-fits of actual engine characteristics obtained from the NERVA nuclear rocket program.




    Full Text Available Liquid rocket engines have variety of propellant combinations which produces very high specific impulses. It is due to this fact; very high heat fluxes are incident on the combustion chamber and the nozzle walls. In order to deal with these heat fluxes, a wide range of cooling techniques have been employed, out of which a combination of film cooling and regenerative cooling promises to be the most effective one. The present study involves the numerical analysis of combustion in a typical film cooled cryogenic rocket engine thrust chamber considering the combustion of the fuel, heat transfer through the chamber walls and the fluid flow simultaneously. Analysis was done for a typical rocket engine thrust chamber with a single coaxial injector which uses gaseous hydrogen as the fuel and liquid oxygen as the oxidizer.

  3. Laser-propelled ram accelerator

    Sasoh, A. [Tohoku Univ., Sendai (Japan). Inst. of Fluid Science


    The concept of 'laser-propelled ram accelerator (L-RAMAC)' is proposed. Theoretically it is capable of achieving a higher launch speed than that by a chemical ram accelerator because a higher specific energy can be input to the propellant gas. The laser beam is supplied through the muzzle, focused as an annulus behind the base of the projectile. The performance of L-RAMAC is analized based on generalized Rankine-Hugoniot relations, suggesting that a superorbital muzzle speed is achievable out of this device. (orig.)

  4. Characteristics of Five Propellers in Flight

    Crowley, J W , Jr; Mixson, R E


    This investigation was made for the purpose of determining the characteristics of five full-scale propellers in flight. The equipment consisted of five propellers in conjunction with a VE-7 airplane and a Wright E-2 engine. The propellers were of the same diameter and aspect ratio. Four of them differed uniformly in thickness and pitch and the fifth propeller was identical with one of the other four with exception of a change of the airfoil section. The propeller efficiencies measured in flight are found to be consistently lower than those obtained in model tests. It is probable that this is mainly a result of the higher tip speeds used in the full-scale tests. The results show also that because of differences in propeller deflections it is difficult to obtain accurate comparisons of propeller characteristics. From this it is concluded that for accurate comparisons it is necessary to know the propeller pitch angles under actual operating conditions. (author)

  5. 78 FR 78290 - Airworthiness Directives; Dowty Propellers


    ..., Boston Aircraft Certification Office, FAA, Engine & Propeller Directorate, 12 New England Executive Park... Aircraft Certification Office, FAA, Engine & Propeller Directorate, 12 New England Executive Park... Federal Aviation Administration 14 CFR Part 39 RIN 2120-AA64 Airworthiness Directives; Dowty...

  6. Experimental determination of the admittances of aluminized propellants by the impedance tube method

    Baum, J. D.; Daniel, B. R.; Zinn, B. T.


    The adaptation of the impedance tube concept for the determination of the pressure coupled admittances and response functions of burning solid propellants is discussed. The results obtained in experiments with UTP-3001 and UTP-19360 aluminized propellants are presented. It is shown that the admittance Y remains constant during the quasi-steady burn period of a test, indicating constant driving of the gas phase disturbance by the burning solid propellant. The measured real part of the admittance is positive, indicating that the burning aluminized propellant is driving the gas phase oscillations. In addition, the measured high gas phase damping, provided by the aluminum oxide particles in the gas phase, suggests that the latter can significantly increase the damping in unstable solid rockets over the investigated frequency range. Finally, it is shown that the wave structure obtained by numerically solving the impedance tube wave equations which utilize the determined propellant admittance as an initial condition and the determined value of G to describe the gas phase losses is in excellent agreement with the measured wave structure.

  7. The Effect of Solid Constituent Particle Size Distributions on TP-H1148 Propellant Slag

    May, Douglas H.; Miles, William L.; Taylor, David S.; Rackham, Jon L.


    Special aluminum and ammonium perchlorate (AP) particle size distributions were prepared for a matrix of five-inch diameter, center-perforated (CP) motor tests to measure the aluminum oxide slag response in Space Shuttle Reusable Solid Rocket Motor (RSRM) propellant. Previous tests of TP-H1148 propellant in five-inch CP spin motors have shown a correlation between aluminum particle size and generated slag. The motors for this study were cast from thirteen five-gallon propellant mixes which used five particle size levels of aluminum powder, five of unground AP and three of ground AP. Aluminum had the greatest effect on slag formation, the more coarse fractions causing greater slag quantities and larger slag particles. Unground AP had about half the effect of aluminum with the coarser fractions again producing more and larger sized slag particles. The variation in ground AP did not have a significant effect on slag formation. Quench bomb tests showed the same trends as the spin motors, that is, larger aluminum and AP particle size distributions generated larger slag particles leaving the propellant surface. Cured propellant mechanical properties were also impacted by particle size variation.

  8. 76 FR 7101 - Airworthiness Directives; Hamilton Sundstrand Propellers Model 247F Propellers


    ... Aircraft Certification Office, FAA, Engine and Propeller Directorate, 12 New England Executive Park... Propellers Model 247F Propellers AGENCY: Federal Aviation Administration (FAA), DOT. ACTION: Final rule... requires removing affected propeller blades from service. This AD was prompted by reports of blades...

  9. Possibilities for decreasing detection limits of analytical methods for determination of transformation products of unsymmetrical dimethylhydrazine in environmental samples

    Bulat Kenessov


    Full Text Available Most rockets of middle and heavy class launched from Kazakhstan, Russia, China and other countries still use highly toxic unsymmetrical dimethylhydrazine (UDMH as a liquid propellant. Study of migration, distribution and accumulation of UDMH transformation products in environment and human health impact assessment of space rocket activity are currently complicated due to the absence of analytical methods allowing detection of trace concentrations of these compounds in analyzed samples. This paper reviews methods and approaches, which can be applied for development of such methods. Detection limits at a part-per-trillion (ppt level may be achieved using most selective and sensitive methods based on gas or liquid chromatography in combination of tandem or high-resolution mass spectrometry. In addition, 1000-fold concentration of samples or integrated sample preparation methods, e.g., dynamic headspace extraction, are required. Special attention during development and application of such methods must be paid to purity of laboratory air, reagents, glassware and analytical instruments.

  10. Thermal decomposition of phase-stabilised ammonium nitrate (PSAM), HTPB based propellants. The effect of iron(III)oxide burning-rate catalyst

    Carvalheira, P.; Gadiot, G.M.H.J.L.; Klerk, W.P.C. de


    Phase-stabilised ammonium nitrate (PSAN) and hydroxyl-terminated polybutadiene (HTPB) are the main ingredients of propellants used with success in some pyrotechnic igniter components of the VULCAIN liquid rocket engine for the ARIANE 5. Small amounts of selected additives play an important role in s

  11. British used Congreve Rockets to Attack Napoleon


    Sir William Congreve developed a rocket with a range of about 9,000 feet. The incendiary rocket used black powder, an iron case, and a 16-foot guide stick. In 1806, British used Congreve rockets to attack Napoleon's headquarters in France. In 1807, Congreve directed a rocket attack against Copenhagen.

  12. Unconventional Propeller Tip Design - Hydrodynamic Study

    Ommundsen, Andreas


    Winglets have been successful in the aircraft industry, but have not yet seen widespread use on marine propellers. Three different propellers (one conventional and two equipped with winglets) have been modelled and analysed with the CFD software STAR-CCM+. The winglet propellers achieved up to 40 % greater thrust than the conventional propeller at the operational propulsion point, but the torque increased even more - meaning that the overall open water efficiency was reduced by as much as...

  13. 14 CFR 25.925 - Propeller clearance.


    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propeller clearance. 25.925 Section 25.925... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant General § 25.925 Propeller clearance. Unless smaller clearances are substantiated, propeller clearances with the airplane at maximum weight, with the most...

  14. Innovative boron nitride-doped propellants

    Thelma Manning


    Full Text Available The U.S. military has a need for more powerful propellants with balanced/stoichiometric amounts of fuel and oxidants. However, balanced and more powerful propellants lead to accelerated gun barrel erosion and markedly shortened useful barrel life. Boron nitride (BN is an interesting potential additive for propellants that could reduce gun wear effects in advanced propellants (US patent pending 2015-026P. Hexagonal boron nitride is a good lubricant that can provide wear resistance and lower flame temperatures for gun barrels. Further, boron can dope steel, which drastically improves its strength and wear resistance, and can block the formation of softer carbides. A scalable synthesis method for producing boron nitride nano-particles that can be readily dispersed into propellants has been developed. Even dispersion of the nano-particles in a double-base propellant has been demonstrated using a solvent-based processing approach. Stability of a composite propellant with the BN additive was verified. In this paper, results from propellant testing of boron nitride nano-composite propellants are presented, including closed bomb and wear and erosion testing. Detailed characterization of the erosion tester substrates before and after firing was obtained by electron microscopy, inductively coupled plasma and x-ray photoelectron spectroscopy. This promising boron nitride additive shows the ability to improve gun wear and erosion resistance without any destabilizing effects to the propellant. Potential applications could include less erosive propellants in propellant ammunition for large, medium and small diameter fire arms.

  15. Innovative boron nitride-doped propellants

    Thelma MANNING; Henry GRAU; Paul MATTER; Michael BEACHY; Christopher HOLT; Samuel SOPOK; Richard FIELD; Kenneth KLINGAMAN; Michael FAIR; John BOLOGNINI; Robin CROWNOVER; Carlton P. ADAM; Viral PANCHAL; Eugene ROZUMOV


    The U.S. military has a need for more powerful propellants with balanced/stoichiometric amounts of fuel and oxidants. However, balanced and more powerful propellants lead to accelerated gun barrel erosion and markedly shortened useful barrel life. Boron nitride (BN) is an interesting potential additive for propellants that could reduce gun wear effects in advanced propellants (US patent pending 2015-026P). Hexagonal boron nitride is a good lubricant that can provide wear resistance and lower flame temperatures for gun barrels. Further, boron can dope steel, which drastically improves its strength and wear resistance, and can block the formation of softer carbides. A scalable synthesis method for producing boron nitride nano-particles that can be readily dispersed into propellants has been developed. Even dispersion of the nano-particles in a double-base propellant has been demonstrated using a solvent-based processing approach. Stability of a composite propellant with the BN additive was verified. In this paper, results from propellant testing of boron nitride nano-composite propellants are presented, including closed bomb and wear and erosion testing. Detailed characterization of the erosion tester substrates before and after firing was obtained by electron microscopy, inductively coupled plasma and x-ray photoelectron spectroscopy. This promising boron nitride additive shows the ability to improve gun wear and erosion resistance without any destabilizing effects to the propellant. Potential applications could include less erosive propellants in propellant ammunition for large, medium and small diameter fire arms.

  16. 14 CFR 25.929 - Propeller deicing.


    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propeller deicing. 25.929 Section 25.929... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant General § 25.929 Propeller deicing. (a) For airplanes... combustible fluid is used for propeller deicing, §§ 25.1181 through 25.1185 and 25.1189 apply....

  17. 14 CFR 25.905 - Propellers.


    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propellers. 25.905 Section 25.905 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant General § 25.905 Propellers. (a) Each propeller must...

  18. 14 CFR 35.2 - Propeller configuration.


    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Propeller configuration. 35.2 Section 35.2 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: PROPELLERS General § 35.2 Propeller configuration. The applicant must provide a list of all...

  19. CS/PVA-GA渗透蒸发膜分离偏二甲肼/水体系的实验研究%The experimental study of the CS/PVA-GA pervaporation membrane in UDMH/H2O system

    张有智; 朱建伟; 曹佳玉; 王煊军


    Chitosan (CS) was blended with poly vinyl alcohol (PVA) by a certain percentage, and then was cross-linked by glutaraldehyde (GA). The CS/PVA-GA mixed associated membrane was prepared. Mixed associated membrane structure,mechanical properties as well as pervaporation performance in UD- MH/H20 system were studied. The results showed that the blending of CS and PVA and GA cross-linked form a dense spatial network structure. With the PVA content increasing, the tensile strength of the membrane was decreased,while the rupturing and extending rate was increased. The feed concentration,temperature , film thickness and other factors have great impact on the membrane pervaporation performance. When UDMH of the feed solution mass fraction reaches 50% , temperature is 20℃, the permeate side pressure is 100 Pa and film thickness is about 25 μm,the CS/PVA-GA membrane separation factor reaches 243,the permeate flux, can be of 167 g/(m2 · h).%将壳聚糖(CS)和聚乙烯醇(PVA)按一定比例共混,再以戊二醛(GA)交联,制备CS/PVA-GA共混交联膜,研究了共混交联膜的结构、力学性能以及在偏二甲肼/水体系中的渗透蒸发脱水性能.结果表明,CS与PVA共混、GA交联,形成了致密的空间网状结构,随PVA含量增加,膜的抗张强度降低,而断裂伸长率增加.料液浓度、温度、膜厚度等因素对膜的渗透蒸发性能有较大影响,当料液中偏二甲肼的质量分数为50%,温度为20℃,渗透侧压力为100 Pa,膜厚约为25μm时,CS/PVA-GA共混交联膜的分离因子达到243,渗透通量可达167g/(m2·h).

  20. Effect of Chamber Pressurization Rate on Combustion and Propagation of Solid Propellant Cracks

    Yuan, Wei-Lan; Wei, Shen; Yuan, Shu-Shen


    area of the propellant grain satisfies the designed value. But cracks in propellant grain can be generated during manufacture, storage, handing and so on. The cracks can provide additional surface area for combustion. The additional combustion may significantly deviate the performance of the rocket motor from the designed conditions, even lead to explosive catastrophe. Therefore a thorough study on the combustion, propagation and fracture of solid propellant cracks must be conducted. This paper takes an isolated propellant crack as the object and studies the effect of chamber pressurization rate on the combustion, propagation and fracture of the crack by experiment and theoretical calculation. deformable, the burning inside a solid propellant crack is a coupling of solid mechanics and combustion dynamics. In this paper, a theoretical model describing the combustion, propagation and fracture of the crack was formulated and solved numerically. The interaction of structural deformation and combustion process was included in the theoretical model. The conservation equations for compressible fluid flow, the equation of state for perfect gas, the heat conducting equation for the solid-phase, constitutive equation for propellant, J-integral fracture criterion and so on are used in the model. The convective burning inside the crack and the propagation and fracture of the crack were numerically studied by solving the set of nonlinear, inhomogeneous gas-phase governing equations and solid-phase equations. On the other hand, the combustion experiments for propellant specimens with a precut crack were conducted by RTR system. Predicted results are in good agreement with experimental data, which validates the reasonableness of the theoretical model. Both theoretical and experimental results indicate that the chamber pressurization rate has strong effects on the convective burning in the crack, crack fracture initiation and fracture pattern.

  1. Not just rocket science

    MacAdam, S.; Anderson, R. [Celan Energy Systems, Rancho Cordova, CA (United States)


    The paper explains a different take on oxyfuel combustion. Clean Energy Systems (CES) has integrated aerospace technology into conventional power systems, creating a zero-emission power generation technology that has some advantages over other similar approaches. When using coal as a feedstock, the CES process burns syngas rather than raw coal. The process uses recycled water and steam to moderate the temperature, instead of recycled CO{sub 2}. With no air ingress, the CES process produces very pure CO{sub 2}. This makes it possible to capture over 99% of the CO{sub 2} resulting from combustion. CES uses the combustion products to drive the turbines, rather than indirectly raising steam for steam turbines, as in the oxyfuel process used by companies such as Vattenfall. The core of the process is a high-pressure oxy-combustor adapted from rocket engine technology. This combustor burns gaseous or liquid fuels with gaseous oxygen in the presence of water. Fuels include natural gas, coal or coke-derived synthesis gas, landfill and biodigester gases, glycerine solutions and oil/water emulsion. 2 figs.

  2. Plasma ignition of LOVA propellants

    Driel, C.A. van; Boluijt, A.G.; Schilt, A.


    Ignition experiments were performed using a gun simulator which is equipped with a burst disk. This equipment facilitates the application of propellant loading densities which are comparable to those applied in regular ammunitions. For this study the gun simulator was equipped with a plasma jet igni

  3. Novel sequences propel familiar folds.

    Jawad, Zahra; Paoli, Massimo


    Recent structure determinations have made new additions to a set of strikingly different sequences that give rise to the same topology. Proteins with a beta propeller fold are characterized by extreme sequence diversity despite the similarity in their three-dimensional structures. Several fold predictions, based in part on sequence repeats thought to match modular beta sheets, have been proved correct.

  4. Plasma ignition of LOVA propellants

    Driel, C.A. van; Boluijt, A.G.; Schilt, A.


    Ignition experiments were performed using a gun simulator which is equipped with a burst disk. This equipment facilitates the application of propellant loading densities which are comparable to those applied in regular ammunitions. For this study the gun simulator was equipped with a plasma jet

  5. The propeller and the frog

    Pan, Margaret


    "Propellers" in planetary rings are disturbances in ring material excited by moonlets that open only partial gaps. We describe a new type of co-orbital resonance that can explain the observed non-Keplerian motions of propellers. The resonance is between the moonlet underlying the propeller, and co-orbiting ring particles downstream of the moonlet where the gap closes. The moonlet librates within the gap about an equilibrium point established by co-orbiting material and stabilized by the Coriolis force. In the limit of small libration amplitude, the libration period scales linearly with the gap azimuthal width and inversely as the square root of the co-orbital mass. The new resonance recalls but is distinct from conventional horseshoe and tadpole orbits; we call it the "frog" resonance, after the relevant term in equine hoof anatomy. For a ring surface density and gap geometry appropriate for the propeller Bl\\'eriot in Saturn's A ring, our theory predicts a libration period of ~4 years, similar to the ~3.7 yea...

  6. A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

    Betts, Erin M.; Frederick, Robert A., Jr.


    This is a comprehensive systems study to examine and evaluate throttling capabilities of liquid rocket engines. The focus of this study is on engine components, and how the interactions of these components are considered for throttling applications. First, an assessment of space mission requirements is performed to determine what applications require engine throttling. A background on liquid rocket engine throttling is provided, along with the basic equations that are used to predict performance. Three engines are discussed that have successfully demonstrated throttling. Next, the engine system is broken down into components to discuss special considerations that need to be made for engine throttling. This study focuses on liquid rocket engines that have demonstrated operational capability on American space launch vehicles, starting with the Apollo vehicle engines and ending with current technology demonstrations. Both deep throttling and shallow throttling engines are discussed. Boost and sustainer engines have demonstrated throttling from 17% to 100% thrust, while upper stage and lunar lander engines have demonstrated throttling in excess of 10% to 100% thrust. The key difficulty in throttling liquid rocket engines is maintaining an adequate pressure drop across the injector, which is necessary to provide propellant atomization and mixing. For the combustion chamber, cooling can be an issue at low thrust levels. For turbomachinery, the primary considerations are to avoid cavitation, stall, surge, and to consider bearing leakage flows, rotordynamics, and structural dynamics. For valves, it is necessary to design valves and actuators that can achieve accurate flow control at all thrust levels. It is also important to assess the amount of nozzle flow separation that can be tolerated at low thrust levels for ground testing.

  7. Environmental fate and transport of nitroglycerin from propellant residues at firing positions in the unsaturated zone

    Bellavance-Godin, A. [Institut National de la Recherche Scientifique, Quebec, PQ (Canada). Eau, Terre et Environnement; Martel, R. [Institut National de la Recherche Scientifique, Varennes, PQ (Canada). Eau, Terre et Environnement, Earth Sciences


    In response to environmental concerns, the Canadian Forces Base (CFB) have initiated studies to better evaluate the impact of various military activities. This paper presented the results of a study in which the fate of propellant residues on large soil columns was investigated. The sites selected for the study were the antitank ranges at Garrison Valcartier, Quebec and those at the CFB Petawawa, Ontario. The shoulder rockets fired on those ranges were propelled by solid propellants based on a nitrocellulose matrix in which nitroglycerine and ammonium perchlorate were dispersed as oxidizer and energetic materials. Propellant residues accumulated in the surface soils because the combustion processes in the rockets was incomplete. This study evaluated the contaminants transport through the unsaturated zone. Sampling was conducted in 2 steps. The first involved collecting uncontaminated soil samples representative of the geological formations of the 2 sites. The second step involved collecting soils containing high levels of propellant residues behind antitank firing positions, which was later spread across the surface of the uncontaminated soil columns and which were representative of the contaminated zone. The soils were watered in the laboratory following the precipitation patterns of the respective regions and interstitial water output of the columns was also sampled. The compounds of interest were nitroglycerine and its degradation metabolites, dinitroglycerine, mononitroglycerine and nitrates as well as perchlorate and bromides. Results presented high concentrations of nitrites, nitrates and perchlorates. Both the NG and its degradation products were monitored using a newly developed analytical method that provides for a better understanding of NG degradation pathways in anaerobic conditions. 12 refs., 3 tabs., 12 figs.

  8. Liquid Bismuth Propellant Flow Sensor

    Polzin, Kurt A.; Stanojev, B. J.; Korman, V.


    Quantifying the propellant mass flow rate in liquid bismuth-fed electric propulsion systems has two challenging facets. First, the flow sensors must be capable of providing a resolvable measurement at propellant mass flow rates on the order of 10 mg/see with and uncertainty of less that 5%. The second challenge has to do with the fact that the materials from which the flow sensors are fabricated must be capable of resisting any of the corrosive effects associated with the high-temperature propellant. The measurement itself is necessary in order to properly assess the performance (thrust efficiency, Isp) of thruster systems in the laboratory environment. The hotspot sensor[I] has been designed to provide the bismuth propellant mass flow rate measurement. In the hotspot sensor, a pulse of thermal energy (derived from a current pulse and associated joule heating) is applied near the inlet of the sensor. The flow is "tagged" with a thermal feature that is convected downstream by the flowing liquid metal. Downstream, a temperature measurement is performed to detect a "ripple" in the local temperature associated with the passing "hotspot" in the propellant. By measuring the time between the upstream generation and downstream detection of the thermal feature, the flow speed can be calculated using a "time of flight" analysis. In addition, the system can be calibrated by measuring the accumulated mass exiting the system as a-function of time and correlating this with the time it takes the hotspot to convect through the sensor. The primary advantage of this technique is that it doesn't depend on an absolute measurement of temperature but, instead, relies on the observation of thermal features. This makes the technique insensitive to other externally generated thermal fluctuations. In this paper, we describe experiments performed using the hotspot flow sensor aimed at quantifying the resolution of the sensor technology. Propellant is expelled onto an electronic scale to

  9. Drag and Torque on Locked Screw Propeller

    Tomasz Tabaczek


    Full Text Available Few data on drag and torque on locked propeller towed in water are available in literature. Those data refer to propellers of specific geometry (number of blades, blade area, pitch and skew of blades. The estimation of drag and torque of an arbitrary propeller considered in analysis of ship resistance or propulsion is laborious. The authors collected and reviewed test data available in the literature. Based on collected data there were developed the empirical formulae for estimation of hydrodynamic drag and torque acting on locked screw propeller. Supplementary CFD computations were carried out in order to prove the applicability of the formulae to modern moderately skewed screw propellers.

  10. High-Speed Propeller for Aircraft

    Sagerser, D. A.; Gatzen, B. S.


    Engine efficiency increased. Propeller blades required to be quite thin and highly swept to minimize compressibility losses and propeller noise during high-speed cruise. Use of 8 or 10 blades with highpropeller-power loading allows overall propeller diameter to be kept relatively small. Area-ruled spinner and integrated nacelle shape reduce compressibility losses in propeller hub region. Finally, large modern turboshaft engine and gearbox provide power to advanced propeller. Fuel savings of 30 to 50 percent over present systems anticipated. Propfan system adaptable to number of applications, such as highspeed (subsonic) business and general-aviation aircraft, and military aircraft including V/STOL.

  11. CFD Simulation of Liquid Rocket Engine Injectors

    Farmer, Richard; Cheng, Gary; Chen, Yen-Sen; Garcia, Roberto (Technical Monitor)


    these investigators to be very valuable for code validation because combustion kinetics, turbulence models and atomization models based on low pressure experiments of hydrogen air combustion do not adequately verify analytical or CFD submodels which are necessary to simulate rocket engine combustion. We wish to emphasize that the simulations which we prepared for this meeting are meant to test the accuracy of the approximations used in our general purpose spray combustion models, rather than represent a definitive analysis of each of the experiments which were conducted. Our goal is to accurately predict local temperatures and mixture ratios in rocket engines; hence predicting individual experiments is used only for code validation. To replace the conventional JANNAF standard axisymmetric finite-rate (TDK) computer code 2 for performance prediction with CFD cases, such codes must posses two features. Firstly, they must be as easy to use and of comparable run times for conventional performance predictions. Secondly, they must provide more detailed predictions of the flowfields near the injector face. Specifically, they must accurately predict the convective mixing of injected liquid propellants in terms of the injector element configurations.

  12. Studies on composite solid propellant with tri-modal ammonium perchlorate containing an ultrafine fraction

    K.V. Suresh Babu


    composite solid propellant is prepared by using burn rate modifiers Copper chromite and Iron oxide. Addition of Copper chromite and Iron oxide has enhanced the burn rate of tri-modal AP based composite solid propellant. The catalytic propensity of copper chromite is higher than that of iron oxide. The pressure exponent increased with the catalyst concentration and the values obtained are compatible for solid rocket motor applications.

  13. Facility for cold flow testing of solid rocket motor models

    Bacchus, D. L.; Hill, O. E.; Whitesides, R. Harold


    A new cold flow test facility was designed and constructed at NASA Marshall Space Flight Center for the purpose of characterizing the flow field in the port and nozzle of solid propellant rocket motors (SRM's). A National Advisory Committee was established to include representatives from industry, government agencies, and universities to guide the establishment of design and instrumentation requirements for the new facility. This facility design includes the basic components of air storage tanks, heater, submicron filter, quiet control valve, venturi, model inlet plenum chamber, solid rocket motor (SRM) model, exhaust diffuser, and exhaust silencer. The facility was designed to accommodate a wide range of motor types and sizes from small tactical motors to large space launch boosters. This facility has the unique capability of testing ten percent scale models of large boosters such as the new Advanced Solid Rocket Motor (ASRM), at full scale motor Reynolds numbers. Previous investigators have established the validity of studying basic features of solid rocket motor development programs include the acquisition of data to (1) directly evaluate and optimize the design configuration of the propellant grain, insulation, and nozzle; and (2) provide data for validation of the computational fluid dynamics, (CFD), analysis codes and the performance analysis codes. A facility checkout model was designed, constructed, and utilized to evaluate the performance characteristics of the new facility. This model consists of a cylindrical chamber and converging/diverging nozzle with appropriate manifolding to connect it to the facility air supply. It was designed using chamber and nozzle dimensions to simulate the flow in a 10 percent scale model of the ASRM. The checkout model was recently tested over the entire range of facility flow conditions which include flow rates from 9.07 to 145 kg/sec (20 to 320 Ibm/sec) and supply pressure from 5.17 x 10 exp 5 to 8.27 x 10 exp 6 Pa. The

  14. Solid Propellant Microthruster Design, Fabrication, and Testing for Nanosatellites

    Sathiyanathan, Kartheephan

    This thesis describes the design, fabrication, and testing of a solid propellant microthruster (SPM), which is a two-dimensional matrix of millimeter-sized rockets each capable of delivering millinewtons of thrust and millinewton-seconds of impulse to perform fine orbit and attitude corrections. The SPM is a potential payload for nanosatellites to increase spacecraft maneuverability and is constrained by strict mass, volume, and power requirements. The dimensions of the SPM in the millimeter-scale result in a number of scaling issues that need consideration such as a low Reynolds number, high heat loss, thermal and radical quenching, and incomplete combustion. The design of the SPM, engineered to address these issues, is outlined. The SPM fabrication using low-cost commercial off-the-shelf materials and standard micromachining is presented. The selection of a suitable propellant and its customization are described. Experimental results of SPM firing to demonstrate successful ignition and sustained combustion are presented for three configurations: nozzleless, sonic nozzle, and supersonic nozzle. The SPM is tested using a ballistic pendulum thrust stand. Impulse and thrust values are calculated and presented. The performance values of the SPM are found to be consistent with existing designs.

  15. The thermal decomposition behavior of ammonium perchlorate and of an ammonium-perchlorate-based composite propellant

    Behrens, R.; Minier, L.


    The thermal decomposition of ammonium perchlorate (AP) and ammonium-perchlorate-based composite propellants is studied using the simultaneous thermogravimetric modulated beam mass spectrometry (STMBMS) technique. The main objective of the present work is to evaluate whether the STMBMS can provide new data on these materials that will have sufficient detail on the reaction mechanisms and associated reaction kinetics to permit creation of a detailed model of the thermal decomposition process. Such a model is a necessary ingredient to engineering models of ignition and slow-cookoff for these AP-based composite propellants. Results show that the decomposition of pure AP is controlled by two processes. One occurs at lower temperatures (240 to 270 C), produces mainly H{sub 2}O, O{sub 2}, Cl{sub 2}, N{sub 2}O and HCl, and is shown to occur in the solid phase within the AP particles. 200{micro} diameter AP particles undergo 25% decomposition in the solid phase, whereas 20{micro} diameter AP particles undergo only 13% decomposition. The second process is dissociative sublimation of AP to NH{sub 3} + HClO{sub 4} followed by the decomposition of, and reaction between, these two products in the gas phase. The dissociative sublimation process occurs over the entire temperature range of AP decomposition, but only becomes dominant at temperatures above those for the solid-phase decomposition. AP-based composite propellants are used extensively in both small tactical rocket motors and large strategic rocket systems.

  16. Development and Short-Range Testing of a 100 kW Side-Illuminated Millimeter-Wave Thermal Rocket

    Bruccoleri, Alexander; Eilers, James A.; Lambot, Thomas; Parkin, Kevin


    The objective of the phase described here of the Millimeter-Wave Thermal Launch System (MTLS) Project was to launch a small thermal rocket into the air using millimeter waves. The preliminary results of the first MTLS flight vehicle launches are presented in this work. The design and construction of a small thermal rocket with a planar ceramic heat exchanger mounted along the axis of the rocket is described. The heat exchanger was illuminated from the side by a millimeter-wave beam and fed propellant from above via a small tank containing high pressure argon or nitrogen. Short-range tests where the rocket was launched, tracked, and heated with the beam are described. The rockets were approximately 1.5 meters in length and 65 millimeters in diameter, with a liftoff mass of 1.8 kilograms. The rocket airframes were coated in aluminum and had a parachute recovery system activated via a timer and Pyrodex. At the rocket heat exchanger, the beam distance was 40 meters with a peak power intensity of 77 watts per square centimeter. and a total power of 32 kilowatts in a 30 centimeter diameter circle. An altitude of approximately 10 meters was achieved. Recommendations for improvements are discussed.

  17. First Stage Solid Propellant Multiply Debris Thermal Analysis

    Toleman, Benjamin M.


    Destruction of a solid rocket stage of a launch vehicle can create a thermal radiation hazard for an aborting crew module. This hazard was assessed for the Constellation Program (Cx) crew and launch vehicle concept. For this concept, if an abort was initiated in first stage flight, the Crew Module (CM) will separate and be pulled away from the malfunctioning launch vehicle via a Launch Abort System (LAS). Having aborted the mission, the launch vehicle will likely be destroyed via a Flight Termination System (FTS) in order to prevent it from errantly traversing back over land and posing a risk to the public. The resulting launch vehicle debris field, composed primarily of first stage solid propellant, poses a threat to the CM. The harsh radiative thermal environment, caused by surrounding burning propellant debris, may lead to CM parachute failure. A methodology, detailed herein, has been developed to address this concern and to quantify the risk of first stage propellant debris leading to the thermal demise of the CM parachutes. Utilizing basic thermal radiation principles, a software program was developed to calculate parachute temperature as a function of time for a given abort trajectory and debris piece trajectory set. Two test cases, considered worst case aborts with regard to launch vehicle debris environments, were analyzed using the simulation: an abort declared at Mach 1 and an abort declared at maximum dynamic pressure (Max Q). For both cases, the resulting temperature profiles indicated that thermal limits for the parachutes were not exceeded. However, short duration close encounters by single debris pieces did have a significant effect on parachute temperature. Therefore while these two test cases did not indicate exceedance of thermal limits, in order to quantify the risk of parachute failure due to radiative effects from the abort environment, a more thorough probability-based analysis using the methodology demonstrated herein must be performed.

  18. First Stage Solid Propellant Multi Debris Thermal Analysis

    Toleman, Benjamin M.


    The crew launch vehicle considered for the Constellation (Cx) Program utilizes a first stage solid rocket motor. If an abort is initiated in first stage flight the Crew Module (CM) will separate and be pulled away from the launch vehicle via a Launch Abort System (LAS) in order to safely and quickly carry the crew away from the malfunction launch vehicle. Having aborted the mission, the launch vehicle will likely be destroyed via a Flight Termination System (FTS) in order to prevent it from errantly traversing back over land and posing a risk to the public. The resulting launch vehicle debris field, composed primarily of first stage solid propellant, poses a threat to the CM. The harsh radiative thermal environment induced by surrounding burning propellant debris may lead to CM parachute failure. A methodology, detailed herein, has been developed to address this concern and quantify the risk of first stage propellant debris leading to radiative thermal demise of the CM parachutes. Utilizing basic thermal radiation principles, a software program was developed to calculate parachute temperature as a function of time for a given abort trajectory and debris piece trajectory set. Two test cases, considered worst-case aborts with regard to launch vehicle debris environments, were analyzed using the simulation: an abort declared at Mach 1 and an abort declared at maximum dynamic pressure (Max Q). For both cases, the resulting temperature profiles indicated that thermal limits for the parachutes were not exceeded. However, short duration close encounters by single debris pieces did have a significant effect on parachute temperature, with magnitudes on the order of 10 s of degrees Fahrenheit. Therefore while these two test cases did not indicate exceedance of thermal limits, in order to quantify the risk of parachute failure due to radiative effects from the abort environment, a more thorough probability-based analysis using the methodology demonstrated herein must be

  19. Design Technique for the High-Boiling Propellant Storage and Preparation Facility at the Cosmodrome «Vostochny»

    O. E. Denisov


    Full Text Available The offered project of storage facility allows us to simplify and unitise the ground-based infrastructure objects. The storage facility implements a full preparatory cycle of the propellant components (PC in all parameters. Another problem the developers of complexes of groundbased equipment face now is bulk receipt of PC from manufacturer. The tanks of launch complexes cannot accept such volumes of propellant. It proves that there is a need to create a storage facility. The facility solves problems concerning the components receipt, temperature preparation, moisture content (drying, gas content, and supply to consumers. For preparation the perspective technologies with low power consumption are used.Receiving the propellant from the dispensing platform is carried out via filters of rough cleaning. Transfer from transport tankage goes using a pump. The received product passes through a gas separator to clean technological gas impurity.To prepare propellant temperature, a technology of cryogenic bubbling by boiling nitrogen is chosen. To improve efficiency of cryogenic bubbling it is advised to use the specialized capacities. Railway dimensions, admissible for the trainload goods across the railroads of Siberia and the Far East, define their sizes.As a drying technology and a gas content preparation the preliminary propellant filtration using vertical electro-separators is chosen to save a space. The chamber vertical electroseparators allow 2 — 3 times increase of dehydration capacity.The article presents calculations to prove that using the chosen cooling and drying technologies is efficient.Prepared PC can be supplied:• to transport-fueling containers (TFC with the subsequent transportation to the launch complexes either by the railway or by road;• to mobile fuelling tanks, which feed rocket-carrier tanks on arrival at the blast-off;• to transport capacities for transportation to the object outside the cosmodrome (spaceport;• directly

  20. The 260: The Largest Solid Rocket Motor Ever Tested

    Crimmins, P.; Cousineau, M.; Rogers, C.; Shell, V.


    Aerojet in the mid 1960s, under contract to NASA, built and static hot fire tested the largest solid rocket motor (SRM) in history for the purpose of demonstrating the feasibility of utilizing large SRMs for space exploration. This program successfully fabricated two high strength steel chambers, loaded each with approximately 1,68 million pounds of propellant, and static test fired these giants with their nozzles up from an underground silo located adjacent to the Florida everglades. Maximum thrust and total impulse in excess of 5,000,000 lbf and 3,470,000,000 lbf-sec were achieved. Flames from the second firing, conducted at night, were seen over eighty miles away. For comparative purposes: the thrust developed was nearly 100 times that of a Minuteman III second stage and the 260 in.-dia cross-section was over 3 times that of the Space Shuttle SRM.

  1. Investigation of the cooling film distribution in liquid rocket engine

    Luís Antonio Silva


    Full Text Available This study presents the results of the investigation of a cooling method widely used in the combustion chambers, which is called cooling film, and it is applied to a liquid rocket engine that uses as propellants liquid oxygen and kerosene. Starting from an engine cooling, whose film is formed through the fuel spray guns positioned on the periphery of the injection system, the film was experimentally examined, it is formed by liquid that seeped through the inner wall of the combustion chamber. The parameter used for validation and refinement of the theoretical penetration of the film was cooling, as this parameter is of paramount importance to obtain an efficient thermal protection inside the combustion chamber. Cold tests confirmed a penetrating cold enough cooling of the film for the length of the combustion chamber of the studied engine.

  2. NERVA-derived rocket module for solar system exploration

    Zweig, Herbert R.; Cooper, Martin H.


    A 50,000 pound thrust nuclear thermal rocket engine module concept based on Rover/NERVA technology is presented. Key engine features selected for reliability and safety have been integrated into this concept to provide 4.5 hours of full-thrust operation at a specific impulse of over 850 seconds. Those features include a single turbopump with an expander turbine, tank-head start in space, open-cycle decay heat removal with minimal loss of propellant, reactor cold end axial reflector, and tie tubes which are insulated end which are not used to provide turbine power. The tie tube configuration complements the open cycle decay heat removal concept. Retractable safety rods for water immersion subcriticality were considered and a design concept was developed. Other important safety issues were identified, and their method for accommodation in the design were considered for future implementation.

  3. VSB-30 sounding rocket: history of flight performance

    Wolfgang Jung


    Full Text Available The VSB-30 vehicle is a two-stage, unguided, rail launched sounding rocket, consisting of two solid propellant motors, payload, with recovery and service system. By the end of 2010, ten vehicles had already been launched, three from Brazil (Alcântara and seven from Sweden (Esrange. The objective of this paper is to give an overview of the main characteristics of the first ten flights of the VSB-30, with emphasis on performance and trajectory data. The circular 3σ dispersion area for payload impact point has around 50 km of radius. In most launchings of such vehicle, the impact of the payload fell within 2 sigma. This provides the possibility for further studies to decrease the area of dispersion from the impact point.

  4. Design and Experimental Study on Spinning Solid Rocket Motor

    Xue, Heng; Jiang, Chunlan; Wang, Zaicheng

    The study on spinning solid rocket motor (SRM) which used as power plant of twice throwing structure of aerial submunition was introduced. This kind of SRM which with the structure of tangential multi-nozzle consists of a combustion chamber, propellant charge, 4 tangential nozzles, ignition device, etc. Grain design, structure design and prediction of interior ballistic performance were described, and problem which need mainly considered in design were analyzed comprehensively. Finally, in order to research working performance of the SRM, measure pressure-time curve and its speed, static test and dynamic test were conducted respectively. And then calculated values and experimental data were compared and analyzed. The results indicate that the designed motor operates normally, and the stable performance of interior ballistic meet demands. And experimental results have the guidance meaning for the pre-research design of SRM.

  5. Materials for advanced rocket engine turbopump turbine blades

    Chandler, W. T.


    A study program was conducted to identify those materials that will provide the greatest benefits as turbine blades for advanced liquid propellant rocket engine turbines and to prepare technology plans for the development of those materials for use in the 1990 through 1995 period. The candidate materials were selected from six classes of materials: single-crystal (SC) superalloys, oxide dispersion-strengthened (ODS) superalloys, rapid solidification processed (RSP) superalloys, directionally solidified eutectic (DSE) superalloys, fiber-reinforced superalloy (FRS) composites, and ceramics. Properties of materials from the six classes were compiled and evaluated and property improvements were projected approximately 5 years into the future for advanced versions of materials in each of the six classes.

  6. Rocket Science 101 Interactive Educational Program

    Armstrong, Dennis; Funkhouse, Deborah; DiMarzio, Donald


    To better educate the public on the basic design of NASA s current mission rockets, Rocket Science 101 software has been developed as an interactive program designed to retain a user s attention and to teach about basic rocket parts. This program also has helped to expand NASA's presence on the Web regarding educating the public about the Agency s goals and accomplishments. The software was designed using Macromedia s Flash 8. It allows the user to select which type of rocket they want to learn about, interact with the basic parts, assemble the parts to create the whole rocket, and then review the basic flight profile of the rocket they have built.

  7. Multi-propeller drive system

    Belenger, Robert V.


    A multipropeller drive system having a single input shaft for connection to an engine system, a differential gear assembly for dividing the driving force from the input drive shaft between a pair of output shafts, and a pair of laterally spaced propellers driven by the output shafts of the differential gear assembly is disclosed. The differential gear assembly operates in a manner wherein one output shaft, if required, is permitted to revolve at a different rate than the other output shaft. A pair of brake mechanisms acting on the output shafts of the differential gear assembly enable an operator to control the rotational speed of the respective propellers without modifying the engine speed or transmission settings.

  8. Characterization and Analyses of Valves, Feed Lines and Tanks used in Propellant Delivery Systems at NASA SSC

    Ryan, Harry M.; Coote, David J.; Ahuja, Vineet; Hosangadi, Ashvin


    Accurate modeling of liquid rocket engine test processes involves assessing critical fluid mechanic and heat and mass transfer mechanisms within a cryogenic environment, and accurately modeling fluid properties such as vapor pressure and liquid and gas densities as a function of pressure and temperature. The Engineering and Science Directorate at the NASA John C. Stennis Space Center has developed and implemented such analytic models and analysis processes that have been used over a broad range of thermodynamic systems and resulted in substantial improvements in rocket propulsion testing services. In this paper, we offer an overview of the analyses techniques used to simulate pressurization and propellant fluid systems associated with the test stands at the NASA John C. Stennis Space Center. More specifically, examples of the global performance (one-dimensional) of a propellant system are provided as predicted using the Rocket Propulsion Test Analysis (RPTA) model. Computational fluid dynamic (CFD) analyses utilizing multi-element, unstructured, moving grid capability of complex cryogenic feed ducts, transient valve operation, and pressurization and mixing in propellant tanks are provided as well.

  9. Kinetics Modeling of Hypergolic Propellants


    reaction OH + NO + M −−→ HONO + M as a function of temperature and pressure in the presence of argon, sf6 , and n2 bath gas . Chemical physics, 171(1-2...understand fundamental processes such as gas phase ignition, vaporization and liquid phase chemistry for characterizing ignition. Such understanding will be...critical for future design efforts targeting rapidly repeatable cyclic ignition of these propellants. Three test cases are considered: gas and liquid

  10. Self-Propelled Leidenfrost Droplets

    Linke, H.; Alemán, B. J.; Melling, L. D.; Taormina, M. J.; Francis, M. J.; Dow-Hygelund, C. C.; Narayanan, V.; Taylor, R. P.; Stout, A.


    We report that liquids perform self-propelled motion when they are placed in contact with hot surfaces with asymmetric (ratchetlike) topology. The pumping effect is observed when the liquid is in the Leidenfrost regime (the film-boiling regime), for many liquids and over a wide temperature range. We propose that liquid motion is driven by a viscous force exerted by vapor flow between the solid and the liquid.

  11. Comparison of super-high-energy-propulsion-systems based on metallic hydrogen propellant for ES to LEO space transportation

    Thierschmann, M.


    The application is studied of metallic H2 as a rocket propellant, which contains a specific energy of about 52 kcal/g in theory yielding a maximum specific impulse of 1700 s. With the convincing advantage of having a density 14 times that of conventional liquid H2/liquid O2 propellants, metallic H2 could satisfy the demands of advanced launch vehicle propulsion for the next millennium. Provided that there is an atomic metallic state of H2, and that this state is metastable at ambient pressure, which still is not proven, the results are given of the study of some important areas, which concern the production of metallic H2, the combustion, chamber cooling, and storage. The results show that the use of metallic H2 as rocket propellant could lead to revolutionary changes in space vehicle philosophy toward small size, small weight, and high performance single stage to orbit systems. The use of high metallic H2 mass fractions results in a dramatic reduction of required propellant volume, while gas temperatures in the combustion chamber exceed 5000 K. Furthermore, it follows, that H2 (liquid or slush) is the most favorable candidate as working fluid. Jet generated noise due to high exhaust velocities could be a problem.

  12. Dynamics of a variable mass system applied to spacecraft rocket attitude theory

    Mudge, Jason Dominic

    This research project is a study of the dynamics of a variable mass system. The scope of this research project is to gain understanding as to how a variable mass system will behave. The intent is to bring the level of understanding of variable mass dynamics higher and closer to the level of constant mass dynamics in the area of spacecrafts in particular. A main contribution is the finding of a set of criteria to minimize or eliminate the deviation of the nutation angle (or cone angle or angle of attack) of spacecraft rockets passively, i.e. without active control. The motivation for this research project is the Star 48 anomaly. The Star 48 is a solid rocket motor which has propelled (boosted) communication satellites from lower earth orbit to a higher one during the 1980's. The anomaly is that when the spacecraft rocket is being propelled, the nutation angle may deviate excessively which is considered undesirable. In the first part of this research project, a variable mass system is described and defined and the governing equations are derived. The type of governing equations derived are those that are most useful for analyzing the motion of a spacecraft rocket. The method of derivation makes use of Leibnitz Theorem, Divergence Theorem and Newton's Second Law of Motion. Next, the governing equations are specialized with several assumptions which are generally accepted assumptions applied in the analysis of spacecraft rockets. With these assumptions, the form governing equations is discussed and then the equations are solved analytically for the system's angular velocity. Having solved for the angular velocity of the system, the attitude of the system is obtained using a unique method which circumvents the nonlinearities that exist using Euler Angles and their kinematical equations. The attitude is approximately found analytically and a set of criteria is discussed which will minimize or eliminate the deviation of the nutation angle of a spacecraft rocket. Finally

  13. Enhanced alkaline hydrolysis and biodegradability studies of nitrocellulose-bearing missile propellant

    Sidhoum, Mohammed; Christodoulatos, Christos; Su, Tsan-Liang; Redis, Mercurios


    Large amounts of energetic materials which have been accumulated over the years in various manufacturing and military installations must be disposed of in an environmentally sound manner. Historically, the method of choice for destruction of obsolete or aging energetic materials has been open burning or open detonation (OB/OD). This destruction approach has become undesirable due to air pollution problems. Therefore, there is a need for new technologies which will effectively and economically deal with the disposal of energetic materials. Along those lines, we have investigated a chemical/biological process for the safe destruction and disposal of a double base solid rocket propellant (AHH), which was used in several 8 inch projectile systems. The solid propellant is made of nitrocellulose and nitroglycerin as energetic components, two lead salts which act as ballistic modifiers, triacetin as a plasticizer and 2-Nitrodiphenylamine (2-NDPA) as a stabilizer. A process train is being developed to convert the organic components of the propellant to biodegradable products and remove the lead from the process stream. The solid propellant is first hydrolyzed through an enhanced alkaline hydrolysis process step. Following lead removal and neutralization, the digested liquor rich in nitrates and nitrites is found to be easily biodegradable. The digestion rate of the intact ground propellant as well as the release of nitrite and nitrate groups were substantially increased when ultrasound were supplied to the alkaline reaction medium compared to the conventional alkaline hydrolysis. The effects of reaction time, temperature, sodium hydroxide concentration and other relevant parameters on the digestion efficiency and biodegradability have been studied. The present work indicates that the AHH propellant can be disposed of safely with a combination of physiochemical and biological processes.

  14. Artist's Concept of Magnetic Launch Assisted Air-Breathing Rocket


    This artist's concept depicts a Magnetic Launch Assist vehicle in orbit. Formerly referred to as the Magnetic Levitation (Maglev) system, the Magnetic Launch Assist system is a launch system developed and tested by engineers at the Marshall Space Flight Center (MSFC) that could levitate and accelerate a launch vehicle along a track at high speeds before it leaves the ground. Using electricity and magnetic fields, a Magnetic Launch Assist system would drive a spacecraft along a horizontal track until it reaches desired speeds. The system is similar to high-speed trains and roller coasters that use high-strength magnets to lift and propel a vehicle a couple of inches above a guideway. A full-scale, operational track would be about 1.5-miles long, capable of accelerating a vehicle to 600 mph in 9.5 seconds, and the vehicle would then shift to rocket engines for launch into orbit. The major advantages of launch assist for NASA launch vehicles is that it reduces the weight of the take-off, the landing gear, the wing size, and less propellant resulting in significant cost savings. The US Navy and the British MOD (Ministry of Defense) are planning to use magnetic launch assist for their next generation aircraft carriers as the aircraft launch system. The US Army is considering using this technology for launching target drones for anti-aircraft training.

  15. Integrated Composite Rocket Nozzle Extension Project

    National Aeronautics and Space Administration — ORBITEC proposes to develop and demonstrate an Integrated Composite Rocket Nozzle Extension (ICRNE) for use in rocket thrust chambers. The ICRNE will utilize an...

  16. Control of Propellant Lead/Lag to the LAE in the AXAF Propulsion System

    Casillas, A. R.; Eninger, J.; Joseph, G.; Kenney, J.; Trinidad, M.


    Control of the rate at which hypergolic propellants are supplied to a rocket engine prior to ignition is critically important. Potentially damaging explosions may result from excessive lead of either propellant into the combustion chamber. Because the injector fill process is governed by the engine as well as the propellant feed system design, proper management of this issue must take both into consideration. This was recognized early in the development of TRW's Advanced Columbium-Liquid Apogee Engine (LAE), which was flight-qualified in 1996 to maneuver the Advanced X-Ray Astrophysics Facility (AXAF) spacecraft into orbit. The LAE runs on hydrazine and nitrogen tetroxide (MON-3) at a nominal mixture ratio of 1.0. This paper describes the comprehensive test program conducted to ensure reliable startup operation of the LAE in the AYAF propulsion system. The most significant factors affecting chamber fuel lead were found to be: (1) engine location, (2) propellant saturation level, (3) amount of undissolved gas in the lines, and (4) off- nominal tank pressures. Hot-fire tests at a chamber fuel lead range over and above that expected for the LAEs in AXAF demonstrated extremely tolerant behavior of the engine. AY-AF is scheduled for launch on NASA's STS-93 in December 1998.

  17. ADN – The new oxidizer around the corner for an environmentally friendly smokeless propellant

    Márcio Y. Nagamachi


    Full Text Available The search for a smokeless propellant has encouraged scientists and engineers to look for a chlorine-free oxidizer as a substitute for AP (ammonium perchlorate. Endeavors seemed to come to an end when ADN (ammonium dinitramide appeared in the West in the early 1990s. Although some drawbacks soon became apparent by that time, the foremost obstacle for its use in rocket-motors came from the patent originally applied for in the United States in 1990. Furthermore, environmental concerns have also increased during these two decades. Ammonium perchlorate is believed to cause thyroid cancer by contaminating soil and water. In addition, AP produces hydrogen chloride during burning which can cause acid rain and ozone layer depletion. Unlike AP, ADN stands for both smokeless and green propellant. Since then, much progress has been made in its development in synthesis, re-shaping, microencapsulation and solid propellant. The high solubility of ADN in water has also allowed its application as liquid monopropellant. Tests have revealed Isp (specific impulse superior to that normally observed with hydrazine, one of the most harmful and hazardous liquid propellants. With constraints of use, along with the patent near to expiry, scientists and engineers are rushing to complete developments and patents until then.

  18. Performance Increase Verification for a Bipropellant Rocket Engine

    Alexander, Leslie; Chapman, Jack; Wilson, Reed; Krismer, David; Lu, Frank; Wilson, Kim; Miller, Scott; England, Chris


    Component performance assessment testing for a, pressure-fed earth storable bipropellant rocket engine was successfully completed at Aerojet's Redmond test facility. The primary goal of the this development project is to increase the specific impulse of an apogee class bi-propellant engine to greater than 330 seconds with nitrogen tetroxide and monomethylhydrazine propellants and greater than 335 seconds with nitrogen tetroxide and hydrazine. The secondary goal of the project is to take greater advantage of the high temperature capabilities of iridium/rhenium chambers. In order to achieve these goals, the propellant feed pressures were increased to 400 psia, nominal, which in turn increased the chamber pressure and temperature, allowing for higher c*. The tests article used a 24-on-24 unlike doublet injector design coupled with a copper heat sink chamber to simulate a flight configuration combustion chamber. The injector is designed to produce a nominal 200 lbf of thrust with a specific impulse of 335 seconds (using hydrazine fuel). Effect of Chamber length on engine C* performance was evaluated with the use of modular, bolt-together test hardware and removable chamber inserts. Multiple short duration firings were performed to characterize injector performance across a range of thrust levels, 180 to 220 lbf, and mixture ratios, from 1.1 to 1.3. During firing, ignition transient, chamber pressure, and various temperatures were measured in order to evaluate the performance of the engine and characterize the thermal conditions. The tests successfully demonstrated the stable operation and performance potential of a full scale engine with a measured c* of XXXX ft/sec (XXXX m/s) under nominal operational conditions.

  19. High Performance Binder for EMCDB Propellants

    V. K. Bhat


    Full Text Available A novel block polymer has been synthesised from caprolactone using hydroxy terminated polybutadiene as ring opening initiator. Usefulness of this polymer as propellant binder has been studied by generating data on physico-chemical properties of the polymer. The polymer exhibited high miscibility with nitrate ester and high solid loading capability. Preliminary data generated on typical propellant formulation indicated higher performance as compared to composite propellant.

  20. An open cycle gas core fusion rocket for space exploration

    Kammash, T.; Godfrey, T.

    A nuclear propulsion system that utilizes fusion reactions to heat a plasma in a magnetically confined device is examined as a potential rocket. It makes use of a high density plasma in a magnetic mirror geometry with a collision mean free path much shorter than its length. Under these conditions the plasma behaves like a fluid with confinement properties dictated by gasdynamic laws. Accordingly, the plasma escape from the device is analogous to the flow of a gas into vacuum from a vessel with a hole. Such a system is capable of producing a very high specific impulse albeit at modest thrust. One approach for enhancing the thrust is to use an auxiliary hydrogen propellant that could be regeneratively heated before it is introduced into the reactor chamber. As is flows past the fusion plasma it will be further heated by the radiation (bremsstrahlung and synchrotron) emanating from the plasma, and upon emergence from the nozzle it will generate the desired thrust. The system thus functions much like an open cycle gas core rocket with very attractive propulsive capabilities. In this paper we present the underlying physics principles of such a concept and assess its capability by applying the results to a round trip mission to Mars. It is shown that the propulsion parameters exceed those of a gas core fission reactor and without many of major hydrodynamic problems confronted by the latter.

  1. More-Accurate Model of Flows in Rocket Injectors

    Hosangadi, Ashvin; Chenoweth, James; Brinckman, Kevin; Dash, Sanford


    An improved computational model for simulating flows in liquid-propellant injectors in rocket engines has been developed. Models like this one are needed for predicting fluxes of heat in, and performances of, the engines. An important part of predicting performance is predicting fluctuations of temperature, fluctuations of concentrations of chemical species, and effects of turbulence on diffusion of heat and chemical species. Customarily, diffusion effects are represented by parameters known in the art as the Prandtl and Schmidt numbers. Prior formulations include ad hoc assumptions of constant values of these parameters, but these assumptions and, hence, the formulations, are inaccurate for complex flows. In the improved model, these parameters are neither constant nor specified in advance: instead, they are variables obtained as part of the solution. Consequently, this model represents the effects of turbulence on diffusion of heat and chemical species more accurately than prior formulations do, and may enable more-accurate prediction of mixing and flows of heat in rocket-engine combustion chambers. The model has been implemented within CRUNCH CFD, a proprietary computational fluid dynamics (CFD) computer program, and has been tested within that program. The model could also be implemented within other CFD programs.

  2. Evaluation of Geopolymer Concrete for Rocket Test Facility Flame Deflectors

    Allgood, Daniel C.; Montes, Carlos; Islam, Rashedul; Allouche, Erez


    The current paper presents results from a combined research effort by Louisiana Tech University (LTU) and NASA Stennis Space Center (SSC) to develop a new alumina-silicate based cementitious binder capable of acting as a high performance refractory material with low heat ablation rate and high early mechanical strength. Such a binder would represent a significant contribution to NASA's efforts to develop a new generation of refractory 'hot face' liners for liquid or solid rocket plume environments. This project was developed as a continuation of on-going collaborations between LTU and SSC, where test sections of a formulation of high temperature geopolymer binder were cast in the floor and walls of Test Stand E-1 Cell 3, an active rocket engine test stand flame trench. Additionally, geopolymer concrete panels were tested using the NASA-SSC Diagnostic Test Facility (DTF) thruster, where supersonic plume environments were generated on a 1ft wide x 2ft long x 6 inch deep refractory panel. The DTF operates on LOX/GH2 propellants producing a nominal thrust of 1,200 lbf and the combustion chamber conditions are Pc=625psig, O/F=6.0. Data collected included high speed video of plume/panel area and surface profiles (depth) of the test panels measured on a 1-inch by 1-inch giving localized erosion rates during the test. Louisiana Tech conducted a microstructure analysis of the geopolymer binder after the testing program to identify phase changes in the material.

  3. Hybrid rocket propulsion systems for outer planet exploration missions

    Jens, Elizabeth T.; Cantwell, Brian J.; Hubbard, G. Scott


    Outer planet exploration missions require significant propulsive capability, particularly to achieve orbit insertion. Missions to explore the moons of outer planets place even more demanding requirements on propulsion systems, since they involve multiple large ΔV maneuvers. Hybrid rockets present a favorable alternative to conventional propulsion systems for many of these missions. They typically enjoy higher specific impulse than solids, can be throttled, stopped/restarted, and have more flexibility in their packaging configuration. Hybrids are more compact and easier to throttle than liquids and have similar performance levels. In order to investigate the suitability of these propulsion systems for exploration missions, this paper presents novel hybrid motor designs for two interplanetary missions. Hybrid propulsion systems for missions to Europa and Uranus are presented and compared to conventional in-space propulsion systems. The hybrid motor design for each of these missions is optimized across a range of parameters, including propellant selection, O/F ratio, nozzle area ratio, and chamber pressure. Details of the design process are described in order to provide guidance for researchers wishing to evaluate hybrid rocket motor designs for other missions and applications.

  4. Low Cost Nuclear Thermal Rocket Cermet Fuel Element Environment Testing

    Bradley, David E.; Mireles, Omar R.; Hickman, Robert R.


    Deep space missions with large payloads require high specific impulse (Isp) and relatively high thrust in order to achieve mission goals in reasonable time frames. Conventional, storable propellants produce average Isp. Nuclear thermal rockets (NTR) capable of high Isp thrust have been proposed. NTR employs heat produced by fission reaction to heat and therefore accelerate hydrogen which is then forced through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high temperature hydrogen exposure on fuel elements is limited. The primary concern is the mechanical failure of fuel elements which employ high-melting-point metals, ceramics or a combination (cermet) as a structural matrix into which the nuclear fuel is distributed. It is not necessary to include fissile material in test samples intended to explore high temperature hydrogen exposure of the structural support matrices. A small-scale test bed designed to heat fuel element samples via non-contact RF heating and expose samples to hydrogen is being developed to assist in optimal material and manufacturing process selection without employing fissile material. This paper details the test bed design and results of testing conducted to date.

  5. Environmental impact statement Space Shuttle advanced solid rocket motor program


    The proposed action is design, development, testing, and evaluation of Advanced Solid Rocket Motors (ASRM) to replace the motors currently used to launch the Space Shuttle. The proposed action includes design, construction, and operation of new government-owned, contractor-operated facilities for manufacturing and testing the ASRM's. The proposed action also includes transport of propellant-filled rocket motor segments from the manufacturing facility to the testing and launch sites and the return of used and/or refurbished segments to the manufacturing site. Sites being considered for the new facilities include John C. Stennis Space Center, Hancock County, Mississippi; the Yellow Creek site in Tishomingo County, Mississippi, which is currently in the custody and control of the Tennessee Valley Authority; and John F. Kennedy Space Center, Brevard County, Florida. TVA proposes to transfer its site to the custody and control of NASA if it is the selected site. All facilities need not be located at the same site. Existing facilities which may provide support for the program include Michoud Assembly Facility, New Orleans Parish, Louisiana; and Slidell Computer Center, St. Tammany Parish, Louisiana. NASA's preferred production location is the Yellow Creek site, and the preferred test location is the Stennis Space Center.

  6. Conceptual Engine System Design for NERVA derived 66.7KN and 111.2KN Thrust Nuclear Thermal Rockets

    Fittje, James E.; Buehrle, Robert J.


    The Nuclear Thermal Rocket concept is being evaluated as an advanced propulsion concept for missions to the moon and Mars. A tremendous effort was undertaken during the 1960's and 1970's to develop and test NERVA derived Nuclear Thermal Rockets in the 111.2 KN to 1112 KN pound thrust class. NASA GRC is leveraging this past NTR investment in their vehicle concepts and mission analysis studies, and has been evaluating NERVA derived engines in the 66.7 KN to the 111.2 KN thrust range. The liquid hydrogen propellant feed system, including the turbopumps, is an essential component of the overall operation of this system. The NASA GRC team is evaluating numerous propellant feed system designs with both single and twin turbopumps. The Nuclear Engine System Simulation code is being exercised to analyze thermodynamic cycle points for these selected concepts. This paper will present propellant feed system concepts and the corresponding thermodynamic cycle points for 66.7 KN and 111.2 KN thrust NTR engine systems. A pump out condition for a twin turbopump concept will also be evaluated, and the NESS code will be assessed against the Small Nuclear Rocket Engine preliminary thermodynamic data.

  7. Propeller Test Facilities Â

    Federal Laboratory Consortium — Description: Three electrically driven whirl test stands are used to determine propeller (or other rotating device) performance at various rotational speeds. These...

  8. High Seed Compressor for Propellant Densification Project

    National Aeronautics and Space Administration — Propellant densification systems particularly for H2 require compression systems developing very large amounts of head. Development of this head requires multiple...

  9. Summarization on variable liquid thrust rocket engines


    The technology actuality and development trend of variable thrust rocket engines at home and abroad are summarized. Key technologies of developing variable thrust rocket engines are analyzed. Development advices on developing variable thrust rocket engines that are adapted to the situation of our country are brought forward.

  10. Nuclear-Thermal Rocket Orbits Mars


    Originally investigated in the 1960's by Marshall Space Flight Center plarners as part of the Nuclear Energy for Rocket Vehicle Applications (NERVA) program, nuclear-thermal rocket propulsion has been more recently considered in spacecraft designs for interplanetary human exploration. This artist's concept illustrates a nuclear-thermal rocket with an aerobrake disk as it orbits Mars.

  11. Measuring Model Rocket Engine Thrust Curves

    Penn, Kim; Slaton, William V.


    This paper describes a method and setup to quickly and easily measure a model rocket engine's thrust curve using a computer data logger and force probe. Horst describes using Vernier's LabPro and force probe to measure the rocket engine's thrust curve; however, the method of attaching the rocket to the force probe is not discussed. We show how a…

  12. Measuring Model Rocket Engine Thrust Curves

    Penn, Kim; Slaton, William V.


    This paper describes a method and setup to quickly and easily measure a model rocket engine's thrust curve using a computer data logger and force probe. Horst describes using Vernier's LabPro and force probe to measure the rocket engine's thrust curve; however, the method of attaching the rocket to the force probe is not discussed. We show how a…

  13. Propellant Management in Microgravity- Further Analysis of an Experiment Flown on REXUS-14

    Strobino, D.; Zumbrunen, E.; Putzu, R.; Pontelandolfo, P.


    This paper is about the further analysis of an experiment named CAESAR (stands for Capillarity-based Experiment for Spatial Advanced Research): a sounding rocket experiment carried out by students of hepia within the REXUS program. The authors have launched on REXUS-14 a propellant management experiment based on capillarity to reliably confirm other ground-based cxperiments. In the framework of the present work, the authors present the comparison of CAESAR experimental data with theoretical profiles provided in literature. The objective of this flight was to place several Propellant Management Devices (PMD) in a microgravity environment and acquire images of the fluid distribution around them. The main element of the experiment, called a sponge, is a PMD for space vehicles, often used in satellites. This radial panel shaped device can be used at the bottom of a satellite tank to keep the propellant near the outlet. It is designed to work even if the vehicle undergoes small accelerations, for example during station-keeping maneuvers. The fluid is eccentric but stays on the sponge and near the outlet, so the injection system of the motor is continuously supplied with the propellant. As previously published, the authors have created a buoyancy test bench and have designed another system by magnetic levitation to perform the same experiment on earth. These systems are easier to use and less expensive than a sounding rocket, a parabolic flight or a drop tower (i.e. other system to obtain microgravity on earth), so they will be very useful to make progress in this particular domain of science. They will also allow universities with small funds to work within this spatial field. A previous publication showed, from a qualitative point of view, a good agreement between experiments and theory; however in this paper quantitative comparisons are given. With this demonstrated, hepia can validate its buoyancy test facility with real flight tests.

  14. The issue of ensuring the safe explosion of the spent orbital stages of a launch vehicle with propulsion rocket engine

    Trushlyakov Valeriy I.


    Full Text Available A method for increasing the safe explosion of the spent orbital stages of a space launch vehicle (SLV with a propulsion rocket engine (PRE based on the gasification of unusable residues propellant and venting fuel tanks. For gasification and ventilation the hot gases used produced by combustion of the specially selected gas generating composition (GGC with a set of physical and chemical properties. Excluding the freezing of the drainage system on reset gasified products (residues propellant+pressurization gas+hot gases in the near-Earth space is achieved by selecting the physical-chemical characteristics of the GGC. Proposed steps to ensure rotation of gasified products due to dumping through the drainage system to ensure the most favorable conditions for propellant gasification residues. For example, a tank with liquid oxygen stays with the orbital spent second stage of the SLV “Zenit”, which shows the effectiveness of the proposed method.

  15. Low Cost Nuclear Thermal Rocket Cermet Fuel Element Environment Testing

    Bradley, D. E.; Mireles, O. R.; Hickman, R. R.


    Deep space missions with large payloads require high specific impulse and relatively high thrust to achieve mission goals in reasonable time frames.1,2 Conventional storable propellants produce average specific impulse. Nuclear thermal rockets capable of producing high specific impulse are proposed. Nuclear thermal rockets employ heat produced by fission reaction to heat and therefore accelerate hydrogen, which is then forced through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000 K), and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high-temperature hydrogen exposure on fuel elements are limited.3 The primary concern is the mechanical failure of fuel elements that employ high-melting-point metals, ceramics, or a combination (cermet) as a structural matrix into which the nuclear fuel is distributed. The purpose of the testing is to obtain data to assess the properties of the non-nuclear support materials, as-fabricated, and determine their ability to survive and maintain thermal performance in a prototypical NTR reactor environment of exposure to hydrogen at very high temperatures. The fission process of the planned fissile material and the resulting heating performance is well known and does not therefore require that active fissile material be integrated in this testing. A small-scale test bed designed to heat fuel element samples via non-contact radio frequency heating and expose samples to hydrogen is being developed to assist in optimal material and manufacturing process selection without employing fissile material. This paper details the test bed design and results of testing conducted to date.

  16. 固体火箭发动机预固化技术及其应用%Pre-cure Technique and Its Applications for Solid Rocket Motors

    苏昌银; 张爱科


    Based on interface properties of HTPB propellant, the cross linking level is regulated through the temperature and time of cure reaction. The chemical reaction takes place gradually in remainder functional groups of the system to form chemical bonds and hydrogen bond, so as to improve the mechanical properties of the resultant. Pre-cure technique and bonding model are described in the paper. They can be used in propellant-liner bonding, propellant loading and integral repairing of the propellant grain of the solid rocket motor (SRM). These test results have been qualified by the successful static firing tests, flight tests of motors and storage tests of specimens for ten-years. The performances of the motor meet the design requirements with good reproducibilities.

  17. Cavitation simulation on marine propellers

    Shin, Keun Woo

    flows on a 2D hydrofoil are compared with the experimental results. In the current implementation, three models with a vapor transport equation show numerical stability and equivalently good accuracy in simulating steady and unsteady sheet cavitation. More validations for cavitating flows on 3D...... hydrofoils and conventional/highly-skewed propellers are performed with one of three cavitation models proven in 2D analysis. 3D cases also show accuracy and robustness of numerical method in simulating steady and unsteady sheet cavitation on complicated geometries. Hydrodynamic characteristics of cavitation...

  18. Cars Spectroscopy of Propellant Flames


    Harris, K. Aron, and J. Fendell "N2 and 00 Vibrational CARS and H2 Rotational CARS Spectroscopy of CHI/N20 Flames," Proceedings of the Nineteenth...JANNAF Combustion Meeting, CIIA Publication No. 366, 1982, p 123. 21. K. Aron, L. E. Harris, and J. Fendell , "N and CO Vibrational CARS and H2 Rotational...9 6 5 . p 3 8 4 . . . . . 23. J. Fendell , L. E, Harris, and K. Aron, "Theoretical Calculation of 11 CARS S-Branches for Propellant Flames

  19. High burn rate solid composite propellants

    Manship, Timothy D.

    High burn rate propellants help maintain high levels of thrust without requiring complex, high surface area grain geometries. Utilizing high burn rate propellants allows for simplified grain geometries that not only make production of the grains easier, but the simplified grains tend to have better mechanical strength, which is important in missiles undergoing high-g accelerations. Additionally, high burn rate propellants allow for a higher volumetric loading which reduces the overall missile's size and weight. The purpose of this study is to present methods of achieving a high burn rate propellant and to develop a composite propellant formulation that burns at 1.5 inches per second at 1000 psia. In this study, several means of achieving a high burn rate propellant were presented. In addition, several candidate approaches were evaluated using the Kepner-Tregoe method with hydroxyl terminated polybutadiene (HTPB)-based propellants using burn rate modifiers and dicyclopentadiene (DCPD)-based propellants being selected for further evaluation. Propellants with varying levels of nano-aluminum, nano-iron oxide, FeBTA, and overall solids loading were produced using the HTPB binder and evaluated in order to determine the effect the various ingredients have on the burn rate and to find a formulation that provides the burn rate desired. Experiments were conducted to compare the burn rates of propellants using the binders HTPB and DCPD. The DCPD formulation matched that of the baseline HTPB mix. Finally, GAP-plasticized DCPD gumstock dogbones were attempted to be made for mechanical evaluation. Results from the study show that nano-additives have a substantial effect on propellant burn rate with nano-iron oxide having the largest influence. Of the formulations tested, the highest burn rate was a 84% solids loading mix using nano-aluminum nano-iron oxide, and ammonium perchlorate in a 3:1(20 micron: 200 micron) ratio which achieved a burn rate of 1.2 inches per second at 1000

  20. Scale Effects on Quasi-Steady Solid Rocket Internal Ballistic Behaviour

    David R. Greatrix


    Full Text Available The ability to predict with some accuracy a given solid rocket motor’s performance before undertaking one or several costly experimental test firings is important. On the numerical prediction side, as various component models evolve, their incorporation into an overall internal ballistics simulation program allows for new motor firing simulations to take place, which in turn allows for updated comparisons to experimental firing data. In the present investigation, utilizing an updated simulation program, the focus is on quasi-steady performance analysis and scale effects (influence of motor size. The predicted effects of negative/positive erosive burning and propellant/casing deflection, as tied to motor size, on a reference cylindrical-grain motor’s internal ballistics, are included in this evaluation. Propellant deflection has only a minor influence on the reference motor’s internal ballistics, regardless of motor size. Erosive burning, on the other hand, is distinctly affected by motor scale.

  1. Performance Studies on Sub-cooling of Cryogenic Liquids Used for Rocket Propulsion Using Helium Bubbling

    Ramesh T


    Full Text Available The sub-cooling of cryogenic propellants contained in tanks is an important and effective method for bringing down the lift-off mass of launch vehicle and thus the performance of the rocket engine is greatly improved. This paper presents the technical and experimental studies conducted on cryogenic liquids such as Liquid Oxygen, Liquid Nitrogen, and Liquid Hydrogen using helium bubbling method. The influence of cooled Helium on the degree of sub-cooling and the variation in flow rate of Helium gas admitted are discussed. The experimental and theoretical studies indicate that the sub-cooling technique using helium injection is a very simple method and can be very well adopted in propellant tanks used for ground and launch vehicle applications.

  2. Mars Sample Return and Flight Test of a Small Bimodal Nuclear Rocket and ISRU Plant

    George, Jeffrey A.; Wolinsky, Jason J.; Bilyeu, Michael B.; Scott, John H.


    A combined Nuclear Thermal Rocket (NTR) flight test and Mars Sample Return mission (MSR) is explored as a means of "jump-starting" NTR development. Development of a small-scale engine with relevant fuel and performance could more affordably and quickly "pathfind" the way to larger scale engines. A flight test with subsequent inflight postirradiation evaluation may also be more affordable and expedient compared to ground testing and associated facilities and approvals. Mission trades and a reference scenario based upon a single expendable launch vehicle (ELV) are discussed. A novel "single stack" spacecraft/lander/ascent vehicle concept is described configured around a "top-mounted" downward firing NTR, reusable common tank, and "bottom-mount" bus, payload and landing gear. Requirements for a hypothetical NTR engine are described that would be capable of direct thermal propulsion with either hydrogen or methane propellant, and modest electrical power generation during cruise and Mars surface insitu resource utilization (ISRU) propellant production.

  3. Status on Technology Development of Optic Fiber-Coupled Laser Ignition System for Rocket Engine Applications

    Trinh, Huu P.; Early, Jim; Osborne, Robin; Thomas, Matthew; Bossard, John


    To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for liquid rocket engine applications. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concept: not only offer system simplicity, but also enhance the combustion performance. Test results have shown that chamber performance is markedly high even at a low chamber length-to-diameter ratio. This incentive can be translated to a convenience in the thrust chamber packaging.

  4. A feasibility study on using inkjet technology, micropumps, and MEMs as fuel injectors for bipropellant rocket engines

    Glynne-Jones, Peter; Coletti, M.; White, N. M.; Gabriel, S. B.; Bramanti, C.


    Control over drop size distributions, injection rates, and geometrical distribution of fuel and oxidizer sprays in bi-propellant rocket engines has the potential to produce more efficient, more stable, less polluting rocket engines. This control also offers the potential of an engine that can be throttled, working efficiently over a wide range of output thrusts. Inkjet printing technologies, MEMS fuel atomizers, and piezoelectric injectors similar in concept to those used in diesel engines are considered for their potential to yield a new, more active injection scheme for a rocket engine. Inkjets are found to be unable to pump at sufficient pressures, and have possibly dangerous failure modes. Active injection is found to be feasible if high pressure drop along the injector plate is used. A conceptual design is presented and its basic behavior assessed.

  5. Effects of propellant composition variables on acceleration-induced burning-rate augmentation of solid propellants

    Northam, G. B.


    This work was conducted to define further the effects of propellant composition variables on the acceleration-induced burning rate augmentation of solid propellants. The rate augmentation at a given acceleration was found to be a nonlinear inverse function of the reference burning rate and not controlled by binder or catalyst type at a given reference rate. A nonaluminized propellant and a low rate double-base propellant exhibited strong transient rate augmentation due to surface pitting resulting from the retention of hot particles on the propellant surface.

  6. NASA's Hydrogen Outpost: The Rocket Systems Area at Plum Brook Station

    Arrighi, Robert S.


    "There was pretty much a general knowledge about hydrogen and its capabilities," recalled former researcher Robert Graham. "The question was, could you use it in a rocket engine? Do we have the technology to handle it? How will it cool? Will it produce so much heat release that we can't cool the engine? These were the questions that we had to address." The National Aeronautics and Space Administration's (NASA) Glenn Research Center, referred to historically as the Lewis Research Center, made a concerted effort to answer these and related questions in the 1950s and 1960s. The center played a critical role transforming hydrogen's theoretical potential into a flight-ready propellant. Since then NASA has utilized liquid hydrogen to send humans and robots to the Moon, propel dozens of spacecraft across the universe, orbit scores of satellite systems, and power 135 space shuttle flights. Rocket pioneers had recognized hydrogen's potential early on, but its extremely low boiling temperature and low density made it impracticable as a fuel. The Lewis laboratory first demonstrated that liquid hydrogen could be safely utilized in rocket and aircraft propulsion systems, then perfected techniques to store, pump, and cleanly burn the fuel, as well as use it to cool the engine. The Rocket Systems Area at Lewis's remote testing area, Plum Brook Station, played a little known, but important role in the center's hydrogen research efforts. This publication focuses on the activities at the Rocket Systems Area, but it also discusses hydrogen's role in NASA's space program and Lewis's overall hydrogen work. The Rocket Systems Area included nine physically modest test sites and three test stands dedicated to liquid-hydrogen-related research. In 1962 Cleveland Plain Dealer reporter Karl Abram claimed, "The rocket facility looks more like a petroleum refinery. Its test rigs sprout pipes, valves and tanks. During the night test runs, excess hydrogen is burned from special stacks in the best

  7. Green plasticizers for multibase gun propellants (Lecture)

    Schoolderman, C.; Driel, C.A. van; Zebregs, M.


    TNO Defence, Security and Safety has a long history of research on gun propellants. Areas investigated are formulating (new ingredients, optimization), manufacturing, charge design and lifetime assessment [1,2,3,4,5]. In conventional propellants inert plasticizers are used to alter performance,

  8. Calculation of Thermochemical Constants of Propellants

    K. P. Rao


    Full Text Available A method for calculation of thermo chemical constants and products of explosion of propellants from the knowledge of molecular formulae and heats of formation of the ingredients is given. A computer programme in AUTOMATH-400 has been established for the method. The results of application of the method for a number of propellants are given.

  9. Design Procedure of 4-Bladed Propeller



    Sep 1, 2013 ... experiments with small–scale propeller mode were made in ... experiments were used to verify the design calculations like ... extent, in their aim to satisfy a particular set of .... 5.5knots. 2. = Brake power B. P from equation 2 is calculated as follow: ..... A step by step design procedure for a 4-bladed propeller ...

  10. Development of hydrazinium nitroformate based solid propellants

    Schöyer, H.F.R.; Schnorhk, A.J.; Korting, P.A.O.G.; Lit, P.J. van; Mul, J.M.; Gadiot, G.; Meulenbrugge, J.J.


    The development of new high-performance propellant combinations requires the establishment of safety and handling characteristics and thermodynamic decomposition and explosive properties. This paper addresses the early development phases of a new composite solid propellant based on HNF as oxidizer a

  11. Green plasticizers for multibase gun propellants (Lecture)

    Schoolderman, C.; Driel, C.A. van; Zebregs, M.


    TNO Defence, Security and Safety has a long history of research on gun propellants. Areas investigated are formulating (new ingredients, optimization), manufacturing, charge design and lifetime assessment [1,2,3,4,5]. In conventional propellants inert plasticizers are used to alter performance, proc

  12. 14 CFR 21.129 - Tests: propellers.


    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Tests: propellers. 21.129 Section 21.129 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT CERTIFICATION PROCEDURES FOR PRODUCTS AND PARTS Production Under Type Certificate Only § 21.129 Tests: propellers....

  13. ASRM radiation and flowfield prediction status. [Advanced Solid Rocket Motor plume radiation prediction

    Reardon, J. E.; Everson, J.; Smith, S. D.; Sulyma, P. R.


    Existing and proposed methods for the prediction of plume radiation are discussed in terms of their application to the NASA Advanced Solid Rocket Motor (ASRM) and Space Shuttle Main Engine (SSME) projects. Extrapolations of the Solid Rocket Motor (SRM) are discussed with respect to preliminary predictions of the primary and secondary radiation environments. The methodology for radiation and initial plume property predictions are set forth, including a new code for scattering media and independent secondary source models based on flight data. The Monte Carlo code employs a reverse-evaluation approach which traces rays back to their point of absorption in the plume. The SRM sea-level plume model is modified to account for the increased radiation in the ASRM plume due to the ASRM's propellant chemistry. The ASRM cycle-1 environment predictions are shown to identify a potential reason for the shutdown spike identified with pre-SRM staging.

  14. Experimental determination of plume properties in full-scale hydrogen-oxygen rockets

    Brown, D. G.; Limbaugh, C. C.; Zaccardi, V. A.; Eskridge, R.


    An IR emission/absorption technique for determining radial profiles of static temperature and species partial pressure for cylindrically symmetric combustion gases typical of the effluent of turbine engines and liquid-propellant rockets is described. In the technique, the IR plume radiance and absorption is measured using a 1 x 256-element platinum silicide detector array which is filtered to obtain plume emission measurements in the H2O band near 3.0 microns. A minicomputer is employed to control data acquisition and reduction.

  15. An injector design model for predicting rocket engine performance and heat transfer

    Calhoon, D. F.; Kors, D. L.; Gordon, L. H.


    A model is formulated for estimating the performance and chamber heat transfer in rocket injectors/chambers operating with gaseous H2-O2 propellants. The model quantifies the combustion performance and chamber heat flux for variables such as chamber length, element type, element area ratio, impingement angle, thrust/element, mixture ratio, moment ratio, element spacing, and physical size. Design equations are given and curves are plotted for evaluation of combustion performance in injectors comprised of F-O-F triplet, premix, coaxial and swirl coaxial element types. Curve plots and equations are also included for estimation of the chamber wall heat fluxes generated by these element types.

  16. Mechanism study of stabilization of double-base propellants by using zeolite stabilizers (nano- and micro-clinoptilolite

    M.A. Zayed


    Full Text Available The mechanism of stabilization of double-base rocket propellants (DBPs using inorganic stabilizer (nano- and micro-clinoptilolite was investigated. The surface structures of the stabilizers, the double-base propellants containing the new stabilizers and the effect of the stabilizers on the surface behavior of propellants and vice versa were checked using XRD and Electron Microscope (AFM and TEM techniques. The results obtained from XRD suggested that the crystalline structure of the new inorganic stabilizers was completely changed when it was introduced into the propellants which may be attributed to the pressing processes of DBPs with stabilizers under very high pressure during their mixture preparation. The results obtained from Atomic Force Electron Microscope (AFM and TEM indicated that nano-clinoptilolite particles become more regularly arranged on the propellant surface than micro-clinoptilolite which gives the stabilizer a higher ability to absorb more nitrogen oxide. The work aimed chiefly to use zeolite stabilizers for DBPs instead of classically used organic compounds; in order to avoid the harmful and carcinogenic organic products coming from the reaction of NOx gases with these organic stabilizers. This is achieved by studying the thermal behavior of these zeolites via investigation of their surface interaction with NOx gases obtained during stabilization process and suggesting possible interaction mechanism.

  17. MAST Propellant and Delivery System Design Methods

    Nadeem, Uzair; Mc Cleskey, Carey M.


    A Mars Aerospace Taxi (MAST) concept and propellant storage and delivery case study is undergoing investigation by NASA's Element Design and Architectural Impact (EDAI) design and analysis forum. The MAST lander concept envisions landing with its ascent propellant storage tanks empty and supplying these reusable Mars landers with propellant that is generated and transferred while on the Mars surface. The report provides an overview of the data derived from modeling between different methods of propellant line routing (or "lining") and differentiate the resulting design and operations complexity of fluid and gaseous paths based on a given set of fluid sources and destinations. The EDAI team desires a rough-order-magnitude algorithm for estimating the lining characteristics (i.e., the plumbing mass and complexity) associated different numbers of vehicle propellant sources and destinations. This paper explored the feasibility of preparing a mathematically sound algorithm for this purpose, and offers a method for the EDAI team to implement.

  18. Burning rate characteristics of CMDB propellants

    Swaminathan, V.; Soosai Marian, M. (Vikram Sarabhai Space Centre, Trivandrum (India). Propellant Engineering Div.)


    The object of this paper is to obtain expressions for the burning rate pressure exponent and the temperature sensitivity of AP-based and HMX-based CMDB propellants in terms of the respective physical constants on the basis of a recently developed model of combustion for CMDB propellants and to examine the effects, if any, on these two parameters, of the changes in propellant composition, AP particle size and pressure. Computer programs were developed for this purpose and the results obtained for typical sets of input data presented in the paper. While the results of the calculation indicate a markedly strong dependence of the pressure exponent and the temperature sensitivity on pressure and composition for both AP-based and HMX-based CMDB propellants, the parameters are characterized by far lesser dependence on AP particle size for the AP-based propellant.

  19. Storage of solid propellants in space

    Udlock, D. E.


    A test program is described which determines the extent of physical property changes that result from extended space exposure. Primary emphasis was placed on determining the effects of space vacuum. Solid propellants were stored and their physical properties tested in a vacuum and in a dry environment. The storage caused significantly greater increases in the propellants' modulus and maximum tensile strength than occurred in parallel ambient stored samples. The data indicate that the loss of trace amounts of residual moisture from cured propellant is the apparent cause of the observed propellant property changes. Therefore, initial screening tests were carried out under dry storage conditions. Upon completion of the dry storage tests, appropriate propellant samples are exposed to an actual space environment using the Long Duration Exposure Facility (LDEF).

  20. Unique nuclear thermal rocket engine

    Culver, D.W. (Aerojet Propulsion Division, P.O. Box 13222, Sacramento, California 95813-6000 (United States)); Rochow, R. (Babcock Wilcox Space Nuclear Systems, P.O. Box 11165, Lynchburg, Virginia 24506-1165 (United States))


    Earlier this year Aerojet Propulsion Division (APD) introduced a new, advanced nuclear thermal rocket engine (NTRE) concept intended for manned missions to the moon and to Mars. This NTRE promises to be both shorter and lighter in weight than conventionally designed engines, because its forward flowing reactor is located within an expansion-deflection (E-D) rocket nozzle. The concept has matured during the year, and this paper discusses a nearer term version that resolves four open issues identified in the initial concept: (1)Reactor design and cooling scheme simplification while retaining a high pressure power balance option; (2)Eliminate need for a new, uncooled nozzle throat material suitable for long life application; (3)Practical provision for reactor power control; and (4)Use near term, long life turbopumps.