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Sample records for inch rocket motor

  1. A study of performance and cost improvement potential of the 120 inch (3.05 m) diameter solid rocket motor. Volume 1: Summary report

    Science.gov (United States)

    Backlund, S. J.; Rossen, J. N.

    1971-01-01

    A parametric study of ballistic modifications to the 120 inch diameter solid propellant rocket engine which forms part of the Air Force Titan 3 system is presented. 576 separate designs were defined and 24 were selected for detailed analysis. Detailed design descriptions, ballistic performance, and mass property data were prepared for each design. It was determined that a relatively simple change in design parameters could provide a wide range of solid propellant rocket engine ballistic characteristics for future launch vehicle applications.

  2. Thiokol Solid Rocket Motors

    Science.gov (United States)

    Graves, S. R.

    2000-01-01

    This paper presents viewgraphs on thiokol solid rocket motors. The topics include: 1) Communications; 2) Military and government intelligence; 3) Positioning satellites; 4) Remote sensing; 5) Space burial; 6) Science; 7) Space manufacturing; 8) Advertising; 9) Space rescue space debris management; 10) Space tourism; 11) Space settlements; 12) Hazardous waste disposal; 13) Extraterrestrial resources; 14) Fast package delivery; and 15) Space utilities.

  3. Measuring the Internal Environment of Solid Rocket Motors During Ignition

    Science.gov (United States)

    Weisenberg, Brent; Smith, Doug; Speas, Kyle; Corliss, Adam

    2003-01-01

    A new instrumentation system has been developed to measure the internal environment of solid rocket test motors during motor ignition. The system leverages conventional, analog gages with custom designed, electronics modules to provide safe, accurate, high speed data acquisition capability. To date, the instrumentation system has been demonstrated in a laboratory environment and on subscale static fire test motors ranging in size from 5-inches to 24-inches in diameter. Ultimately, this system is intended to be installed on a full-scale Reusable Solid Rocket Motor. This paper explains the need for the data, the components and capabilities of the system, and the test results.

  4. The Advanced Solid Rocket Motor

    Science.gov (United States)

    Mitchell, Royce E.

    1992-01-01

    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.

  5. Parameters Affecting the Erosive Burning of Solid Rocket Motor

    Directory of Open Access Journals (Sweden)

    Abdelaziz Almostafa

    2018-01-01

    Full Text Available Increasing the velocity of gases inside solid rocket motors with low port-to-throat area ratios, leading to increased occurrence and severity of burning rate augmentation due to flow of propellant products across burning propellant surfaces (erosive burning, erosive burning of high energy composite propellant was investigated to supply rocket motor design criteria and to supplement knowledge of combustion phenomena, pressure, burning rate and high velocity of gases all of these are parameters affect on erosive burning. Investigate the phenomena of the erosive burning by using the 2’inch rocket motor and modified one. Different tests applied to fulfil all the parameters that calculated out from the experiments and by studying the pressure time curve and erosive burning phenomena.

  6. Experimental investigation of solid rocket motors for small sounding rockets

    Science.gov (United States)

    Suksila, Thada

    2018-01-01

    Experimentation and research of solid rocket motors are important subjects for aerospace engineering students. However, many institutes in Thailand rarely include experiments on solid rocket motors in research projects of aerospace engineering students, mainly because of the complexity of mixing the explosive propellants. This paper focuses on the design and construction of a solid rocket motor for total impulse in the class I-J that can be utilised as a small sounding rocket by researchers in the near future. Initially, the test stands intended for measuring the pressure in the combustion chamber and the thrust of the solid rocket motor were designed and constructed. The basic design of the propellant configuration was evaluated. Several formulas and ratios of solid propellants were compared for achieving the maximum thrust. The convenience of manufacturing and casting of the fabricated solid rocket motors were a critical consideration. The motor structural analysis such as the combustion chamber wall thickness was also discussed. Several types of nozzles were compared and evaluated for ensuring the maximum thrust of the solid rocket motors during the experiments. The theory of heat transfer analysis in the combustion chamber was discussed and compared with the experimental data.

  7. DEVELOPMENT OF A NINE INCH DIAMETER, MACH 5.5, MONORAIL, ROCKET SLED.

    Science.gov (United States)

    A nine inch diameter monorail rocket sled was designed, fabricated and tested at Holloman Air Force Base. The vehicle was designed to allow easy...replacement of appendages which were subject to severe aerodynamic heating and/or high wear rates. The monorail vehicle as described was shown to be

  8. Design methods in solid rocket motors

    Energy Technology Data Exchange (ETDEWEB)

    1987-03-01

    A compilation of lectures summarizing the current state-of-the-art in designing solid rocket motors and and their components is presented. The experience of several countries in the use of new technologies and methods is represented. Specific sessions address propellant grains, cases, nozzles, internal thermal insulation, and the general optimization of solid rocket motor designs.

  9. Solid rocket motor cost model

    Science.gov (United States)

    Harney, A. G.; Raphael, L.; Warren, S.; Yakura, J. K.

    1972-01-01

    A systematic and standardized procedure for estimating life cycle costs of solid rocket motor booster configurations. The model consists of clearly defined cost categories and appropriate cost equations in which cost is related to program and hardware parameters. Cost estimating relationships are generally based on analogous experience. In this model the experience drawn on is from estimates prepared by the study contractors. Contractors' estimates are derived by means of engineering estimates for some predetermined level of detail of the SRM hardware and program functions of the system life cycle. This method is frequently referred to as bottom-up. A parametric cost analysis is a useful technique when rapid estimates are required. This is particularly true during the planning stages of a system when hardware designs and program definition are conceptual and constantly changing as the selection process, which includes cost comparisons or trade-offs, is performed. The use of cost estimating relationships also facilitates the performance of cost sensitivity studies in which relative and comparable cost comparisons are significant.

  10. Flow-Structural Interaction in Solid Rocket Motors

    National Research Council Canada - National Science Library

    Murdock, John

    2004-01-01

    .... The static test failure of the Titan solid rocket motor upgrade (SRMU) that occurred on 1 April, 1991, demonstrated the importance of flow-structural modeling in the design of large, solid rocket motors...

  11. Motor actuated vacuum door. [for photography from sounding rockets

    Science.gov (United States)

    Hanagud, A. V.

    1986-01-01

    Doors that allow scientific instruments to record and retrieve the observed data are often required to be designed and installed as a part of sounding rocket hardware. The motor-actuated vacuum door was designed to maintain a medium vacuum of the order of 0.0001 torr or better while closed, and to provide an opening 15 inches long x 8.5 inches wide while open for cameras to image Halley's comet. When the electric motor receives the instruction to open the door through the payload battery, timer, and relay circuit, the first operation is to unlock the door. After unlatching, the torque transmitted by the motor to the main shaft through the links opens the door. A microswitch actuator, which rides on the linear motion conversion mechanism, is adjusted to trip the limit switch at the end of the travel. The process is repeated in the reverse order to close the door. 'O' rings are designed to maintain the seal. Door mechanisms similar to the one described have flown on Aerobee 17.018 and Black Brant 27.047 payloads.

  12. Solid Rocket Motor Design Using Hybrid Optimization

    Directory of Open Access Journals (Sweden)

    Kevin Albarado

    2012-01-01

    Full Text Available A particle swarm/pattern search hybrid optimizer was used to drive a solid rocket motor modeling code to an optimal solution. The solid motor code models tapered motor geometries using analytical burn back methods by slicing the grain into thin sections along the axial direction. Grains with circular perforated stars, wagon wheels, and dog bones can be considered and multiple tapered sections can be constructed. The hybrid approach to optimization is capable of exploring large areas of the solution space through particle swarming, but is also able to climb “hills” of optimality through gradient based pattern searching. A preliminary method for designing tapered internal geometry as well as tapered outer mold-line geometry is presented. A total of four optimization cases were performed. The first two case studies examines designing motors to match a given regressive-progressive-regressive burn profile. The third case study studies designing a neutrally burning right circular perforated grain (utilizing inner and external geometry tapering. The final case study studies designing a linearly regressive burning profile for right circular perforated (tapered grains.

  13. Rotational flow in tapered slab rocket motors

    Science.gov (United States)

    Saad, Tony; Sams, Oliver C.; Majdalani, Joseph

    2006-10-01

    Internal flow modeling is a requisite for obtaining critical parameters in the design and fabrication of modern solid rocket motors. In this work, the analytical formulation of internal flows particular to motors with tapered sidewalls is pursued. The analysis employs the vorticity-streamfunction approach to treat this problem assuming steady, incompressible, inviscid, and nonreactive flow conditions. The resulting solution is rotational following the analyses presented by Culick for a cylindrical motor. In an extension to Culick's work, Clayton has recently managed to incorporate the effect of tapered walls. Here, an approach similar to that of Clayton is applied to a slab motor in which the chamber is modeled as a rectangular channel with tapered sidewalls. The solutions are shown to be reducible, at leading order, to Taylor's inviscid profile in a porous channel. The analysis also captures the generation of vorticity at the surface of the propellant and its transport along the streamlines. It is from the axial pressure gradient that the proper form of the vorticity is ascertained. Regular perturbations are then used to solve the vorticity equation that prescribes the mean flow motion. Subsequently, numerical simulations via a finite volume solver are carried out to gain further confidence in the analytical approximations. In illustrating the effects of the taper on flow conditions, comparisons of total pressure and velocity profiles in tapered and nontapered chambers are entertained. Finally, a comparison with the axisymmetric flow analog is presented.

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

    Science.gov (United States)

    2016-02-16

    Laboratory, Edwards, CA Abstract In a solid rocket motor (SRM), when the aluminum based propellant combusts, the fuel is oxidized into alumina (Al2O3...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...34 Solid Rocket Nozzle Design Summary," in 4th AIAA Propulsion Joint Specialist Conference, Cleveland, OH, 1968. [5] Nayfeh, A. H.; Saric, W. S

  15. Linear stability analysis in a solid-propellant rocket motor

    Energy Technology Data Exchange (ETDEWEB)

    Kim, K.M.; Kang, K.T.; Yoon, J.K. [Agency for Defense Development, Taejon (Korea, Republic of)

    1995-10-01

    Combustion instability in solid-propellant rocket motors depends on the balance between acoustic energy gains and losses of the system. The objective of this paper is to demonstrate the capability of the program which predicts the standard longitudinal stability using acoustic modes based on linear stability analysis and T-burner test results of propellants. Commercial ANSYS 5.0A program can be used to calculate the acoustic characteristic of a rocket motor. The linear stability prediction was compared with the static firing test results of rocket motors. (author). 11 refs., 17 figs.

  16. Aluminum Agglomeration and Trajectory in Solid Rocket Motors

    National Research Council Canada - National Science Library

    Coats, Douglas; Hylin, E. C; Babbitt, Deborah; Tullos, James A; Beckstead, Merrill; Webb, Michael; Davis, I. L; Dang, Anthony

    2007-01-01

    Report developed under STTR contract for Topic AF06-T012. The demand for higher performance rocket motors at a reduced cost requires continuous improvements in understanding and controlling propellant combustion...

  17. Hybrid rocket motor testing at Nammo Raufoss A/S

    Science.gov (United States)

    Rønningen, Jan-Erik; Kubberud, Nils

    2005-08-01

    Hybrid rocket motor technology and the use of hybrid rockets have gained increased interest in recent years in many countries. A typical hybrid rocket consists of a tank containing the oxidizer in either liquid or gaseous state connected to the combustion chamber containing an injector, inert solid fuel grain and nozzle. Nammo Raufoss A/S has for almost 40 years designed and produced high-performance solid propellant rocket motors for many military missile systems as well as solid propellant rocket motors for civil space use. In 2003 an in-house technology program was initiated to investigate and study hybrid rocket technology. On 23 September 2004 the first in-house designed hybrid test rocket motor was static test fired at Nammo Raufoss Test Center. The oxidizer was gaseous oxygen contained in a tank pressurized to 10MPa, flow controlled through a sonic orifice into the combustion chamber containing a multi port radial injector and six bore cartridge-loaded fuel grain containing a modified HTPB fuel composition. The motor was ignited using a non-explosive heated wire. This paper will present what has been achieved at Nammo Raufoss since the start of the program.

  18. Parameters Affecting the Erosive Burning of Solid Rocket Motor

    OpenAIRE

    Abdelaziz Almostafa; Guozhu Liang; Elsayed Anwer

    2018-01-01

    Increasing the velocity of gases inside solid rocket motors with low port-to-throat area ratios, leading to increased occurrence and severity of burning rate augmentation due to flow of propellant products across burning propellant surfaces (erosive burning), erosive burning of high energy composite propellant was investigated to supply rocket motor design criteria and to supplement knowledge of combustion phenomena, pressure, burning rate and high velocity of gases all of these are parameter...

  19. Integral performance optimum design for multistage solid propellant rocket motors

    Energy Technology Data Exchange (ETDEWEB)

    Zhang, Hongtao (Shaanxi Power Machinery Institute (China))

    1989-04-01

    A mathematical model for integral performance optimization of multistage solid propellant rocket motors is presented. A calculation on a three-stage, volume-fixed, solid propellant rocket motor is used as an example. It is shown that the velocity at burnout of intermediate-range or long-range ballistic missile calculated using this model is four percent greater than that using the usual empirical method.

  20. Structural design of an in-line bolted joint for the space shuttle solid rocket motor case segments

    Science.gov (United States)

    Dorsey, John T.; Stein, Peter A.; Bush, Harold G.

    1987-01-01

    Results of a structural design study of an in-line bolted joint concept which can be used to assemble Space Shuttle Solid Rocket Motor (SRM) case segments are presented. Numerous parametric studies are performed to characterize the in-line bolted joint behavior as major design variables are altered, with the primary objective always being to keep the inside of the joint (where the O-rings are located) closed during the SRM firing. The resulting design has 180 1-inch studs, an eccentricity of -0.5 inch, a flange thickness of 3/4 inch, a bearing plate thickness of 1/4 inch, and the studs are subjected to a preload which is 70% of ultimate. The mass penalty per case segment joint for the in-line design is 346 lbm more than the weight penalty for the proposed capture tang fix.

  1. The XQC microcalorimeter sounding rocket: a stable LTD platform 30 seconds after rocket motor burnout

    International Nuclear Information System (INIS)

    Porter, F.S.; Almy, R.; Apodaca, E.; Figueroa-Feliciano, E.; Galeazzi, M.; Kelley, R.; McCammon, D.; Stahle, C.K.; Szymkowiak, A.E.; Sanders, W.T.

    2000-01-01

    The XQC microcalorimeter sounding rocket experiment is designed to provide a stable thermal environment for an LTD detector system within 30 s of the burnout of its second stage rocket motor. The detector system used for this instrument is a 36-pixel microcalorimeter array operated at 60 mK with a single-stage adiabatic demagnetization refrigerator (ADR). The ADR is mounted on a space-pumped liquid helium tank with vapor cooled shields which is vibration isolated from the rocket structure. We present here some of the design and performance details of this mature LTD instrument, which has just completed its third suborbital flight

  2. The XQC microcalorimeter sounding rocket: a stable LTD platform 30 seconds after rocket motor burnout

    Energy Technology Data Exchange (ETDEWEB)

    Porter, F.S. E-mail: frederick.s.porter@gsfc.nasa.gov; Almy, R.; Apodaca, E.; Figueroa-Feliciano, E.; Galeazzi, M.; Kelley, R.; McCammon, D.; Stahle, C.K.; Szymkowiak, A.E.; Sanders, W.T

    2000-04-07

    The XQC microcalorimeter sounding rocket experiment is designed to provide a stable thermal environment for an LTD detector system within 30 s of the burnout of its second stage rocket motor. The detector system used for this instrument is a 36-pixel microcalorimeter array operated at 60 mK with a single-stage adiabatic demagnetization refrigerator (ADR). The ADR is mounted on a space-pumped liquid helium tank with vapor cooled shields which is vibration isolated from the rocket structure. We present here some of the design and performance details of this mature LTD instrument, which has just completed its third suborbital flight.

  3. Rocket motors incorporating basalt fiber and nanoclay compositions and methods of insulating a rocket motor with the same

    Science.gov (United States)

    Gajiwala, Himansu M. (Inventor)

    2011-01-01

    An insulation composition that comprises at least one nitrile butadiene rubber, basalt fibers, and nanoclay is disclosed. Further disclosed is an insulation composition that comprises polybenzimidazole fibers, basalt fibers, and nanoclay. The basalt fibers may be present in the insulation compositions in a range of from approximately 1% by weight to approximately 6% by weight of the total weight of the insulation composition. The nanoclay may be present in the insulation compositions in a range of from approximately 5% by weight to approximately 10% by weight of the total weight of the insulation composition. Rocket motors including the insulation compositions and methods of insulating a rocket motor are also disclosed.

  4. Development and Performance of the 10 kN Hybrid Rocket Motor for the Stratos II Sounding Rocket

    NARCIS (Netherlands)

    Werner, R.M.; Knop, T.R.; Wink, J; Ehlen, J; Huijsman, R; Powell, S; Florea, R.; Wieling, W; Cervone, A.; Zandbergen, B.T.C.

    2016-01-01

    This paper presents the development work of the 10 kN hybrid rocket motor DHX-200 Aurora. The DHX-200 Aurora was developed by Delft Aerospace Rocket Engineering (DARE) to power the Stratos II and Stratos II+ sounding rocket, with the later one being launched in October 2015. Stratos II and Stratos

  5. Hydrodynamic Stability Analysis of Particle-Laden Solid Rocket Motors

    Science.gov (United States)

    Elliott, T. S.; Majdalani, J.

    2014-11-01

    Fluid-wall interactions within solid rocket motors can result in parietal vortex shedding giving rise to hydrodynamic instabilities, or unsteady waves, that translate into pressure oscillations. The oscillations can result in vibrations observed by the rocket, rocket subsystems, or payload, which can lead to changes in flight characteristics, design failure, or other undesirable effects. For many years particles have been embedded in solid rocket propellants with the understanding that their presence increases specific impulse and suppresses fluctuations in the flowfield. This study utilizes a two dimensional framework to understand and quantify the aforementioned two-phase flowfield inside a motor case with a cylindrical grain perforation. This is accomplished through the use of linearized Navier-Stokes equations with the Stokes drag equation and application of the biglobal ansatz. Obtaining the biglobal equations for analysis requires quantification of the mean flowfield within the solid rocket motor. To that end, the extended Taylor-Culick form will be utilized to represent the gaseous phase of the mean flowfield while the self-similar form will be employed for the particle phase. Advancing the mean flowfield by quantifying the particle mass concentration with a semi-analytical solution the finalized mean flowfield is combined with the biglobal equations resulting in a system of eight partial differential equations. This system is solved using an eigensolver within the framework yielding the entire spectrum of eigenvalues, frequency and growth rate components, at once. This work will detail the parametric analysis performed to demonstrate the stabilizing and destabilizing effects of particles within solid rocket combustion.

  6. Hydrodynamic Stability Analysis of Particle-Laden Solid Rocket Motors

    International Nuclear Information System (INIS)

    Elliott, T S; Majdalani, J

    2014-01-01

    Fluid-wall interactions within solid rocket motors can result in parietal vortex shedding giving rise to hydrodynamic instabilities, or unsteady waves, that translate into pressure oscillations. The oscillations can result in vibrations observed by the rocket, rocket subsystems, or payload, which can lead to changes in flight characteristics, design failure, or other undesirable effects. For many years particles have been embedded in solid rocket propellants with the understanding that their presence increases specific impulse and suppresses fluctuations in the flowfield. This study utilizes a two dimensional framework to understand and quantify the aforementioned two-phase flowfield inside a motor case with a cylindrical grain perforation. This is accomplished through the use of linearized Navier-Stokes equations with the Stokes drag equation and application of the biglobal ansatz. Obtaining the biglobal equations for analysis requires quantification of the mean flowfield within the solid rocket motor. To that end, the extended Taylor-Culick form will be utilized to represent the gaseous phase of the mean flowfield while the self-similar form will be employed for the particle phase. Advancing the mean flowfield by quantifying the particle mass concentration with a semi-analytical solution the finalized mean flowfield is combined with the biglobal equations resulting in a system of eight partial differential equations. This system is solved using an eigensolver within the framework yielding the entire spectrum of eigenvalues, frequency and growth rate components, at once. This work will detail the parametric analysis performed to demonstrate the stabilizing and destabilizing effects of particles within solid rocket combustion

  7. An example of successful international cooperation in rocket motor technology

    Science.gov (United States)

    Ellis, Russell A.; Berdoyes, Michel

    2002-07-01

    The history of over 25 years of cooperation between Pratt & Whitney, San Jose, CA, USA and Snecma Moteurs, Le Haillan, France in solid rocket motor and, in one case, liquid rocket engine technology is presented. Cooperative efforts resulted in achievements that likely would not have been realized individually. The combination of resources and technologies resulted in synergistic benefits and advancement of the state of the art in rocket motors and components. Discussions begun between the two companies in the early 1970's led to the first cooperative project, demonstration of an advanced apogee motor nozzle, during the mid 1970's. Shortly thereafter advanced carboncarbon (CC) throat materials from Snecma were comparatively tested with other materials in a P&W program funded by the USAF. Use of Snecma throat materials in CSD Tomahawk boosters followed. Advanced space motors were jointly demonstrated in company-funded joint programs in the late 1970's and early 1980's: an advanced space motor with an extendible exit cone and an all-composite advanced space motor that included a composite chamber polar adapter. Eight integral-throat entrances (ITEs) of 4D and 6D construction were tested by P&W for Snecma in 1982. Other joint programs in the 1980's included test firing of a "membrane" CC exit cone, and integral throat and exit cone (ITEC) nozzle incorporating NOVOLTEX® SEPCARB® material. A variation of this same material was demonstrated as a chamber aft polar boss in motor firings that included demonstration of composite material hot gas valve thrust vector control (TVC). In the 1990's a supersonic splitline flexseal nozzle was successfully demonstrated by the two companies as part of a US Integrated High Payoff Rocket Propulsion Technology (IHPRPT) program effort. Also in the mid-1990s the NOVOLTEX® SEPCARB® material, so successful in solid rocket motor application, was successfully applied to a liquid engine nozzle extension. The first cooperative

  8. Star-grain rocket motor - nonsteady internal ballistics

    Energy Technology Data Exchange (ETDEWEB)

    Loncaric, S.; Greatrix, D.R.; Fawaz, Z. [Ryerson University, Dept. of Aerospace Engineering, Toronto (Canada)

    2004-01-01

    The nonsteady internal ballistics of a star-grain solid-propellant rocket motor are investigated through a numerical simulation model that incorporates both the internal flow and surrounding structure. The effects of structural vibration on burning rate augmentation and wave development in nonsteady operation are demonstrated. The amount of damping plays a role in influencing the predicted axial combustion instability symptoms of the motor. The variation in oscillation frequencies about a given star grain section periphery, and along the grain with different levels of burn-back, also influences the means by which the local acceleration drives the combustion and flow behaviour. (authors)

  9. Five-Segment Solid Rocket Motor Development Status

    Science.gov (United States)

    Priskos, Alex S.

    2012-01-01

    In support of the National Aeronautics and Space Administration (NASA), Marshall Space Flight Center (MSFC) is developing a new, more powerful solid rocket motor for space launch applications. To minimize technical risks and development costs, NASA chose to use the Space Shuttle s solid rocket boosters as a starting point in the design and development. The new, five segment motor provides a greater total impulse with improved, more environmentally friendly materials. To meet the mass and trajectory requirements, the motor incorporates substantial design and system upgrades, including new propellant grain geometry with an additional segment, new internal insulation system, and a state-of-the art avionics system. Significant progress has been made in the design, development and testing of the propulsion, and avionics systems. To date, three development motors (one each in 2009, 2010, and 2011) have been successfully static tested by NASA and ATK s Launch Systems Group in Promontory, UT. These development motor tests have validated much of the engineering with substantial data collected, analyzed, and utilized to improve the design. This paper provides an overview of the development progress on the first stage propulsion system.

  10. On Nonlinear Combustion Instability in Liquid Propellant Rocket Motors

    Science.gov (United States)

    Sims, J. D. (Technical Monitor); Flandro, Gary A.; Majdalani, Joseph; Sims, Joseph D.

    2004-01-01

    All liquid propellant rocket instability calculations in current use have limited value in the predictive sense and serve mainly as a correlating framework for the available data sets. The well-known n-t model first introduced by Crocco and Cheng in 1956 is still used as the primary analytical tool of this type. A multitude of attempts to establish practical analytical methods have achieved only limited success. These methods usually produce only stability boundary maps that are of little use in making critical design decisions in new motor development programs. Recent progress in understanding the mechanisms of combustion instability in solid propellant rockets"' provides a firm foundation for a new approach to prediction, diagnosis, and correction of the closely related problems in liquid motor instability. For predictive tools to be useful in the motor design process, they must have the capability to accurately determine: 1) time evolution of the pressure oscillations and limit amplitude, 2) critical triggering pulse amplitude, and 3) unsteady heat transfer rates at injector surfaces and chamber walls. The method described in this paper relates these critical motor characteristics directly to system design parameters. Inclusion of mechanisms such as wave steepening, vorticity production and transport, and unsteady detonation wave phenomena greatly enhance the representation of key features of motor chamber oscillatory behavior. The basic theoretical model is described and preliminary computations are compared to experimental data. A plan to develop the new predictive method into a comprehensive analysis tool is also described.

  11. Using PDV to Understand Damage in Rocket Motor Propellants

    Science.gov (United States)

    Tear, Gareth; Chapman, David; Ottley, Phillip; Proud, William; Gould, Peter; Cullis, Ian

    2017-06-01

    There is a continuing requirement to design and manufacture insensitive munition (IM) rocket motors for in-service use under a wide range of conditions, particularly due to shock initiation and detonation of damaged propellant spalled across the central bore of the rocket motor (XDT). High speed photography has been crucial in determining this behaviour, however attempts to model the dynamic behaviour are limited by the lack of precision particle and wave velocity data with which to validate against. In this work Photonic Doppler Velocimetery (PDV) has been combined with high speed video to give accurate point velocity and timing measurements of the rear surface of a propellant block impacted by a fragment travelling upto 1.4 km s-1. By combining traditional high speed video with PDV through a dichroic mirror, the point of velocity measurement within the debris cloud has been determined. This demonstrates a new capability to characterise the damage behaviour of a double base rocket motor propellant and hence validate the damage and fragmentation algorithms used in the numerical simulations.

  12. Integration of Flex Nozzle System and Electro Hydraulic Actuators to Solid Rocket Motors

    Science.gov (United States)

    Nayani, Kishore Nath; Bajaj, Dinesh Kumar

    2017-10-01

    A rocket motor assembly comprised of solid rocket motor and flex nozzle system. Integration of flex nozzle system and hydraulic actuators to the solid rocket motors are done after transportation to the required place where integration occurred. The flex nozzle system is integrated to the rocket motor in horizontal condition and the electro hydraulic actuators are assembled to the flex nozzle systems. The electro hydraulic actuators are connected to the hydraulic power pack to operate the actuators. The nozzle-motor critical interface are insulation diametrical compression, inhibition resin-28, insulation facial compression, shaft seal `O' ring compression and face seal `O' ring compression.

  13. Direct electrical arc ignition of hybrid rocket motors

    Science.gov (United States)

    Judson, Michael I., Jr.

    Hybrid rockets motors provide distinct safety advantages when compared to traditional liquid or solid propellant systems, due to the inherent stability and relative inertness of the propellants prior to established combustion. As a result of this inherent propellant stability, hybrid motors have historically proven difficult to ignite. State of the art hybrid igniter designs continue to require solid or liquid reactants distinct from the main propellants. These ignition methods however, reintroduce to the hybrid propulsion system the safety and complexity disadvantages associated with traditional liquid or solid propellants. The results of this study demonstrate the feasibility of a novel direct electrostatic arc ignition method for hybrid motors. A series of small prototype stand-alone thrusters demonstrating this technology were successfully designed and tested using Acrylonitrile Butadiene Styrene (ABS) plastic and Gaseous Oxygen (GOX) as propellants. Measurements of input voltage and current demonstrated that arc-ignition will occur using as little as 10 watts peak power and less than 5 joules total energy. The motor developed for the stand-alone small thruster was adapted as a gas generator to ignite a medium-scale hybrid rocket motor using nitrous oxide /and HTPB as propellants. Multiple consecutive ignitions were performed. A large data set as well as a collection of development `lessons learned' were compiled to guide future development and research. Since the completion of this original groundwork research, the concept has been developed into a reliable, operational igniter system for a 75mm hybrid motor using both gaseous oxygen and liquid nitrous oxide as oxidizers. A development map of the direct spark ignition concept is presented showing the flow of key lessons learned between this original work and later follow on development.

  14. Nutation instability of spinning solid rocket motor spacecraft

    Directory of Open Access Journals (Sweden)

    Dan YANG

    2017-08-01

    Full Text Available The variation of mass, and moment of inertia of a spin-stabilized spacecraft leads to concern about the nutation instability. Here a careful analysis on the nutation instability is performed on a spacecraft propelled by solid rocket booster (SRB. The influences of specific solid propellant designs on transversal angular velocity are discussed. The results show that the typical SRB of End Burn suppresses the non-principal axial angular velocity. On the contrary, the frequently used SRB of Radial Burn could amplify the transversal angular velocity. The nutation instability caused by a design of Radial Burn could be remedied by the addition of End Burn at the same time based on the study of the combination design of both End Burn and Radial Burn. The analysis of the results proposes the design conception of how to control the nutation motion. The method is suitable to resolve the nutation instability of solid rocket motor with complex propellant patterns.

  15. Solid propellant processing factor in rocket motor design

    Science.gov (United States)

    1971-01-01

    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.

  16. The Potential for Ozone Depletion in Solid Rocket Motor Plumes by Heterogeneous Chemistry

    National Research Council Canada - National Science Library

    Hanning-Lee, M

    1996-01-01

    ... (hydroxylated alumina), respectively, over the temperature range -60 to 200 degrees C. This work addresses the potential for stratospheric ozone depletion by launch vehicle solid rocket motor exhaust...

  17. Design and Experimental Study on Spinning Solid Rocket Motor

    Science.gov (United States)

    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.

  18. Molecular beam sampling from a rocket-motor combustion chamber

    International Nuclear Information System (INIS)

    Houseman, John; Young, W.S.

    1974-01-01

    A molecular-beam mass-spectrometer sampling apparatus has been developed to study the reactive species concentrations as a function of position in a rocket-motor combustion chamber. Unique design features of the sampling system include (a) the use of a multiple-nozzle end plate for preserving the nonuniform properties of the flow field inside the combustion chamber, (b) the use of a water-injection heat shield, and (c) the use of a 300 CFM mechanical pump for the first vacuum stage (eliminating the use of a huge conventional oil booster pump). Preliminary rocket-motor tests have been performed using the highly reactive propellants nitrogen tetroxide/hydrazine (N 2 O 4 /N 2 H 4 ) at an oxidizer/fuel ratio of 1.2 by weight. The combustion-chamber pressure is approximately 60psig. Qualitative results on unreacted oxidizer/fuel ratio, relative abundance of oxidizer and fuel fragments, and HN 3 distribution across the chamber are presented

  19. Maturation of Structural Health Management Systems for Solid Rocket Motors

    Science.gov (United States)

    Quing, Xinlin; Beard, Shawn; Zhang, Chang

    2011-01-01

    Concepts of an autonomous and automated space-compliant diagnostic system were developed for conditioned-based maintenance (CBM) of rocket motors for space exploration vehicles. The diagnostic system will provide real-time information on the integrity of critical structures on launch vehicles, improve their performance, and greatly increase crew safety while decreasing inspection costs. Using the SMART Layer technology as a basis, detailed procedures and calibration techniques for implementation of the diagnostic system were developed. The diagnostic system is a distributed system, which consists of a sensor network, local data loggers, and a host central processor. The system detects external impact to the structure. The major functions of the system include an estimate of impact location, estimate of impact force at impacted location, and estimate of the structure damage at impacted location. This system consists of a large-area sensor network, dedicated multiple local data loggers with signal processing and data analysis software to allow for real-time, in situ monitoring, and longterm tracking of structural integrity of solid rocket motors. Specifically, the system could provide easy installation of large sensor networks, onboard operation under harsh environments and loading, inspection of inaccessible areas without disassembly, detection of impact events and impact damage in real-time, and monitoring of a large area with local data processing to reduce wiring.

  20. Reusable Solid Rocket Motor - V(RSRMV)Nozzle Forward Nose Ring Thermo-Structural Modeling

    Science.gov (United States)

    Clayton, J. Louie

    2012-01-01

    During the developmental static fire program for NASAs Reusable Solid Rocket Motor-V (RSRMV), an anomalous erosion condition appeared on the nozzle Carbon Cloth Phenolic nose ring that had not been observed in the space shuttle RSRM program. There were regions of augmented erosion located on the bottom of the forward nose ring (FNR) that measured nine tenths of an inch deeper than the surrounding material. Estimates of heating conditions for the RSRMV nozzle based on limited char and erosion data indicate that the total heat loading into the FNR, for the new five segment motor, is about 40-50% higher than the baseline shuttle RSRM nozzle FNR. Fault tree analysis of the augmented erosion condition has lead to a focus on a thermomechanical response of the material that is outside the existing experience base of shuttle CCP materials for this application. This paper provides a sensitivity study of the CCP material thermo-structural response subject to the design constraints and heating conditions unique to the RSRMV Forward Nose Ring application. Modeling techniques are based on 1-D thermal and porous media calculations where in-depth interlaminar loading conditions are calculated and compared to known capabilities at elevated temperatures. Parameters such as heat rate, in-depth pressures and temperature, degree of char, associated with initiation of the mechanical removal process are quantified and compared to a baseline thermo-chemical material removal mode. Conclusions regarding postulated material loss mechanisms are offered.

  1. Laser Shearography Inspection of TPS (Thermal Protection System) Cork on RSRM (Reusable Solid Rocket Motors)

    Science.gov (United States)

    Lingbloom, Mike; Plaia, Jim; Newman, John

    2006-01-01

    Laser Shearography is a viable inspection method for detection of de-bonds and voids within the external TPS (thermal protection system) on to the Space Shuttle RSRM (reusable solid rocket motors). Cork samples with thicknesses up to 1 inch were tested at the LTI (Laser Technology Incorporated) laboratory using vacuum-applied stress in a vacuum chamber. The testing proved that the technology could detect cork to steel un-bonds using vacuum stress techniques in the laboratory environment. The next logical step was to inspect the TPS on a RSRM. Although detailed post flight inspection has confirmed that ATK Thiokol's cork bonding technique provides a reliable cork to case bond, due to the Space Shuttle Columbia incident there is a great interest in verifying bond-lines on the external TPS. This interest provided and opportunity to inspect a RSRM motor with Laser Shearography. This paper will describe the laboratory testing and RSRM testing that has been performed to date. Descriptions of the test equipment setup and techniques for data collection and detailed results will be given. The data from the test show that Laser Shearography is an effective technology and readily adaptable to inspect a RSRM.

  2. Development of a new generation solid rocket motor ignition computer code

    Science.gov (United States)

    Foster, Winfred A., Jr.; Jenkins, Rhonald M.; Ciucci, Alessandro; Johnson, Shelby D.

    1994-01-01

    This report presents the results of experimental and numerical investigations of the flow field in the head-end star grain slots of the Space Shuttle Solid Rocket Motor. This work provided the basis for the development of an improved solid rocket motor ignition transient code which is also described in this report. The correlation between the experimental and numerical results is excellent and provides a firm basis for the development of a fully three-dimensional solid rocket motor ignition transient computer code.

  3. Combustion Stability Assessments of the Black Brant Solid Rocket Motor

    Science.gov (United States)

    Fischbach, Sean

    2014-01-01

    The Black Brant variation of the Standard Brant developed in the 1960's has been a workhorse motor of the NASA Sounding Rocket Project Office (SRPO) since the 1970's. In March 2012, the Black Brant Mk1 used on mission 36.277 experienced combustion instability during a flight at White Sands Missile Range, the third event in the last four years, the first occurring in November, 2009, the second in April 2010. After the 2010 event the program has been increasing the motor's throat diameter post-delivery with the goal of lowering the chamber pressure and increasing the margin against combustion instability. During the most recent combustion instability event, the vibrations exceeded the qualification levels for the Flight Termination System. The present study utilizes data generated from T-burner testing of multiple Black Brant propellants at the Naval Air Warfare Center at China Lake, to improve the combustion stability predictions for the Black Brant Mk1 and to generate new predictions for the Mk2. Three unique one dimensional (1-D) stability models were generated, representing distinct Black Brant flights, two of which experienced instabilities. The individual models allowed for comparison of stability characteristics between various nozzle configurations. A long standing "rule of thumb" states that increased stability margin is gained by increasing the throat diameter. In contradiction to this experience based rule, the analysis shows that little or no margin is gained from a larger throat diameter. The present analysis demonstrates competing effects resulting from an increased throat diameter accompanying a large response function. As is expected, more acoustic energy was expelled through the nozzle, but conversely more acoustic energy was generated due to larger gas velocities near the propellant surfaces.

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

    Science.gov (United States)

    Cai, Guobiao; Li, Chengen; Tian, Hui

    2016-11-01

    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.

  5. Applied algorithm in the liner inspection of solid rocket motors

    Science.gov (United States)

    Hoffmann, Luiz Felipe Simões; Bizarria, Francisco Carlos Parquet; Bizarria, José Walter Parquet

    2018-03-01

    In rocket motors, the bonding between the solid propellant and thermal insulation is accomplished by a thin adhesive layer, known as liner. The liner application method involves a complex sequence of tasks, which includes in its final stage, the surface integrity inspection. Nowadays in Brazil, an expert carries out a thorough visual inspection to detect defects on the liner surface that may compromise the propellant interface bonding. Therefore, this paper proposes an algorithm that uses the photometric stereo technique and the K-nearest neighbor (KNN) classifier to assist the expert in the surface inspection. Photometric stereo allows the surface information recovery of the test images, while the KNN method enables image pixels classification into two classes: non-defect and defect. Tests performed on a computer vision based prototype validate the algorithm. The positive results suggest that the algorithm is feasible and when implemented in a real scenario, will be able to help the expert in detecting defective areas on the liner surface.

  6. Cooperative Threat Reduction: Solid Rocket Motor Disposition Facility Project (D-2003-131)

    National Research Council Canada - National Science Library

    2003-01-01

    .... DoD contracted with Lockheed Martin Advanced Environmental Systems for $52.4 million to design, develop, fabricate, and test a closed burn, solid rocket motor disposition facility for the Russian Federation in April 1997...

  7. Experimental validation of solid rocket motor damping models

    Science.gov (United States)

    Riso, Cristina; Fransen, Sebastiaan; Mastroddi, Franco; Coppotelli, Giuliano; Trequattrini, Francesco; De Vivo, Alessio

    2017-12-01

    In design and certification of spacecraft, payload/launcher coupled load analyses are performed to simulate the satellite dynamic environment. To obtain accurate predictions, the system damping properties must be properly taken into account in the finite element model used for coupled load analysis. This is typically done using a structural damping characterization in the frequency domain, which is not applicable in the time domain. Therefore, the structural damping matrix of the system must be converted into an equivalent viscous damping matrix when a transient coupled load analysis is performed. This paper focuses on the validation of equivalent viscous damping methods for dynamically condensed finite element models via correlation with experimental data for a realistic structure representative of a slender launch vehicle with solid rocket motors. A second scope of the paper is to investigate how to conveniently choose a single combination of Young's modulus and structural damping coefficient—complex Young's modulus—to approximate the viscoelastic behavior of a solid propellant material in the frequency band of interest for coupled load analysis. A scaled-down test article inspired to the Z9-ignition Vega launcher configuration is designed, manufactured, and experimentally tested to obtain data for validation of the equivalent viscous damping methods. The Z9-like component of the test article is filled with a viscoelastic material representative of the Z9 solid propellant that is also preliminarily tested to investigate the dependency of the complex Young's modulus on the excitation frequency and provide data for the test article finite element model. Experimental results from seismic and shock tests performed on the test configuration are correlated with numerical results from frequency and time domain analyses carried out on its dynamically condensed finite element model to assess the applicability of different equivalent viscous damping methods to describe

  8. Experimental validation of solid rocket motor damping models

    Science.gov (United States)

    Riso, Cristina; Fransen, Sebastiaan; Mastroddi, Franco; Coppotelli, Giuliano; Trequattrini, Francesco; De Vivo, Alessio

    2018-06-01

    In design and certification of spacecraft, payload/launcher coupled load analyses are performed to simulate the satellite dynamic environment. To obtain accurate predictions, the system damping properties must be properly taken into account in the finite element model used for coupled load analysis. This is typically done using a structural damping characterization in the frequency domain, which is not applicable in the time domain. Therefore, the structural damping matrix of the system must be converted into an equivalent viscous damping matrix when a transient coupled load analysis is performed. This paper focuses on the validation of equivalent viscous damping methods for dynamically condensed finite element models via correlation with experimental data for a realistic structure representative of a slender launch vehicle with solid rocket motors. A second scope of the paper is to investigate how to conveniently choose a single combination of Young's modulus and structural damping coefficient—complex Young's modulus—to approximate the viscoelastic behavior of a solid propellant material in the frequency band of interest for coupled load analysis. A scaled-down test article inspired to the Z9-ignition Vega launcher configuration is designed, manufactured, and experimentally tested to obtain data for validation of the equivalent viscous damping methods. The Z9-like component of the test article is filled with a viscoelastic material representative of the Z9 solid propellant that is also preliminarily tested to investigate the dependency of the complex Young's modulus on the excitation frequency and provide data for the test article finite element model. Experimental results from seismic and shock tests performed on the test configuration are correlated with numerical results from frequency and time domain analyses carried out on its dynamically condensed finite element model to assess the applicability of different equivalent viscous damping methods to describe

  9. Transient simulation of chamber flowfield in a rod-and-tube configuration solid rocket motor

    International Nuclear Information System (INIS)

    Weaver, J.T.; Stowe, R.A.

    2004-01-01

    Currently, DRDC Valcartier of the Canadian Department of National Defence is designing a prototype rod-and-tube configuration solid propellant rocket motor that will propel a hypersonic velocity missile. This configuration will incorporate a very low port-to-throat area ratio, which in turn results in very high velocity propellant gas traveling across burning propellant surfaces, particularly near the nozzle end of the rocket. This causes an augmentation in the propellant burning rate. While numerical and lumped parameter models are available to design and analyze solid propellant rocket motors and nozzles, many of them provide solutions based on the assumption of quasi-steady flow. Due to the high pressure, high velocity and highly transient nature of the flows expected in the motor under design, it is believed that a CFD simulation will better model the time-dependent phenomena that occur during the functioning of a motor of this type. This simulation couples the fluid dynamics and heat transfer of the gas flowfield within the rocket port to the nozzle and the regression rate of the propellant. By incorporating the regression of the propellant surfaces into the model, the information provided by the resulting time-accurate solution will enable a much improved understanding of the flow phenomena within this rod-and-tube grain motor and a better prediction of the internal ballistics of the motor, which in turn will help in the design of both the motor and the nozzle. (author)

  10. Transient simulation of chamber flowfield in a rod-and-tube configuration solid rocket motor

    Energy Technology Data Exchange (ETDEWEB)

    Weaver, J.T. [Carleton Univ., Ottawa, Ontario (Canada)]. E-mail: jrweaver@storm.ca; Stowe, R.A. [Defence R and D Canada - Valcartier, Val-Belair, Quebec (Canada)

    2004-07-01

    Currently, DRDC Valcartier of the Canadian Department of National Defence is designing a prototype rod-and-tube configuration solid propellant rocket motor that will propel a hypersonic velocity missile. This configuration will incorporate a very low port-to-throat area ratio, which in turn results in very high velocity propellant gas traveling across burning propellant surfaces, particularly near the nozzle end of the rocket. This causes an augmentation in the propellant burning rate. While numerical and lumped parameter models are available to design and analyze solid propellant rocket motors and nozzles, many of them provide solutions based on the assumption of quasi-steady flow. Due to the high pressure, high velocity and highly transient nature of the flows expected in the motor under design, it is believed that a CFD simulation will better model the time-dependent phenomena that occur during the functioning of a motor of this type. This simulation couples the fluid dynamics and heat transfer of the gas flowfield within the rocket port to the nozzle and the regression rate of the propellant. By incorporating the regression of the propellant surfaces into the model, the information provided by the resulting time-accurate solution will enable a much improved understanding of the flow phenomena within this rod-and-tube grain motor and a better prediction of the internal ballistics of the motor, which in turn will help in the design of both the motor and the nozzle. (author)

  11. Electric Propellant Solid Rocket Motor Thruster Results Enabling Small Satellites

    OpenAIRE

    Koehler, Frederick; Langhenry, Mark; Summers, Matt; Villarreal, James; Villarreal, Thomas

    2017-01-01

    Raytheon Missile Systems has developed and tested true on/off/restart solid propellant thrusters which are controlled only by electrical current. This new patented class of energetic rocket propellant is safe, controllable and simple. The range of applications for this game changing technology includes attitude control systems and a safe alternative to higher impulse space satellite thrusters. Described herein are descriptions and performance data for several small electric propellant solid r...

  12. Study of the Deposition of Ammonium Perchlorate Following the Static Firing of MK-58 Rocket Motors

    Science.gov (United States)

    2008-10-01

    hyperthyroidism , gas generators, electrolytes for lithium cells, and as chemical reagents. The occurrence of perchlorate in the environment is...and prevent their movement by the rocket motor plume (Fig. 5). The water in the traps was collected using 1-l amber glass containers and the exact...them. On day one, after the firing of the second motor, heavy rain and lightning prevented the collection of samples from the witness plates. Only

  13. Simulation of Axial Combustion Instability Development and Suppression in Solid Rocket Motors

    OpenAIRE

    David R. Greatrix

    2009-01-01

    In the design of solid-propellant rocket motors, the ability to understand and predict the expected behaviour of a given motor under unsteady conditions is important. Research towards predicting, quantifying, and ultimately suppressing undesirable strong transient axial combustion instability symptoms necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. An updated numerical model incorporating recent developments in predi...

  14. Regarding the perturbed operating process of DB propellant rocket motor at extreme initial grain temperatures

    Directory of Open Access Journals (Sweden)

    Ioan ION

    2012-03-01

    Full Text Available Despite many decades of study, the combustion instability of several DB propellants is still of particular concern, especially at extreme grain temperature conditions of rocket motor operating. The purpose of the first part of the paper is to give an overview of our main experimental results on combustion instabilities and pressure oscillations in DB propellant segmented grain rocket motors (SPRM-01, large L/D ratio, working at extreme initial grain temperatures. Thus, we recorded some particular pressure-time traces with significant perturbed pressure signal that was FFT analysed. An updated mathematical model incorporating transient frequency-dependent combustion response, in conjunction with pressure-dependent burning, is applied to investigate and predict the DB propellant combustion instability phenomenon. The susceptibility of the tested motor SPRM-01 with DB propellant to get a perturbed working and to go unstable with pressure was evidenced and this risk has to be evaluated. In the last part of our paper we evaluated the influence of recorded perturbed thrust on the rocket behaviour on the trajectory. The study revealed that at firing-table initial conditions, this kind of perturbed motor operating may not lead to an unstable rocket flight, but the ballistic parameters would be influenced in an unacceptable manner.

  15. Numerical Simulations of Flow and Fuel Regression Rate Coupling in Hybrid Rocket Motors

    Directory of Open Access Journals (Sweden)

    Marius STOIA-DJESKA

    2017-03-01

    Full Text Available The hybrid propulsion offers some remarkable advantages like high safety and high specific impulse and thus it is considered a promising technology for the next generation launchers and space systems. The purpose of this work is to validate a design tool for hybrid rocket motors (HRM through numerical simulations.

  16. Reusable Solid Rocket Motor - Accomplishments, Lessons, and a Culture of Success

    Science.gov (United States)

    Moore, Dennis R.; Phelps, Willie J.

    2011-01-01

    The Reusable Solid Rocket Motor represents the largest solid rocket motor ever flown and the only human rated solid motor. Each Reusable Solid Rocket Motor (RSRM) provides approximately 3-million lb of thrust to lift the integrated Space Shuttle vehicle from the launch pad. The motors burn out approximately 2 minutes later, separate from the vehicle and are recovered and refurbished. The size of the motor and the need for high reliability were challenges. Thrust shaping, via shaping of the propellant grain, was needed to limit structural loads during ascent. The motor design evolved through several block upgrades to increase performance and to increase safety and reliability. A major redesign occurred after STS-51L with the Redesigned Solid Rocket Motor. Significant improvements in the joint sealing systems were added. Design improvements continued throughout the Program via block changes with a number of innovations including development of low temperature o-ring materials and incorporation of a unique carbon fiber rope thermal barrier material. Recovery of the motors and post flight inspection improved understanding of hardware performance, and led to key design improvements. Because of the multidecade program duration material obsolescence was addressed, and requalification of materials and vendors was sometimes needed. Thermal protection systems and ablatives were used to protect the motor cases and nozzle structures. Significant understanding of design and manufacturing features of the ablatives was developed during the program resulting in optimization of design features and processing parameters. The project advanced technology in eliminating ozone-depleting materials in manufacturing processes and the development of an asbestos-free case insulation. Manufacturing processes for the large motor components were unique and safety in the manufacturing environment was a special concern. Transportation and handling approaches were also needed for the large

  17. Research on combustion instability and application to solid propellant rocket motors. II.

    Science.gov (United States)

    Culick, F. E. C.

    1972-01-01

    Review of the current state of analyses of combustion instability in solid-propellant rocket motors, citing appropriate measurements and observations. The work discussed has become increasingly important, both for the interpretation of laboratory data and for predicting the transient behavior of disturbances in full-scale motors. Two central questions are considered - namely, linear stability and nonlinear behavior. Several classes of problems are discussed as special cases of a general approach to the analysis of combustion instability. Application to motors, and particularly the limitations presently understood, are stressed.

  18. Laser welding of maraging steel rocket motor casing

    CSIR Research Space (South Africa)

    Van Rooyen, C

    2009-11-01

    Full Text Available This presentation looks at the experimental procedure and results of laser welding of maraging steel rocker motor casing. It concludes that a fracture occurred in weld metal of autogenous welding and that a fracture occurred in base material when...

  19. Numerical simulation of a liquid propellant rocket motor

    Science.gov (United States)

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

    2001-03-01

    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 external 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.

  20. Internal Flow Analysis of Large L/D Solid Rocket Motors

    Science.gov (United States)

    Laubacher, Brian A.

    2000-01-01

    Traditionally, Solid Rocket Motor (SRM) internal ballistic performance has been analyzed and predicted with either zero-dimensional (volume filling) codes or one-dimensional ballistics codes. One dimensional simulation of SRM performance is only necessary for ignition modeling, or for motors that have large length to port diameter ratios which exhibit an axial "pressure drop" during the early burn times. This type of prediction works quite well for many types of motors, however, when motor aspect ratios get large, and port to throat ratios get closer to one, two dimensional effects can become significant. The initial propellant grain configuration for the Space Shuttle Reusable Solid Rocket Motor (RSRM) was analyzed with 2-D, steady, axi-symmetric computational fluid dynamics (CFD). The results of the CFD analysis show that the steady-state performance prediction at the initial burn geometry, in general, agrees well with 1-D transient prediction results at an early time, however, significant features of the 2-D flow are captured with the CFD results that would otherwise go unnoticed. Capturing these subtle differences gives a greater confidence to modeling accuracy, and additional insight with which to model secondary internal flow effects like erosive burning. Detailed analysis of the 2-D flowfield has led to the discovery of its hidden 1-D isentropic behavior, and provided the means for a thorough and simplified understanding of internal solid rocket motor flow. Performance parameters such as nozzle stagnation pressure, static pressure drop, characteristic velocity, thrust and specific impulse are discussed in detail and compared for different modeling and prediction methods. The predicted performance using both the 1-D codes and the CFD results are compared with measured data obtained from static tests of the RSRM. The differences and limitations of predictions using ID and 2-D flow fields are discussed and some suggestions for the design of large L/D motors and

  1. Near noise field characteristics of Nike rocket motors for application to space vehicle payload acoustic qualification

    Science.gov (United States)

    Hilton, D. A.; Bruton, D.

    1977-01-01

    Results of a series of noise measurements that were made under controlled conditions during the static firing of two Nike solid propellant rocket motors are presented. The usefulness of these motors as sources for general spacecraft noise testing was assessed, and the noise expected in the cargo bay of the orbiter was reproduced. Brief descriptions of the Nike motor, the general procedures utilized for the noise tests, and representative noise data including overall sound pressure levels, one third octave band spectra, and octave band spectra were reviewed. Data are presented on two motors of different ages in order to show the similarity between noise measurements made on motors having different loading dates. The measured noise from these tests is then compared to that estimated for the space shuttle orbiter cargo bay.

  2. Development of Displacement Gages Exposed to Solid Rocket Motor Internal Environments

    Science.gov (United States)

    Bolton, D. E.; Cook, D. J.

    2003-01-01

    The Space Shuttle Reusable Solid Rocket Motor (RSRM) has three non-vented segment-to-segment case field joints. These joints use an interference fit J-joint that is bonded at assembly with a Pressure Sensitive Adhesive (PSA) inboard of redundant O-ring seals. Full-scale motor and sub-scale test article experience has shown that the ability to preclude gas leakage past the J-joint is a function of PSA type, joint moisture from pre-assembly humidity exposure, and the magnitude of joint displacement during motor operation. To more accurately determine the axial displacements at the J-joints, two thermally durable displacement gages (one mechanical and one electrical) were designed and developed. The mechanical displacement gage concept was generated first as a non-electrical, self-contained gage to capture the maximum magnitude of the J-joint motion. When it became feasible, the electrical displacement gage concept was generated second as a real-time linear displacement gage. Both of these gages were refined in development testing that included hot internal solid rocket motor environments and simulated vibration environments. As a result of this gage development effort, joint motions have been measured in static fired RSRM J-joints where intentional venting was produced (Flight Support Motor #8, FSM-8) and nominal non-vented behavior occurred (FSM-9 and FSM-10). This data gives new insight into the nominal characteristics of the three case J-joint positions (forward, center and aft) and characteristics of some case J-joints that became vented during motor operation. The data supports previous structural model predictions. These gages will also be useful in evaluating J-joint motion differences in a five-segment Space Shuttle solid rocket motor.

  3. Study of organic ablative thermal-protection coating for solid rocket motor

    Science.gov (United States)

    Hua, Zenggong

    1992-06-01

    A study is conducted to find a new interior thermal-protection material that possesses good thermal-protection performance and simple manufacturing possibilities. Quartz powder and Cr2O3 are investigated using epoxy resin as a binder and Al2O3 as the burning inhibitor. Results indicate that the developed thermal-protection coating is suitable as ablative insulation material for solid rocket motors.

  4. Numerical Evaluation of the Use of Aluminum Particles for Enhancing Solid Rocket Motor Combustion Stability

    OpenAIRE

    David Greatrix

    2015-01-01

    The ability to 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 typically necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. On the mitigation side, one in practice sees the use of inert or reactive particles for the suppression of pressure wave ...

  5. Particle size determination in small solid propellant rocket motors using the diffractively scattered light method.

    OpenAIRE

    Cramer, Robert Grewelle.

    1982-01-01

    Approved for public release; distribution unlimited A dual beam apparatus was developed which simultaneously measured particle size (D32) at the entrance and exit of an exhaust nozzle of a small solid propellant rocket motor. The diameters were determined using measurements of dif fractiveiy scattered laser power spectra. The apparatus was calibrated by using spherical glass beads and aluminum oxide powder. Measurements were successfully made at both locations. Because of...

  6. Development and Characterization of Fast Burning Solid Fuels/Propellants for Hybrid Rocket Motors with High Volumetric Efficiency

    Data.gov (United States)

    National Aeronautics and Space Administration — The objective of this proposed work is to develop several fast burning solid fuels/fuel-rich solid propellants for hybrid rocket motor applications. In the...

  7. Finite element analysis of propellant of solid rocket motor during ship motion

    Directory of Open Access Journals (Sweden)

    Kai Qu

    2013-03-01

    Full Text Available In order to simulate the stress and strain of solid rocket motors (SRMs, a finite element analysis model was established. The stress spectra of the SRM elements with respect to time in the case that the vessel cruises under a certain shipping condition were obtained by simulation. According to the analysis of the simulation results, a critical zone was confirmed, and the Mises stress amplitudes of the different critical zones were acquired. The results show that the maximum stress and strain of SRM are less than the maximum tensile strength and elongation, respectively, of the propellant. The cumulative damage of the motor must also be evaluated by random fatigue loading.

  8. Reusable Solid Rocket Motor - Accomplishment, Lessons, and a Culture of Success

    Science.gov (United States)

    Moore, D. R.; Phelps, W. J.

    2011-01-01

    The Reusable Solid Rocket Motor (RSRM) represents the largest solid rocket motor (SRM) ever flown and the only human-rated solid motor. High reliability of the RSRM has been the result of challenges addressed and lessons learned. Advancements have resulted by applying attention to process control, testing, and postflight through timely and thorough communication in dealing with all issues. A structured and disciplined approach was taken to identify and disposition all concerns. Careful consideration and application of alternate opinions was embraced. Focus was placed on process control, ground test programs, and postflight assessment. Process control is mandatory for an SRM, because an acceptance test of the delivered product is not feasible. The RSRM maintained both full-scale and subscale test articles, which enabled continuous improvement of design and evaluation of process control and material behavior. Additionally RSRM reliability was achieved through attention to detail in post flight assessment to observe any shift in performance. The postflight analysis and inspections provided invaluable reliability data as it enables observation of actual flight performance, most of which would not be available if the motors were not recovered. RSRM reusability offered unique opportunities to learn about the hardware. NASA is moving forward with the Space Launch System that incorporates propulsion systems that takes advantage of the heritage Shuttle and Ares solid motor programs. These unique challenges, features of the RSRM, materials and manufacturing issues, and design improvements will be discussed in the paper.

  9. Numerical investigation on the regression rate of hybrid rocket motor with star swirl fuel grain

    Science.gov (United States)

    Zhang, Shuai; Hu, Fan; Zhang, Weihua

    2016-10-01

    Although hybrid rocket motor is prospected to have distinct advantages over liquid and solid rocket motor, low regression rate and insufficient efficiency are two major disadvantages which have prevented it from being commercially viable. In recent years, complex fuel grain configurations are attractive in overcoming the disadvantages with the help of Rapid Prototyping technology. In this work, an attempt has been made to numerically investigate the flow field characteristics and local regression rate distribution inside the hybrid rocket motor with complex star swirl grain. A propellant combination with GOX and HTPB has been chosen. The numerical model is established based on the three dimensional Navier-Stokes equations with turbulence, combustion, and coupled gas/solid phase formulations. The calculated fuel regression rate is compared with the experimental data to validate the accuracy of numerical model. The results indicate that, comparing the star swirl grain with the tube grain under the conditions of the same port area and the same grain length, the burning surface area rises about 200%, the spatially averaged regression rate rises as high as about 60%, and the oxidizer can combust sufficiently due to the big vortex around the axis in the aft-mixing chamber. The combustion efficiency of star swirl grain is better and more stable than that of tube grain.

  10. Parametric Study of Design Options aecting Solid Rocket Motor Start-up and Onset of Pressure Oscillations

    OpenAIRE

    Di Giacinto, M.; Cavallini, E.; Favini, B.; Steelant, Johan

    2014-01-01

    The start-up represents a very critical phase during the whole operational life of solid rocket motors. This paper provides a detailed study of the eects on the ignition transient of the main design parameters of solid propellant motors. The analysis is made with the use of a Q1D unsteady model of solid rocket ignition transient, extensively validated in the frame of the VEGA program, for ignition transient predictions and reconstructions, during the last ten years. Two baseline soli...

  11. Validation of a Solid Rocket Motor Internal Environment Model

    Science.gov (United States)

    Martin, Heath T.

    2017-01-01

    In a prior effort, a thermal/fluid model of the interior of Penn State University's laboratory-scale Insulation Test Motor (ITM) was constructed to predict both the convective and radiative heat transfer to the interior walls of the ITM with a minimum of empiricism. These predictions were then compared to values of total and radiative heat flux measured in a previous series of ITM test firings to assess the capabilities and shortcomings of the chosen modeling approach. Though the calculated fluxes reasonably agreed with those measured during testing, this exercise revealed means of improving the fidelity of the model to, in the case of the thermal radiation, enable direct comparison of the measured and calculated fluxes and, for the total heat flux, compute a value indicative of the average measured condition. By replacing the P1-Approximation with the discrete ordinates (DO) model for the solution of the gray radiative transfer equation, the radiation intensity field in the optically thin region near the radiometer is accurately estimated, allowing the thermal radiation flux to be calculated on the heat-flux sensor itself, which was then compared directly to the measured values. Though the fully coupling the wall thermal response with the flow model was not attempted due to the excessive computational time required, a separate wall thermal response model was used to better estimate the average temperature of the graphite surfaces upstream of the heat flux gauges and improve the accuracy of both the total and radiative heat flux computations. The success of this modeling approach increases confidence in the ability of state-of-the-art thermal and fluid modeling to accurately predict SRM internal environments, offers corrections to older methods, and supplies a tool for further studies of the dynamics of SRM interiors.

  12. Real-Time Inhibitor Recession Measurements in Two Space Shuttle Reusable Solid Rocket Motors

    Science.gov (United States)

    McWhorter, B. B.; Ewing, M. E.; Bolton, D. E.; Albrechtsen, K. U.; Earnest, T. E.; Noble, T. C.; Longaker, M.

    2003-01-01

    Real-time internal motor insulation char line recession measurements have been evaluated for two full-scale static tests of the Space Shuttle Reusable Solid Rocket Motor (RSRM). These char line recession measurements were recorded on the forward facing propellant grain inhibitors to better understand the thermal performance of these inhibitors. The RSRM propellant grain inhibitors are designed to erode away during motor operation, thus making it difficult to use post-fire observations to determine inhibitor thermal performance. Therefore, this new internal motor instrumentation is invaluable in establishing an accurate understanding of inhibitor recession versus motor operation time. The data for the first test was presented at the 37th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit (AIAA 2001-3280) in July 2001. Since that time, a second full scale static test has delivered additional real-time data on inhibitor thermal performance. The evaluation of this data is presented in this paper. The second static test, in contrast to the first test, used a slightly different arrangement of instrumentation in the inhibitors. This instrumentation has yielded a better understanding of the inhibitor time dependent inboard tip recession. Graphs of inhibitor recession profiles with time are presented. Inhibitor thermal ablation models have been created from theoretical principals. The model predictions compare favorably with data from both tests. This verified modeling effort is important to support new inhibitor designs for a five segment Space Shuttle solid rocket motor. The internal instrumentation project on RSRM static tests is providing unique opportunities for other real-time internal motor measurements that could not otherwise be directly quantified.

  13. Experimental Study of the Swirling Oxidizer Flow in HTPB/N2O Hybrid Rocket Motor

    Directory of Open Access Journals (Sweden)

    Mohammad Mahdi Heydari

    2017-01-01

    Full Text Available Effects of swirling oxidizer flow on the performance of a HTPB/N2O Hybrid rocket motor were studied. A hybrid propulsion laboratory has been developed, to characterize internal ballistics characteristics of swirl flow hybrid motors and to define the operating parameters, like fuel regression rate, specific impulse, and characteristics velocity and combustion efficiency. Primitive variables, like pressure, thrust, temperature, and the oxidizer mass flow rate, were logged. A modular motor with 70 mm outer diameter and variable chamber length is designed for experimental analysis. The injector module has four tangential injectors and one axial injector. Liquid nitrous oxide (N2O as an oxidizer is injected at the head of combustion chamber into the motor. The feed system uses pressurized air as the pressurant. Two sets of tests have been performed. Some tests with axial and tangential oxidizer injection and a test with axial oxidizer injection were done. The test results show that the fuel grain regression rate has been improved by applying tangential oxidizer injection at the head of the motor. Besides, it was seen that combustion efficiency of motors with the swirl flow was about 10 percent more than motors with axial flow.

  14. Plume particle collection and sizing from static firing of solid rocket motors

    Science.gov (United States)

    Sambamurthi, Jay K.

    1995-01-01

    A unique dart system has been designed and built at the NASA Marshall Space Flight Center to collect aluminum oxide plume particles from the plumes of large scale solid rocket motors, such as the space shuttle RSRM. The capability of this system to collect clean samples from both the vertically fired MNASA (18.3% scaled version of the RSRM) motors and the horizontally fired RSRM motor has been demonstrated. The particle mass averaged diameters, d43, measured from the samples for the different motors, ranged from 8 to 11 mu m and were independent of the dart collection surface and the motor burn time. The measured results agreed well with those calculated using the industry standard Hermsen's correlation within the standard deviation of the correlation . For each of the samples analyzed from both MNASA and RSRM motors, the distribution of the cumulative mass fraction of the plume oxide particles as a function of the particle diameter was best described by a monomodal log-normal distribution with a standard deviation of 0.13 - 0.15. This distribution agreed well with the theoretical prediction by Salita using the OD3P code for the RSRM motor at the nozzle exit plane.

  15. Thermo-Structural Response Caused by Structure Gap and Gap Design for Solid Rocket Motor Nozzles

    Directory of Open Access Journals (Sweden)

    Lin Sun

    2016-06-01

    Full Text Available The thermo-structural response of solid rocket motor nozzles is widely investigated in the design of modern rockets, and many factors related to the material properties have been considered. However, little work has been done to evaluate the effects of structure gaps on the generation of flame leaks. In this paper, a numerical simulation was performed by the finite element method to study the thermo-structural response of a typical nozzle with consideration of the structure gap. Initial boundary conditions for thermo-structural simulation were defined by a quasi-1D model, and then coupled simulations of different gap size matching modes were conducted. It was found that frictional interface treatment could efficiently reduce the stress level. Based on the defined flame leak criteria, gap size optimization was carried out, and the best gap matching mode was determined for designing the nozzle. Testing experiment indicated that the simulation results from the proposed method agreed well with the experimental results. It is believed that the simulation method is effective for investigating thermo-structural responses, as well as designing proper gaps for solid rocket motor nozzles.

  16. Injection and swirl driven flowfields in solid and liquid rocket motors

    Science.gov (United States)

    Vyas, Anand B.

    In this work, we seek approximate analytical solutions to describe the bulk flow motion in certain types of solid and liquid rocket motors. In the case of an idealized solid rocket motor, a cylindrical double base propellant grain with steady regression rate is considered. The well known inviscid profile determined by Culick is extended here to include the effects of viscosity and steady grain regression. The approximate analytical solution for the cold flow is obtained from similarity principles, perturbation methods and the method of variation of parameters. The velocity, vorticity, pressure gradient and the shear stress distributions are determined and interpreted for different rates of wall regression and injection Reynolds number. The liquid propellant rocket engine considered here is based on a novel design that gives rise to a cyclonic flow. The resulting bidirectional motion is triggered by the tangential injection of an oxidizer just upstream of the chamber nozzle. Velocity, vorticity and pressure gradient distributions are determined for the bulk gas dynamics using a non-reactive inviscid model. Viscous corrections are then incorporated to explain the formation of a forced vortex near the core. Our results compare favorably with numerical simulations and experimental measurements obtained by other researchers. They also indicate that the bidirectional vortex in a cylindrical chamber is a physical solution of the Euler equations. In closing, we investigate the possibility of multi-directional flow behavior as predicted by Euler's equation and as reported recently in laboratory experiments.

  17. Simulation of Axial Combustion Instability Development and Suppression in Solid Rocket Motors

    Directory of Open Access Journals (Sweden)

    David R. Greatrix

    2009-03-01

    Full Text Available In the design of solid-propellant rocket motors, the ability to understand and predict the expected behaviour of a given motor under unsteady conditions is important. Research towards predicting, quantifying, and ultimately suppressing undesirable strong transient axial combustion instability symptoms necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. An updated numerical model incorporating recent developments in predicting negative and positive erosive burning, and transient, frequency-dependent combustion response, in conjunction with pressure-dependent and acceleration-dependent burning, is applied to the investigation of instability-related behaviour in a small cylindrical-grain motor. Pertinent key factors, like the initial pressure disturbance magnitude and the propellant's net surface heat release, are evaluated with respect to their influence on the production of instability symptoms. Two traditional suppression techniques, axial transitions in grain geometry and inert particle loading, are in turn evaluated with respect to suppressing these axial instability symptoms.

  18. Failure mode and effects analysis (FMEA) for the Space Shuttle solid rocket motor

    Science.gov (United States)

    Russell, D. L.; Blacklock, K.; Langhenry, M. T.

    1988-01-01

    The recertification of the Space Shuttle Solid Rocket Booster (SRB) and Solid Rocket Motor (SRM) has included an extensive rewriting of the Failure Mode and Effects Analysis (FMEA) and Critical Items List (CIL). The evolution of the groundrules and methodology used in the analysis is discussed and compared to standard FMEA techniques. Especially highlighted are aspects of the FMEA/CIL which are unique to the analysis of an SRM. The criticality category definitions are presented and the rationale for assigning criticality is presented. The various data required by the CIL and contribution of this data to the retention rationale is also presented. As an example, the FMEA and CIL for the SRM nozzle assembly is discussed in detail. This highlights some of the difficulties associated with the analysis of a system with the unique mission requirements of the Space Shuttle.

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

    OpenAIRE

    S. Saha; D. Chakraborty

    2016-01-01

    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 ...

  20. Indirect and direct methods for measuring a dynamic throat diameter in a solid rocket motor

    Science.gov (United States)

    Colbaugh, Lauren

    In a solid rocket motor, nozzle throat erosion is dictated by propellant composition, throat material properties, and operating conditions. Throat erosion has a significant effect on motor performance, so it must be accurately characterized to produce a good motor design. In order to correlate throat erosion rate to other parameters, it is first necessary to know what the throat diameter is throughout a motor burn. Thus, an indirect method and a direct method for determining throat diameter in a solid rocket motor are investigated in this thesis. The indirect method looks at the use of pressure and thrust data to solve for throat diameter as a function of time. The indirect method's proof of concept was shown by the good agreement between the ballistics model and the test data from a static motor firing. The ballistics model was within 10% of all measured and calculated performance parameters (e.g. average pressure, specific impulse, maximum thrust, etc.) for tests with throat erosion and within 6% of all measured and calculated performance parameters for tests without throat erosion. The direct method involves the use of x-rays to directly observe a simulated nozzle throat erode in a dynamic environment; this is achieved with a dynamic calibration standard. An image processing algorithm is developed for extracting the diameter dimensions from the x-ray intensity digital images. Static and dynamic tests were conducted. The measured diameter was compared to the known diameter in the calibration standard. All dynamic test results were within +6% / -7% of the actual diameter. Part of the edge detection method consists of dividing the entire x-ray image by an average pixel value, calculated from a set of pixels in the x-ray image. It was found that the accuracy of the edge detection method depends upon the selection of the average pixel value area and subsequently the average pixel value. An average pixel value sensitivity analysis is presented. Both the indirect

  1. Modeling the Gas Dynamics Environment in a Subscale Solid Rocket Test Motor

    Science.gov (United States)

    Eaton, Andrew M.; Ewing, Mark E.; Bailey, Kirk M.; McCool, Alex (Technical Monitor)

    2001-01-01

    Subscale test motors are often used for the evaluation of solid rocket motor component materials such as internal insulation. These motors are useful for characterizing insulation performance behavior, screening insulation material candidates and obtaining material thermal and ablative property design data. One of the primary challenges associated with using subscale motors however, is the uncertainty involved when extrapolating the results to full-scale motor conditions. These uncertainties are related to differences in such phenomena as turbulent flow behavior and boundary layer development, propellant particle interactions with the wall, insulation off-gas mixing and thermochemical reactions with the bulk flow, radiation levels, material response to the local environment, and other anomalous flow conditions. In addition to the need for better understanding of physical mechanisms, there is also a need to better understand how to best simulate these phenomena using numerical modeling approaches such as computational fluid dynamics (CFD). To better understand and model interactions between major phenomena in a subscale test motor, a numerical study of the internal flow environment of a representative motor was performed. Simulation of the environment included not only gas dynamics, but two-phase flow modeling of entrained alumina particles like those found in an aluminized propellant, and offgassing from wall surfaces similar to an ablating insulation material. This work represents a starting point for establishing the internal environment of a subscale test motor using comprehensive modeling techniques, and lays the groundwork for improving the understanding of the applicability of subscale test data to full-scale motors. It was found that grid resolution, and inclusion of phenomena in addition to gas dynamics, such as two-phase and multi-component gas composition are all important factors that can effect the overall flow field predictions.

  2. Radiometric probe design for the measurement of heat flux within a solid rocket motor nozzle

    Science.gov (United States)

    Goldey, Charles L.; Laughlin, William T.; Popper, Leslie A.

    1996-11-01

    Improvements to solid rocket motor (SRM) nozzle designs and material performance is based on the ability to instrument motors during test firings to understand the internal combustion processes and the response of nozzle components to the severe heating environment. Measuring the desired parameters is very difficult because the environment inside of an SRM is extremely severe. Instrumentation can be quickly destroyed if exposed to the internal rocket motor environment. An optical method is under development to quantify the heating of the internal nozzle surface. A radiometric probe designed for measuring the thermal response and material surface recession within a nozzle while simultaneously confining the combustion products has been devised and demonstrated. As part of the probe design, optical fibers lead to calibrated detectors that measure the interior nozzle thermal response. This two color radiometric measurement can be used for a direct determination of the total heat flux impinging on interior nozzle surfaces. This measurement has been demonstrated using a high power CO2 laser to simulate SRM nozzle heating conditions on carbon phenolic and graphite phenolic materials.

  3. Numerical study on similarity of plume infrared radiation between reduced-scale solid rocket motors

    Directory of Open Access Journals (Sweden)

    Zhang Xiaoying

    2016-08-01

    Full Text Available This study seeks to determine the similarities in plume radiation between reduced and full-scale solid rocket models in ground test conditions through investigation of flow and radiation for a series of scale ratios ranging from 0.1 to 1. The radiative transfer equation (RTE considering gas and particle radiation in a non-uniform plume has been adopted and solved by the finite volume method (FVM to compute the three dimensional, spectral and directional radiation of a plume in the infrared waveband 2–6 μm. Conditions at wavelengths 2.7 μm and 4.3 μm are discussed in detail, and ratios of plume radiation for reduced-scale through full-scale models are examined. This work shows that, with increasing scale ratio of a computed rocket motor, area of the high-temperature core increases as a 2 power function of the scale ratio, and the radiation intensity of the plume increases with 2–2.5 power of the scale ratio. The infrared radiation of plume gases shows a strong spectral dependency, while that of Al2O3 particles shows spectral continuity of gray media. Spectral radiation intensity of a computed solid rocket plume’s high temperature core increases significantly in peak radiation spectra of plume gases CO and CO2. Al2O3 particles are the major radiation component in a rocket plume. There is good similarity between contours of plume spectral radiance from different scale models of computed rockets, and there are two peak spectra of radiation intensity at wavebands 2.7–3.0 μm and 4.2–4.6 μm. Directed radiation intensity of the entire plume volume will rise with increasing elevation angle.

  4. Coupled Fluid-Structure Interaction Analysis of Solid Rocket Motor with Flexible Inhibitors

    Science.gov (United States)

    Yang, H. Q.; West, Jeff; Harris, Robert E.

    2014-01-01

    Flexible inhibitors are generally used in solid rocket motors (SRMs) as a means to control the burning of propellant. Vortices generated by the flow of propellant around the flexible inhibitors have been identified as a driving source of instabilities that can lead to thrust oscillations in launch vehicles. Potential coupling between the SRM thrust oscillations and structural vibration modes is an important risk factor in launch vehicle design. As a means to predict and better understand these phenomena, a multidisciplinary simulation capability that couples the NASA production CFD code, Loci/CHEM, with CFDRC's structural finite element code, CoBi, has been developed. This capability is crucial to the development of NASA's new space launch system (SLS). This paper summarizes the efforts in applying the coupled software to demonstrate and investigate fluid-structure interaction (FSI) phenomena between pressure waves and flexible inhibitors inside reusable solid rocket motors (RSRMs). The features of the fluid and structural solvers are described in detail, and the coupling methodology and interfacial continuity requirements are then presented in a general Eulerian-Lagrangian framework. The simulations presented herein utilize production level CFD with hybrid RANS/LES turbulence modeling and grid resolution in excess of 80 million cells. The fluid domain in the SRM is discretized using a general mixed polyhedral unstructured mesh, while full 3D shell elements are utilized in the structural domain for the flexible inhibitors. Verifications against analytical solutions for a structural model under a steady uniform pressure condition and under dynamic modal analysis show excellent agreement in terms of displacement distribution and eigenmode frequencies. The preliminary coupled results indicate that due to acoustic coupling, the dynamics of one of the more flexible inhibitors shift from its first modal frequency to the first acoustic frequency of the solid rocket motor

  5. Parametric study and performance analysis of hybrid rocket motors with double-tube configuration

    Science.gov (United States)

    Yu, Nanjia; Zhao, Bo; Lorente, Arnau Pons; Wang, Jue

    2017-03-01

    The practical implementation of hybrid rocket motors has historically been hampered by the slow regression rate of the solid fuel. In recent years, the research on advanced injector designs has achieved notable results in the enhancement of the regression rate and combustion efficiency of hybrid rockets. Following this path, this work studies a new configuration called double-tube characterized by injecting the gaseous oxidizer through a head end injector and an inner tube with injector holes distributed along the motor longitudinal axis. This design has demonstrated a significant potential for improving the performance of hybrid rockets by means of a better mixing of the species achieved through a customized injection of the oxidizer. Indeed, the CFD analysis of the double-tube configuration has revealed that this design may increase the regression rate over 50% with respect to the same motor with a conventional axial showerhead injector. However, in order to fully exploit the advantages of the double-tube concept, it is necessary to acquire a deeper understanding of the influence of the different design parameters in the overall performance. In this way, a parametric study is carried out taking into account the variation of the oxidizer mass flux rate, the ratio of oxidizer mass flow rate injected through the inner tube to the total oxidizer mass flow rate, and injection angle. The data for the analysis have been gathered from a large series of three-dimensional numerical simulations that considered the changes in the design parameters. The propellant combination adopted consists of gaseous oxygen as oxidizer and high-density polyethylene as solid fuel. Furthermore, the numerical model comprises Navier-Stokes equations, k-ε turbulence model, eddy-dissipation combustion model and solid-fuel pyrolysis, which is computed through user-defined functions. This numerical model was previously validated by analyzing the computational and experimental results obtained for

  6. Real-Time Inhibitor Recession Measurements in the Space Shuttle Reusable Solid Rocket Motors

    Science.gov (United States)

    McWhorter, Bruce B.; Ewing, Mark E.; McCool, Alex (Technical Monitor)

    2001-01-01

    Real-time char line recession measurements were made on propellant inhibitors of the Space Shuttle Reusable Solid Rocket Motor (RSRM). The RSRM FSM-8 static test motor propellant inhibitors (composed of a rubber insulation material) were successfully instrumented with eroding potentiometers and thermocouples. The data was used to establish inhibitor recession versus time relationships. Normally, pre-fire and post-fire insulation thickness measurements establish the thermal performance of an ablating insulation material. However, post-fire inhibitor decomposition and recession measurements are complicated by the fact that most of the inhibitor is back during motor operation. It is therefore a difficult task to evaluate the thermal protection offered by the inhibitor material. Real-time measurements would help this task. The instrumentation program for this static test motor marks the first time that real-time inhibitors. This report presents that data for the center and aft field joint forward facing inhibitors. The data was primarily used to measure char line recession of the forward face of the inhibitors which provides inhibitor thickness reduction versus time data. The data was also used to estimate the inhibitor height versus time relationship during motor operation.

  7. Plume Particle Collection and Sizing from Static Firing of Solid Rocket Motors

    Science.gov (United States)

    Sambamurthi, Jay K.

    1995-01-01

    Thermal radiation from the plume of any solid rocket motor, containing aluminum as one of the propellant ingredients, is mainly from the microscopic, hot aluminum oxide particles in the plume. The plume radiation to the base components of the flight vehicle is primarily determined by the plume flowfield properties, the size distribution of the plume particles, and their optical properties. The optimum design of a vehicle base thermal protection system is dependent on the ability to accurately predict this intense thermal radiation using validated theoretical models. This article describes a successful effort to collect reasonably clean plume particle samples from the static firing of the flight simulation motor (FSM-4) on March 10, 1994 at the T-24 test bed at the Thiokol space operations facility as well as three 18.3% scaled MNASA motors tested at NASA/MSFC. Prior attempts to collect plume particles from the full-scale motor firings have been unsuccessful due to the extremely hostile thermal and acoustic environment in the vicinity of the motor nozzle.

  8. Space Shuttle Redesigned Solid Rocket Motor nozzle natural frequency variations with burn time

    Science.gov (United States)

    Lui, C. Y.; Mason, D. R.

    1991-01-01

    The effects of erosion and thermal degradation on the Space Shuttle Redesigned Solid Rocket Motor (RSRM) nozzle's structural dynamic characteristics were analytically evaluated. Also considered was stiffening of the structure due to internal pressurization. A detailed NASTRAN finite element model of the nozzle was developed and used to evaluate the influence of these effects at several discrete times during motor burn. Methods were developed for treating erosion and thermal degradation, and a procedure was developed to account for internal pressure stiffening using differential stiffness matrix techniques. Results were verified using static firing test accelerometer data. Fast Fourier Transform and Maximum Entropy Method techniques were applied to the data to generate waterfall plots which track modal frequencies with burn time. Results indicate that the lower frequency nozzle 'vectoring' modes are only slightly affected by erosion, thermal effects and internal pressurization. The higher frequency shell modes of the nozzle are, however, significantly reduced.

  9. Nozzle erosion characterization and minimization for high-pressure rocket motor applications

    Science.gov (United States)

    Evans, Brian

    Understanding of the processes that cause nozzle throat erosion and developing methods for mitigation of erosion rate can allow higher operating pressures for advanced rocket motors. However, erosion of the nozzle throat region, which is a strong function of operating pressure, must be controlled to realize the performance gains of higher operating pressures. The objective of this work was the study the nozzle erosion rates at a broad range of pressures from 7 to 34.5 MPa (1,000 to 5,000 psia) using two different rocket motors. The first is an instrumented solidpropellant motor (ISPM), which uses two baseline solid propellants; one is a non-metallized propellant called Propellant S and the other is a metallized propellant called Propellant M. The second test rig is a non-metallized solid-propellant rocket motor simulator (RMS). The RMS is a gas rocket with the ability to vary the combustion-product species composition by systematically varying the flow rates of gaseous reactants. Several reactant mixtures were utilized in the study to determine the relative importance of different oxidizing species (such as H2O, OH, and CO2). Both test rigs are equipped with a windowed nozzle section for real-time X-ray radiography diagnostics of the instantaneous throat variations for deducing the instantaneous erosion rates. The nozzle test section for both motors can also incorporate a nozzle boundary-layer control system (NBLCS) as a means of nozzle erosion mitigation. The effectiveness of the NBLCS at preventing nozzle throat erosion was demonstrated for both the RMS and the ISPM motors at chamber pressures up to 34 MPa (4930 psia). All tests conducted with the NBLCS showed signs of coning of the propellant surface, leading to increased mass burning rate and resultant chamber pressure. Two correlations were developed for the nozzle erosion rates from solid propellant testing, one for metallized propellant and one for non-metallized propellants. The non-metallized propellant

  10. Feasibility study of palm-based fuels for hybrid rocket motor applications

    Science.gov (United States)

    Tarmizi Ahmad, M.; Abidin, Razali; Taha, A. Latif; Anudip, Amzaryi

    2018-02-01

    This paper describes the combined analysis done in pure palm-based wax that can be used as solid fuel in a hybrid rocket engine. The measurement of pure palm wax calorific value was performed using a bomb calorimeter. An experimental rocket engine and static test stand facility were established. After initial measurement and calibration, repeated procedures were performed. Instrumentation supplies carried out allow fuel regression rate measurements, oxidizer mass flow rates and stearic acid rocket motors measurements. Similar tests are also carried out with stearate acid (from palm oil by-products) dissolved with nitrocellulose and bee solution. Calculated data and experiments show that rates and regression thrust can be achieved even in pure-tested palm-based wax. Additionally, palm-based wax is mixed with beeswax characterized by higher nominal melting temperatures to increase moisturizing points to higher temperatures without affecting regression rate values. Calorie measurements and ballistic experiments were performed on this new fuel formulation. This new formulation promises driving applications in a wide range of temperatures.

  11. Elastomeric Thermal Insulation Design Considerations in Long, Aluminized Solid Rocket Motors

    Science.gov (United States)

    Martin, Heath T.

    2017-01-01

    An all-new sounding rocket was designed at NASA's Marshall Space Flight Center that featured an aft finocyl, aluminized solid propellant grain and silica-filled ethylene-propylene-diene monomer (SFEPDM) internal insulation. Upon the initial static firing of the first of this new design, the solid rocket motor (SRM) case failed thermally just upstream of the aft closure early in the burn time. Subsequent fluid modeling indicated that the high-velocity combustion-product jets emanating from the fin-slots in the propellant grain were likely inducing a strongly swirling flow, thus substantially increasing the severity of the convective environment on the exposed portion of the SFEPDM insulation in this region. The aft portion of the fin-slots in another of the motors were filled with propellant to eliminate the possibility of both direct jet impingement on the exposed SFEPDM and the appearance of strongly swirling flow in the aft region of the motor. When static-fired, this motor's case still failed in the same axial location, and, though somewhat later than for the first static firing, still in less than 1/3rd of the desired burn duration. These results indicate that the extreme material decomposition rates of the SFEPDM in this application are not due to gas-phase convection or shear but rather to interactions with burning aluminum or alumina slag. Further comparisons with between SFEPDM performance in this design and that in other hot-fire tests provide insight into the mechanisms of SFEPDM decomposition in SRM aft domes that can guide the upcoming redesign effort, as well as other future SRM designs. These data also highlight the current limitations of modeling elastomeric insulators solely with diffusion-controlled, gas-phase thermochemistry in SRM regions with significant viscous shear and/or condense-phase impingement or flow.

  12. Test Report for MSFC Test No. 83-2: Pressure scaled water impact test of a 12.5 inch diameter model of the Space Shuttle solid rocket booster filament wound case and external TVC PCD

    Science.gov (United States)

    1983-01-01

    Water impact tests using a 12.5 inch diameter model representing a 8.56 percent scale of the Space Shuttle Solid Rocket Booster configuration were conducted. The two primary objectives of this SRB scale model water impact test program were: 1. Obtain cavity collapse applied pressure distributions for the 8.56 percent rigid body scale model FWC pressure magnitudes as a function of full-scale initial impact conditions at vertical velocities from 65 to 85 ft/sec, horizontal velocities from 0 to 45 ft/sec, and angles from -10 to +10 degrees. 2. Obtain rigid body applied pressures on the TVC pod and aft skirt internal stiffener rings at initial impact and cavity collapse loading events. In addition, nozzle loads were measured. Full scale vertical velocities of 65 to 85 ft/sec, horizontal velocities of 0 to 45 ft/sec, and impact angles from -10 to +10 degrees simulated.

  13. Modified computation of the nozzle damping coefficient in solid rocket motors

    Science.gov (United States)

    Liu, Peijin; Wang, Muxin; Yang, Wenjing; Gupta, Vikrant; Guan, Yu; Li, Larry K. B.

    2018-02-01

    In solid rocket motors, the bulk advection of acoustic energy out of the nozzle constitutes a significant source of damping and can thus influence the thermoacoustic stability of the system. In this paper, we propose and test a modified version of a historically accepted method of calculating the nozzle damping coefficient. Building on previous work, we separate the nozzle from the combustor, but compute the acoustic admittance at the nozzle entry using the linearized Euler equations (LEEs) rather than with short nozzle theory. We compute the combustor's acoustic modes also with the LEEs, taking the nozzle admittance as the boundary condition at the combustor exit while accounting for the mean flow field in the combustor using an analytical solution to Taylor-Culick flow. We then compute the nozzle damping coefficient via a balance of the unsteady energy flux through the nozzle. Compared with established methods, the proposed method offers competitive accuracy at reduced computational costs, helping to improve predictions of thermoacoustic instability in solid rocket motors.

  14. A Coupled Fluid-Structure Interaction Analysis of Solid Rocket Motor with Flexible Inhibitors

    Science.gov (United States)

    Yang, H. Q.; West, Jeff

    2014-01-01

    A capability to couple NASA production CFD code, Loci/CHEM, with CFDRC's structural finite element code, CoBi, has been developed. This paper summarizes the efforts in applying the installed coupling software to demonstrate/investigate fluid-structure interaction (FSI) between pressure wave and flexible inhibitor inside reusable solid rocket motor (RSRM). First a unified governing equation for both fluid and structure is presented, then an Eulerian-Lagrangian framework is described to satisfy the interfacial continuity requirements. The features of fluid solver, Loci/CHEM and structural solver, CoBi, are discussed before the coupling methodology of the solvers is described. The simulation uses production level CFD LES turbulence model with a grid resolution of 80 million cells. The flexible inhibitor is modeled with full 3D shell elements. Verifications against analytical solutions of structural model under steady uniform pressure condition and under dynamic condition of modal analysis show excellent agreements in terms of displacement distribution and eigen modal frequencies. The preliminary coupled result shows that due to acoustic coupling, the dynamics of one of the more flexible inhibitors shift from its first modal frequency to the first acoustic frequency of the solid rocket motor.

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

    Science.gov (United States)

    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

  16. Closed-loop thrust and pressure profile throttling of a nitrous oxide/hydroxyl-terminated polybutadiene hybrid rocket motor

    Science.gov (United States)

    Peterson, Zachary W.

    Hybrid motors that employ non-toxic, non-explosive components with a liquid oxidizer and a solid hydrocarbon fuel grain have inherently safe operating characteristics. The inherent safety of hybrid rocket motors offers the potential to greatly reduce overall operating costs. Another key advantage of hybrid rocket motors is the potential for in-flight shutdown, restart, and throttle by controlling the pressure drop between the oxidizer tank and the injector. This research designed, developed, and ground tested a closed-loop throttle controller for a hybrid rocket motor using nitrous oxide and hydroxyl-terminated polybutadiene as propellants. The research simultaneously developed closed-loop throttle algorithms and lab scale motor hardware to evaluate the fidelity of the throttle simulations and algorithms. Initial open-loop motor tests were performed to better classify system parameters and to validate motor performance values. Deep-throttle open-loop tests evaluated limits of stable thrust that can be achieved on the test hardware. Open-loop tests demonstrated the ability to throttle the motor to less than 10% of maximum thrust with little reduction in effective specific impulse and acoustical stability. Following the open-loop development, closed-loop, hardware-in-the-loop tests were performed. The closed-loop controller successfully tracked prescribed step and ramp command profiles with a high degree of fidelity. Steady-state accuracy was greatly improved over uncontrolled thrust.

  17. Qualification Status of Non-Asbestos Internal Insulation in the Reusable Solid Rocket Motor Program

    Science.gov (United States)

    Clayton, Louie

    2011-01-01

    This paper provides a status of the qualification efforts associated with NASA's RSRMV non-asbestos internal insulation program. For many years, NASA has been actively engaged in removal of asbestos from the shuttle RSRM motors due to occupation health concerns where technicians are working with an EPA banned material. Careful laboratory and subscale testing has lead to the downselect of a organic fiber known as Polybenzimidazol to replace the asbestos fiber filler in the existing synthetic rubber copolymer Nitrile Butadiene - now named PBI/NBR. Manufacturing, processing, and layup of the new material has been a challenge due to the differences in the baseline shuttle RSRM internal insulator properties and PBI/NBR material properties. For this study, data gathering and reduction procedures for thermal and chemical property characterization for the new candidate material are discussed. Difficulties with test procedures, implementation of properties into the Charring Material Ablator (CMA) codes, and results correlation with static motor fire data are provided. After two successful five segment motor firings using the PBI/NBR insulator, performance results for the new material look good and the material should eventually be qualified for man rated use in large solid rocket motor applications.

  18. Nonlinear rocket motor stability prediction: Limit amplitude, triggering, and mean pressure shifta)

    Science.gov (United States)

    Flandro, Gary A.; Fischbach, Sean R.; Majdalani, Joseph

    2007-09-01

    High-amplitude pressure oscillations in solid propellant rocket motor combustion chambers display nonlinear effects including: (1) limit cycle behavior in which the fluctuations may dwell for a considerable period of time near their peak amplitude, (2) elevated mean chamber pressure (DC shift), and (3) a triggering amplitude above which pulsing will cause an apparently stable system to transition to violent oscillations. Along with the obvious undesirable vibrations, these features constitute the most damaging impact of combustion instability on system reliability and structural integrity. The physical mechanisms behind these phenomena and their relationship to motor geometry and physical parameters must, therefore, be fully understood if instability is to be avoided in the design process, or if effective corrective measures must be devised during system development. Predictive algorithms now in use have limited ability to characterize the actual time evolution of the oscillations, and they do not supply the motor designer with information regarding peak amplitudes or the associated critical triggering amplitudes. A pivotal missing element is the ability to predict the mean pressure shift; clearly, the designer requires information regarding the maximum chamber pressure that might be experienced during motor operation. In this paper, a comprehensive nonlinear combustion instability model is described that supplies vital information. The central role played by steep-fronted waves is emphasized. The resulting algorithm provides both detailed physical models of nonlinear instability phenomena and the critically needed predictive capability. In particular, the origin of the DC shift is revealed.

  19. Development of Erosive Burning Models for CFD Predictions of Solid Rocket Motor Internal Environments

    Science.gov (United States)

    Wang, Qun-Zhen

    2003-01-01

    Four erosive burning models, equations (11) to (14). are developed in this work by using a power law relationship to correlate (1) the erosive burning ratio and the local velocity gradient at propellant surfaces; (2) the erosive burning ratio and the velocity gradient divided by centerline velocity; (3) the erosive burning difference and the local velocity gradient at propellant surfaces; and (4) the erosive burning difference and the velocity gradient divided by centerline velocity. These models depend on the local velocity gradient at the propellant surface (or the velocity gradient divided by centerline velocity) only and, unlike other empirical models, are independent of the motor size. It was argued that, since the erosive burning is a local phenomenon occurring near the surface of the solid propellant, the erosive burning ratio should be independent of the bore diameter if it is correlated with some local flow parameters such as the velocity gradient at the propellant surface. This seems to be true considering the good results obtained by applying these models, which are developed from the small size 5 inch CP tandem motor testing, to CFD simulations of much bigger motors.

  20. Performance characteristics of conventional X-ray generator isotope source and high energy accelerator in rocket motor evaluation

    International Nuclear Information System (INIS)

    Viswanathan, K.; Rao, K.V.; Subbalah, C.; Uttam, M.C.

    1985-01-01

    Final qualification of solid rocket motors and other related components in the Indian Space Programme is carried out using radiographic sources of different energies. The necessity to have different sources of varying energies arises from the fact that the components in the space programme vary from small fastners to gigantic solid rocket motors. In order to achieve the best radiographic quality with the optimised exposure time different X-ray sources are used. To have 100% coverage and to reduce the inspection time, a Real Time Radiography for the high energy LINAC is also planned

  1. Basalt fiber and nanoclay compositions, articles incorporating the same, and methods of insulating a rocket motor with the same

    Science.gov (United States)

    Gajiwala, Himansu M. (Inventor)

    2010-01-01

    An insulation composition that comprises at least one nitrile butadiene rubber, basalt fibers, and nanoclay is disclosed. Further disclosed is an insulation composition that comprises polybenzimidazole fibers, basalt fibers, and nanoclay. The basalt fibers may be present in the insulation compositions in a range of from approximately 1% by weight to approximately 6% by weight of the total weight of the insulation composition. The nanoclay may be present in the insulation compositions in a range of from approximately 5% by weight to approximately 10% by weight of the total weight of the insulation composition. Rocket motors including the insulation compositions and methods of insulating a rocket motor are also disclosed.

  2. High performance Solid Rocket Motor (SRM) submerged nozzle/combustion cavity flowfield assessment

    Science.gov (United States)

    Freeman, J. A.; Chan, J. S.; Murph, J. E.; Xiques, K. E.

    1987-01-01

    Two and three dimensional internal flowfield solutions for critical points in the Space Shuttle solid rocket booster burn time were developed using the Lockheed Huntsville GIM/PAID Navier-Stokes solvers. These perfect gas, viscous solutions for the high performance motor characterize the flow in the aft segment and nozzle of the booster. Two dimensional axisymmetric solutions were developed at t = 20 and t = 85 sec motor burn times. The t = 85 sec solution indicates that the aft segment forward inhibitor stub produces vortices with are shed and convected downwards. A three dimensional 3.5 deg gimbaled nozzle flowfield solution was developed for the aft segment and nozzle at t = 9 sec motor burn time. This perfect gas, viscous analysis, provided a steady state solution for the core region and the flow through the nozzle, but indicated that unsteady flow exists in the region under the nozzle nose and near the flexible boot and nozzle/case joint. The flow in the nozzle/case joint region is characterized by low magnitude pressure waves which travel in the circumferential direction. From the two and three dimensional flowfield calculations presented it can be concluded that there is no evidence from these results that steady state gas dynamics is the primary mechanism resulting in the nozzle pocketing erosion experienced on SRM nozzles 8A or 17B. The steady state flowfield results indicate pocketing erosion is not directly initiated by a steady state gas dynamics phenomenon.

  3. Numerical Evaluation of the Use of Aluminum Particles for Enhancing Solid Rocket Motor Combustion Stability

    Directory of Open Access Journals (Sweden)

    David Greatrix

    2015-02-01

    Full Text Available The ability to 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 typically necessitates a comprehensive numerical model for internal ballistic simulation under dynamic flow and combustion conditions. On the mitigation side, one in practice sees the use of inert or reactive particles for the suppression of pressure wave development in the motor chamber flow. With the focus of the present study placed on reactive particles, a numerical internal ballistic model incorporating relevant elements, such as a transient, frequency-dependent combustion response to axial pressure wave activity above the burning propellant surface, is applied to the investigation of using aluminum particles within the central internal flow (particles whose surfaces nominally regress with time, as a function of current particle size, as they move downstream as a means of suppressing instability-related symptoms in a cylindrical-grain motor. The results of this investigation reveal that the loading percentage and starting size of the aluminum particles have a significant influence on reducing the resulting transient pressure wave magnitude.

  4. Evaluation of Solid Rocket Motor Component Data Using a Commercially Available Statistical Software Package

    Science.gov (United States)

    Stefanski, Philip L.

    2015-01-01

    Commercially available software packages today allow users to quickly perform the routine evaluations of (1) descriptive statistics to numerically and graphically summarize both sample and population data, (2) inferential statistics that draws conclusions about a given population from samples taken of it, (3) probability determinations that can be used to generate estimates of reliability allowables, and finally (4) the setup of designed experiments and analysis of their data to identify significant material and process characteristics for application in both product manufacturing and performance enhancement. This paper presents examples of analysis and experimental design work that has been conducted using Statgraphics®(Registered Trademark) statistical software to obtain useful information with regard to solid rocket motor propellants and internal insulation material. Data were obtained from a number of programs (Shuttle, Constellation, and Space Launch System) and sources that include solid propellant burn rate strands, tensile specimens, sub-scale test motors, full-scale operational motors, rubber insulation specimens, and sub-scale rubber insulation analog samples. Besides facilitating the experimental design process to yield meaningful results, statistical software has demonstrated its ability to quickly perform complex data analyses and yield significant findings that might otherwise have gone unnoticed. One caveat to these successes is that useful results not only derive from the inherent power of the software package, but also from the skill and understanding of the data analyst.

  5. Flight demonstration of flight termination system and solid rocket motor ignition using semiconductor laser initiated ordnance

    Science.gov (United States)

    Schulze, Norman R.; Maxfield, B.; Boucher, C.

    1995-01-01

    Solid State Laser Initiated Ordnance (LIO) offers new technology having potential for enhanced safety, reduced costs, and improved operational efficiency. Concerns over the absence of programmatic applications of the technology, which has prevented acceptance by flight programs, should be abated since LIO has now been operationally implemented by the Laser Initiated Ordnance Sounding Rocket Demonstration (LOSRD) Program. The first launch of solid state laser diode LIO at the NASA Wallops Flight Facility (WFF) occurred on March 15, 1995 with all mission objectives accomplished. This project, Phase 3 of a series of three NASA Headquarters LIO demonstration initiatives, accomplished its objective by the flight of a dedicated, all-LIO sounding rocket mission using a two-stage Nike-Orion launch vehicle. LIO flight hardware, made by The Ensign-Bickford Company under NASA's first Cooperative Agreement with Profit Making Organizations, safely initiated three demanding pyrotechnic sequence events, namely, solid rocket motor ignition from the ground and in flight, and flight termination, i.e., as a Flight Termination System (FTS). A flight LIO system was designed, built, tested, and flown to support the objectives of quickly and inexpensively putting LIO through ground and flight operational paces. The hardware was fully qualified for this mission, including component testing as well as a full-scale system test. The launch accomplished all mission objectives in less than 11 months from proposal receipt. This paper concentrates on accomplishments of the ordnance aspects of the program and on the program's implementation and results. While this program does not generically qualify LIO for all applications, it demonstrated the safety, technical, and operational feasibility of those two most demanding applications, using an all solid state safe and arm system in critical flight applications.

  6. Uncertainty analysis and design optimization of hybrid rocket motor powered vehicle for suborbital flight

    Directory of Open Access Journals (Sweden)

    Zhu Hao

    2015-06-01

    Full Text Available In this paper, we propose an uncertainty analysis and design optimization method and its applications on a hybrid rocket motor (HRM powered vehicle. The multidisciplinary design model of the rocket system is established and the design uncertainties are quantified. The sensitivity analysis of the uncertainties shows that the uncertainty generated from the error of fuel regression rate model has the most significant effect on the system performances. Then the differences between deterministic design optimization (DDO and uncertainty-based design optimization (UDO are discussed. Two newly formed uncertainty analysis methods, including the Kriging-based Monte Carlo simulation (KMCS and Kriging-based Taylor series approximation (KTSA, are carried out using a global approximation Kriging modeling method. Based on the system design model and the results of design uncertainty analysis, the design optimization of an HRM powered vehicle for suborbital flight is implemented using three design optimization methods: DDO, KMCS and KTSA. The comparisons indicate that the two UDO methods can enhance the design reliability and robustness. The researches and methods proposed in this paper can provide a better way for the general design of HRM powered vehicles.

  7. Test data from small solid propellant rocket motor plume measurements (FA-21)

    Science.gov (United States)

    Hair, L. M.; Somers, R. E.

    1976-01-01

    A program is described for obtaining a reliable, parametric set of measurements in the exhaust plumes of solid propellant rocket motors. Plume measurements included pressures, temperatures, forces, heat transfer rates, particle sampling, and high-speed movies. Approximately 210,000 digital data points and 15,000 movie frames were acquired. Measurements were made at points in the plumes via rake-mounted probes, and on the surface of a large plate impinged by the exhaust plume. Parametric variations were made in pressure altitude, propellant aluminum loading, impinged plate incidence angle and distance from nozzle exit to plate or rake. Reliability was incorporated by continual use of repeat runs. The test setup of the various hardware items is described along with an account of test procedures. Test results and data accuracy are discussed. Format of the data presentation is detailed. Complete data are included in the appendix.

  8. Multiple time scale analysis of pressure oscillations in solid rocket motors

    Science.gov (United States)

    Ahmed, Waqas; Maqsood, Adnan; Riaz, Rizwan

    2018-03-01

    In this study, acoustic pressure oscillations for single and coupled longitudinal acoustic modes in Solid Rocket Motor (SRM) are investigated using Multiple Time Scales (MTS) method. Two independent time scales are introduced. The oscillations occur on fast time scale whereas the amplitude and phase changes on slow time scale. Hopf bifurcation is employed to investigate the properties of the solution. The supercritical bifurcation phenomenon is observed for linearly unstable system. The amplitude of the oscillations result from equal energy gain and loss rates of longitudinal acoustic modes. The effect of linear instability and frequency of longitudinal modes on amplitude and phase of oscillations are determined for both single and coupled modes. For both cases, the maximum amplitude of oscillations decreases with the frequency of acoustic mode and linear instability of SRM. The comparison of analytical MTS results and numerical simulations demonstrate an excellent agreement.

  9. Lightweight structural design of a bolted case joint for the space shuttle solid rocket motor

    Science.gov (United States)

    Dorsey, John T.; Stein, Peter A.; Bush, Harold G.

    1988-01-01

    The structural design of a bolted joint with a static face seal which can be used to join Space Shuttle Solid Rocket Motor (SRM) case segments is given. Results from numerous finite element parametric studies indicate that the bolted joint meets the design requirement of preventing joint opening at the O-ring locations during SRM pressurization. A final design recommended for further development has the following parameters: 180 one-in.-diam. studs, stud centerline offset of 0.5 in radially inward from the shell wall center line, flange thickness of 0.75 in, bearing plate thickness of 0.25 in, studs prestressed to 70 percent of ultimate load, and the intermediate alcove. The design has a mass penalty of 1096 lbm, which is 164 lbm greater than the currently proposed capture tang redesign.

  10. Qualification of Magnesium/Teflon/Viton Pyrotechnic Composition Used in Rocket Motors Ignition System

    Directory of Open Access Journals (Sweden)

    Luciana de Barros

    2016-04-01

    Full Text Available The application of fluoropolymers in high-energy-release pyrotechnic compositions is common in the space and defense areas. Pyrotechnic compositions of magnesium/Teflon/Viton are widely used in military flares and pyrogen igniters for igniting the solid propellant of a rocket motor. Pyrotechnic components are considered high-risk products as they may cause catastrophic accidents if initiated or ignited inadvertently. To reduce the hazards involved in the handling, storage and transportation of these devices, the magnesium/Teflon/Viton composition was subjected to various sensitivity tests, DSC and had its stability and compatibility tested with other materials. This composition obtained satisfactory results in all the tests, which qualifies it as safe for production, handling, use, storage and transportation.

  11. Numerical and experimental study of liquid breakup process in solid rocket motor nozzle

    Science.gov (United States)

    Yen, Yi-Hsin

    Rocket propulsion is an important travel method for space exploration and national defense, rockets needs to be able to withstand wide range of operation environment and also stable and precise enough to carry sophisticated payload into orbit, those engineering requirement makes rocket becomes one of the state of the art industry. The rocket family have been classified into two major group of liquid and solid rocket based on the fuel phase of liquid or solid state. The solid rocket has the advantages of simple working mechanism, less maintenance and preparing procedure and higher storage safety, those characters of solid rocket make it becomes popular in aerospace industry. Aluminum based propellant is widely used in solid rocket motor (SRM) industry due to its avalibility, combusion performance and economical fuel option, however after aluminum react with oxidant of amonimum perchrate (AP), it will generate liquid phase alumina (Al2O3) as product in high temperature (2,700˜3,000 K) combustion chamber enviornment. The liquid phase alumina particles aggromorate inside combustion chamber into larger particle which becomes major erosion calprit on inner nozzle wall while alumina aggromorates impinge on the nozzle wall surface. The erosion mechanism result nozzle throat material removal, increase the performance optimized throat diameter and reduce nozzle exit to throat area ratio which leads to the reduction of exhaust gas velocity, Mach number and lower the propulsion thrust force. The approach to avoid particle erosion phenomenon taking place in SRM's nozzle is to reduce the alumina particle size inside combustion chamber which could be done by further breakup of the alumina droplet size in SRM's combustion chamber. The study of liquid breakup mechanism is an important means to smaller combustion chamber alumina droplet size and mitigate the erosion tack place on rocket nozzle region. In this study, a straight two phase air-water flow channel experiment is set up

  12. PCV Solid Rocket Motor: Design Status of the Motor Case Structure

    Science.gov (United States)

    Mataloni, A.; Zallo, A.; Perugini, P.; Di Cosmo, A.; Pasquale, N.; Mucci, R.

    2014-06-01

    For the VEGA Launch system new developments are running in order to allow: a) performances increase b) cost reduction c) introduction of new technologies.In the VEGA C configuration the PCV SRM replace the P80 in the first stage.The PCV design is based on the consolidate AVIO heritage with important improvements both from the material and from the technological side.Important improvements in skirts manufacturing will be tested as well, with the development of a customized automatic tape laying machine.From the material side a top class fiber will be selected on the bases of extensive trade-off plan which is under completion.The pre-preg material is based on an in-house resin formulation tailored to the specific motor case process requirements.

  13. Study of solid rocket motor for space shuttle booster, volume 2, book 1

    Science.gov (United States)

    1972-01-01

    The technical requirements for the solid propellant rocket engine to be used with the space shuttle orbiter are presented. The subjects discussed are: (1) propulsion system definition, (2) solid rocket engine stage design, (3) solid rocket engine stage recovery, (4) environmental effects, (5) manrating of the solid rocket engine stage, (6) system safety analysis, and (7) ground support equipment.

  14. Launch Vehicles Based on Advanced Hybrid Rocket Motors: An Enabling Technology for the Commercial Small and Micro Satellite Planetary Science

    Science.gov (United States)

    Karabeyoglu, Arif; Tuncer, Onur; Inalhan, Gokhan

    2016-07-01

    Mankind is relient on chemical propulsion systems for space access. Nevertheless, this has been a stagnant area in terms of technological development and the technology base has not changed much almost for the past forty years. This poses a vicious circle for launch applications such that high launch costs constrain the demand and low launch freqencies drive costs higher. This also has been a key limiting factor for small and micro satellites that are geared towards planetary science. Rather this be because of the launch frequencies or the costs, the access of small and micro satellites to orbit has been limited. With today's technology it is not possible to escape this circle. However the emergence of cost effective and high performance propulsion systems such as advanced hybrid rockets can decrease launch costs by almost an order or magnitude. This paper briefly introduces the timeline and research challenges that were overcome during the development of advanced hybrid LOX/paraffin based rockets. Experimental studies demonstrated effectiveness of these advanced hybrid rockets which incorporate fast burning parafin based fuels, advanced yet simple internal balistic design and carbon composite winding/fuel casting technology that enables the rocket motor to be built from inside out. A feasibility scenario is studied using these rocket motors as building blocks for a modular launch vehicle capable of delivering micro satellites into low earth orbit. In addition, the building block rocket motor can be used further solar system missions providing the ability to do standalone small and micro satellite missions to planets within the solar system. This enabling technology therefore offers a viable alternative in order to escape the viscous that has plagued the space launch industry and that has limited the small and micro satellite delivery for planetary science.

  15. Ispitivanje piropatrona i raketnog motora pilotskog sedišta / Testing pyrocartridges and the rocket motor of the ejection seat

    Directory of Open Access Journals (Sweden)

    Milorad Savković

    2008-04-01

    Full Text Available Raketni motor pilotskog sedišta ima složen geometrijski oblik, tako da njegov potisak deluje pod određenim uglom u odnosu na ravan simetrije pilotskog sedišta. Radi određivanja intenziteta i napadne linije potiska izvršen je veći broj eksperimenata. Meren je potisak raketnog motora na višekomponentnom opitnom stolu. Letno ispitivanje pilotskog sedišta obavljeno je pomoću lutke koja simulira masu pilota. Takođe, analizirano je letno ispitivanje pilotskog sedišta u početnom periodu katapultiranja za vreme rada raketnog motora. Obrađeni su i rezultati merenja ubrzanja, koji su korišćeni za određivanje karakteristika leta pilotskog sedišta. U radu je prikazan teorijski model kretanja sedišta. / Due to a complex geometrical shape of the rocket motor of the ejection seat, the rocket motor thrust occurs under certain angle in relation to the plane of symmetry of the ejection seat. A number of tests were carried out in order to determine thrust intensity and angle of attack. The rocket motor thrust was measured on the multicomponent test stand. The ejection seat whit a dummy simulating a mass of a pilot was tested during ejection. The paper presents an analysis of the ejection seat flight in the initial phase of ejection, during the rocket motor running. The results of the acceleration read-outs were processed and then used for the determination of the characteristics of the ejection seat flight. A theoretical model of the ejection seat flight is given in the paper.

  16. IR radiation characteristics of rocket exhaust plumes under varying motor operating conditions

    Directory of Open Access Journals (Sweden)

    Qinglin NIU

    2017-06-01

    Full Text Available The infrared (IR irradiance signature from rocket motor exhaust plumes is closely related to motor type, propellant composition, burn time, rocket geometry, chamber parameters and flight conditions. In this paper, an infrared signature analysis tool (IRSAT was developed to understand the spectral characteristics of exhaust plumes in detail. Through a finite volume technique, flow field properties were obtained through the solution of axisymmetric Navier-Stokes equations with the Reynolds-averaged approach. A refined 13-species, 30-reaction chemistry scheme was used for combustion effects and a k-ε-Rt turbulence model for entrainment effects. Using flowfield properties as input data, the spectrum was integrated with a line of sight (LOS method based on a single line group (SLG model with Curtis-Godson approximation. The model correctly predicted spectral distribution in the wavelengths of 1.50–5.50 μm and had good agreement for its location with imaging spectrometer data. The IRSAT was then applied to discuss the effects of three operating conditions on IR signatures: (a afterburning; (b chamber pressure from ignition to cutoff; and (c minor changes in the ratio of hydroxyl-terminated polybutadiene (HTPB binder to ammonium perchlorate (AP oxidizer in propellant. Results show that afterburning effects can increase the size and shape of radiance images with enhancement of radiation intensity up to 40%. Also, the total IR irradiance in different bands can be characterized by a non-dimensional chamber pressure trace in which the maximum discrepancy is less than 13% during ignition and engine cutoff. An increase of chamber pressure can lead to more distinct diamonds, whose distance intervals are extended, and the position of the first diamond moving backwards. In addition, an increase in HTPB/AP causes a significant jump in spectral intensity. The incremental rates of radiance intensity integrated in each band are linear with the increase of HTPB

  17. Experimental determination of convective heat transfer coefficients in the separated flow region of the Space Shuttle Solid Rocket Motor

    Science.gov (United States)

    Whitesides, R. Harold; Majumdar, Alok K.; Jenkins, Susan L.; Bacchus, David L.

    1990-01-01

    A series of cold flow heat transfer tests was conducted with a 7.5-percent scale model of the Space Shuttle Rocket Motor (SRM) to measure the heat transfer coefficients in the separated flow region around the nose of the submerged nozzle. Modifications were made to an existing 7.5 percent scale model of the internal geometry of the aft end of the SRM, including the gimballed nozzle in order to accomplish the measurements. The model nozzle nose was fitted with a stainless steel shell with numerous thermocouples welded to the backside of the thin wall. A transient 'thin skin' experimental technique was used to measure the local heat transfer coefficients. The effects of Reynolds number, nozzle gimbal angle, and model location were correlated with a Stanton number versus Reynolds number correlation which may be used to determine the convective heating rates for the full scale Space Shuttle Solid Rocket Motor nozzle.

  18. An analysis of the orbital distribution of solid rocket motor slag

    Science.gov (United States)

    Horstman, Matthew F.; Mulrooney, Mark

    2009-01-01

    The contribution by solid rocket motors (SRMs) to the orbital debris environment is potentially significant and insufficiently studied. Design and combustion processes can lead to the emission of enough by-products to warrant assessment of their contribution to orbital debris. These particles are formed during SRM tail-off, or burn termination, by the rapid solidification of molten Al2O3 slag accumulated during the burn. The propensity of SRMs to generate particles larger than 100μm raises concerns regarding the debris environment. Sizes as large as 1 cm have been witnessed in ground tests, and comparable sizes have been estimated via observations of sub-orbital tail-off events. Utilizing previous research we have developed more sophisticated size distributions and modeled the time evolution of resultant orbital populations using a historical database of SRM launches, propellant, and likely location and time of tail-off. This analysis indicates that SRM ejecta is a significant component of the debris environment.

  19. Analysis of pressure blips in aft-finocyl solid rocket motor

    Science.gov (United States)

    Di Giacinto, M.; Favini, B.; Cavallini, E.

    2016-07-01

    Ballistic anomalies have frequently occurred during the firing of several solid rocket motors (SRMs) (Inertial Upper Stage, Space Shuttle Redesigned SRM (RSRM) and Titan IV SRM Upgrade (SRMU)), producing even relevant and unexpected variations of the SRM pressure trace from its nominal profile. This paper has the purpose to provide a numerical analysis of the following possible causes of ballistic anomalies in SRMs: an inert object discharge, a slag ejection, and an unexpected increase in the propellant burning rate or in the combustion surface. The SRM configuration under investigation is an aft-finocyl SRM with a first-stage/small booster design. The numerical simulations are performed with a quasi-one-dimensional (Q1D) unsteady model of the SRM internal ballistics, properly tailored to model each possible cause of the ballistic anomalies. The results have shown that a classification based on the head-end pressure (HEP) signature, relating each other the HEP shape and the ballistic anomaly cause, can be made. For each cause of ballistic anomalies, a deepened discussion of the parameters driving the HEP signatures is provided, as well as qualitative and quantitative assessments of the resultant pressure signals.

  20. Development of efficient finite elements for structural integrity analysis of solid rocket motor propellant grains

    International Nuclear Information System (INIS)

    Marimuthu, R.; Nageswara Rao, B.

    2013-01-01

    Solid propellant rocket motors (SRM) are regularly used in the satellite launch vehicles which consist of mainly three different structural materials viz., solid propellant, liner, and casing materials. It is essential to assess the structural integrity of solid propellant grains under the specified gravity, thermal and pressure loading conditions. For this purpose finite elements developed following the Herrmann formulation are: twenty node brick element (BH20), eight node quadrilateral plane strain element (PH8) and, eight node axi-symmetric solid of revolution element (AH8). The time-dependent nature of the solid propellant grains is taken into account utilizing the direct inverse method of Schepary to specify the effective Young's modulus and Poisson's ratio. The developed elements are tested considering various problems prior to implementation in the in-house software package (viz., Finite Element Analysis of STructures, FEAST). Several SRM configurations are analyzed to assess the structural integrity under different loading conditions. Finite element analysis results are found to be in good agreement with those obtained earlier from MARC software. -- Highlights: • Developed efficient Herrmann elements. • Accuracy of finite elements demonstrated solving several known solution problems. • Time dependent structural response obtained using the direct inverse method of Schepary. • Performed structural analysis of grains under gravity, thermal and pressure loads

  1. Design and Fabrication of a 200N Thrust Rocket Motor Based on NH4ClO4+Al+HTPB as Solid Propellant

    Science.gov (United States)

    Wahid, Mastura Ab; Ali, Wan Khairuddin Wan

    2010-06-01

    The development of rocket motor using potassium nitrate, carbon and sulphur mixture has successfully been developed by researchers and students from UTM and recently a new combination for solid propellant is being created. The new solid propellant will combine a composition of Ammonium perchlorate, NH4ClO4 with aluminium, Al and Hydroxyl Terminated Polybutadiene, HTPB as the binder. It is the aim of this research to design and fabricate a new rocket motor that will produce a thrust of 200N by using this new solid propellant. A static test is done to obtain the thrust produced by the rocket motor and analyses by observation and also calculation will be done. The experiment for the rocket motor is successful but the thrust did not achieve its required thrust.

  2. A multilayered thick cylindrical shell under internal pressure and thermal loads applicable to solid propellant rocket motors

    Energy Technology Data Exchange (ETDEWEB)

    Renganathan, K.; Nageswara Rao, B.; Jana, M.K. [Vikram Sarabhai Space Centre, Trivandrum (India). Structural Engineering Group

    2000-09-01

    A solid propellant rocket motor can be considered to be made of various circumferential layers of different properties. A simple procedure is described here to obtain an analytical solution for the general case of multilayered thick cyclindrical shell for internal pressure and thermal loads. This analytical procedure is useful in the preliminary design analysis of solid propellant rocket motors. Since solid propellant material is of viscoelastic behaviour an approximate viscoelastic solution methodology for the multilayered shell is described for estimation of time dependent solutions of propellant grain in a rocket motor. The analytical solution for a two layer reinforced thick cylindrical shell available in the literature is shown to be a special case of the present analytical solution. The results from the present analytical solution for multilayers is found to be in good agreement with FEA results. (orig.) [German] Der grundlegende Aufbau von Feststoffraketenmotoren kann auf einen Zylinder aus mehreren Schichten mit unterschiedlichen Eigenschaften zurueckgefuehrt werden. Eine einfache Berechnungsprozedur fuer die analytische Loesung des allgemeinen Falles eines mehrschichtigen Zylinders unter innerem Druck und thermischer Belastung wird hier vorgestellt. Diese analytische Methodik ist fuer den Auslegungsprozess von Feststoffraketenmotoren von grundlegender Bedeutung. Das viskoelastische Fliessverhalten des festen Brennstoffes, das den zeitlichen Ablauf des Verbrennungsprozesses wesentlich bestimmt, wird durch ein Naeherungsverfahren gut erfasst. Ein in der Literatur enthaltenes spezielles Ergebnis fuer einen zweischaligen verstaerkten Zylinder ergibt sich als Sonderfall der hier vorgestellten Methodik. Die analytisch erhaltenen Loesungen fuer mehrschichtige Aufbauten sind in guter Uebereinstimmung mit mittels der FEM ermittelten Ergebnisse. (orig.)

  3. An Internal Thermal Environment Model of an Aluminized Solid Rocket Motor with Experimental Validation

    Science.gov (United States)

    Martin, Heath T.

    2015-01-01

    Due to the severity of the internal solid rocket motor (SRM) environment, very few direct measurements of that environment exist; therefore, the appearance of such data provides a unique opportunity to assess current thermal/fluid modeling capabilities. As part of a previous study of SRM internal insulation performance, the internal thermal environment of a laboratory-scale SRM featuring aluminized propellant was characterized with two types of custom heat-flux calorimeters: one that measured the total heat flux to a graphite slab within the SRM chamber and another that measured the thermal radiation flux. Therefore, in the current study, a thermal/fluid model of this lab-scale SRM was constructed using ANSYS Fluent to predict not only the flow field structure within the SRM and the convective heat transfer to the interior walls, but also the resulting dispersion of alumina droplets and the radiative heat transfer to the interior walls. The dispersion of alumina droplets within the SRM chamber was determined by employing the Lagrangian discrete phase model that was fully coupled to the Eulerian gas-phase flow. The P1-approximation was engaged to model the radiative heat transfer through the SRM chamber where the radiative contributions of the gas phase were ignored and the aggregate radiative properties of the alumina dispersion were computed from the radiative properties of its individual constituent droplets, which were sourced from literature. The convective and radiative heat fluxes computed from the thermal/fluid model were then compared with those measured in the lab-scale SRM test firings and the modeling approach evaluated.

  4. Real-Time X-ray Radiography Diagnostics of Components in Solid Rocket Motors

    Science.gov (United States)

    Cortopassi, A. C.; Martin, H. T.; Boyer, E.; Kuo, K. K.

    2012-01-01

    Solid rocket motors (SRMs) typically use nozzle materials which are required to maintain their shape as well as insulate the underlying support structure during the motor operation. In addition, SRMs need internal insulation materials to protect the motor case from the harsh environment resulting from the combustion of solid propellant. In the nozzle, typical materials consist of high density graphite, carbon-carbon composites and carbon phenolic composites. Internal insulation of the motor cases is typically a composite material with carbon, asbestos, Kevlar, or silica fibers in an ablative matrix such as EPDM or NBR. For both nozzle and internal insulation materials, the charring process occurs when the hot combustion products heat the material intensely. The pyrolysis of the matrix material takes away a portion of the thermal energy near the wall surface and leaves behind a char layer. The fiber reinforcement retains the porous char layer which provides continued thermal protection from the hot combustion products. It is of great interest to characterize both the total erosion rates of the material and the char layer thickness. By better understanding of the erosion process for a particular ablative material in a specific flow environment, the required insulation material thickness can be properly selected. The recession rates of internal insulation and nozzle materials of SRMs are typically determined by testing in some sort of simulated environment; either arc-jet testing, flame torch testing, or subscale SRMs of different size. Material recession rates are deduced by comparison of pre- and post-test measurements and then averaging over the duration of the test. However, these averaging techniques cannot be used to determine the instantaneous recession rates of the material. Knowledge of the variation in recession rates in response to the instantaneous flow conditions during the motor operation is of great importance. For example, in many SRM configurations

  5. The development of an erosive burning model for solid rocket motors using direct numerical simulation

    Science.gov (United States)

    McDonald, Brian A.

    A method for developing an erosive burning model for use in solid propellant design-and-analysis interior ballistics codes is described and evaluated. Using Direct Numerical Simulation, the primary mechanisms controlling erosive burning (turbulent heat transfer, and finite rate reactions) have been studied independently through the development of models using finite rate chemistry, and infinite rate chemistry. Both approaches are calibrated to strand burn rate data by modeling the propellant burning in an environment with no cross-flow, and adjusting thermophysical properties until the predicted regression rate matches test data. Subsequent runs are conducted where the cross-flow is increased from M = 0.0 up to M = 0.8. The resulting relationship of burn rate increase versus Mach Number is used in an interior ballistics analysis to compute the chamber pressure of an existing solid rocket motor. The resulting predictions are compared to static test data. Both the infinite rate model and the finite rate model show good agreement when compared to test data. The propellant considered is an AP/HTPB with an average AP particle size of 37 microns. The finite rate model shows that as the cross-flow increases, near wall vorticity increases due to the lifting of the boundary caused by the side injection of gases from the burning propellant surface. The point of maximum vorticity corresponds to the outer edge of the APd-binder flame. As the cross-flow increases, the APd-binder flame thickness becomes thinner; however, the point of highest reaction rate moves only slightly closer to the propellant surface. As such, the net increase of heat transfer to the propellant surface due to finite rate chemistry affects is small. This leads to the conclusion that augmentation of thermal transport properties and the resulting heat transfer increase due to turbulence dominates over combustion chemistry in the erosive burning problem. This conclusion is advantageous in the development of

  6. Analysis and control of the compaction force in the composite prepreg tape winding process for rocket motor nozzles

    Directory of Open Access Journals (Sweden)

    Xiaodong He

    2017-04-01

    Full Text Available In the process of composite prepreg tape winding, the compaction force could influence the quality of winding products. According to the analysis and experiments, during the winding process of a rocket motor nozzle aft exit cone with a winding angle, there would be an error between the deposition speed of tape layers and the feeding speed of the compaction roller, which could influence the compaction force. Both a lack of compaction and overcompaction related to the feeding of the compaction roller could result in defects of winding nozzles. Thus, a flexible winding system has been developed for rocket motor nozzle winding. In the system, feeding of the compaction roller could be adjusted in real time to achieve an invariable compaction force. According to experiments, the force deformation model of the winding tape is a time-varying system. Thus, a forgetting factor recursive least square based parameter estimation proportional-integral-differential (PID controller has been developed, which could estimate the time-varying parameter and control the compaction force by adjusting the feeding of the compaction roller during the winding process. According to the experimental results, a winding nozzle with fewer voids and a smooth surface could be wounded by the invariable compaction force in the flexible winding system.

  7. Extension of a simplified computer program for analysis of solid-propellant rocket motors

    Science.gov (United States)

    Sforzini, R. H.

    1973-01-01

    A research project to develop a computer program for the preliminary design and performance analysis of solid propellant rocket engines is discussed. The following capabilities are included as computer program options: (1) treatment of wagon wheel cross sectional propellant configurations alone or in combination with circular perforated grains, (2) calculation of ignition transients with the igniter treated as a small rocket engine, (3) representation of spherical circular perforated grain ends as an alternative to the conical end surface approximation used in the original program, and (4) graphical presentation of program results using a digital plotter.

  8. Solid propellant ignition motors for LH_2/LOX rocket engine system

    OpenAIRE

    ARAKI, Tetsuo; AKIBA, Ryojiro; HASHIMOTO, Yasunari; AIHARA, Kenji; TOMITA, Etsu; YASUDA, Seiichi; 荒木, 哲夫; 秋葉, 鐐二郎; 橋本, 保成; 相原, 賢二; 富田, 悦; 安田, 誠一

    1983-01-01

    Solid propellant ignition motors are used in the series of experiments of the 10 ton LH_2/LOX engine featured by the channel wall thrust chamber, This paper presents design specification, experiments and results obtained by actual applications of those ignition motors.

  9. Design and Testing of Digitally Manufactured Paraffin Acrylonitrile-Butadiene-Styrene Hybrid Rocket Motors

    OpenAIRE

    McCulley, Jonathan M.

    2013-01-01

    This research investigates the application of additive manufacturing techniques for fabricating hybrid rocket fuel grains composed of porous Acrylonitrile-butadiene-styrene impregnated with paraffin wax. The digitally manufactured ABS substrate provides mechanical support for the paraffin fuel material and serves as an additional fuel component. The embedded paraffin provides an enhanced fuel regression rate while having no detrimental effect on the thermodynamic burn properties of the fuel g...

  10. Finite element method for viscoelastic medium with damage and the application to structural analysis of solid rocket motor grain

    Science.gov (United States)

    Deng, Bin; Shen, ZhiBin; Duan, JingBo; Tang, GuoJin

    2014-05-01

    This paper studies the damage-viscoelastic behavior of composite solid propellants of solid rocket motors (SRM). Based on viscoelastic theories and strain equivalent hypothesis in damage mechanics, a three-dimensional (3-D) nonlinear viscoelastic constitutive model incorporating with damage is developed. The resulting viscoelastic constitutive equations are numerically discretized by integration algorithm, and a stress-updating method is presented by solving nonlinear equations according to the Newton-Raphson method. A material subroutine of stress-updating is made up and embedded into commercial code of Abaqus. The material subroutine is validated through typical examples. Our results indicate that the finite element results are in good agreement with the analytical ones and have high accuracy, and the suggested method and designed subroutine are efficient and can be further applied to damage-coupling structural analysis of practical SRM grain.

  11. Hybrid uncertainty-based design optimization and its application to hybrid rocket motors for manned lunar landing

    Directory of Open Access Journals (Sweden)

    Hao Zhu

    2017-04-01

    Full Text Available Design reliability and robustness are getting increasingly important for the general design of aerospace systems with many inherently uncertain design parameters. This paper presents a hybrid uncertainty-based design optimization (UDO method developed from probability theory and interval theory. Most of the uncertain design parameters which have sufficient information or experimental data are classified as random variables using probability theory, while the others are defined as interval variables with interval theory. Then a hybrid uncertainty analysis method based on Monte Carlo simulation and Taylor series interval analysis is developed to obtain the uncertainty propagation from the design parameters to system responses. Three design optimization strategies, including deterministic design optimization (DDO, probabilistic UDO and hybrid UDO, are applied to the conceptual design of a hybrid rocket motor (HRM used as the ascent propulsion system in Apollo lunar module. By comparison, the hybrid UDO is a feasible method and can be effectively applied to the general design of aerospace systems.

  12. Hybrid uncertainty-based design optimization and its application to hybrid rocket motors for manned lunar landing

    Institute of Scientific and Technical Information of China (English)

    Zhu Hao; Tian Hui; Cai Guobiao

    2017-01-01

    Design reliability and robustness are getting increasingly important for the general design of aerospace systems with many inherently uncertain design parameters. This paper presents a hybrid uncertainty-based design optimization (UDO) method developed from probability theory and interval theory. Most of the uncertain design parameters which have sufficient information or experimental data are classified as random variables using probability theory, while the others are defined as interval variables with interval theory. Then a hybrid uncertainty analysis method based on Monte Carlo simulation and Taylor series interval analysis is developed to obtain the uncer-tainty propagation from the design parameters to system responses. Three design optimization strategies, including deterministic design optimization (DDO), probabilistic UDO and hybrid UDO, are applied to the conceptual design of a hybrid rocket motor (HRM) used as the ascent propulsion system in Apollo lunar module. By comparison, the hybrid UDO is a feasible method and can be effectively applied to the general design of aerospace systems.

  13. A parallel solution-adaptive scheme for predicting multi-phase core flows in solid propellant rocket motors

    International Nuclear Information System (INIS)

    Sachdev, J.S.; Groth, C.P.T.; Gottlieb, J.J.

    2003-01-01

    The development of a parallel adaptive mesh refinement (AMR) scheme is described for solving the governing equations for multi-phase (gas-particle) core flows in solid propellant rocket motors (SRM). An Eulerian formulation is used to described the coupled motion between the gas and particle phases. A cell-centred upwind finite-volume discretization and the use of limited solution reconstruction, Riemann solver based flux functions for the gas and particle phases, and explicit multi-stage time-stepping allows for high solution accuracy and computational robustness. A Riemann problem is formulated for prescribing boundary data at the burning surface. Efficient and scalable parallel implementations are achieved with domain decomposition on distributed memory multiprocessor architectures. Numerical results are described to demonstrate the capabilities of the approach for predicting SRM core flows. (author)

  14. An Automated Fluid-Structural Interaction Analysis of a Large Segmented Solid Rocket Motor

    National Research Council Canada - National Science Library

    Rex, Brian

    2003-01-01

    .... The fluid-structural interaction (FSI) analysis of the ETM-3 motor used PYTHON, a powerful programming language, and FEM BUILDER, a pre- and post processor developed by ATK Thiokol Propulsion under contract to the AFRL, to automatically...

  15. Off Like a Rocket: A Media Discourse Analysis of Tesla Motor Corporation

    OpenAIRE

    McKay, Jordan

    2016-01-01

    Energy and transportation are topics of great importance to global sustainable development.  Tesla Motor Corporation is an electric vehicle company with the objective to “accelerate the world’s transition to sustainable energy” (Musk, 2016).  This thesis, a media discourse analysis, examines media texts concerning Tesla Motors to provide a better understanding of the company’s hitherto success in penetrating the automotive market.  Qualitative analyses of text were utilized to first define th...

  16. Study of solid rocket motors for a space shuttle booster. Volume 1: Executive summary

    Science.gov (United States)

    1972-01-01

    The design, development, production, and launch support analysis for determining the solid propellant rocket engine to be used with the space shuttle are discussed. Specific program objectives considered were: (1) definition of engine designs to satisfy the performance and configuration requirements of the various vehicle/booster concepts, (2) definition of requirements to produce booster stages at rates of 60, 40, 20, and 10 launches per year in a man-rated system, and (3) estimation of costs for the defined SRM booster stages.

  17. Throttleable GOX/ABS launch assist hybrid rocket motor for small scale air launch platform

    Science.gov (United States)

    Spurrier, Zachary S.

    Aircraft-based space-launch platforms allow operational flexibility and offer the potential for significant propellant savings for small-to-medium orbital payloads. The NASA Armstrong Flight Research Center's Towed Glider Air-Launch System (TGALS) is a small-scale flight research project investigating the feasibility for a remotely-piloted, towed, glider system to act as a versatile air launch platform for nano-scale satellites. Removing the crew from the launch vehicle means that the system does not have to be human rated, and offers a potential for considerable cost savings. Utah State University is developing a small throttled launch-assist system for the TGALS platform. This "stage zero" design allows the TGALS platform to achieve the required flight path angle for the launch point, a condition that the TGALS cannot achieve without external propulsion. Throttling is required in order to achieve and sustain the proper launch attitude without structurally overloading the airframe. The hybrid rocket system employs gaseous-oxygen and acrylonitrile butadiene styrene (ABS) as propellants. This thesis summarizes the development and testing campaign, and presents results from the clean-sheet design through ground-based static fire testing. Development of the closed-loop throttle control system is presented.

  18. Effect of ITE and nozzle exit cone erosion on specific impulse of solid rocket motors

    Science.gov (United States)

    Smith-Kent, Randall; Ridder, Jeffrey P.; Loh, Hai-Tien; Abel, Ralph

    1993-06-01

    Specific impulse loss due to the use of a slowly eroding integral throat entrance, or a throat insert, with a faster eroding nozzle exit cone is studied. It is suggested that an oblique shock wave produced by step-off erosion results in loss of specific impulse. This is studied by use of a shock capturing CFD method. The shock loss predictions for first-stage Peacekeeper and Castor 25 motors are found to match the trends of the test data. This work suggests that a loss mechanism, previously unaccounted, should be considered in the specific impulse prediction procedure for nozzles with step-off exit cone erosion.

  19. Space Launch System Base Heating Test: Sub-Scale Rocket Engine/Motor Design, Development and Performance Analysis

    Science.gov (United States)

    Mehta, Manish; Seaford, Mark; Kovarik, Brian; Dufrene, Aaron; Solly, Nathan; Kirchner, Robert; Engel, Carl D.

    2014-01-01

    The Space Launch System (SLS) base heating test is broken down into two test programs: (1) Pathfinder and (2) Main Test. The Pathfinder Test Program focuses on the design, development, hot-fire test and performance analyses of the 2% sub-scale SLS core-stage and booster element propulsion systems. The core-stage propulsion system is composed of four gaseous oxygen/hydrogen RS-25D model engines and the booster element is composed of two aluminum-based model solid rocket motors (SRMs). The first section of the paper discusses the motivation and test facility specifications for the test program. The second section briefly investigates the internal flow path of the design. The third section briefly shows the performance of the model RS-25D engines and SRMs for the conducted short duration hot-fire tests. Good agreement is observed based on design prediction analysis and test data. This program is a challenging research and development effort that has not been attempted in 40+ years for a NASA vehicle.

  20. Three-dimensional multi-physics coupled simulation of ignition transient in a dual pulse solid rocket motor

    Science.gov (United States)

    Li, Yingkun; Chen, Xiong; Xu, Jinsheng; Zhou, Changsheng; Musa, Omer

    2018-05-01

    In this paper, numerical investigation of ignition transient in a dual pulse solid rocket motor has been conducted. An in-house code has been developed in order to solve multi-physics governing equations, including unsteady compressible flow, heat conduction and structural dynamic. The simplified numerical models for solid propellant ignition and combustion have been added. The conventional serial staggered algorithm is adopted to simulate the fluid structure interaction problems in a loosely-coupled manner. The accuracy of the coupling procedure is validated by the behavior of a cantilever panel subjected to a shock wave. Then, the detailed flow field development, flame propagation characteristics, pressure evolution in the combustion chamber, and the structural response of metal diaphragm are analyzed carefully. The burst-time and burst-pressure of the metal diaphragm are also obtained. The individual effects of the igniter's mass flow rate, metal diaphragm thickness and diameter on the ignition transient have been systemically compared. The numerical results show that the evolution of the flow field in the combustion chamber, the temperature distribution on the propellant surface and the pressure loading on the metal diaphragm surface present a strong three-dimensional behavior during the initial ignition stage. The rupture of metal diaphragm is not only related to the magnitude of pressure loading on the diaphragm surface, but also to the history of pressure loading. The metal diaphragm thickness and diameter have a significant effect on the burst-time and burst-pressure of metal diaphragm.

  1. Effect of propellant morphology on acoustics in a planar rocket motor

    Energy Technology Data Exchange (ETDEWEB)

    Daimon, Y.; Jackson, T.L. [University of Illinois at Urbana-Champaign, Center for Simulation of Advanced Rockets, Urbana, IL (United States); Topalian, V. [University of Illinois at Urbana-Champaign, Mechanical Science and Engineering, Urbana, IL (United States); Freund, J.B. [University of Illinois at Urbana-Champaign, Mechanical Science and Engineering, Aerospace Engineering, Urbana, IL (United States); Buckmaster, J. [Buckmaster Research, Urbana, IL (United States)

    2009-03-15

    This paper reports the results of numerical simulations of the acoustics in a two-dimensional (plane) motor using a high-order accurate, low-dissipation numerical solver. For verification we compare solutions to Culick's (AIAA J 4(8):1462-1464, 1966) asymptotic solution for constant injection, and to recent results of Hegab and Kassoy (AIAA J 44(4):812-826, 2006) for a space- and time-dependent mass injection. We present results when the injection boundary condition is described by propellant morphology and by white noise. Morphology strongly affects the amplitude of the longitudinal acoustic modes, and in this connection white noise is not a suitable surrogate. (orig.)

  2. Thrust imbalance of solid rocket motor pairs on Space Shuttle flights

    Science.gov (United States)

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

    1986-01-01

    This analysis extends the investigation presented at the 17th Joint Propulsion Conference in 1981 to include fifteen sets of Space Shuttle flight data. The previous report dealt only with static test data and the first flight pair. The objective is to compare the authors' previous theoretical analysis of thrust imbalance with actual Space Shuttle performance. The theoretical prediction method, which involves a Monte Carlo technique, is reviewed briefly as are salient features of the flight instrumentation system and the statistical analysis. A scheme for smoothing flight data is discussed. The effects of changes in design parameters are discussed with special emphasis on the filament wound motor case being developed to replace the steel case. Good agreement between the predictions and the flight data is demonstrated.

  3. Design and Evaluation of a Turbojet Exhaust Simulator, Utilizing a Solid-Propellant Rocket Motor, for use in Free-Flight Aerodynamic Research Models

    Science.gov (United States)

    deMoraes, Carlos A.; Hagginbothom, William K., Jr.; Falanga, Ralph A.

    1954-01-01

    A method has been developed for modifying a rocket motor so that its exhaust characteristics simulate those of a turbojet engine. The analysis necessary to the design is presented along with tests from which the designs are evaluated. Simulation was found to be best if the exhaust characteristics to be duplicated were those of a turbojet engine at high altitudes and with the afterburner operative.

  4. Transient simulation of regression rate on thrust regulation process in hybrid rocket motor

    Directory of Open Access Journals (Sweden)

    Tian Hui

    2014-12-01

    Full Text Available The main goal of this paper is to study the characteristics of regression rate of solid grain during thrust regulation process. For this purpose, an unsteady numerical model of regression rate is established. Gas–solid coupling is considered between the solid grain surface and combustion gas. Dynamic mesh is used to simulate the regression process of the solid fuel surface. Based on this model, numerical simulations on a H2O2/HTPB (hydroxyl-terminated polybutadiene hybrid motor have been performed in the flow control process. The simulation results show that under the step change of the oxidizer mass flow rate condition, the regression rate cannot reach a stable value instantly because the flow field requires a short time period to adjust. The regression rate increases with the linear gain of oxidizer mass flow rate, and has a higher slope than the relative inlet function of oxidizer flow rate. A shorter regulation time can cause a higher regression rate during regulation process. The results also show that transient calculation can better simulate the instantaneous regression rate in the operation process.

  5. The sky is falling III: The effect of deposition from static solid rocket motor tests on juvenile crops.

    Science.gov (United States)

    Doucette, William J; Curry, Eric; McNeill, Laurie S; Heavilin, Justin

    2017-12-01

    A mixture of combustion products (mainly hydrogen chloride, aluminum oxide, and water) and entrained soil, referred to as Test Fire Soil (TFS), can be deposited on crops during static solid rocket motor tests. The impact of a reported worst-case event was previously evaluated by exposing corn and alfalfa to 3200-gTFS/m 2 at 54days after emergence. Exposures via soil and leaves were evaluated separately. Reduced growth (soil exposure) and leaf "scorch" (leaf exposure) were attributed mainly to the high chloride concentrations in the TFS (56,000mg/kg). A follow-up study was conducted to evaluate the effect of a typical deposition event (70-gTFS/m 2 , estimated by radar during several tests) and exposure (soil and leaves simultaneously) on juvenile corn, alfalfa, and winter wheat. Younger crops were used to examine potential age sensitivity differences. Impact was evaluated by comparing the growth, elemental composition, and leaf chlorophyll content of treated and untreated plants. The relationship between deposition exposure and response was also addressed. Growth of corn, alfalfa, and winter wheat exposed to a typical TFS loading was not impacted, although slightly elevated concentrations of aluminum and iron were found in the leaves. At the highest loadings used for the exposure-response experiment, concentrations of chloride and calcium were higher in TFS-exposed corn leaves than in the untreated leaves. Overall results indicate that exposure to a typical deposition event does not adversely impact juvenile crops and that younger plants may be less vulnerable to TFS. However, higher TFS loadings can cause leaf scorch and increase the leaf concentrations of some elements. Copyright © 2017 Elsevier B.V. All rights reserved.

  6. Thermal-Flow Code for Modeling Gas Dynamics and Heat Transfer in Space Shuttle Solid Rocket Motor Joints

    Science.gov (United States)

    Wang, Qunzhen; Mathias, Edward C.; Heman, Joe R.; Smith, Cory W.

    2000-01-01

    A new, thermal-flow simulation code, called SFLOW. has been developed to model the gas dynamics, heat transfer, as well as O-ring and flow path erosion inside the space shuttle solid rocket motor joints by combining SINDA/Glo, a commercial thermal analyzer. and SHARPO, a general-purpose CFD code developed at Thiokol Propulsion. SHARP was modified so that friction, heat transfer, mass addition, as well as minor losses in one-dimensional flow can be taken into account. The pressure, temperature and velocity of the combustion gas in the leak paths are calculated in SHARP by solving the time-dependent Navier-Stokes equations while the heat conduction in the solid is modeled by SINDA/G. The two codes are coupled by the heat flux at the solid-gas interface. A few test cases are presented and the results from SFLOW agree very well with the exact solutions or experimental data. These cases include Fanno flow where friction is important, Rayleigh flow where heat transfer between gas and solid is important, flow with mass addition due to the erosion of the solid wall, a transient volume venting process, as well as some transient one-dimensional flows with analytical solutions. In addition, SFLOW is applied to model the RSRM nozzle joint 4 subscale hot-flow tests and the predicted pressures, temperatures (both gas and solid), and O-ring erosions agree well with the experimental data. It was also found that the heat transfer between gas and solid has a major effect on the pressures and temperatures of the fill bottles in the RSRM nozzle joint 4 configuration No. 8 test.

  7. An evaluation of the total quality management implementation strategy for the advanced solid rocket motor project at NASA's Marshall Space Flight Center. M.S. Thesis - Tennessee Univ.

    Science.gov (United States)

    Schramm, Harry F.; Sullivan, Kenneth W.

    1991-01-01

    An evaluation of the NASA's Marshall Space Flight Center (MSFC) strategy to implement Total Quality Management (TQM) in the Advanced Solid Rocket Motor (ASRM) Project is presented. The evaluation of the implementation strategy reflected the Civil Service personnel perspective at the project level. The external and internal environments at MSFC were analyzed for their effects on the ASRM TQM strategy. Organizational forms, cultures, management systems, problem solving techniques, and training were assessed for their influence on the implementation strategy. The influence of ASRM's effort was assessed relative to its impact on mature projects as well as future projects at MSFC.

  8. Shuttle Rocket Motor Program: NASA should delay awarding some construction contracts. Report to the Chair, Subcommittee on Government Activities and Transportation, Committee on Government Operations, House of Representatives

    Science.gov (United States)

    1992-01-01

    Even though the executive branch has proposed terminating the Advanced Solid Rocket Motor (ASRM) program, NASA is proceeding with all construction activity planned for FY 1992 to avoid schedule slippage if the program is reinstated by Congress. However, NASA could delay some construction activities for at least a few months without affecting the current launch data schedule. For example, NASA could delay Yellow Creek's motor storage and dock projects, Stennis' dock project, and Kennedy's rotation processing and surge facility and dock projects. Starting all construction activities as originally planned could result in unnecessarily incurring additional costs and termination liability if the funding for FY 1993 is not provided. If Congress decides to continue the program, construction could still be completed in time to avoid schedule slippage.

  9. Simulation of reactive polydisperse sprays strongly coupled to unsteady flows in solid rocket motors: Efficient strategy using Eulerian Multi-Fluid methods

    Science.gov (United States)

    Sibra, A.; Dupays, J.; Murrone, A.; Laurent, F.; Massot, M.

    2017-06-01

    In this paper, we tackle the issue of the accurate simulation of evaporating and reactive polydisperse sprays strongly coupled to unsteady gaseous flows. In solid propulsion, aluminum particles are included in the propellant to improve the global performances but the distributed combustion of these droplets in the chamber is suspected to be a driving mechanism of hydrodynamic and acoustic instabilities. The faithful prediction of two-phase interactions is a determining step for future solid rocket motor optimization. When looking at saving computational ressources as required for industrial applications, performing reliable simulations of two-phase flow instabilities appears as a challenge for both modeling and scientific computing. The size polydispersity, which conditions the droplet dynamics, is a key parameter that has to be accounted for. For moderately dense sprays, a kinetic approach based on a statistical point of view is particularly appropriate. The spray is described by a number density function and its evolution follows a Williams-Boltzmann transport equation. To solve it, we use Eulerian Multi-Fluid methods, based on a continuous discretization of the size phase space into sections, which offer an accurate treatment of the polydispersion. The objective of this paper is threefold: first to derive a new Two Size Moment Multi-Fluid model that is able to tackle evaporating polydisperse sprays at low cost while accurately describing the main driving mechanisms, second to develop a dedicated evaporation scheme to treat simultaneously mass, moment and energy exchanges with the gas and between the sections. Finally, to design a time splitting operator strategy respecting both reactive two-phase flow physics and cost/accuracy ratio required for industrial computations. Using a research code, we provide 0D validations of the new scheme before assessing the splitting technique's ability on a reference two-phase flow acoustic case. Implemented in the industrial

  10. 8-inch IBM floppy disk

    CERN Multimedia

    1971-01-01

    The 8-inch floppy disk was a magnetic storage disk for the data introduced commercially by IBM in 1971. It was designed by an IBM team as an inexpensive way to load data into the IBM System / 370. Plus it was a read-only bare disk containing 80 KB of data. The first read-write version was introduced in 1972 by Memorex and could contain 175 KB on 50 tracks (with 8 sectors per track). Other improvements have led to various coatings and increased capacities. Finally, it was surpassed by the mini diskette of 5.25 inches introduced in 1976.

  11. Space Launch System Base Heating Test: Sub-Scale Rocket Engine/Motor Design, Development & Performance Analysis

    Science.gov (United States)

    Mehta, Manish; Seaford, Mark; Kovarik, Brian; Dufrene, Aaron; Solly, Nathan

    2014-01-01

    ATA-002 Technical Team has successfully designed, developed, tested and assessed the SLS Pathfinder propulsion systems for the Main Base Heating Test Program. Major Outcomes of the Pathfinder Test Program: Reach 90% of full-scale chamber pressure Achieved all engine/motor design parameter requirements Reach steady plume flow behavior in less than 35 msec Steady chamber pressure for 60 to 100 msec during engine/motor operation Similar model engine/motor performance to full-scale SLS system Mitigated nozzle throat and combustor thermal erosion Test data shows good agreement with numerical prediction codes Next phase of the ATA-002 Test Program Design & development of the SLS OML for the Main Base Heating Test Tweak BSRM design to optimize performance Tweak CS-REM design to increase robustness MSFC Aerosciences and CUBRC have the capability to develop sub-scale propulsion systems to meet desired performance requirements for short-duration testing.

  12. On the importance of reduced scale Ariane 5 P230 solid rocket motor models in the comprehension and prevention of thrust oscillations

    Science.gov (United States)

    Hijlkema, J.; Prévost, M.; Casalis, G.

    2011-09-01

    Down-scaled solid propellant motors are a valuable tool in the study of thrust oscillations and the underlying, vortex-shedding-induced, pressure instabilities. These fluctuations, observed in large segmented solid rocket motors such as the Ariane 5 P230, impose a serious constraint on both structure and payload. This paper contains a survey of the numerous configurations tested at ONERA over the last 20 years. Presented are the phenomena searched to reproduce and the successes and failures of the different approaches tried. The results of over 130 experiments have contributed to numerous studies aimed at understanding the complicated physics behind this thorny problem, in order to pave the way to pressure instability reduction measures. Slowly but surely our understanding of what makes large segmented solid boosters exhibit this type of instabilities will lead to realistic modifications that will allow for a reduction of pressure oscillations. A "quieter" launcher will be an important advantage in an ever more competitive market, giving a easier ride to payload and designers alike.

  13. Rocket Flight.

    Science.gov (United States)

    Van Evera, Bill; Sterling, Donna R.

    2002-01-01

    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)

  14. Rocket science

    International Nuclear Information System (INIS)

    Upson Sandra

    2011-01-01

    Expanding across the Solar System will require more than a simple blast off, a range of promising new propulsion technologies are being investigated by ex- NASA shuttle astronaut Chang Diaz. He is developing an alternative to chemical rockets, called VASIMR -Variable Specific Impulse Magnetoplasm Rocket. In 2012 Ad Astra plans to test a prototype, using solar power rather than nuclear, on the International Space Station. Development of this rocket for human space travel is discussed. The nuclear reactor's heat would be converted into electricity in an electric rocket such as VASIMR, and at the peak of nuclear rocket research thrust levels of almost one million newtons were reached.

  15. This Is Rocket Science!

    Science.gov (United States)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-09-01

    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 than that offered by Estes Industries, but more basic than the analysis of Nelson et al. In particular, drag is neglected until the very end of the exercise, which allows the concept of conservation of energy to be shown when predicting the rocket's flight. Also, the variable mass of the rocket motor is assumed to decrease linearly during the flight (while the propulsion charge and recovery delay charge are burning) and handled simplistically by using an average mass value. These changes greatly simplify the equations needed to predict the times and heights at various stages of flight, making it more useful as a review of basic physics. Details about model rocket motors, range safety, and other supplemental information may be found online at Apogee Components4 and the National Association of Rocketry.5

  16. The sky is falling II: Impact of deposition produced during the static testing of solid rocket motors on corn and alfalfa.

    Science.gov (United States)

    Doucette, William J; Mendenhall, Scout; McNeill, Laurie S; Heavilin, Justin

    2014-06-01

    Tests of horizontally restrained rocket motors at the ATK facility in Promontory, Utah, USA result in the deposition of an estimated 1.5million kg of entrained soil and combustion products (mainly aluminum oxide, gaseous hydrogen chloride and water) on the surrounding area. The deposition is referred to as test fire soil (TFS). Farmers observing TFS deposited on their crops expressed concerns regarding the impact of this material. To address these concerns, we exposed corn and alfalfa to TFS collected during a September 2009 test. The impact was evaluated by comparing the growth and tissue composition of controls relative to the treatments. Exposure to TFS, containing elevated levels of chloride (1000 times) and aluminum (2 times) relative to native soils, affected the germination, growth and tissue concentrations of various elements, depending on the type and level of exposure. Germination was inhibited by high concentrations of TFS in soil, but the impact was reduced if the TFS was pre-leached with water. Biomass production was reduced in the TFS amended soils and corn grown in TFS amended soils did not develop kernels. Chloride concentrations in corn and alfalfa grown in TFS amended soils were two orders of magnitude greater than controls. TFS exposed plants contained higher concentrations of several cations, although the concentrations were well below livestock feed recommendations. Foliar applications of TFS had no impact on biomass, but some differences in the elemental composition of leaves relative to controls were observed. Washing the TFS off the leaves lessened the impact. Results indicate that the TFS deposition could have an effect, depending on the amount and growth stage of the crops, but the impact could be mitigated with rainfall or the application of additional irrigation water. The high level of chloride associated with the TFS is the main cause of the observed impacts. Copyright © 2014 Elsevier B.V. All rights reserved.

  17. Nuclear rockets

    International Nuclear Information System (INIS)

    Sarram, M.

    1972-01-01

    Nuclear energy has found many applications in space projects. This article deals with these applications. The first application is the use of nuclear energy for the production of electricity in space and the second main application is the use of nuclear energy for propulsion purposes in space flight. The main objective is to develop a 75000 pound thrust flight engine call NERVA by heating liquid hydrogen, in a nuclear reactor, from 420F to 4000 0 F. The paper describes in detail the salient features of the NERVA rocket as well as its comparison with the conventional chemical rockets. It is shown that a nuclear rocket using liquid hydrogen as medium is at least 85% more efficient as compared with the chemical rockets such as those used for the APOLLO moon flight

  18. Nuclear rockets

    Energy Technology Data Exchange (ETDEWEB)

    Sarram, M [Teheran Univ. (Iran). Inst. of Nuclear Science and Technology

    1972-02-01

    Nuclear energy has found many applications in space projects. This article deals with these applications. The first application is the use of nuclear energy for the production of electricity in space and the second main application is the use of nuclear energy for propulsion purposes in space flight. The main objective is to develop a 75000 pound thrust flight engine called NERVA by heating liquid hydrogen in a nuclear reactor. The paper describes in detail the salient features of the NERVA rocket as well as its comparison with the conventional chemical rockets. It is shown that a nuclear rocket using liquid hydrogen as medium is at least 85% more efficient as compared with the chemical rockets such as those used for the APOLLO moon flight.

  19. High-speed schlieren imaging of rocket exhaust plumes

    Science.gov (United States)

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

    2016-11-01

    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.

  20. Hybrid rocket engine, theoretical model and experiment

    Science.gov (United States)

    Chelaru, Teodor-Viorel; Mingireanu, Florin

    2011-06-01

    The purpose of this paper is to build a theoretical model for the hybrid rocket engine/motor and to validate it using experimental results. The work approaches the main problems of the hybrid motor: the scalability, the stability/controllability of the operating parameters and the increasing of the solid fuel regression rate. At first, we focus on theoretical models for hybrid rocket motor and compare the results with already available experimental data from various research groups. A primary computation model is presented together with results from a numerical algorithm based on a computational model. We present theoretical predictions for several commercial hybrid rocket motors, having different scales and compare them with experimental measurements of those hybrid rocket motors. Next the paper focuses on tribrid rocket motor concept, which by supplementary liquid fuel injection can improve the thrust controllability. A complementary computation model is also presented to estimate regression rate increase of solid fuel doped with oxidizer. Finally, the stability of the hybrid rocket motor is investigated using Liapunov theory. Stability coefficients obtained are dependent on burning parameters while the stability and command matrixes are identified. The paper presents thoroughly the input data of the model, which ensures the reproducibility of the numerical results by independent researchers.

  1. Technology for low cost solid rocket boosters.

    Science.gov (United States)

    Ciepluch, C.

    1971-01-01

    A review of low cost large solid rocket motors developed at the Lewis Research Center is given. An estimate is made of the total cost reduction obtainable by incorporating this new technology package into the rocket motor design. The propellant, case material, insulation, nozzle ablatives, and thrust vector control are discussed. The effect of the new technology on motor cost is calculated for a typical expandable 260-in. booster application. Included in the cost analysis is the influence of motor performance variations due to specific impulse and weight changes. It is found for this application that motor costs may be reduced by up to 30% and that the economic attractiveness of future large solid rocket motors will be improved when the new technology is implemented.

  2. Rocket observations

    Science.gov (United States)

    1984-05-01

    The Institute of Space and Astronautical Science (ISAS) sounding rocket experiments were carried out during the periods of August to September, 1982, January to February and August to September, 1983 and January to February, 1984 with sounding rockets. Among 9 rockets, 3 were K-9M, 1 was S-210, 3 were S-310 and 2 were S-520. Two scientific satellites were launched on February 20, 1983 for solar physics and on February 14, 1984 for X-ray astronomy. These satellites were named as TENMA and OHZORA and designated as 1983-011A and 1984-015A, respectively. Their initial orbital elements are also described. A payload recovery was successfully carried out by S-520-6 rocket as a part of MINIX (Microwave Ionosphere Non-linear Interaction Experiment) which is a scientific study of nonlinear plasma phenomena in conjunction with the environmental assessment study for the future SPS project. Near IR observation of the background sky shows a more intense flux than expected possibly coming from some extragalactic origin and this may be related to the evolution of the universe. US-Japan cooperative program of Tether Experiment was done on board US rocket.

  3. Determination of the availability of appropriate aged flight rocket motors. [captive tests to determine case bond separation and grain bore cracking

    Science.gov (United States)

    Martin, P. J.

    1974-01-01

    A program to identify surplus solid rocket propellant engines which would be available for a program of functional integrity testing was conducted. The engines are classified as: (1) upper stage and apogee engines, (2) sounding rocket and launch vehicle engines, and (3) jato, sled, and tactical engines. Nearly all the engines were available because their age exceeds the warranted shelf life. The preference for testing included tests at nominal flight conditions, at design limits, and to establish margin limits. The principal failure modes of interest were case bond separation and grain bore cracking. Data concerning the identification and characteristics of each engine are tabulated. Methods for conducting the tests are described.

  4. Air-Powered Rockets.

    Science.gov (United States)

    Rodriguez, Charley; Raynovic, Jim

    This document describes methods for designing and building two types of rockets--rockets from paper and rockets from bottles. Devices used for measuring the heights that the rockets obtain are also discussed. (KHR)

  5. Celiac Disease: Four Inches and Seven Pounds...

    Science.gov (United States)

    ... Disease" Articles Celiac Disease Changes Everything / What is Celiac Disease? / Symptoms, Diagnosis & Treatment / Four Inches and Seven Pounds… / Learning to Live Well with Celiac Disease / Living Gluten-Free Spring 2015 Issue: Volume 10 ...

  6. Improving of Hybrid Rocket Engine on the Basis of Optimizing Design Fuel Grain

    Science.gov (United States)

    Oriekov, K. M.; Ushkin, M. P.

    2015-09-01

    This article examines the processes intrachamber in hybrid rocket engine (HRE) and the comparative assessment of the use of solid rocket motors (SRM) and HRE for meteorological rockets with a mass of payload of the 364 kg. Results of the research showed the possibility of a significant increase in the ballistic effectiveness of meteorological rocket.

  7. Rocket Tablet,

    Science.gov (United States)

    1984-09-12

    not accustomed to Chinese food, he ran off directly to the home of the Mayor of Beijing and requested two Western cuisine cooks from a hotel. At the...played out by our Chinese sons and daughters of ancient times. The famous Han dynasty general Li Guang was quickly cured of disease and led an army...Union) of China. This place was about to become the birthplace of the Chinese people’s first rocket baby. Section One In this eternal wasteland called

  8. An investigation of the aerodynamic characteristics of a 0.00548 scale model (model no. 486) of the space shuttle 146-inch diameter solid rocket booster at angels of attack from 113 deg to 180 deg in the AEDC PWT 4-foot transonic wind tunnel (SA16F)

    Science.gov (United States)

    Ramsey, P. E.

    1976-01-01

    An experimental investigation (SA16F) was conducted in the AEDC PWT 4T to determine the entry static stability of a 0.00548 scale space shuttle solid rocket booster (SRB). The primary objective was to improve the definition of the aerodynamic characteristics in the angle of attack range beyond 90 deg in the vicinity of the entry trim point. The SRB scale model consisted of the reentry configuration with all major protuberances. A simulated heat shield around the engine nozzle was also included. Data were obtained for a 60 deg side mounted sting and a straight nose mounted sting. The angle of attack range for the side mounted sting was 113 deg to 147 deg and for the nose mounted sting 152 deg to 187 deg. The Mach number range consisted of 0.4 to 1.2 at roll angles of 0 and 90 deg. The resulting 6-component aerodynamic force data was presented as the variation of coefficients with angle of attack for each Mach number and roll angle.

  9. 83-inch cyclotron research program. Final report

    International Nuclear Information System (INIS)

    Parkinson, W.C.

    1983-07-01

    In June of 1960 the US Atomic Energy Commission authorized the construction of a modern variable energy cyclotron facility at The University of Michigan to be used for research in nuclear spectroscopy. The Legislature of the State of Michigan made available funds for construction of a building to house the 83-inch cyclotron and auxiliary equipment as well as the University's remodeled 42-inch cyclotron. The research program centered around the 83-inch cyclotron was funded by the AEC and its successors, the Energy Research and Development Administration (ERDA) and the Department of Energy (DOE), from September 1964 through March 1977. The program represented a continuation of the research effort using the 42-inch cyclotron facility which had been supported continuously by the AEC since February 1950. This final report to DOE briefly describes the research facility, the research program, and highlights the principal accomplishments of the effort. It begins with a historical note to place this effort within the context of nuclear physics research in the Department of Physics of the University of Michigan

  10. Infrared Imagery of Solid Rocket Exhaust Plumes

    Science.gov (United States)

    Moran, Robert P.; Houston, Janice D.

    2011-01-01

    The Ares I Scale Model Acoustic Test program consisted of a series of 18 solid rocket motor static firings, simulating the liftoff conditions of the Ares I five-segment Reusable Solid Rocket Motor Vehicle. Primary test objectives included acquiring acoustic and pressure data which will be used to validate analytical models for the prediction of Ares 1 liftoff acoustics and ignition overpressure environments. The test article consisted of a 5% scale Ares I vehicle and launch tower mounted on the Mobile Launch Pad. The testing also incorporated several Water Sound Suppression Systems. Infrared imagery was employed during the solid rocket testing to support the validation or improvement of analytical models, and identify corollaries between rocket plume size or shape and the accompanying measured level of noise suppression obtained by water sound suppression systems.

  11. Design criteria of launching rockets for burst aerial shells

    Energy Technology Data Exchange (ETDEWEB)

    Kuwahara, T.; Takishita, Y.; Onda, T.; Shibamoto, H.; Hosaya, F. [Hosaya Kako Co. Ltd (Japan); Kubota, N. [Mitsubishi Electric Corporation (Japan)

    2000-04-01

    Rocket motors attached to large-sized aerial shells are proposed to compensate for the increase in the lifting charge in the mortar and the thickness of the shell wall. The proposal is the result of an evaluation of the performance of solid propellants to provide information useful in designing launch rockets for large-size shells. The propellants composed of ammonium perchlorate and hydroxy-terminated polybutadiene were used to evaluate the ballistic characteristics such as the relationship between propellant mass and trajectories of shells and launch rockets. In order to obtain an optimum rocket design, the evaluation also included a study of the velocity and height of the rocket motor and shell separation. A launch rocket with a large-sized shell (84.5 cm in diameter) was designed to verify the effectiveness of this class of launch system. 2 refs., 6 figs.

  12. Design of a Subscale Propellant Slag Evaluation Motor Using Two-Phase Fluid Dynamic Analysis

    Science.gov (United States)

    Whitesides, R. Harold; Dill, Richard A.; Purinton, David C.; Sambamurthi, Jay K.

    1996-01-01

    Small pressure perturbations in the Space Shuttle Reusable Solid Rocket Motor (RSRM) are caused by the periodic expulsion of molten aluminum oxide slag from a pool that collects in the aft end of the motor around the submerged nozzle nose during the last half of motor operation. It is suspected that some motors produce more slag than others due to differences in aluminum oxide agglomerate particle sizes that may relate to subtle differences in propellant ingredient characteristics such as particle size distributions or processing variations. A subscale motor experiment was designed to determine the effect of propellant ingredient characteristics on the propensity for slag production. An existing 5 inch ballistic test motor was selected as the basic test vehicle. The standard converging/diverging nozzle was replaced with a submerged nose nozzle design to provide a positive trap for the slag that would increase the measured slag weights. Two-phase fluid dynamic analyses were performed to develop a nozzle nose design that maintained similitude in major flow field features with the full scale RSRM. The 5 inch motor was spun about its longitudinal axis to further enhance slag collection and retention. Two-phase flow analysis was used to select an appropriate spin rate along with other considerations, such as avoiding bum rate increases due to radial acceleration effects. Aluminum oxide particle distributions used in the flow analyses were measured in a quench bomb for RSRM type propellants with minor variations in ingredient characteristics. Detailed predictions for slag accumulation weights during motor bum compared favorably with slag weight data taken from defined zones in the subscale motor and nozzle. The use of two-phase flow analysis proved successful in gauging the viability of the experimental program during the planning phase and in guiding the design of the critical submerged nose nozzle.

  13. On use of hybrid rocket propulsion for suborbital vehicles

    Science.gov (United States)

    Okninski, Adam

    2018-04-01

    While the majority of operating suborbital rockets use solid rocket propulsion, recent advancements in the field of hybrid rocket motors lead to renewed interest in their use in sounding rockets. This paper presents results of optimisation of sounding rockets using hybrid propulsion. An overview of vehicles under development during the last decade, as well as heritage systems is provided. Different propellant combinations are discussed and their performance assessment is given. While Liquid Oxygen, Nitrous Oxide and Nitric Acid have been widely tested with various solid fuels in flight, Hydrogen Peroxide remains an oxidiser with very limited sounding rocket applications. The benefits of hybrid propulsion for sounding rockets are given. In case of hybrid rocket motors the thrust curve can be optimised for each flight, using a flow regulator, depending on the payload and mission. Results of studies concerning the optimal burn duration and nozzle selection are given. Specific considerations are provided for the Polish ILR-33 "Amber" sounding rocket. Low regression rates, which up to date were viewed as a drawback of hybrid propulsion may be used to the benefit of maximising rocket performance if small solid rocket boosters are used during the initial flight period. While increased interest in hybrid propulsion is present, no up-to-date reference concerning use of hybrid rocket propulsion for sounding rockets is available. The ultimate goal of the paper is to provide insight into the sensitivity of different design parameters on performance of hybrid sounding rockets and delve into the potential and challenges of using hybrid rocket technology for expendable suborbital applications.

  14. ELIMINATION OF ROCKET IGNITION SIDE LOADS, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — This proposal is responsive to Topic H10: Ground Processing and in particular to Subtopic H10.02. When a rocket motor/engine is ignited at low altitude its...

  15. Space Shuttle solid rocket booster

    Science.gov (United States)

    Hardy, G. B.

    1979-01-01

    Details of the design, operation, testing and recovery procedures of the reusable solid rocket boosters (SRB) are given. Using a composite PBAN propellant, they will provide the primary thrust (six million pounds maximum at 20 s after ignition) within a 3 g acceleration constraint, as well as thrust vector control for the Space Shuttle. The drogues were tested to a load of 305,000 pounds, and the main parachutes to 205,000. Insulation in the solid rocket motor (SRM) will be provided by asbestos-silica dioxide filled acrylonitrile butadiene rubber ('asbestos filled NBR') except in high erosion areas (principally in the aft dome), where a carbon-filled ethylene propylene diene monomer-neopreme rubber will be utilized. Furthermore, twenty uses for the SRM nozzle will be allowed by its ablative materials, which are principally carbon cloth and silica cloth phenolics.

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

    Science.gov (United States)

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

    2017-09-01

    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.

  17. Size of silicon strip sensor from 6 inch wafer (right) compared to that from a 4 inch wafer (left).

    CERN Multimedia

    Honma, Alan

    1999-01-01

    Silicon strip sensors made from 6 inch wafers will allow for much larger surface area coverage at a reduced cost per unit surface area. A prototype sensor of size 8cm x 11cm made by Hamamatsu from a 6 inch wafer is shown next to a traditional 6cm x 6cm sensor from a 4 inch wafer.

  18. Rockets two classic papers

    CERN Document Server

    Goddard, Robert

    2002-01-01

    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

  19. History of Solid Rockets

    Science.gov (United States)

    Green, Rebecca

    2017-01-01

    Solid rockets are of interest to the space program because they are commonly used as boosters that provide the additional thrust needed for the space launch vehicle to escape the gravitational pull of the Earth. Larger, more advanced solid rockets allow for space launch vehicles with larger payload capacities, enabling mankind to reach new depths of space. This presentation will discuss, in detail, the history of solid rockets. The history begins with the invention and origin of the solid rocket, and then goes into the early uses and design of the solid rocket. The evolution of solid rockets is depicted by a description of how solid rockets changed and improved and how they were used throughout the 16th, 17th, 18th, and 19th centuries. Modern uses of the solid rocket include the Solid Rocket Boosters (SRBs) on the Space Shuttle and the solid rockets used on current space launch vehicles. The functions and design of the SRB and the advancements in solid rocket technology since the use of the SRB are discussed as well. Common failure modes and design difficulties are discussed as well.

  20. Eddie Rocket's Franchise

    OpenAIRE

    Vahter, Jenni

    2008-01-01

    Eddie Rocket's Franchise - Setting up a franchise restaurant in Helsinki. TIIVISTELMÄ: Eddie Rocket's on menestynyt amerikkalaistyylinen 1950-luvun ”diner” franchiseravintolaketju Irlannista. Ravintoloita on perustettu viimeisen 18 vuoden aikana 28 kappaletta Irlantiin ja Isoon Britanniaan sekä yksi Espanjaan. Tämän tutkimuksen tarkoitus on tutkia onko Eddie Rocket'silla potentiaalia menestyä Helsingissä, Suomessa. Tutkimuskysymystä on lähestytty toimiala-analyysin, markkinatutkimuksen j...

  1. Liquid Rocket Engine Testing

    Science.gov (United States)

    2016-10-21

    Briefing Charts 3. DATES COVERED (From - To) 17 October 2016 – 26 October 2016 4. TITLE AND SUBTITLE Liquid Rocket Engine Testing 5a. CONTRACT NUMBER...298 (Rev. 8-98) Prescribed by ANSI Std. 239.18 Liquid Rocket Engine Testing SFTE Symposium 21 October 2016 Jake Robertson, Capt USAF AFRL...Distribution Unlimited. PA Clearance 16493 Liquid Rocket Engine Testing • Engines and their components are extensively static-tested in development • This

  2. The flight of uncontrolled rockets

    CERN Document Server

    Gantmakher, F R; Dryden, H L

    1964-01-01

    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

  3. Substituição de amianto por silicato de alumínio e grafite expansível em compósito de poliuretano utilizado em motor-foguete Substitution of asbestos for aluminosiliacate and expandable graphite in polyurethane composites used in rocket motors

    Directory of Open Access Journals (Sweden)

    Henrique Crespim

    2007-09-01

    Full Text Available Compósitos de poliuretano e amianto (liner são utilizados como revestimento interno em paredes de motor-foguete, conferindo proteção térmica e garantindo a adesão entre o propelente e as paredes do motor. No entanto, o uso do amianto tem sido restringido devido à sua toxidade. No presente trabalho, o amianto foi substituído por um silicato de alumínio hidratado (SA e pelo grafite expansível (GE em diferentes teores no liner. Resultados de análise termogravimétrica (TG mostraram que a estabilidade térmica do liner praticamente não é afetada pela substituição das cargas, embora a energia de ativação (Ea obtida para a decomposição tenha mudado, mostrando maiores valores para as amostras contendo as cargas SA e GE. A análise termomecânica (TMA mostrou que o coeficiente de expansão térmica linear do liner contendo SA foi menor que aquele encontrado para o liner contendo amianto. O liner contendo a carga SA também apresentou os maiores valores de tensão nos testes mecânicos de tração.Composites of polyurethane (PU and asbestos (liner are used as internal coating of rocket motors, providing thermal protection and assuring the adhesion between propellant and the motor walls. However, the use of asbestos has been restricted due to its hazardous nature. In the present work, asbestos was replaced by hydrated alumina silicate (SA and expandable graphite (GE in different contents. Thermogravimetric analysis (TG showed that the thermal stability of liners was practically unaffected by the filler replacement although the activation energy obtained for the decomposition has changed. Thermomechanical analysis (TMA showed that coefficients of thermal expansion of SA/liners were lower than asbestos/liner. SA/liners also presented the highest tension values in mechanical tests.

  4. 30 CFR 18.34 - Motors.

    Science.gov (United States)

    2010-07-01

    ... houses the bearing. (9) If unavoidable, holes may be made through motor casings for bolts, studs, or... narrow interpoles, the distance from the edge of the pole piece to any bolt hole in the frame shall be... bolt shall not exceed 1/64 inch for not less than 1/2 inch through the frame. Furthermore, the pole...

  5. Application of Two-Phase CFD to the Design and Analysis of a Subscale Motor Experiment to Evaluate Propellant Slag Production

    Science.gov (United States)

    Whitesides, R. Harold; Dill, Richard A.

    1996-01-01

    The redesigned solid rocket motor (RSRM) Pressure Perturbation Investigation Team concluded that the cause of recent pressure spikes during both static and flight motor burns was the expulsion of molten aluminum oxide slag from a pool which collects in the aft end of the motor around the submerged nozzle nose during the last half of motor operation. It is suspected that some motors produce more slag than others due to differences in aluminum oxide agglomerate particle sizes which may relate to subtle differences in propellant ingredient characteristics such as particle size distribution, contaminants, or processing variations. In order to determine the effect of suspect propellant ingredient characteristics on the propensity for slag production in a real motor environment, a subscale motor experiment was designed. An existing 5 inch ballistic test motor was selected as the basic test vehicle due to low cost and quick turn around times. The standard converging/diverging nozzle was replaced with a submerged nozzle nose design to provide a positive trap for the slag which would increase both the quantity and repeatability of measured slag weights. Computational fluid dynamics (CFD) was used to assess a variety of submerged nose configurations to identify the design which possessed the best capability to reliably collect slag. CFD also was used to assure that the final selected nozzle design would result in flow field characteristics such as dividing streamline location, nose attach point, and separated flow structure which would have similtude with the RSRM submerged nozzle nose flow field. It also was decided to spin the 5 inch motor about its longitudinal axis to further enhance slag collection quantities. Again, CFD was used to select an appropriate spin rate along with other considerations, including the avoidance of burn rate enhancement from radial acceleration effects.

  6. South Pole rockets, (1)

    International Nuclear Information System (INIS)

    Kimura, Iwane

    1977-01-01

    Wave-particle interaction was observed, using three rockets, S-210 JA-20, -21 and S-310 JA-2, launched from the South Pole into aurora. Electron density and temperature were measured with these rockets. Simultaneous observations of waves were also made from a satellite (ISIS-II) and at two ground bases (Showa base and Mizuho base). Observed data are presented in this paper. These include electron density and temperature in relation to altitude; variation of electron (60 - 80 keV) count rate with altitude; VLF spectra measured by the PWL of S-210 JA-20 and -21 rockets and the corresponding VLF spectra at the ground bases; low-energy (<10 keV) electron flux measured by S-310 JA-2 rocket; and VLF spectrum measured with S-310 JA-2 rocket. Scheduled measurements for the next project are also briefly described. (Aoki, K.)

  7. Pegasus Rocket Model

    Science.gov (United States)

    1996-01-01

    A small, desk-top model of Orbital Sciences Corporation's Pegasus winged rocket booster. Pegasus is an air-launched space booster produced by Orbital Sciences Corporation and Hercules Aerospace Company (initially; later, Alliant Tech Systems) to provide small satellite users with a cost-effective, flexible, and reliable method for placing payloads into low earth orbit. Pegasus has been used to launch a number of satellites and the PHYSX experiment. That experiment consisted of a smooth glove installed on the first-stage delta wing of the Pegasus. The glove was used to gather data at speeds of up to Mach 8 and at altitudes approaching 200,000 feet. The flight took place on October 22, 1998. The PHYSX experiment focused on determining where boundary-layer transition occurs on the glove and on identifying the flow mechanism causing transition over the glove. Data from this flight-research effort included temperature, heat transfer, pressure measurements, airflow, and trajectory reconstruction. Hypersonic flight-research programs are an approach to validate design methods for hypersonic vehicles (those that fly more than five times the speed of sound, or Mach 5). Dryden Flight Research Center, Edwards, California, provided overall management of the glove experiment, glove design, and buildup. Dryden also was responsible for conducting the flight tests. Langley Research Center, Hampton, Virginia, was responsible for the design of the aerodynamic glove as well as development of sensor and instrumentation systems for the glove. Other participating NASA centers included Ames Research Center, Mountain View, California; Goddard Space Flight Center, Greenbelt, Maryland; and Kennedy Space Center, Florida. Orbital Sciences Corporation, Dulles, Virginia, is the manufacturer of the Pegasus vehicle, while Vandenberg Air Force Base served as a pre-launch assembly facility for the launch that included the PHYSX experiment. NASA used data from Pegasus launches to obtain considerable

  8. US Rocket Propulsion Industrial Base Health Metrics

    Science.gov (United States)

    Doreswamy, Rajiv

    2013-01-01

    The number of active liquid rocket engine and solid rocket motor development programs has severely declined since the "space race" of the 1950s and 1960s center dot This downward trend has been exacerbated by the retirement of the Space Shuttle, transition from the Constellation Program to the Space launch System (SLS) and similar activity in DoD programs center dot In addition with consolidation in the industry, the rocket propulsion industrial base is under stress. To Improve the "health" of the RPIB, we need to understand - The current condition of the RPIB - How this compares to past history - The trend of RPIB health center dot This drives the need for a concise set of "metrics" - Analogous to the basic data a physician uses to determine the state of health of his patients - Easy to measure and collect - The trend is often more useful than the actual data point - Can be used to focus on problem areas and develop preventative measures The nation's capability to conceive, design, develop, manufacture, test, and support missions using liquid rocket engines and solid rocket motors that are critical to its national security, economic health and growth, and future scientific needs. center dot The RPIB encompasses US government, academic, and commercial (including industry primes and their supplier base) research, development, test, evaluation, and manufacturing capabilities and facilities. center dot The RPIB includes the skilled workforce, related intellectual property, engineering and support services, and supply chain operations and management. This definition touches the five main segments of the U.S. RPIB as categorized by the USG: defense, intelligence community, civil government, academia, and commercial sector. The nation's capability to conceive, design, develop, manufacture, test, and support missions using liquid rocket engines and solid rocket motors that are critical to its national security, economic health and growth, and future scientific needs

  9. Optimization of Construction of the rocket-assisted projectile

    Directory of Open Access Journals (Sweden)

    Arkhipov Vladimir

    2017-01-01

    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.

  10. Development of 52 inches last stage blade for steam turbines

    International Nuclear Information System (INIS)

    Suzuki, Atsuhide; Hisa, Shoichi; Nagao, Shinichiro; Ogata, Hisao

    1986-01-01

    The last stage blades of steam turbines are the important component controlling the power output and performance of plants. In order to realize a unit of large capacity and high efficiency, the proper exhaust area and the last stage blades having good performance are indispensable. Toshiba Corp. has completed the development of the 52 inch last stage blades for 1500 and 1800 rpm steam turbines. The 52 inch last stage blades are the longest in the world, which have the annular exhaust area nearly 1.5 times as much as that of 41 inch blades used for 1100 MW, 1500 rpm turbines in nuclear power stations. By adopting these 52 inch blades, the large capacity nuclear power plants up to 1800 MW can be economically constructed, the rate of heat consumption of 1350 MW plants is improved by 3 ∼ 4 % as compared with 41 inch blades, and in the plants up to 1100 MW, LP turbines can be reduced from three sets to two. The features of 52 inch blades, the flow pattern and blade form design, the structural strength analysis and the erosion withstanding property, and the verification by the rotation test of the actual blades, the performance test using a test turbine, the vibration analysis of the actually loaded blades and the analysis of wet steam behavior are reported. (Kako, I.)

  11. Another Look at Rocket Thrust

    Science.gov (United States)

    Hester, Brooke; Burris, Jennifer

    2012-01-01

    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…

  12. The History of Rockets.

    Science.gov (United States)

    Newby, J. C.

    1988-01-01

    Discusses the origins and development of rockets mainly from the perspective of warfare. Includes some early enthusiasts, such as Congreve, Tsiolkovosky, Goddard, and Oberth. Describes developments from World War II, and during satellite development. (YP)

  13. Space shuttle booster separation motor design

    Science.gov (United States)

    Smith, G. W.; Chase, C. A.

    1976-01-01

    The separation characteristics of the space shuttle solid rocket boosters (SRBs) are introduced along with the system level requirements for the booster separation motors (BSMs). These system requirements are then translated into specific motor requirements that control the design of the BSM. Each motor component is discussed including its geometry, material selection, and fabrication process. Also discussed is the propellant selection, grain design, and performance capabilities of the motor. The upcoming test program to develop and qualify the motor is outlined.

  14. Evaluation of Geopolymer Concrete for Rocket Test Facility Flame Deflectors

    Science.gov (United States)

    Allgood, Daniel C.; Montes, Carlos; Islam, Rashedul; Allouche, Erez

    2014-01-01

    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.

  15. Scaled Rocket Testing in Hypersonic Flow

    Science.gov (United States)

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

    2015-01-01

    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.

  16. SAFT 4{1/2} inch nickel hydrogen battery cells

    Energy Technology Data Exchange (ETDEWEB)

    Duquesne, D.; Lacout, B.; Sennet, A. [SAFT Advanced Batteries, Poitiers (France)

    1995-12-31

    SAFT Advanced Batteries has now produced over 400 high capacity 4{1/2} inch Nickel Hydrogen Battery Cells for flight programs. The 4.5 inch diameter, rabbit-ear cell design is designed to provide the anticipated energy required at the lowest practical weight. SAFT has incorporated into the design of the dry-powder nickel electrode, truly hermetic ceramic to metal seals, qualified terminal feedthroughs, high reliability mechanical design, composite pure platinum negative electrode, and zircar separator, plus more than 25 years experience in aerospace nickel cell technology, resulting in a 4{1/2} inch configuration with the 3{1/2} inch cell design carryover heritage. General performance requirements for GEO missions that SAFT cells meet are 15 years in orbit lifetime, 80% DOD, low mass to energy ratios, and flexible capacity by modifying number of electrodes in the stack. This design is qualified for geostationary orbits based on SAFT`s 3{1/2} inch qualification heritage, design verification, and cycling performed by customer Space Systems/LORAL in support of the INTELSAT VIIA and N-STAR flight programs.

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

    Science.gov (United States)

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

    2004-08-01

    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

  18. Rocket Flight Path

    Directory of Open Access Journals (Sweden)

    Jamie Waters

    2014-09-01

    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.

  19. Cryogenic rocket engine development at Delft aerospace rocket engineering

    NARCIS (Netherlands)

    Wink, J; Hermsen, R.; Huijsman, R; Akkermans, C.; Denies, L.; Barreiro, F.; Schutte, A.; Cervone, A.; Zandbergen, B.T.C.

    2016-01-01

    This paper describes the current developments regarding cryogenic rocket engine technology at Delft Aerospace Rocket Engineering (DARE). DARE is a student society based at Delft University of Technology with the goal of being the first student group in the world to launch a rocket into space. After

  20. Scale Effects on Solid Rocket Combustion Instability Behaviour

    OpenAIRE

    David R. Greatrix

    2011-01-01

    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 combusti...

  1. The Relativistic Rocket

    Science.gov (United States)

    Antippa, Adel F.

    2009-01-01

    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…

  2. This "Is" Rocket Science!

    Science.gov (United States)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-01-01

    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…

  3. ROCKETS: Soar to Success

    Science.gov (United States)

    Brett, Christine E. W.; O'Merle, Mary Jane; White, Gene

    2017-01-01

    This article describes ROCKETS, an after-school program for at-risk youth, and how the university students became involved in this service-learning project. The article discusses the steps that were taken to start the program, what is being done to continue the program, and the challenges that faculty have faced. This program is an authentic…

  4. Liquid Rocket Engine Testing

    Science.gov (United States)

    Rahman, Shamim

    2005-01-01

    Comprehensive Liquid Rocket Engine testing is essential to risk reduction for Space Flight. Test capability represents significant national investments in expertise and infrastructure. Historical experience underpins current test capabilities. Test facilities continually seek proactive alignment with national space development goals and objectives including government and commercial sectors.

  5. Baking Soda and Vinegar Rockets

    Science.gov (United States)

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

    2009-01-01

    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…

  6. Multi-Parameter Wireless Monitoring and Telecommand of a Rocket Payload: Design and Implementation

    Science.gov (United States)

    Pamungkas, Arga C.; Putra, Alma A.; Puspitaningayu, Pradini; Fransisca, Yulia; Widodo, Arif

    2018-04-01

    A rocket system generally consists of two parts, the rocket motor and the payload. The payload system is built of several sensors such as accelerometer, gyroscope, magnetometer, and also a surveillance camera. These sensors are used to monitor the rocket in a three-dimensional axis which determine its attitude. Additionally, the payload must be able to perform image capturing in a certain distance using telecommand. This article is intended to describe the design and also the implementation of a rocket payload which has attitude monitoring and telecommand ability from the ground control station using a long-range wireless module Digi XBee Pro 900 HP.

  7. NASA Sounding Rocket Program Educational Outreach

    Science.gov (United States)

    Rosanova, G.

    2013-01-01

    Sat-C elements of the "pipeline" have been successfully demonstrated by five annual flights thus far from Wallops Flight Facility. RockSat-X has successfully flown twice, also from Wallops. The NSRP utilizes launch vehicles comprised of military surplus rocket motors (Terrier-Improved Orion and Terrier-Improved Malemute) to execute these missions. The NASA Sounding Rocket Program is proud of its role in inspiring the "next generation of explorers" and is working to expand its reach to all regions of the United States and the international community as well.

  8. Solid Rocket Testing at AFRL (Briefing Charts)

    Science.gov (United States)

    2016-10-21

    Distribution Unlimited. PA#16492 2 Agenda • Solid Rocket Motors • History of Sea Level Testing • Small Component Testing • Full-scale Testing • Altitude...Facility • History of Testing • Questions -Distribution A: Approved for Public Release; Distribution Unlimited. PA#16492 3 RQ-West • AFRL/RQ...INTEGRATION FACILITY NATIONAL HOVER TEST FACILITY TITAN SRM TEST FACILITY TS-1C1-125 LARGE ENGINE/COMPONENT TEST FACILITY TS-1A 1-120 1-115 X-33 LAUNCH

  9. Rocket-Powered Parachutes Rescue Entire Planes

    Science.gov (United States)

    2010-01-01

    Small Business Innovation Research (SBIR) contracts with Langley Research Center helped BRS Aerospace, of Saint Paul, Minnesota, to develop technology that has saved 246 lives to date. The company s whole aircraft parachute systems deploy in less than 1 second thanks to solid rocket motors and are capable of arresting the descent of a small aircraft, lowering it safely to the ground. BRS has sold more than 30,000 systems worldwide, and the technology is now standard equipment on many of the world s top-selling aircraft. Parachutes for larger airplanes are in the works.

  10. Vibration Disturbance Damping System Design to Protect Payload of the Rocket

    Directory of Open Access Journals (Sweden)

    Sutisno Sutisno

    2012-12-01

    Full Text Available Rocket motor generates vibrations acting on whole rocket body including its contents. Part of the body which is sensitive to disturbance is the rocket payload. The payload consists of various electronic instruments including: transmitter, various sensors, accelerometer, gyro, the embedded controller system, and others. This paper presents research on rocket vibration influence to the payload and the method to avoid disturbance. Avoiding influence of vibration disturbance can be done using silicone gel material whose typical damping factors are relatively high. The rocket vibration was simulated using electromagnetic motor, and the vibrations were measured using an accelerometer sensor. The measurement results were displayed in the form of curve, indicating the vibration level on some parts of the tested material. Some measurement results can be applied to determine the good material to attenuate vibration disturbance on the instruments of the payload.

  11. Development of the Astrobee F sounding rocket system.

    Science.gov (United States)

    Jenkins, R. B.; Taylor, J. P.; Honecker, H. J., Jr.

    1973-01-01

    The development of the Astrobee F sounding rocket vehicle through the first flight test at NASA-Wallops Station is described. Design and development of a 15 in. diameter, dual thrust, solid propellant motor demonstrating several new technology features provided the basis for the flight vehicle. The 'F' motor test program described demonstrated the following advanced propulsion technology: tandem dual grain configuration, low burning rate HTPB case-bonded propellant, and molded plastic nozzle. The resultant motor integrated into a flight vehicle was successfully flown with extensive diagnostic instrumentation.-

  12. The relativistic rocket

    Energy Technology Data Exchange (ETDEWEB)

    Antippa, Adel F [Departement de Physique, Universite du Quebec a Trois-Rivieres, Trois-Rivieres, Quebec G9A 5H7 (Canada)

    2009-05-15

    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 method that can be applied to a wide range of special relativistic problems of linear acceleration.

  13. National Report on the NASA Sounding Rocket and Balloon Programs

    Science.gov (United States)

    Eberspeaker, Philip; Fairbrother, Debora

    2013-01-01

    The U. S. National Aeronautics and Space Administration (NASA) Sounding Rockets and Balloon Programs conduct a total of 30 to 40 missions per year in support of the NASA scientific community and other users. The NASA Sounding Rockets Program supports the science community by integrating their experiments into the sounding rocket payloads, and providing both the rocket vehicle and launch operations services. Activities since 2011 have included two flights from Andoya Rocket Range, more than eight flights from White Sands Missile Range, approximately sixteen flights from Wallops Flight Facility, two flights from Poker Flat Research Range, and four flights from Kwajalein Atoll. Other activities included the final developmental flight of the Terrier-Improved Malemute launch vehicle, a test flight of the Talos-Terrier-Oriole launch vehicle, and a host of smaller activities to improve program support capabilities. Several operational missions have utilized the new Terrier-Malemute vehicle. The NASA Sounding Rockets Program is currently engaged in the development of a new sustainer motor known as the Peregrine. The Peregrine development effort will involve one static firing and three flight tests with a target completion data of August 2014. The NASA Balloon Program supported numerous scientific and developmental missions since its last report. The program conducted flights from the U.S., Sweden, Australia, and Antarctica utilizing standard and experimental vehicles. Of particular note are the successful test flights of the Wallops Arc Second Pointer (WASP), the successful demonstration of a medium-size Super Pressure Balloon (SPB), and most recently, three simultaneous missions aloft over Antarctica. NASA continues its successful incremental design qualification program and will support a science mission aboard WASP in late 2013 and a science mission aboard the SPB in early 2015. NASA has also embarked on an intra-agency collaboration to launch a rocket from a balloon to

  14. Space shuttle solid rocket booster water entry cavity collapse loads

    Science.gov (United States)

    Keefe, R. T.; Rawls, E. A.; Kross, D. A.

    1982-01-01

    Solid rocket booster cavity collapse flight measurements included external pressures on the motor case and aft skirt, internal motor case pressures, accelerometers located in the forward skirt, mid-body area, and aft skirt, as well as strain gages located on the skin of the motor case. This flight data yielded applied pressure longitudinal and circumferential distributions which compare well with model test predictions. The internal motor case ullage pressure, which is below atmospheric due to the rapid cooling of the hot internal gas, was more severe (lower) than anticipated due to the ullage gas being hotter than predicted. The structural dynamic response characteristics were as expected. Structural ring and wall damage are detailed and are considered to be attributable to the direct application of cavity collapse pressure combined with the structurally destabilizing, low internal motor case pressure.

  15. Liquid Rocket Engine Testing Overview

    Science.gov (United States)

    Rahman, Shamim

    2005-01-01

    Contents include the following: Objectives and motivation for testing. Technology, Research and Development Test and Evaluation (RDT&E), evolutionary. Representative Liquid Rocket Engine (LRE) test compaigns. Apollo, shuttle, Expandable Launch Vehicles (ELV) propulsion. Overview of test facilities for liquid rocket engines. Boost, upper stage (sea-level and altitude). Statistics (historical) of Liquid Rocket Engine Testing. LOX/LH, LOX/RP, other development. Test project enablers: engineering tools, operations, processes, infrastructure.

  16. Rocket + Science = Dialogue

    Science.gov (United States)

    Morris,Bruce; Sullivan, Greg; Burkey, Martin

    2010-01-01

    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.

  17. Rocket Assembly and Checkout Facility

    Data.gov (United States)

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

  18. Nuclear rocket propulsion

    International Nuclear Information System (INIS)

    Clark, J.S.; Miller, T.J.

    1991-01-01

    NASA has initiated planning for a technology development project for nuclear rocket propulsion systems for Space Exploration Initiative (SEI) human and robotic missions to the Moon and to Mars. An Interagency project is underway that includes the Department of Energy National Laboratories for nuclear technology development. This paper summarizes the activities of the project planning team in FY 1990 and FY 1991, discusses the progress to date, and reviews the project plan. Critical technology issues have been identified and include: nuclear fuel temperature, life, and reliability; nuclear system ground test; safety; autonomous system operation and health monitoring; minimum mass and high specific impulse

  19. Two-Dimensional Motions of Rockets

    Science.gov (United States)

    Kang, Yoonhwan; Bae, Saebyok

    2007-01-01

    We analyse the two-dimensional motions of the rockets for various types of rocket thrusts, the air friction and the gravitation by using a suitable representation of the rocket equation and the numerical calculation. The slope shapes of the rocket trajectories are discussed for the three types of rocket engines. Unlike the projectile motions, the…

  20. Mean Flow Augmented Acoustics in Rocket Systems

    Science.gov (United States)

    Fischbach, Sean R.

    2014-01-01

    Oscillatory motion in solid rocket motors and liquid engines has long been a subject of concern. Many rockets display violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process and gas dynamics. The customary approach to modeling acoustic waves inside a rocket chamber is to apply the classical inhomogeneous wave equation to the combustion gas. The assumption of a linear, non-dissipative wave in a quiescent fluid remains valid while the acoustic amplitudes are small and local gas velocities stay below Mach 0.2. The converging section of a rocket nozzle, where gradients in pressure, density, and velocity become large, is a notable region where this approach is not applicable. The expulsion of unsteady energy through the nozzle of a rocket is identified as the predominate source of acoustic damping for most rocket systems. An accurate model of the acoustic behavior within this region where acoustic modes are influenced by the presence of a steady mean flow is required for reliable stability predictions. Recently, an approach to address nozzle damping with mean flow effects was implemented by French [1]. This new approach extends the work originated by Sigman and Zinn [2] by solving the acoustic velocity potential equation (AVPE) formulated by perturbing the Euler equations [3]. The acoustic velocity potential (psi) describing the acoustic wave motion in the presence of an inhomogeneous steady high-speed flow is defined by, (del squared)(psi) - (lambda/c)(exp 2)(psi) - M(dot)[M(dot)(del)(del(psi))] - 2(lambda(M/c) + (M(dot)del(M))(dot)del(psi)-2(lambda)(psi)[M(dot)del(1/c)]=0 (1) with M as the Mach vector, c as the speed of sound, and lambda as the complex eigenvalue. French apply the finite volume method to solve the steady flow field within the combustion chamber and nozzle with inviscid walls. The complex eigenvalues and eigenvector are determined with the use of the ARPACK eigensolver. The

  1. Pressure scaled water impact test of a 12.5 inch diameter model of the Space Shuttle solid rocket booster

    Science.gov (United States)

    1982-01-01

    A total of 59 tail first drops were made. Model entry conditions simulated full scale vertical velocities of approximately 75 to 110 ft/sec with horizontal velocities up to 45 ft/sec and impact angles to + or - 10 deg. These tests were conducted at scaled atmospheric pressures (1.26 psia or 65 mm.Hg). The model, test program, test facility, test equipment, instrumentation system, data reduction procedures, and test results are described.

  2. NO PLIF imaging in the CUBRC 48-inch shock tunnel

    Science.gov (United States)

    Jiang, N.; Bruzzese, J.; Patton, R.; Sutton, J.; Yentsch, R.; Gaitonde, D. V.; Lempert, W. R.; Miller, J. D.; Meyer, T. R.; Parker, R.; Wadham, T.; Holden, M.; Danehy, P. M.

    2012-12-01

    Nitric oxide planar laser-induced fluorescence (NO PLIF) imaging is demonstrated at a 10-kHz repetition rate in the Calspan University at Buffalo Research Center's (CUBRC) 48-inch Mach 9 hypervelocity shock tunnel using a pulse burst laser-based high frame rate imaging system. Sequences of up to ten images are obtained internal to a supersonic combustor model, located within the shock tunnel, during a single ~10-millisecond duration run of the ground test facility. Comparison with a CFD simulation shows good overall qualitative agreement in the jet penetration and spreading observed with an average of forty individual PLIF images obtained during several facility runs.

  3. The 52-inch last-stage blades for steam turbines

    International Nuclear Information System (INIS)

    Suzuki, Atsuhide; Hisa, Shoichi; Nagao, Shin-ichiro; Ogata, Hisao

    1986-01-01

    The last-stage blades (LSB) of steam turbines are one of the most important components determining the plant's maximum capacity and efficiency. The development of LSBs necessitates high-technology including advanced methods of analyses and verifications as well as ample accumulation of technical data. The 52-inch LSB recently developed by Toshiba has raised nuclear power plant's capacity up to 1,300 ∼ 1,800 MW, has effected compact design of turbine units, and has improved thermal efficiency, keeping high reliability. (author)

  4. Cryogenic vacuum pumping at the LBL 88-inch cyclotron

    International Nuclear Information System (INIS)

    Elo, D.; Morris, D.; Clark, D.J.; Gough, R.A.

    1978-09-01

    A cryogenic vacuum pumping panel has been in operation at the 88-inch cyclotron since 1974. The nude pumping panel is located in the acceleration chamber. The pumping surface consists of tubing cooled to 20 0 K by a closed loop helium refrigeration system. The pumping surfaces are shielded from radiation heat loads and water vapors by liquid nitrogen cooled baffles. The panel was designed for an average pumping speed of 14,000 liters/sec. for air. This approximately tripled the total effective pumping on the acceleration chamber from the existing diffusion pumped system, significantly reducing charge exchange losses of heavy ions during acceleration. Design, installation and performance characteristics are described

  5. Metallurgical Evaluation of the Five-Inch Cylindrical Induction Melter

    International Nuclear Information System (INIS)

    Imrich, K.J.

    2000-01-01

    A metallurgical evaluation of the 5-inch cylindrical induction melter (CIM) vessel was performed by the Materials Technology Section to evaluate the metallurgical condition after operating for approximately 375 hours at 1400 to 1500 Degrees Celsius during a 2 year period. Results indicate that wall thinning and significant grain growth occurred in the lower portion of the conical section and the drain tube. No through-wall penetrations were found in the cylindrical and conical sections of the CIM vessel and only one leak site was identified in the drain tube. Failure of the drain tube was associated with a localized over heating and intercrystalline fracture

  6. Rhenium Rocket Manufacturing Technology

    Science.gov (United States)

    1997-01-01

    The NASA Lewis Research Center's On-Board Propulsion Branch has a research and technology program to develop high-temperature (2200 C), iridium-coated rhenium rocket chamber materials for radiation-cooled rockets in satellite propulsion systems. Although successful material demonstrations have gained much industry interest, acceptance of the technology has been hindered by a lack of demonstrated joining technologies and a sparse materials property data base. To alleviate these concerns, we fabricated rhenium to C-103 alloy joints by three methods: explosive bonding, diffusion bonding, and brazing. The joints were tested by simulating their incorporation into a structure by welding and by simulating high-temperature operation. Test results show that the shear strength of the joints degrades with welding and elevated temperature operation but that it is adequate for the application. Rhenium is known to form brittle intermetallics with a number of elements, and this phenomena is suspected to cause the strength degradation. Further bonding tests with a tantalum diffusion barrier between the rhenium and C-103 is planned to prevent the formation of brittle intermetallics.

  7. Micro-Rockets for the Classroom.

    Science.gov (United States)

    Huebner, Jay S.; Fletcher, Alice S.; Cato, Julia A.; Barrett, Jennifer A.

    1999-01-01

    Compares micro-rockets to commercial models and water rockets. Finds that micro-rockets are more advantageous because they are constructed with inexpensive and readily available materials and can be safely launched indoors. (CCM)

  8. Effect of Six Missile-Bay Baffle Configurations and a Rocket End Plate on Ejection Releases of an MB-1 Rocket from a 0.05 Scale Model of the Convair F-106A Airplane

    Science.gov (United States)

    Hinson, William F.; Lee, John B.

    1959-01-01

    As a continuation of an investigation of the release characteristics of an MB-1 rocket carried internally by the Convair F-106A airplane, six missile-bay baffle configurations and a rocket end plate have been investigated in the 27- by 27-inch preflight jet of the NASA Wallops Station. The MB-1 rocket used had retractable fins and was ejected from a missile bay modified by the addition of six different baffle configurations. For some tests a rocket end plate was added to the model. Dynamically scaled models (0.04956 scale) were tested at a simulated altitude of 22,450 feet and Mach numbers of 0.86, 1.59, and 1.98, and at a simulated altitude of 29,450 feet and a Mach number of 1.98. The results of this investigation indicate that the missile-bay baffle configurations and the rocket end plate may be used to reduce the positive pitch amplitude of the MB-1 rocket after release. The initial negative pitching velocity applied to the MB-1 rocket might then be reduced in order to maintain a near-level-flight attitude after release. As the fuselage angle of attack is increased, the negative pitch amplitude of the rocket is decreased.

  9. Combustion of metal agglomerates in a solid rocket core flow

    Science.gov (United States)

    Maggi, Filippo; Dossi, Stefano; DeLuca, Luigi T.

    2013-12-01

    The need for access to space may require the use of solid propellants. High thrust and density are appealing features for different applications, spanning from boosting phase to other service applications (separation, de-orbiting, orbit insertion). Aluminum is widely used as a fuel in composite solid rocket motors because metal oxidation increases enthalpy release in combustion chamber and grants higher specific impulse. Combustion process of metal particles is complex and involves aggregation, agglomeration and evolution of reacting particulate inside the core flow of the rocket. It is always stated that residence time should be enough in order to grant complete metal oxidation but agglomerate initial size, rocket grain geometry, burning rate, and other factors have to be reconsidered. New space missions may not require large rocket systems and metal combustion efficiency becomes potentially a key issue to understand whether solid propulsion embodies a viable solution or liquid/hybrid systems are better. A simple model for metal combustion is set up in this paper. Metal particles are represented as single drops trailed by the core flow and reacted according to Beckstead's model. The fluid dynamics is inviscid, incompressible, 1D. The paper presents parametric computations on ideal single-size particles as well as on experimental agglomerate populations as a function of operating rocket conditions and geometries.

  10. Nuclear Rocket Engine Reactor

    CERN Document Server

    Lanin, Anatoly

    2013-01-01

    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. Scale effects on solid rocket combustion instability behaviour

    Energy Technology Data Exchange (ETDEWEB)

    Greatrix, D. R. [Ryerson University, Department of Aerospace Engineering, Toronto, Ontario (Canada)

    2011-07-01

    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. (author)

  12. Leak in spiral weld in a 16 inches gas pipeline

    Energy Technology Data Exchange (ETDEWEB)

    Fazzini, Pablo G; Bona, Jeremias de [GIE S.A., Mar del Plata (Argentina); Otegui, Jose L [University of Mar del Plata (Argentina)

    2009-07-01

    This paper discusses a failure analysis after a leak in the spiral weld of a 16 inches natural gas pipeline, in service since 1974. The leak was the result of the coalescence of two different defects, on each surface of the pipe wall, located in the center of the inner cord of the helical DSAW weld. Fractographic and metallographic studies revealed that the leak was a combination of three conditions. During fabrication of the pipe, segregation in grain boundary grouped in mid weld. During service, these segregations underwent a process of selective galvanic corrosion. One of these volumetric defects coincided with a tubular pore in the outer weld. Pigging of the pipeline in 2005 for cleaning likely contributed to the increase of the leak flow, when eliminating corrosion product plugs. Although these defects are likely to repeat, fracture mechanics shows that a defect of this type is unlikely to cause a blowout. (author)

  13. NO PLIF imaging in the CUBRC 48-inch shock tunnel

    Energy Technology Data Exchange (ETDEWEB)

    Jiang, N.; Bruzzese, J.; Patton, R.; Sutton, J.; Yentsch, R.; Gaitonde, D.V.; Lempert, W.R. [The Ohio State University, Departments of Mechanical and Aerospace Engineering, Columbus, OH (United States); Miller, J.D.; Meyer, T.R. [Iowa State University, Department of Mechanical Engineering, Ames, IA (United States); Parker, R.; Wadham, T.; Holden, M. [CUBRC, Buffalo, NY (United States); Danehy, P.M. [NASA Langley Research Center, Hampton, VA (United States)

    2012-12-15

    Nitric oxide planar laser-induced fluorescence (NO PLIF) imaging is demonstrated at a 10-kHz repetition rate in the Calspan University at Buffalo Research Center's (CUBRC) 48-inch Mach 9 hypervelocity shock tunnel using a pulse burst laser-based high frame rate imaging system. Sequences of up to ten images are obtained internal to a supersonic combustor model, located within the shock tunnel, during a single {proportional_to}10-millisecond duration run of the ground test facility. Comparison with a CFD simulation shows good overall qualitative agreement in the jet penetration and spreading observed with an average of forty individual PLIF images obtained during several facility runs. (orig.)

  14. NO PLIF Imaging in the CUBRC 48 Inch Shock Tunnel

    Science.gov (United States)

    Jiang, N.; Bruzzese, J.; Patton, R.; Sutton J.; Lempert W.; Miller, J. D.; Meyer, T. R.; Parker, R.; Wadham, T.; Holden, M.; hide

    2011-01-01

    Nitric Oxide Planar Laser-Induced Fluorescence (NO PLIF) imaging is demonstrated at a 10 kHz repetition rate in the Calspan-University at Buffalo Research Center s (CUBRC) 48-inch Mach 9 hypervelocity shock tunnel using a pulse burst laser-based high frame rate imaging system. Sequences of up to ten images are obtained internal to a supersonic combustor model, located within the shock tunnel, during a single approx.10-millisecond duration run of the ground test facility. This represents over an order of magnitude improvement in data rate from previous PLIF-based diagnostic approaches. Comparison with a preliminary CFD simulation shows good overall qualitative agreement between the prediction of the mean NO density field and the observed PLIF image intensity, averaged over forty individual images obtained during several facility runs.

  15. Development of a 10-inch HPD with integrated readout electronics

    CERN Document Server

    Braem, André; Joram, C; Séguinot, Jacques; Weilhammer, Peter; Giunta, M; Malakhov, N; Menzione, A; Pegna, R; Piccioli, A; Raffaelli, F; Sartori, G

    2003-01-01

    A round 10-in. diameter Hybrid Photodiode (HPD) with spherical entrance window is under development for Cherenkov imaging applications in cosmic ray astronomy. The HPD adopts the fountain focusing electron optics, which, as already demonstrated in the 5 inch Pad HPD, allows for a linear demagnification of the image over practically the full tube diameter. Self-triggering front-end electronics providing also sparse readout capability, has been tested. High-efficiency Rb//2Te cathodes have been produced on a UV extended borosilicate glass windows with very thin conductive underlayers of Indium Tin Oxide. We report on the design of the 10- in. HPD, the fabrication procedure and first tests of a 5-in. HPD with Rb//2Te photocathode and 2048 channels.

  16. Radiation effects testing at the 88-Inch Cyclotron at LBNL

    International Nuclear Information System (INIS)

    McMahan, Margaret A.; Koga, Rokotura

    2002-01-01

    The effects of ionizing particles on sensitive microelectronics is an important component of the design of systems as diverse as satellites and space probes, detectors for high energy physics experiments and even internet server farms. Understanding the effects of radiation on human cells is an equally important endeavor directed towards future manned missions in space and towards cancer therapy. At the 88-Inch Cyclotron at the Berkeley Laboratory, facilities are available for radiation effects testing (RET) with heavy ions and with protons. The techniques for doing these measurements and the advantages of using a cyclotron will be discussed, and the Cyclotron facilities will be compared with other facilities worldwide. RET of the same part at several facilities of varying beam energy can provide tests of the simple models used in this field and elucidate the relative importance of atomic and nuclear effects. The results and implications of such measurements will be discussed

  17. Remote target removal for the Oak Ridge 86-inch Cyclotron

    International Nuclear Information System (INIS)

    Walls, A.A.

    1982-01-01

    A remotely operated target remover has been plaed in operation at the 86-Inch Cyclotron located in Oak Ridge. The system provides for the remote removal of a target from inside the cyclotron, loading it into a cask, and the removal of the cask from the 1.5 m (5-ft) shielding walls. The remote system consists of multiple electrical and pneumatically operated equipment which is designed for controlled step-by-step operation, operated with an electrical control panel, and monitored by a television system. The target remover has reduced the radiation exposures to operating personnel at the facility and has increased the effective operating time. The system is fast, requires a minimum of skill to operate, and has demonstrated both reliability and durability

  18. Seismic fragility test of a 6-inch diameter pipe system

    International Nuclear Information System (INIS)

    Chen, W.P.; Onesto, A.T.; DeVita, V.

    1987-02-01

    This report contains the test results and assessments of seismic fragility tests performed on a 6-inch diameter piping system. The test was funded by the US Nuclear Regulatory Commission (NRC) and conducted by ETEC. The objective of the test was to investigate the ability of a representative nuclear piping system to withstand high level dynamic seismic and other loadings. Levels of loadings achieved during seismic testing were 20 to 30 times larger than normal elastic design evaluations to ASME Level D limits would permit. Based on failure data obtained during seismic and other dynamic testing, it was concluded that nuclear piping systems are inherently able to withstand much larger dynamic seismic loadings than permitted by current design practice criteria or predicted by the probabilistic risk assessment (PRA) methods and several proposed nonlinear methods of failure analysis

  19. Electromechanical Dynamics Simulations of Superconducting LSM Rocket Launcher System in Attractive-Mode

    Science.gov (United States)

    Yoshida, Kinjiro; Hayashi, Kengo; Takami, Hiroshi

    1996-01-01

    Further feasibility study on a superconducting linear synchronous motor (LSM) rocket launcher system is presented on the basis of dynamic simulations of electric power, efficiency and power factor as well as the ascending motions of the launcher and rocket. The advantages of attractive-mode operation are found from comparison with repulsive-mode operation. It is made clear that the LSM rocket launcher system, of which the long-stator is divided optimally into 60 sections according to launcher speeds, can obtain high efficiency and power factor.

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

    Science.gov (United States)

    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.

  1. Two-Rockets Thought Experiment

    Science.gov (United States)

    Smarandache, Florentin

    2014-03-01

    Let n>=2 be identical rockets: R1 ,R2 , ..., Rn. Each of them moving at constant different velocities respectively v1, v2, ..., vn on parallel directions in the same sense. In each rocket there is a light clock, the observer on earth also has a light clock. All n + 1 light clocks are identical and synchronized. The proper time Δt' in each rocket is the same. Let's focus on two arbitrary rockets Ri and Rjfrom the previous n rockets. Let's suppose, without loss of generality, that their speeds verify virocket Rj is contracted with the factor C(vj -vi) , i.e. Lj =Lj' C(vj -vi) .(2) But in the reference frame of the astronaut in Rjit is like rocket Rjis stationary andRi moves with the speed vj -vi in opposite direction. Therefore, similarly, the non-proper time interval as measured by the astronaut inRj with respect to the event inRi is dilated with the same factor D(vj -vi) , i.e. Δtj . i = Δt' D(vj -vi) , and rocketRi is contracted with the factor C(vj -vi) , i.e. Li =Li' C(vj -vi) .But it is a contradiction to have time dilations in both rockets. (3) Varying i, j in {1, 2, ..., n} in this Thought Experiment we get again other multiple contradictions about time dilations. Similarly about length contractions, because we get for a rocket Rj, n-2 different length contraction factors: C(vj -v1) , C(vj -v2) , ..., C(vj -vj - 1) , C(vj -vj + 1) , ..., C(vj -vn) simultaneously! Which is abnormal.

  2. The impact of wire caliber on ERCP outcomes: a multicenter randomized controlled trial of 0.025-inch and 0.035-inch guidewires.

    Science.gov (United States)

    Bassan, Milan S; Sundaralingam, Praka; Fanning, Scott B; Lau, James; Menon, Jayaram; Ong, Evan; Rerknimitr, Rungsun; Seo, Dong-Wan; Teo, Eng Kiong; Wang, Hsiu-Po; Reddy, D Nageshwar; Goh, Khean Lee; Bourke, Michael J

    2018-06-01

    Wire-guided biliary cannulation has been demonstrated to improve cannulation rates and reduce post-ERCP pancreatitis (PEP), but the impact of wire caliber has not been studied. This study compares successful cannulation rates and ERCP adverse events by using a 0.025-inch and 0.035-inch guidewire. A randomized, single blinded, prospective, multicenter trial at 9 high-volume tertiary-care referral centers in the Asia-Pacific region was performed. Patients with an intact papilla and conventional anatomy who did not have malignancy in the head of the pancreas or ampulla and were undergoing ERCP were recruited. ERCP was performed by using a standardized cannulation algorithm, and patients were randomized to either a 0.025-inch or 0.035-inch guidewire. The primary outcomes of the study were successful wire-guided cannulation and the incidence of PEP. Overall successful cannulation and ERCP adverse events also were studied. A total of 710 patients were enrolled in the study. The primary wire-guided biliary cannulation rate was similar in 0.025-inch and 0.035-inch wire groups (80.7% vs 80.3%; P = .90). The rate of PEP between the 0.025-inch and the 0.035-inch wire groups did not differ significantly (7.8% vs 9.3%; P = .51). No differences were noted in secondary outcomes. Similar rates of successful cannulation and PEP were demonstrated in the use of 0.025-inch and 0.035-inch guidewires. (Clinical trial registration number: NCT01408264.). Copyright © 2018. Published by Elsevier Inc.

  3. The Chameleon Solid Rocket Propulsion Model

    International Nuclear Information System (INIS)

    Robertson, Glen A.

    2010-01-01

    The Khoury and Weltman (2004a and 2004b) Chameleon Model presents an addition to the gravitation force and was shown by the author (Robertson, 2009a and 2009b) to present a new means by which one can view other forces in the Universe. The Chameleon Model is basically a density-dependent model and while the idea is not new, this model is novel in that densities in the Universe to include the vacuum of space are viewed as scalar fields. Such an analogy gives the Chameleon scalar field, dark energy/dark matter like characteristics; fitting well within cosmological expansion theories. In respect to this forum, in this paper, it is shown how the Chameleon Model can be used to derive the thrust of a solid rocket motor. This presents a first step toward the development of new propulsion models using density variations verse mass ejection as the mechanism for thrust. Further, through the Chameleon Model connection, these new propulsion models can be tied to dark energy/dark matter toward new space propulsion systems utilizing the vacuum scalar field in a way understandable by engineers, the key toward the development of such systems. This paper provides corrections to the Chameleon rocket model in Robertson (2009b).

  4. The Swedish sounding rocket programme

    International Nuclear Information System (INIS)

    Bostroem, R.

    1980-01-01

    Within the Swedish Sounding Rocket Program the scientific groups perform experimental studies of magnetospheric and ionospheric physics, upper atmosphere physics, astrophysics, and material sciences in zero g. New projects are planned for studies of auroral electrodynamics using high altitude rockets, investigations of noctilucent clouds, and active release experiments. These will require increased technical capabilities with respect to payload design, rocket performance and ground support as compared with the current program. Coordination with EISCAT and the planned Viking satellite is essential for the future projects. (Auth.)

  5. Modal Survey of ETM-3, A 5-Segment Derivative of the Space Shuttle Solid Rocket Booster

    Science.gov (United States)

    Nielsen, D.; Townsend, J.; Kappus, K.; Driskill, T.; Torres, I.; Parks, R.

    2005-01-01

    The complex interactions between internal motor generated pressure oscillations and motor structural vibration modes associated with the static test configuration of a Reusable Solid Rocket Motor have potential to generate significant dynamic thrust loads in the 5-segment configuration (Engineering Test Motor 3). Finite element model load predictions for worst-case conditions were generated based on extrapolation of a previously correlated 4-segment motor model. A modal survey was performed on the largest rocket motor to date, Engineering Test Motor #3 (ETM-3), to provide data for finite element model correlation and validation of model generated design loads. The modal survey preparation included pretest analyses to determine an efficient analysis set selection using the Effective Independence Method and test simulations to assure critical test stand component loads did not exceed design limits. Historical Reusable Solid Rocket Motor modal testing, ETM-3 test analysis model development and pre-test loads analyses, as well as test execution, and a comparison of results to pre-test predictions are discussed.

  6. Scale effects on quasi-steady solid rocket internal ballistic behaviour

    Energy Technology Data Exchange (ETDEWEB)

    Greatrix, D. R. [Department of Aerospace Engineering, Ryerson University, 350 Victoria Street, Toronto, Ontario, M5B2K3 (Canada)

    2010-11-15

    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. (author)

  7. Theodore von Karman - Rocket Scientist

    Indian Academy of Sciences (India)

    seminal contributions to several areas of fluid and solid mechanics, as the first head of ... nent position in Aeronautics research, as a pioneer of rocket science in America ... toral work, however, was on the theory of buckling of large structures.

  8. Sounding rockets explore the ionosphere

    International Nuclear Information System (INIS)

    Mendillo, M.

    1990-01-01

    It is suggested that small, expendable, solid-fuel rockets used to explore ionospheric plasma can offer insight into all the processes and complexities common to space plasma. NASA's sounding rocket program for ionospheric research focuses on the flight of instruments to measure parameters governing the natural state of the ionosphere. Parameters include input functions, such as photons, particles, and composition of the neutral atmosphere; resultant structures, such as electron and ion densities, temperatures and drifts; and emerging signals such as photons and electric and magnetic fields. Systematic study of the aurora is also conducted by these rockets, allowing sampling at relatively high spatial and temporal rates as well as investigation of parameters, such as energetic particle fluxes, not accessible to ground based systems. Recent active experiments in the ionosphere are discussed, and future sounding rocket missions are cited

  9. EUVS Sounding Rocket Payload

    Science.gov (United States)

    Stern, Alan S.

    1996-01-01

    During the first half of this year (CY 1996), the EUVS project began preparations of the EUVS payload for the upcoming NASA sounding rocket flight 36.148CL, slated for launch on July 26, 1996 to observe and record a high-resolution (approx. 2 A FWHM) EUV spectrum of the planet Venus. These preparations were designed to improve the spectral resolution and sensitivity performance of the EUVS payload as well as prepare the payload for this upcoming mission. The following is a list of the EUVS project activities that have taken place since the beginning of this CY: (1) Applied a fresh, new SiC optical coating to our existing 2400 groove/mm grating to boost its reflectivity; (2) modified the Ranicon science detector to boost its detective quantum efficiency with the addition of a repeller grid; (3) constructed a new entrance slit plane to achieve 2 A FWHM spectral resolution; (4) prepared and held the Payload Initiation Conference (PIC) with the assigned NASA support team from Wallops Island for the upcoming 36.148CL flight (PIC held on March 8, 1996; see Attachment A); (5) began wavelength calibration activities of EUVS in the laboratory; (6) made arrangements for travel to WSMR to begin integration activities in preparation for the July 1996 launch; (7) paper detailing our previous EUVS Venus mission (NASA flight 36.117CL) published in Icarus (see Attachment B); and (8) continued data analysis of the previous EUVS mission 36.137CL (Spica occultation flight).

  10. Not just rocket science

    Energy Technology Data Exchange (ETDEWEB)

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

    2007-10-15

    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.

  11. Nuclear rocket engine reactor

    Energy Technology Data Exchange (ETDEWEB)

    Lanin, Anatoly

    2013-07-01

    Covers a new technology of nuclear reactors and the related materials aspects. Integrates physics, materials science and engineering Serves as a basic book for nuclear engineers and nuclear physicists. 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.

  12. Easier Analysis With Rocket Science

    Science.gov (United States)

    2003-01-01

    Analyzing rocket engines is one of Marshall Space Flight Center's specialties. When Marshall engineers lacked a software program flexible enough to meet their needs for analyzing rocket engine fluid flow, they overcame the challenge by inventing the Generalized Fluid System Simulation Program (GFSSP), which was named the co-winner of the NASA Software of the Year award in 2001. This paper describes the GFSSP in a wide variety of applications

  13. The development of 8 inch roll-to-plate nanoimprint lithography (8-R2P-NIL) system

    Science.gov (United States)

    Lee, Lai Seng; Mohamed, Khairudin; Ooi, Su Guan

    2017-07-01

    Growth in semiconductor and integrated circuit industry was observed in the past decennium of years for industrial technology which followed Moore's law. The line width of nanostructure to be exposed was influenced by the essential technology of photolithography. Thus, it is crucial to have a low cost and high throughput manufacturing process for nanostructures. Nanoimprint Lithography technique invented by Stephen Y. Chou was considered as major nanolithography process to be used in future integrated circuit and integrated optics. The drawbacks of high imprint pressure, high imprint temperature, air bubbles formation, resist sticking to mold and low throughput of thermal nanoimprint lithography on silicon wafer have yet to be solved. Thus, the objectives of this work is to develop a high throughput, low imprint force, room temperature UV assisted 8 inch roll to plate nanoimprint lithography system capable of imprinting nanostructures on 200 mm silicon wafer using roller imprint with flexible mold. A piece of resist spin coated silicon wafer was placed onto vacuum chuck drives forward by a stepper motor. A quartz roller wrapped with a piece of transparent flexible mold was used as imprint roller. The imprinted nanostructures were cured by 10 W, 365 nm UV LED which situated inside the quartz roller. Heat generated by UV LED was dissipated by micro heat pipe. The flexible mold detaches from imprinted nanostructures in a 'line peeling' pattern and imprint pressure was measured by ultra-thin force sensors. This system has imprinting speed capability ranging from 0.19 mm/s to 5.65 mm/s, equivalent to imprinting capability of 3 to 20 pieces of 8 inch wafers per hour. Speed synchronization between imprint roller and vacuum chuck was achieved by controlling pulse rate supplied to stepper motor which drive the vacuum chuck. The speed different ranging from 2 nm/s to 98 nm/s is achievable. Vacuum chuck height was controlled by stepper motor with displacement of 5 nm/step.

  14. SAFE testing nuclear rockets economically

    International Nuclear Information System (INIS)

    Howe, Steven D.; Travis, Bryan; Zerkle, David K.

    2003-01-01

    Several studies over the past few decades have recognized the need for advanced propulsion to explore the solar system. As early as the 1960s, Werner Von Braun and others recognized the need for a nuclear rocket for sending humans to Mars. The great distances, the intense radiation levels, and the physiological response to zero-gravity all supported the concept of using a nuclear rocket to decrease mission time. These same needs have been recognized in later studies, especially in the Space Exploration Initiative in 1989. One of the key questions that has arisen in later studies, however, is the ability to test a nuclear rocket engine in the current societal environment. Unlike the Rover/NERVA programs in the 1960s, the rocket exhaust can no longer be vented to the open atmosphere. As a consequence, previous studies have examined the feasibility of building a large-scale version of the Nuclear Furnace Scrubber that was demonstrated in 1971. We have investigated an alternative that would deposit the rocket exhaust along with any entrained fission products directly into the ground. The Subsurface Active Filtering of Exhaust, or SAFE, concept would allow variable sized engines to be tested for long times at a modest expense. A system overview, results of preliminary calculations, and cost estimates of proof of concept demonstrations are presented. The results indicate that a nuclear rocket could be tested at the Nevada Test Site for under $20 M

  15. Flight performance summary for three NASA Terrier-Malemute II sounding rockets

    Science.gov (United States)

    Patterson, R. A.

    1982-01-01

    The subject of this paper is the presentation of flight data for three Terrier-Malemute II sounding rocket vehicles. The Malemute motor was modified by adding insulation and using a propellant that produced less Al2O3 agglomerate in the chamber. This modification, designated Malemute II, reduced the sensitivity of the motor to the roll rate induced motor case burnthrough experienced on some earlier Malemute flights. Two flight tests, including a single stage Malemute II and a Terrier-Malemute II, were made by Sandia to qualify this modification. The three NASA operational flights that are the subject of this paper were made using the modified Malemute II motors.

  16. Scale Effects on Solid Rocket Combustion Instability Behaviour

    Directory of Open Access Journals (Sweden)

    David R. Greatrix

    2011-01-01

    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.

  17. 3000 Horsepower super conductive field acyclic motor

    International Nuclear Information System (INIS)

    Marshall, R.

    1983-01-01

    A 3000 hp acyclic motor was assembled and tested utilizing superconducting field coils. The magnet assembly is designed as a quadrupole magnet, utilizing a multifilamentary niobium titanium superconductor. Each magnet coil is 18 inches in diameter and 10 inches long, and operates at rated current of 200 amperes, providing 5.8 tesla in the bore of the coils in the motor configuration. The average winding current density is 10,600 A/cm 2 . The acyclic motor is of a drum-type design with liquid metal current collectors, and is designed to model full-scale machinery for ship propulsion applications. Laboratory test data verified the electrical and electromagnetic design to be within three percent of the calculated values

  18. Yes--This is Rocket Science: MMCs for Liquid Rocket Engines

    National Research Council Canada - National Science Library

    Shelley, J

    2001-01-01

    The Air Force's Integrated High-Payoff Rocket Propulsion Technologies (IHPRPT) Program has established aggressive goals for both improved performance and reduced cost of rocket engines and components...

  19. Wake effect in rocket observation

    International Nuclear Information System (INIS)

    Matsumoto, Haruya; Kaya, Nobuyuki; Yamanaka, Akira; Hayashi, Tomomasa

    1975-01-01

    The mechanism of the wake phenomena due to a probe and in rocket observation is discussed on the basis of experimental data. In the low energy electron measurement performed with the L-3H-5 rocket, the electron count rate changed synchronously with the rocket spin. This seems to be a wake effect. It is also conceivable that the probe itself generates the wake of ion beam. The latter problem is considered in the first part. Experiment was performed with laboratory plasma, in which a portion of the electron component of the probe current was counted with a CEM (a channel type multiplier). The change of probe voltage-count rate charactersitics due to the change of relative position of the ion source was observed. From the measured angular distributions of electron density and electron temperature around the probe, it is concluded that anisotropy exists around the probe, which seems to be a kinds of wake structure. In the second part, the wake effect due to a rocket is discussed on the basis of the measurement of leaking electrons with L-3H-5 rocket. Comparison between the theory of wake formation and the measured results is also shortly made in the final part. (Aoki, K.)

  20. Multi-Rocket Thought Experiment

    Science.gov (United States)

    Smarandache, Florentin

    2014-03-01

    We consider n>=2 identical rockets: R1 ,R2 , ..., Rn. Each of them moving at constant different velocities respectively v1 ,v2 , ..., vn on parallel directions in the same sense. In each rocket there is a light clock, the observer on earth also has a light clock. All n + 1 light clocks are identical and synchronized. The proper time Δt' in each rocket is the same. (1) If we consider the observer on earth and the first rocket R1, then the non-proper time Δt of the observer on earth is dilated with the factor D(v1) : or Δt = Δt' D(v1) (1) But if we consider the observer on earth and the second rocket R2 , then the non-proper time Δt of the observer on earth is dilated with a different factor D(v2) : or Δt = Δt' D(v2) And so on. Therefore simultaneously Δt is dilated with different factors D(v1) , D(v2), ..., D(vn) , which is a multiple contradiction.

  1. Advanced Ultra-High Speed Motor for Drilling

    Energy Technology Data Exchange (ETDEWEB)

    Impact Technologies LLC; University of Texas at Arlington

    2007-03-31

    Three (3) designs have been made for two sizes, 6.91 cm (2.72 inch) and 4.29 cm (1.69 inch) outer diameters, of a patented inverted configured Permanent Magnet Synchronous Machines (PMSM) electric motor specifically for drilling at ultra-high rotational speeds (10,000 rpm) and that can utilize advanced drilling methods. Benefits of these motors are stackable power sections, full control (speed and direction) of downhole motors, flow hydraulics independent of motor operation, application of advanced drilling methods (water jetting and abrasive slurry jetting), and the ability of signal/power electric wires through motor(s). Key features of the final designed motors are: fixed non-rotating shaft with stator coils attached; rotating housing with permanent magnet (PM) rotor attached; bit attached to rotating housing; internal channel(s) in a nonrotating shaft; electric components that are hydrostatically isolated from high internal pressure circulating fluids ('muds') by static metal to metal seals; liquid filled motor with smoothed features for minimized turbulence in the motor during operation; and new inverted coated metal-metal hydrodynamic bearings and seals. PMSM, Induction and Switched Reluctance Machines (SRM), all pulse modulated, were considered, but PMSM were determined to provide the highest power density for the shortest motors. Both radial and axial electric PMSM driven motors were designed with axial designs deemed more rugged for ultra-high speed, drilling applications. The 6.91 cm (2.72 inch) OD axial inverted motor can generate 4.18KW (5.61 Hp) power at 10,000 rpm with a 4 Nm (2.95 ft-lbs) of torque for every 30.48 cm (12 inches) of power section. The 6.91 cm (2.72 inch) OD radial inverted motor can generate 5.03 KW (6.74 Hp) with 4.8 Nm (3.54 ft-lb) torque at 10,000 rpm for every 30.48 cm (12 inches) of power section. The 4.29 cm (1.69 inch) OD radial inverted motor can generate 2.56 KW (3.43 Hp) power with 2.44 Nm (1.8 ft-lb) torque at

  2. High quality single atomic layer deposition of hexagonal boron nitride on single crystalline Rh(111) four-inch wafers

    Energy Technology Data Exchange (ETDEWEB)

    Hemmi, A.; Bernard, C.; Cun, H.; Roth, S.; Klöckner, M.; Kälin, T.; Osterwalder, J.; Greber, T., E-mail: greber@physik.uzh.ch [Physik-Institut, Universität Zürich, CH-8057 Zürich (Switzerland); Weinl, M.; Gsell, S.; Schreck, M. [Institut für Physik, Universität Augsburg, D-86135 Augsburg (Germany)

    2014-03-15

    The setup of an apparatus for chemical vapor deposition (CVD) of hexagonal boron nitride (h-BN) and its characterization on four-inch wafers in ultra high vacuum (UHV) environment is reported. It provides well-controlled preparation conditions, such as oxygen and argon plasma assisted cleaning and high temperature annealing. In situ characterization of a wafer is accomplished with target current spectroscopy. A piezo motor driven x-y stage allows measurements with a step size of 1 nm on the complete wafer. To benchmark the system performance, we investigated the growth of single layer h-BN on epitaxial Rh(111) thin films. A thorough analysis of the wafer was performed after cutting in atmosphere by low energy electron diffraction, scanning tunneling microscopy, and ultraviolet and X-ray photoelectron spectroscopies. The apparatus is located in a clean room environment and delivers high quality single layers of h-BN and thus grants access to large area UHV processed surfaces, which had been hitherto restricted to expensive, small area single crystal substrates. The facility is versatile enough for customization to other UHV-CVD processes, e.g., graphene on four-inch wafers.

  3. Investigation of Exhaust Backflow From a Simulated Cluster of Three Wide-Spaced Rocket Nozzles in a Near-Space Environment

    National Research Council Canada - National Science Library

    Cubbage, James M

    1965-01-01

    ... and to determine pressure and heat- transfer coefficients in the region washed by the backflow. Experiments were conducted in a 61-foot-diameter vacuum sphere using a sine solid-propellant rocket motor and a reflection plate...

  4. Lowell Observatory's 24-inch Clark Refractor: Its History and Renovation

    Science.gov (United States)

    Schindler, Kevin; Nye, Ralph; Rosenthal, Peter

    2016-01-01

    In 1895, Percival Lowell hired eminent telescope maker Alvan G. Clark to build a 24-inch refractor. Lowell intended the telescope intitally for observing Mars in support of his controversial theories about life on that planet. Clark finished the telescope within a year and at a cost of $20,000. Lowell and his staff of assistants and astronomers began observing through it on July 23, 1896, setting off a long and productive career for the telescope.While Lowell's Mars studies dominated early work with the Clark, V.M. Slipher by the 1910s was using it to observe planetary rotations and atmospheric compositions. He soon revolutionized spectroscopic studies, gathering excruciatingly long spectra - some in excess of 40 hours - of the so-called white nebula and determining startling radial velocities, evidence of an expanding universe. In the 1960s, scientists and artists teamed up on the Clark and created detailed lunar maps in support of the Apollo program.In recent decades, the Clark has played a central role in the education programs at Lowell, with general public audiences, students, and private groups all taking advantage of this unique resource.With this nearly 120 years of contant use, the Clark had been wearing down in recent years. The telescope was becoming more difficult to move, old electrical wiring in the dome was a fire hazard, and many of the telescope's parts needed to be repaired or replaced.In 2013, Lowell Observatory began a fundraising campaign, collecting $291,000 to cover the cost of dome and telescope renovation. Workers removed the entire telescope mount and tube assembly from the dome, examining every part from tube sections to individuals screws. They also stabilized the dome, adding a water vapor barrier and new outer wall while reinforcing the upper dome. The project lasted from January, 2014 through August, 2015. The facility reopened for daytime tours in September, 2015 and evening viewing the following month.

  5. Rocket Science 101 Interactive Educational Program

    Science.gov (United States)

    Armstrong, Dennis; Funkhouse, Deborah; DiMarzio, Donald

    2007-01-01

    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.

  6. Rocket Science at the Nanoscale.

    Science.gov (United States)

    Li, Jinxing; Rozen, Isaac; Wang, Joseph

    2016-06-28

    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.

  7. Motor Neurons

    DEFF Research Database (Denmark)

    Hounsgaard, Jorn

    2017-01-01

    Motor neurons translate synaptic input from widely distributed premotor networks into patterns of action potentials that orchestrate motor unit force and motor behavior. Intercalated between the CNS and muscles, motor neurons add to and adjust the final motor command. The identity and functional...... in in vitro preparations is far from complete. Nevertheless, a foundation has been provided for pursuing functional significance of intrinsic response properties in motoneurons in vivo during motor behavior at levels from molecules to systems....

  8. Lymphocytes on sounding rocket flights.

    Science.gov (United States)

    Cogoli-Greuter, M; Pippia, P; Sciola, L; Cogoli, A

    1994-05-01

    Cell-cell interactions and the formation of cell aggregates are important events in the mitogen-induced lymphocyte activation. The fact that the formation of cell aggregates is only slightly reduced in microgravity suggests that cells are moving and interacting also in space, but direct evidence was still lacking. Here we report on two experiments carried out on a flight of the sounding rocket MAXUS 1B, launched in November 1992 from the base of Esrange in Sweden. The rocket reached the altitude of 716 km and provided 12.5 min of microgravity conditions.

  9. Consort 1 sounding rocket flight

    Science.gov (United States)

    Wessling, Francis C.; Maybee, George W.

    1989-01-01

    This paper describes a payload of six experiments developed for a 7-min microgravity flight aboard a sounding rocket Consort 1, in order to investigate the effects of low gravity on certain material processes. The experiments in question were designed to test the effect of microgravity on the demixing of aqueous polymer two-phase systems, the electrodeposition process, the production of elastomer-modified epoxy resins, the foam formation process and the characteristics of foam, the material dispersion, and metal sintering. The apparatuses designed for these experiments are examined, and the rocket-payload integration and operations are discussed.

  10. Comparison of silicon strip tracker module size using large sensors from 6 inch wafers

    CERN Multimedia

    Honma, Alan

    1999-01-01

    Two large silicon strip sensor made from 6 inch wafers are placed next to each other to simulate the size of a CMS outer silicon tracker module. On the left is a prototype 2 sensor CMS inner endcap silicon tracker module made from 4 inch wafers.

  11. Torque expression of 0.018 and 0.022 inch conventional brackets

    NARCIS (Netherlands)

    Sifakakis, I.; Pandis, N.; Makou, M.; Eliades, T.; Katsaros, C.; Bourauel, C.

    2013-01-01

    The aim of this study was to assess the effect of the moments generated with low- and high-torque brackets. Four different bracket prescription-slot combinations of the same bracket type (Mini Diamond(R) Twin) were evaluated: high-torque 0.018 and 0.022 inch and low-torque 0.018 and 0.022 inch.

  12. Finite element based electric motor design optimization

    Science.gov (United States)

    Campbell, C. Warren

    1993-01-01

    The purpose of this effort was to develop a finite element code for the analysis and design of permanent magnet electric motors. These motors would drive electromechanical actuators in advanced rocket engines. The actuators would control fuel valves and thrust vector control systems. Refurbishing the hydraulic systems of the Space Shuttle after each flight is costly and time consuming. Electromechanical actuators could replace hydraulics, improve system reliability, and reduce down time.

  13. Test procedure for the Master-Lee and the modified Champion four inch hydraulic cutters

    International Nuclear Information System (INIS)

    Crystal, J.B.

    1995-01-01

    The Master-Lee and the modified Champion 4 Inch hydraulic cutters are being retested to gather and document information related to the following: determine if the Master-Lee cutters will cut the trunnions of an Aluminum fuel canister and a Stainless Steel fuel canister; determine if the Master-Lee cutters will cut 1 1/2 inch diameter fire hose; determine if the modified Champion 4 inch blade will cut sections of piping; and determine the effectiveness of the centering device for the Champion 4 Inch cutters. Determining the limitations of the hydraulic cutter will aid in the process of debris removal in the K-Basin. Based on a previous test, the cutters were returned to the manufacturer for modifications. The modifications to the Champion 4 Inch Cutter and further testing of the Master-Lee Cutter are the subjects of these feature tests

  14. Hybrid rocket propulsion systems for outer planet exploration missions

    Science.gov (United States)

    Jens, Elizabeth T.; Cantwell, Brian J.; Hubbard, G. Scott

    2016-11-01

    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.

  15. Integrated Composite Rocket Nozzle Extension, Phase I

    Data.gov (United States)

    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. Design study of laser fusion rocket

    International Nuclear Information System (INIS)

    Nakashima, Hideki; Shoyama, Hidetoshi; Kanda, Yukinori

    1991-01-01

    A design study was made on a rocket powered by laser fusion. Dependence of its flight performance on target gain, driver repetition rate and fuel composition was analyzed to obtain optimal design parameters of the laser fusion rocket. The results indicate that the laser fusion rocket fueled with DT or D 3 He has the potential advantages over other propulsion systems such as fission rocket for interplanetary travel. (author)

  17. Launch Excitement with Water Rockets

    Science.gov (United States)

    Sanchez, Juan Carlos; Penick, John

    2007-01-01

    Explosions and fires--these are what many students are waiting for in science classes. And when they do occur, students pay attention. While we can't entertain our students with continual mayhem, we can catch their attention and cater to their desires for excitement by saying, "Let's make rockets." In this activity, students make simple, reusable…

  18. Measuring Model Rocket Engine Thrust Curves

    Science.gov (United States)

    Penn, Kim; Slaton, William V.

    2010-01-01

    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…

  19. The SERTS-97 Rocket Experiment on Study Activity on the Sun: Flight 36.167-GS on 1997 November 18

    Science.gov (United States)

    Swartz, Marvin; Condor, Charles E.; Davila, Joseph M.; Haas, J. Patrick; Jordan, Stuart D.; Linard, David L.; Miko, Joseph J.; Nash, I. Carol; Novello, Joseph; Payne, Leslie J.; hide

    1999-01-01

    This paper describes mainly the 1997 version of the Solar EUV Rocket Telescope and Spectrograph (SERTS-97), a scientific experiment that operated on NASA's suborbital rocket flight 36.167-GS. Its function was to study activity on the Sun and to provide a cross calibration for the CDS instrument on the SOHO satellite. The experiment was designed, built, and tested by the Solar Physics Branch of the Laboratory for Astronomy and Solar Physics at the Goddard Space Flight Center (GSFC). Other essential sections of the rocket were built under the management of the Sounding Rockets Program Office. These sections include the electronics, timers, IGN despin, the SPARCS pointing controls, the S-19 flight course correction section, the rocket motors, the telemetry, ORSA, and OGIVE.

  20. Theoretical Rocket Performance of Liquid Methane with Several Fluorine-Oxygen Mixtures Assuming Frozen Composition

    Science.gov (United States)

    Gordon, Sanford; Kastner, Michael E

    1958-01-01

    Theoretical rocket performance for frozen composition during expansion was calculated for liquid methane with several fluorine-oxygen mixtures for a range of pressure ratios and oxidant-fuel ratios. The parameters included are specific impulse, combustion-chamber temperature, nozzle-exit temperature molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, and thermal conductivity. The maximum calculated value of specific impulse for a chamber pressure of 600 pounds per square inch absolute (40.827atm) and an exit pressure of 1 atmosphere is 315.3 for 79.67 percent fluorine in the oxidant.

  1. The problem of space travel: The rocket motor

    Science.gov (United States)

    Noordung, Hermann; Stuhlinger, Ernst (Editor); Hunley, J. D. (Editor); Garland, Jennifer (Editor)

    1995-01-01

    This is an English translation of Noordung's 'Das Problem der Befahrung des Weltraums'. It is a part of the NASA History Series. The book provides Noordung's insight as to how a space station might be constructed. Keep in mind that the author passed away in 1929. It contains ideas that were criticized for faults and those that were acclaimed for their potentially successful ingenuity.

  2. Design Methods in Solid Rocket Motors. Revised Version 1988

    Science.gov (United States)

    1988-04-01

    La necessite de faire Ie point de maniere periodique sur ces prob1emes est absolument evidente. On doit ajouter qu’il existe tres peu d’ouvrages de...structurales froides . Cependant, de nombreux problemes techniques et economiques restent a resoudre : assurance de la qua- lite, controle non destructif

  3. Aluminum Agglomeration and Trajectory in Solid Rocket Motors

    Science.gov (United States)

    2007-08-30

    34the stepwise oxidation of aluminum (that) is caused by the sequence of polymorphic phase transitions occurring in the growing oxide film",2 5 . 25...C. and Yang, V., "Analysis of RDX Monopropellant Combustion with Two-Phase Subsurface Reactions", Journal of Propulsion and Power, Vol. 11, No. 4...temperature. Generalized mechanisms have been developed and applied to many ingredients such as HMX , GAP, NG, BTTN, ADN and AP.10 The burning rates of

  4. Nuclear Rocket Test Facility Decommissioning Including Controlled Explosive Demolition of a Neutron-Activated Shield Wall

    International Nuclear Information System (INIS)

    Michael Kruzic

    2007-01-01

    Located in Area 25 of the Nevada Test Site, the Test Cell A Facility was used in the 1960s for the testing of nuclear rocket engines, as part of the Nuclear Rocket Development Program. The facility was decontaminated and decommissioned (D and D) in 2005 using the Streamlined Approach For Environmental Restoration (SAFER) process, under the Federal Facilities Agreement and Consent Order (FFACO). Utilities and process piping were verified void of contents, hazardous materials were removed, concrete with removable contamination decontaminated, large sections mechanically demolished, and the remaining five-foot, five-inch thick radiologically-activated reinforced concrete shield wall demolished using open-air controlled explosive demolition (CED). CED of the shield wall was closely monitored and resulted in no radiological exposure or atmospheric release

  5. 14 CFR 101.25 - Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets.

    Science.gov (United States)

    2010-01-01

    ... Power Rockets and Class 3-Advanced High Power Rockets. 101.25 Section 101.25 Aeronautics and Space... OPERATING RULES MOORED BALLOONS, KITES, AMATEUR ROCKETS AND UNMANNED FREE BALLOONS Amateur Rockets § 101.25 Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets. When operating...

  6. Piezoelectric/magnetostrictive resonant inchworm motor

    Science.gov (United States)

    Miesner, John E.; Teter, Joseph P.

    1994-05-01

    Magnetostrictive and piezoelectric materials were used to create a linear motor operating on the inchworm principle. This motor operates at an electrical resonance, switching power internally between inductive and capacitive components. Magnetic coils surrounding the two Terfenol-D rods which drive the inchworm's center expanding element form the inductive component. Piezoelectric stacks that control the end clamping action are the capacitive components. The normal electrical phase relationship between these components provides natural drive timing for the inchworm. The motor direction can be easily reversed by changing the magnetic bias on the Terfenol. A prototype motor was built that achieved a stall load of 26 lb and no-load speed of 1 inch/sec vs the design of 30 lb and 1.3 inch/sec. A new type of power supply that switches power from a dc source was built for the motor. This power supply uses a small number of components to exactly supply the energy used in each inchworm cycle. It tracks the motor circuit resonance and is not affected by frequency shifts.

  7. Development of high performance hybrid rocket fuels

    Science.gov (United States)

    Zaseck, Christopher R.

    . In order to examine paraffin/additive combustion in a motor environment, I conducted experiments on well characterized aluminum based additives. In particular, I investigate the influence of aluminum, unpassivated aluminum, milled aluminum/polytetrafluoroethylene (PTFE), and aluminum hydride on the performance of paraffin fuels for hybrid rocket propulsion. I use an optically accessible combustor to examine the performance of the fuel mixtures in terms of characteristic velocity efficiency and regression rate. Each combustor test consumes a 12.7 cm long, 1.9 cm diameter fuel strand under 160 kg/m 2s of oxygen at up to 1.4 MPa. The experimental results indicate that the addition of 5 wt.% 30 mum or 80 nm aluminum to paraffin increases the regression rate by approximately 15% compared to neat paraffin grains. At higher aluminum concentrations and nano-scale particles sizes, the increased melt layer viscosity causes slower regression. Alane and Al/PTFE at 12.5 wt.% increase the regression of paraffin by 21% and 32% respectively. Finally, an aging study indicates that paraffin can protect air and moisture sensitive particles from oxidation. The opposed burner and aluminum/paraffin hybrid rocket experiments show that additives can alter bulk fuel properties, such as viscosity, that regulate entrainment. The general effect of melt layer properties on the entrainment and regression rate of paraffin is not well understood. Improved understanding of how solid additives affect the properties and regression of paraffin is essential to maximize performance. In this document I investigate the effect of melt layer properties on paraffin regression using inert additives. Tests are performed in the optical cylindrical combustor at ˜1 MPa under a gaseous oxygen mass flux of ˜160 kg/m2s. The experiments indicate that the regression rate is proportional to mu0.08rho 0.38kappa0.82. In addition, I explore how to predict fuel viscosity, thermal conductivity, and density prior to testing

  8. Unique nuclear thermal rocket engine

    International Nuclear Information System (INIS)

    Culver, D.W.; Rochow, R.

    1993-06-01

    In January, 1992, a new, advanced nuclear thermal rocket engine (NTRE) concept intended for manned missions to the moon and to Mars was introduced (Culver, 1992). 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 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) the reactor design and cooling scheme simplification while retaining a high pressure power balance option; (2) elimination need for a new, uncooled nozzle throat material suitable for long life application; (3) a practical provision for reactor power control; and (4) use of near-term, long-life turbopumps

  9. Description and Flight Performance Results of the WASP Sounding Rocket

    Science.gov (United States)

    De Pauw, J. F.; Steffens, L. E.; Yuska, J. A.

    1968-01-01

    A general description of the design and construction of the WASP sounding rocket and of the performance of its first flight are presented. The purpose of the flight test was to place the 862-pound (391-kg) spacecraft above 250 000 feet (76.25 km) on free-fall trajectory for at least 6 minutes in order to study the effect of "weightlessness" on a slosh dynamics experiment. The WASP sounding rocket fulfilled its intended mission requirements. The sounding rocket approximately followed a nominal trajectory. The payload was in free fall above 250 000 feet (76.25 km) for 6.5 minutes and reached an apogee altitude of 134 nautical miles (248 km). Flight data including velocity, altitude, acceleration, roll rate, and angle of attack are discussed and compared to nominal performance calculations. The effect of residual burning of the second stage motor is analyzed. The flight vibration environment is presented and analyzed, including root mean square (RMS) and power spectral density analysis.

  10. Development of the Hawk/Nike Hawk sounding rocket vehicles

    Science.gov (United States)

    Flowers, B. J.

    1976-01-01

    A new sounding rocket family, the Hawk and Nike-Hawk Vehicles, have been developed, flight tested and added to the NASA Sounding Rocket Vehicle Stable. The Hawk is a single-stage vehicle that will carry 35.6 cm diameter payloads weighing 45.5 kg to 91 kg to altitudes of 78 km to 56 km, respectively. The two-stage Nike-Hawk will carry payloads weighing 68 kg to 136 kg to altitudes of 118 km to 113 km, respectively. Both vehicles utilize the XM22E8 Hawk rocket motor which is available in large numbers as a surplus item from the U.S. Army. The Hawk fin and tail can hardware were designed in-house. The Nike tail can and fin hardware are surplus Nike-Ajax booster hardware. Development objectives were to provide a vehicle family with a larger diameter, larger volume payload capability than the Nike-Apache and Nike-Tomahawk vehicles at comparable cost. Both vehicles performed nominally in flight tests.

  11. Project W320 52-inch diameter equipment container load test: Test report

    International Nuclear Information System (INIS)

    Bellomy, J.R.

    1995-01-01

    This test report summarizes testing activities and documents the results of the load tests performed on-site and off-site to structural qualify the 52-inch equipment containers designed and fabricated under Project W-320

  12. Preparatory Study of Photomultiplier Tubes of 10-inch and 3-inch Diameter for KM3NeT Underwater Neutrino Telescope

    International Nuclear Information System (INIS)

    Aiello, S.; Giordano, V.; Leonora, E.

    2015-01-01

    Large area photomultipliers are widely used in neutrino and astro-particle detectors to measure Cherenkov light in media like water or ice. The key element of these detectors are the so-called 'optical module', which consists of a photodetector enclosed in a transparent pressure-resistant container to protect it and ensure good light transmission. KM3NeT collaboration aims to construct an underwater 'hybrid' neutrino telescope by using two models detection unit. The 'tower' detection unit will be composed of large area 10-inch photomultipliers tube enclosed into 13-inch glass vessel sphere. In the 'string' detection unit instead, the light detector will be the 'digital optical module' (DOM) a glass vessel of 17-inch with 31 photomultipliers of 3- inch diameter looking upwards and downwards. The choice of two different kinds of photomultipliers, obliges us to investigate their main characteristics. Noise pulses at the anode of each photomultiplier strongly affect the performance of the detector. A large study was conducted on noise pulses of large area photomultipliers, considering time and charge distributions of dark pulses, pre-pulses, delayed pulses, and after-pulses. The contribution to noise pulses due to the presence of the external glass vessels was also studied. Moreover the presence of the Earth's magnetic field should modify quantities like gain and transit time spread in photomultipliers and we will deeply investigate on this. (authors)

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

    International Nuclear Information System (INIS)

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

    2016-01-01

    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. (paper)

  14. A Green, Safe, Dual-pulse Solid Motor for CubeSat Orbit Changing, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — Small satellites such as CubeSats are in need of responsive propulsion, but are limited due to their size. Though single pulse, AP/HTPB fueled solid rocket motors...

  15. MODELING A ROCKET ELASTIC STRUCTURE AS A BECK’S COLUMN UNDER FOLLOWER FORCE

    OpenAIRE

    Brejão, Leandro Forne; Brasil, Reyolando M. L. R. F.

    2017-01-01

    It is intended, in this paper, to develop a mathematical model of an elastic space rocket structure as a Beck’s column excited by a follower (or circulatory) force. This force represents the rocket motor thrust that should be always in the direction of the tangent to the structure deformed axis at the base of the vehicle. We present a simplified two degree of freedom rigid bars discrete model. Its system of two second order nonlinear ordinary differential equations of motion are derived via L...

  16. Current and Future Critical Issues in Rocket Propulsion Systems

    Science.gov (United States)

    Navaz, Homayun K.; Dix, Jeff C.

    1998-01-01

    The objective of this research was to tackle several problems that are currently of great importance to NASA. In a liquid rocket engine several complex processes take place that are not thoroughly understood. Droplet evaporation, turbulence, finite rate chemistry, instability, and injection/atomization phenomena are some of the critical issues being encountered in a liquid rocket engine environment. Pulse Detonation Engines (PDE) performance, combustion chamber instability analysis, 60K motor flowfield pattern from hydrocarbon fuel combustion, and 3D flowfield analysis for the Combined Cycle engine were of special interest to NASA. During the summer of 1997, we made an attempt to generate computational results for all of the above problems and shed some light on understanding some of the complex physical phenomena. For this purpose, the Liquid Thrust Chamber Performance (LTCP) code, mainly designed for liquid rocket engine applications, was utilized. The following test cases were considered: (1) Characterization of a detonation wave in a Pulse Detonation Tube; (2) 60K Motor wall temperature studies; (3) Propagation of a pressure pulse in a combustion chamber (under single and two-phase flow conditions); (4) Transonic region flowfield analysis affected by viscous effects; (5) Exploring the viscous differences between a smooth and a corrugated wall; and (6) 3D thrust chamber flowfield analysis of the Combined Cycle engine. It was shown that the LTCP-2D and LTCP-3D codes are capable of solving complex and stiff conservation equations for gaseous and droplet phases in a very robust and efficient manner. These codes can be run on a workstation and personal computers (PC's).

  17. Two-dimensional motions of rockets

    International Nuclear Information System (INIS)

    Kang, Yoonhwan; Bae, Saebyok

    2007-01-01

    We analyse the two-dimensional motions of the rockets for various types of rocket thrusts, the air friction and the gravitation by using a suitable representation of the rocket equation and the numerical calculation. The slope shapes of the rocket trajectories are discussed for the three types of rocket engines. Unlike the projectile motions, the descending parts of the trajectories tend to be gentler and straighter slopes than the ascending parts for relatively large launching angles due to the non-vanishing thrusts. We discuss the ranges, the maximum altitudes and the engine performances of the rockets. It seems that the exponential fuel exhaustion can be the most potent engine for the longest and highest flights

  18. Risk assessment of K Basin twelve-inch and four-inch drain valve failure from a postulated seismic initiating event

    Energy Technology Data Exchange (ETDEWEB)

    MORGAN, R.G.

    1999-06-23

    The Spent Nuclear Fuel (SNF) Project will transfer metallic SNF from the Hanford 105 K-East and 105 K-West Basins to safe interim storage in the Canister Storage Building in the 200 Area. The initial basis for design, fabrication, installation, and operation of the fuel removal systems was that the basin leak rate which could result from a postulated accident condition would not be excessive relative to reasonable recovery operations. However, an additional potential K Basin water leak path is through the K Basin drain valves. Three twelve-inch drain valves are located in the main basin bays along the north wall. Five four-inch drain valves are located in the north and south loadout pits (NLOP and SLOP), the weasel pit, the technical viewing pit, and the discharge chute pit. The sumps containing the valves are filled with concrete which covers the drain valve body. Visual observations indicate that only the valve's bonnet and stem are exposed above the basin concrete floor for the twelve-inch drain valve and that much less of the valve's bonnet and stem are exposed above the basin concrete floor for the five four-inch drain valves. It was recognized, however, that damage of the drain valve bonnet or stem during a seismic initiating event could provide a potential K Basin water leak path. The objectives of this analysis are to: (1) evaluate the likelihood of damaging the three twelve-inch drain valves located along the north wall of the main basin and the five four-inch drain valves located in the pits from a seismic initiating event, and (2) determine the likelihood of exceeding a specific consequence (initial leak rate) from a damaged valve. The analysis process is a risk-based uncertainty analysis where each variable is modeled using available information and engineering judgement. The uncertainty associated with each variable is represented by a probability distribution (probability density function). Uncertainty exists because of the inherent

  19. Am/Cm TTR testing - 3/8-inch glass beads evaluation in CIM5

    International Nuclear Information System (INIS)

    Witt, D. C.

    2000-01-01

    To facilitate the procurement and handling of the glass former for Am/Cm vitrification in the F-Canyon MPPF, 1/4 inch and 3/8 inch diameter glass beads were purchased from Corning for evaluation in the 5 inch Cylindrical Induction Melter (CIM5). Prior to evaluating the beads in the CIM5, tests were conducted in the Drain Tube Test Stand (DTTS) with 1/4 inch beads, 3/8 inch beads, and a 50/50 mixture to identify any process concerns. Results of the DTTS tests are summarized in Attachment 1. A somewhat larger volume expansion was experienced in all three DTTS runs as compared to a standard run using cullet. Further testing of the use of glass beads in the CIM5 was requested by the Design Authority as Task 1.02 of Technical Task Request 99-MNSS/SE-006. Since the Technical Task Plan was not yet approved, the completion of this task was conducted under an authorization request approved by the SRTC Laboratory Director, S. Wood. This request is included as Attachment 2

  20. Operational maneuvers and pipelines activities repairs for the 32 inches scraper tool recovering

    Energy Technology Data Exchange (ETDEWEB)

    Valdivia, Jose; Salguero, Luis; Villanueva, Pedro [Compania Operadora del Gas Amazonas, Lima (Peru)

    2009-07-01

    Transportadora de Gas del Peru and the Compania Operadora de Gas del Amazonas, responsible companies of the transport, operation and maintenance of the pipelines who transport natural gas and natural gas liquids respectively of the Camisea Project - Peru, following the internal policies and the maintenance plan of the pipeline transportation system was planned the activities for the internal pipeline inspection of these activities for 729.3 Km of natural gas pipeline covering diameters of 32 inches, 24 inches and 18 inches. After the first run of the cleaning tool, was scheduled the launch of the dummy scraper (scraper tool) along to the first 210 Km of the 32 inches natural gas pipeline , given changes in elevation along the trace and the low flow of transport. This scraper tool could not reach the final destination. After many series operational maneuvers as venting, creation of differential pressure in valves, the scraper tool only reach the first 75 Km of the trace. After an exhaustive analysis of trending pressure variations, it was concluded that this scraper showed intermittent progress of short durations, concluding that this scraper had not reach the next check point. In this way was decided to conduct operational maneuvers in order to locate, relocated and retrieve the scrapper tool from de 32 inches natural gas pipeline. (author)

  1. SSTO rockets. A practical possibility

    Science.gov (United States)

    Bekey, Ivan

    1994-07-01

    Most experts agree that single-stage-to-orbit (SSTO) rockets would become feasible if more advanced technologies were available to reduce the vehicle dry weight, increase propulsion system performance, or both. However, these technologies are usually judged to be very ambitious and very far off. This notion persists despite major advances in technology and vehicle design in the past decade. There appears to be four major misperceptions about SSTOs, regarding their mass fraction, their presumed inadequate performance margin, their supposedly small payloads, and their extreme sensitivity to unanticipated vehicle weight growth. These misperceptions can be dispelled for SSTO rockets using advanced technologies that could be matured and demonstrated in the near term. These include a graphite-composite primary structure, graphite-composite and Al-Li propellant tanks with integral reusable thermal protection, long-life tripropellant or LOX-hydrogen engines, and several technologies related to operational effectiveness, including vehicle health monitoring, autonomous avionics/flight control, and operable launch and ground handling systems.

  2. The Development of Rocketry Capability in New Zealand—World Record Rocket and First of Its Kind Rocketry Course

    Directory of Open Access Journals (Sweden)

    George Buchanan

    2015-02-01

    Full Text Available The University of Canterbury has developed a rocket research group, UC Rocketry, which recently broke the world altitude record for an I-class motor (impulse of 320–640 Ns and has run a rocketry course for the first time in New Zealand. This paper discusses the development and results of the world record rocket “Milly” and details all the fundamental elements of the rocketry final year engineering course, including the manufacturing processes, wind tunnel testing, avionics, control and the final rocket launch of “Smokey”. The rockets Milly and Smokey are an example of the design, implementation and testing methodologies that have significantly contributed to research and graduates for New Zealand’s space program.

  3. Progress in Electron Beam Mastering of 100 Gbit/inch2 Density Disc

    Science.gov (United States)

    Takeda, Minoru; Furuki, Motohiro; Yamamoto, Masanobu; Shinoda, Masataka; Saito, Kimihiro; Aki, Yuichi; Kawase, Hiroshi; Koizumi, Mitsuru; Miyokawa, Toshiaki; Mutou, Masao; Handa, Nobuo

    2004-07-01

    We developed an electron beam recorder (EBR) capable of recording master discs under atmospheric conditions using a novel differential pumping head. Using the EBR and optimized fabrication process for Si-etched discs with reactive ion etching (RIE), a bottom signal jitter of 9.6% was obtained from a 36 Gbit/inch2 density disc, readout using a near-field optical pickup with an effective numerical aperture (NA) of 1.85 and a wavelength of 405 nm. We also obtained the eye patterns from a 70 Gbit/inch2 density disc readout using an optical pickup with a 2.05 NA and the same wavelength, and showed almost the same modulation ratio as the simulation value. Moreover, the capability of producing pit patterns corresponding to a 104 Gbit/inch2 density is demonstrated.

  4. Comparision of Incidental Reflection From Containerized Maintenance/Housekeeping Solutions and One Inch of Water

    Energy Technology Data Exchange (ETDEWEB)

    Chapman, Bryan Scott [Los Alamos National Lab. (LANL), Los Alamos, NM (United States); MacQuigg, Michael Robert [Los Alamos National Lab. (LANL), Los Alamos, NM (United States); Wysong, Andrew Russell [Los Alamos National Lab. (LANL), Los Alamos, NM (United States)

    2016-12-15

    This document addresses the incidental reflector reactivity worth of containerized maintenance/housekeeping fluids for use in PF-4 at Los Alamos National Laboratory (LANL). The intent of the document is to analyze containerized maintenance/housekeeping fluids which will be analyzed as water that may be present under normal conditions of an operation. The reactivity worth is compared to the reactivity worth due to I-inch of close-fitting 4n water reflection and I-inch of close-fitting radial water reflection. Both have been used to bound incidental reflection by 2-liter bottles in criticality safety evaluations. The conclusion is that, when the maintenance/housekeeping fluids are containerized the reactivity increase from a configuration which is bounding of normal conditions (up to eight bottles modeled with 2-liters of solution at varying diameter) is bound by I-inch of close fitting 4n water relection.

  5. Maneuver of Spinning Rocket in Flight

    OpenAIRE

    HAYAKAWA, Satio; ITO, Koji; MATSUI, Yutaka; NOGUCHI, Kunio; UESUGI, Kuninori; YAMASHITA, Kojun

    1980-01-01

    A Yo-despin device successfully functioned to change in flight the precession axis of a sounding rocket for astronomical observation. The rocket attitudes before and after yodespin were measured with a UV star sensor, an infrared horizon sensor and an infrared telescope. Instrumentation and performance of these devices as well as the attitude data during flight are described.

  6. Ionospheric shock waves triggered by rockets

    Directory of Open Access Journals (Sweden)

    C. H. Lin

    2014-09-01

    Full Text Available This paper presents a two-dimensional structure of the shock wave signatures in ionospheric electron density resulting from a rocket transit using the rate of change of the total electron content (TEC derived from ground-based GPS receivers around Japan and Taiwan for the first time. From the TEC maps constructed for the 2009 North Korea (NK Taepodong-2 and 2013 South Korea (SK Korea Space Launch Vehicle-II (KSLV-II rocket launches, features of the V-shaped shock wave fronts in TEC perturbations are prominently seen. These fronts, with periods of 100–600 s, produced by the propulsive blasts of the rockets appear immediately and then propagate perpendicularly outward from the rocket trajectory with supersonic velocities between 800–1200 m s−1 for both events. Additionally, clear rocket exhaust depletions of TECs are seen along the trajectory and are deflected by the background thermospheric neutral wind. Twenty minutes after the rocket transits, delayed electron density perturbation waves propagating along the bow wave direction appear with phase velocities of 800–1200 m s−1. According to their propagation character, these delayed waves may be generated by rocket exhaust plumes at earlier rocket locations at lower altitudes.

  7. Aerodynamics and flow characterisation of multistage rockets

    Science.gov (United States)

    Srinivas, G.; Prakash, M. V. S.

    2017-05-01

    The main objective of this paper is to conduct a systematic flow analysis on single, double and multistage rockets using ANSYS software. Today non-air breathing propulsion is increasing dramatically for the enhancement of space exploration. The rocket propulsion is playing vital role in carrying the payload to the destination. Day to day rocket aerodynamic performance and flow characterization analysis has becoming challenging task to the researchers. Taking this task as motivation a systematic literature is conducted to achieve better aerodynamic and flow characterization on various rocket models. The analyses on rocket models are very little especially in numerical side and experimental area. Each rocket stage analysis conducted for different Mach numbers and having different flow varying angle of attacks for finding the critical efficiency performance parameters like pressure, density and velocity. After successful completion of the analysis the research reveals that flow around the rocket body for Mach number 4 and 5 best suitable for designed payload. Another major objective of this paper is to bring best aerodynamics flow characterizations in both aero and mechanical features. This paper also brings feature prospectus of rocket stage technology in the field of aerodynamic design.

  8. A Flight Demonstration of Plasma Rocket Propulsion

    Science.gov (United States)

    Petro, Andrew

    1999-01-01

    The Advanced Space Propulsion Laboratory at the Johnson Space Center has been engaged in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One advantageous feature of this rocket concept is the ability to vary its specific impulse so that it can be operated in a mode which maximizes propellant efficiency or a mode which maximizes thrust. This presentation will describe a proposed flight experiment in which a simple version of the rocket will be tested in space. In addition to the plasma rocket, the flight experiment will also demonstrate the use of a superconducting electromagnet, extensive use of heat pipes, and possibly the transfer of cryogenic propellant in space.

  9. Subsonic Glideback Rocket Demonstrator Flight Testing

    Science.gov (United States)

    DeTurris, Dianne J.; Foster, Trevor J.; Barthel, Paul E.; Macy, Daniel J.; Droney, Christopher K.; Talay, Theodore A. (Technical Monitor)

    2001-01-01

    For the past two years, Cal Poly's rocket program has been aggressively exploring the concept of remotely controlled, fixed wing, flyable rocket boosters. This program, embodied by a group of student engineers known as Cal Poly Space Systems, has successfully demonstrated the idea of a rocket design that incorporates a vertical launch pattern followed by a horizontal return flight and landing. Though the design is meant for supersonic flight, CPSS demonstrators are deployed at a subsonic speed. Many steps have been taken by the club that allowed the evolution of the StarBooster prototype to reach its current size: a ten-foot tall, one-foot diameter, composite material rocket. Progress is currently being made that involves multiple boosters along with a second stage, third rocket.

  10. Performances Study of a Hybrid Rocket Engine

    Directory of Open Access Journals (Sweden)

    Adrian-Nicolae BUTURACHE

    2018-06-01

    Full Text Available This paper presents a study which analyses the functioning and performances optimization of a hybrid rocket engine based on gaseous oxygen and polybutadiene polymer (HTPB. Calculations were performed with NASA CEA software in order to obtain the parameters resulted following the combustion process. Using these parameters, the main parameters of the hybrid rocket engine were optimized. Using the calculus previously stated, an experimental rocket engine producing 100 N of thrust was pre-dimensioned, followed by an optimization of the rocket engine as a function of several parameters. Having the geometry and the main parameters of the hybrid rocket engine combustion process, numerical simulations were performed in the CFX – ANSYS commercial software, which allowed visualizing the flow field and the jet expansion. Finally, the analytical calculus was validated through numerical simulations.

  11. Studies on Flame Spread with Sudden Expansions of Ports of Solid Propellant Rockets under Elevated Pressure.

    OpenAIRE

    B.N. Raghunandan; N.S. Madhavan; C. Sanjeev; V.R.S. Kumar

    1996-01-01

    A detailed experimental study on flame spread over non-uniform ports of solid propellant rockets has been carried out. An idealised. 2-dimensional laboratory motor was used for the experimental study with the aid of cinephotography. Freshly prepared rectangular HTPB propellant with backward facing step was used as the specimenfor this study. It has been shown conclusively that under certain conditions of step location. step height and port height which govern the velocity of gases at the step...

  12. 2005 40th Annual Armament Systems: Guns - Ammunition - Rockets - Missiles Conference and Exhibition. Volume 2: Wednesday

    Science.gov (United States)

    2005-04-28

    PM] Abraham Overview, Mr. Robert Daunfeldt, Bofors Defence Summary Overview of an Advanced 2.75 Hypervelocity Weapon, Mr. Larry Bradford , CAT Flight...Substantially Improves 2.75 Rocket Lethality, Safety, Survivability Mr. Larry Bradford , CAT Flight Services, Inc. APKWS Flight Test Results Mr. Larry S...Company Lead: Larry Bradford Atlantic Research Propellant Mixing/Loading, Nozzle Manufacturing, Corporation Motor Static Testing Company Lead: Steve

  13. Terrier Black Brant VC design characteristics and program status. [rocket development

    Science.gov (United States)

    Payne, B. R.; Mayo, E. E.

    1979-01-01

    In the present paper, the design analysis of the Terrier-Black Brant VC, representing the latest addition to the Black Brant rocket family, is discussed, including the aerodynamic, structural, thermal, and operational aspects. An appreciable increase in apogee, as compared to the BBVC and Nike/BBVC, is achieved without any modifications to the well-proven BBV motor or degradation of the thermal or dynamic flight environment.

  14. The electric motor handbook

    Energy Technology Data Exchange (ETDEWEB)

    Hurst, R.W.; Feltham, P. (eds.)

    2004-05-01

    This handbook outlines the important role that electric motors play in modern society. It covers the field of motor applications from various motor types to their use and repair. It also presents practical applications of electric motors and methods on motor efficiency. More than half of all electricity generated, and 75 per cent of all industrial electricity consumption is consumed by electric motors. Electrical personnel must be aware of all factors involved in electric motors in order to choose and apply the appropriate size of electric motor. These factors include efficiency, sizing and proper application. The efficient use and maximum life expectancy of electric motors depends on proper motor protection, control and maintenance. This handbook includes articles from leading experts on electric motors in modern electrical systems. The content includes: design considerations; proper electric motor sizing techniques; optimal electric motor application; electric motor protection technology; electric motor control principles; electric motor maintenance and troubleshooting; induction electric motors; electric motor bearing currents; electric motor bearing lubrication; electromagnetism; electric motor enclosures; electric motor testing; electric motor repair; DC electric motor; electric motor starters; electric motor brushes; industrial electric motors; electric motor diagrams; AC electric motors; electric motor wiring; electric motor service; electric motor rewinding; electric motor winding; diagram of electric motor wiring; electric motor kit; and, troubleshooting electric motors. A directory of motor manufacturers and suppliers was also included. refs., tabs., figs.

  15. Interim results: fines recycle testing using the 4-inch diameter primary graphite burner

    International Nuclear Information System (INIS)

    Palmer, W.B.

    1975-05-01

    The results of twenty-two HTGR primary burner runs in which graphite fines were recycled pneumatically to the 4-inch diameter pilot-plant primary fluidized-bed burner are described. The result of the tests showed that zero fines accumulation can easily be achieved while operating at plant equivalent burn rates. (U.S.)

  16. Evaluation of 3 Inch SN-219 Failure and S and SX Tank Farm Saltwell Piping

    International Nuclear Information System (INIS)

    ELSEN, J.J.

    2000-01-01

    Evaluation of direct buried piping currently in use or designated for future Saltwell pumping in S and SX Farms. Documented evaluation of failed S-103 saltwell pumping transfer line 3 inch SN-219. This evaluation is intended to reflect current status of Saltwell piping, when taken in context with referenced documents

  17. Child seat belt guidelines: Examining the 4 feet 9 inches rule as the standard.

    Science.gov (United States)

    Morse, Amber M; Aitken, Mary E; Mullins, Samantha H; Miller, Beverly K; Pomtree, Mindy M; Ulloa, Erin M; Montgomery, Jeffrey S; Saylors, Marie E

    2017-08-01

    Current American Academy of Pediatrics recommendations regarding transition from child safety/booster seat to adult safety belt use indicate that children should be at least 4 feet 9 inches, 8 years old, or 80 pounds. Proper fit in the vehicle seat, assessed with a five-point fit test, should also be met. Although most children reach 4 feet 9 inches around age 8 years, each child and vehicle presents a unique combination; thus a child may not fit appropriately in all vehicle types using only the 4 feet 9 inches requirement. We enrolled children, aged 7 years to 12 years, into our study. Height, weight, and demographic data were obtained. A Child Passenger Safety Technician then performed the five-point fit test in each of a uniform lineup of five vehicles. Data were collected on fit in the standard vehicle seat and also in a booster seat. We set 90% as the threshold proportion of children who meet all criteria for proper fit to validate current recommendations of a height of 4 feet 9 inches. Data were collected on 388 children. The percentage of 90% proper fit was met in the compact car and small sport-utility vehicle (SUV). However, only 80 (77%) of 104 students (p vehicles (large SUVs and trucks). This emphasizes the need for evaluation of fit by a trained personnel and/or development of standard back seat dimensions in all vehicles for maximum safety. Epidemiologic study, level III; Therapeutic study, level V.

  18. Full scale ambient water flow tests of a 10-inch emergency release coupling for LNG transfer

    NARCIS (Netherlands)

    Putte, L.J. van der; Webber, T.; Bokhorst, E. van; Revell, C.

    2016-01-01

    For LNG transfer in ship-to-ship and ship-to-shore configurations emergency release couplings (F.RC) in combination with loading arms and multi-composite hoses are applied In view of a demand for increasing transfer flow rates in offshore LNG applications a 10-inch ERC has been developed intended

  19. 50 CFR Figure 12 to Part 223 - Escape Opening & Cover Dimensions for 71-inch TED

    Science.gov (United States)

    2010-10-01

    ... 50 Wildlife and Fisheries 7 2010-10-01 2010-10-01 false Escape Opening & Cover Dimensions for 71-inch TED 12 Figure 12 to Part 223 Wildlife and Fisheries NATIONAL MARINE FISHERIES SERVICE, NATIONAL OCEANIC AND ATMOSPHERIC ADMINISTRATION, DEPARTMENT OF COMMERCE MARINE MAMMALS THREATENED MARINE AND ANADROMOUS SPECIES Pt. 223, Fig. 12 Figure 12 to...

  20. Worcester 1 Inch Solenoid-Actuated Gas-Operated VPS System Ball Valve

    International Nuclear Information System (INIS)

    VAN KATWIJK, C.

    2000-01-01

    1 inch Gas-operated full-port ball valve incorporates a solenoid and limit switches as integral parts of the actuator. The valve is normally open and fails safe to the closed position. The associated valve position switch is class GS

  1. Rocket Ozone Data Recovery for Digital Archival

    Science.gov (United States)

    Hwang, S. H.; Krueger, A. J.; Hilsenrath, E.; Haffner, D. P.; Bhartia, P. K.

    2014-12-01

    Ozone distributions in the photochemically-controlled upper stratosphere and mesosphere were first measured using spectrometers on V-2 rockets after WWII. The IGY(1957-1958) spurred development of new optical and chemical instruments for flight on meteorological and sounding rockets. In the early 1960's, the US Navy developed an Arcas rocket-borne optical ozonesonde and NASA GSFC developed chemiluminescent ozonesonde onboard Nike_Cajun and Arcas rocket. The Navy optical ozone program was moved in 1969 to GSFC where rocket ozone research was expanded and continued until 1994 using Super Loki-Dart rocket at 11 sites in the range of 0-65N and 35W-160W. Over 300 optical ozone soundings and 40 chemiluminescent soundings were made. The data have been used to produce the US Standard Ozone Atmosphere, determine seasonal and diurnal variations, and validate early photochemical models. The current effort includes soundings conducted by Australia, Japan, and Korea using optical techniques. New satellite ozone sounding techniques were initially calibrated and later validated using the rocket ozone data. As satellite techniques superseded the rocket methods, the sponsoring agencies lost interest in the data and many of those records have been discarded. The current task intends to recover as much of the data as possible from the private records of the experimenters and their publications, and to archive those records in the WOUDC (World Ozone and Ultraviolet Data Centre). The original data records are handwritten tabulations, computer printouts that are scanned with OCR techniques, and plots digitized from publications. This newly recovered digital rocket ozone profile data from 1965 to 2002 could make significant contributions to the Earth science community in atmospheric research including long-term trend analysis.

  2. VSB-30 sounding rocket: history of flight performance

    Directory of Open Access Journals (Sweden)

    Wolfgang Jung

    2011-09-01

    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.

  3. Hydrocarbon Rocket Technology Impact Forecasting

    Science.gov (United States)

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

    2012-01-01

    Forecasting method is a normative forecasting technique that allows the designer to quantify the effects of adding new technologies on a given design. This method can be used to assess and identify the necessary technological improvements needed to close the gap that exists between the current design and one that satisfies all constraints imposed on the design. The TIF methodology allows for more design knowledge to be brought to the earlier phases of the design process, making use of tools such as Quality Function Deployments, Morphological Matrices, Response Surface Methodology, and Monte Carlo Simulations.2 This increased knowledge allows for more informed decisions to be made earlier in the design process, resulting in shortened design cycle time. This paper will investigate applying the TIF method, which has been widely used in aircraft applications, to the conceptual design of a hydrocarbon rocket engine. In order to reinstate a manned presence in space, the U.S. must develop an affordable and sustainable launch capability. Hydrocarbon-fueled rockets have drawn interest from numerous major government and commercial entities because they offer a low-cost heavy-lift option that would allow for frequent launches1. However, the development of effective new hydrocarbon rockets would likely require new technologies in order to overcome certain design constraints. The use of advanced design methods, such as the TIF method, enables the designer to identify key areas in need of improvement, allowing one to dial in a proposed technology and assess its impact on the system. Through analyses such as this one, a conceptual design for a hydrocarbon-fueled vehicle that meets all imposed requirements can be achieved.

  4. Motor homopolar

    OpenAIRE

    Martín Muñoz, Agustín

    2007-01-01

    Mostramos la construcción de un modelo de motor homopolar, uno de los más antiguos tipos de motores eléctricos. Se caracterizan porque el campo magnético del imán mantiene siempre la misma polaridad (de ahí su nombre, del griego homos, igual), de modo que, cuando una corriente eléctrica atraviesa el campo magnético, aparece una fuerza que hace girar los elementos no fijados mecánicamente. En el sencillísimo motor homopolar colgado (Schlichting y Ucke 2004), el imán puede girar ...

  5. The Spanish national programme of balloons and sounding rockets

    International Nuclear Information System (INIS)

    Casas, J.; Pueyo, L.

    1978-01-01

    The main points of the Spanish scientific programme are briefly described: CONIE/NASA cooperative project on meteorological sounding rocket launchings; ozonospheric programme; CONIE/NASA/CNES cooperative ionospheric sounding rocket project; D-layer research; rocket infrared dayglow measurements; ultraviolet astronomy research; cosmic ray research. The schedule of sounding rocket launchings at El Arenosillo station during 1977 is given

  6. RX LAPAN Rocket data Program With Dbase III Plus

    International Nuclear Information System (INIS)

    Sauman

    2001-01-01

    The components data rocket RX LAPAN are taken from workshop product and assembling rocket RX. In this application software, the test data are organized into two data files, i.e. test file and rocket file. Besides [providing facilities to add, edit and delete data, this software provides also data manipulation facility to support analysis and identification of rocket RX failures and success

  7. 16 CFR 1507.10 - Rockets with sticks.

    Science.gov (United States)

    2010-01-01

    ... 16 Commercial Practices 2 2010-01-01 2010-01-01 false Rockets with sticks. 1507.10 Section 1507.10... FIREWORKS DEVICES § 1507.10 Rockets with sticks. Rockets with sticks (including skyrockets and bottle rockets) shall utilize a straight and rigid stick to provide a direct and stable flight. Such sticks shall...

  8. Alternate Propellant Thermal Rocket, Phase I

    Data.gov (United States)

    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...

  9. The electromagnetic rocket gun impact fusion driver

    International Nuclear Information System (INIS)

    Winterberg, F.

    1984-01-01

    A macroparticle accelerator to be used as an impact fusion driver is discussed and which can accelerate a small projectile to --200 km/sec over a distance of a few 100 meters. The driver which we have named electromagnetic rocket gun, accelerates a small rocket-like projectile by a travelling magnetic wave. The rocket propellant not only serves as a sink to absorb the heat produced in the projectile by resistive energy losses, but at the same time is also the source of additional thrust through the heating of the propellant to high temperatures by the travelling magnetic wave. The total thrust on the projectile is the sum of the magnetic and recoil forces. In comparison to a rocket, the efficiency is here much larger, with the momentum transferred to the gun barrel of the gun rather than to a tenuous jet. (author)

  10. Ceremony celebrates 50 years of rocket launches

    Science.gov (United States)

    2000-01-01

    Ceremony celebrates 50 years of rocket launches PL00C-10364.12 At the 50th anniversary ceremony celebrating the first rocket launch from pad 3 on what is now Cape Canaveral Air Force Station, Norris Gray waves to the audience. Gray was part of the team who successfully launched the first rocket, known as Bumper 8. The ceremony was hosted by the Air Force Space & Missile Museum Foundation, Inc. , and included launch of a Bumper 8 model rocket, presentation of a Bumper Award to Florida Sen. George Kirkpatrick by the National Space Club; plus remarks by Sen. Kirkpatrick, KSC's Center Director Roy Bridges, and the Commander of the 45th Space Wing, Brig. Gen. Donald Pettit. Also attending the ceremony were other members of the original Bumper 8 team. A reception followed at Hangar C. Since 1950 there have been a total of 3,245 launches from Cape Canaveral.

  11. Fundamentals of aircraft and rocket propulsion

    CERN Document Server

    El-Sayed, Ahmed F

    2016-01-01

    This book provides a comprehensive basics-to-advanced course in an aero-thermal science vital to the design of engines for either type of craft. The text classifies engines powering aircraft and single/multi-stage rockets, and derives performance parameters for both from basic aerodynamics and thermodynamics laws. Each type of engine is analyzed for optimum performance goals, and mission-appropriate engines selection is explained. Fundamentals of Aircraft and Rocket Propulsion provides information about and analyses of: thermodynamic cycles of shaft engines (piston, turboprop, turboshaft and propfan); jet engines (pulsejet, pulse detonation engine, ramjet, scramjet, turbojet and turbofan); chemical and non-chemical rocket engines; conceptual design of modular rocket engines (combustor, nozzle and turbopumps); and conceptual design of different modules of aero-engines in their design and off-design state. Aimed at graduate and final-year undergraduate students, this textbook provides a thorough grounding in th...

  12. Space Power Experiments Aboard Rockets SPEAR-3

    National Research Council Canada - National Science Library

    Raitt, W. J

    1997-01-01

    The SPEAR-3 program was a sounding rocket payload designed to study the interaction of a charged body with the Earth's upper atmosphere with particular reference to the discharging ability of selected...

  13. Application of stepping motor

    International Nuclear Information System (INIS)

    1980-10-01

    This book is divided into three parts, which is about practical using of stepping motor. The first part has six chapters. The contents of the first part are about stepping motor, classification of stepping motor, basic theory og stepping motor, characteristic and basic words, types and characteristic of stepping motor in hybrid type and basic control of stepping motor. The second part deals with application of stepping motor with hardware of stepping motor control, stepping motor control by microcomputer and software of stepping motor control. The last part mentions choice of stepping motor system, examples of stepping motor, measurement of stepping motor and practical cases of application of stepping motor.

  14. NASA Space Rocket Logistics Challenges

    Science.gov (United States)

    Neeley, James R.; Jones, James V.; Watson, Michael D.; Bramon, Christopher J.; Inman, Sharon K.; Tuttle, Loraine

    2014-01-01

    The Space Launch System (SLS) is the new NASA heavy lift launch vehicle and is scheduled for its first mission in 2017. The goal of the first mission, which will be uncrewed, is to demonstrate the integrated system performance of the SLS rocket and spacecraft before a crewed flight in 2021. SLS has many of the same logistics challenges as any other large scale program. Common logistics concerns for SLS include integration of discreet programs geographically separated, multiple prime contractors with distinct and different goals, schedule pressures and funding constraints. However, SLS also faces unique challenges. The new program is a confluence of new hardware and heritage, with heritage hardware constituting seventy-five percent of the program. This unique approach to design makes logistics concerns such as commonality especially problematic. Additionally, a very low manifest rate of one flight every four years makes logistics comparatively expensive. That, along with the SLS architecture being developed using a block upgrade evolutionary approach, exacerbates long-range planning for supportability considerations. These common and unique logistics challenges must be clearly identified and tackled to allow SLS to have a successful program. This paper will address the common and unique challenges facing the SLS programs, along with the analysis and decisions the NASA Logistics engineers are making to mitigate the threats posed by each.

  15. Two stage turbine for rockets

    Science.gov (United States)

    Veres, Joseph P.

    1993-01-01

    The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The advanced turbine airfoils were designed and tested by Pratt & Whitney. The technology represented by this turbine is being developed for a turbopump to be used in an advanced upper stage rocket engine. The advanced engine will use a hydrogen expander cycle and achieve high performance through efficient combustion of hydrogen/oxygen propellants, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low mass flow rates and high operating pressures result in very small airfoil heights and diameters. The high efficiency and small size requirements present a challenging turbine design problem. The shrouded axial turbine blades are 50 percent reaction with a maximum thickness to chord ratio near 1. At 6 deg from the tangential direction, the nozzle and blade exit flow angles are well below the traditional design minimum limits. The blade turning angle of 160 deg also exceeds the maximum limits used in traditional turbine designs.

  16. Focused RBCC Experiments: Two-Rocket Configuration Experiments and Hydrocarbon/Oxygen Rocket Ejector Experiments

    Science.gov (United States)

    Santoro, Robert J.; Pal, Sibtosh

    2003-01-01

    This addendum report documents the results of two additional efforts for the Rocket Based Combined Cycle (RBCC) rocket-ejector mode research work carried out at the Penn State Propulsion Engineering Research Center in support of NASA s technology development efforts for enabling 3 d generation Reusable Launch Vehicles (RLV). The tasks reported here build on an earlier NASA MSFC funded research program on rocket ejector investigations. The first task investigated the improvements of a gaseous hydrogen/oxygen twin thruster RBCC rocket ejector system over a single rocket system. The second task investigated the performance of a hydrocarbon (liquid JP-7)/gaseous oxygen single thruster rocket-ejector system. To gain a systematic understanding of the rocket-ejector s internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static diffusion and afterburning (DAB) configurations for a range of rocket operating conditions. For all experimental conditions, overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust. Detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (gaseous oxygen, hydrogen, nitrogen and water vapor) for the gaseous hydrogen/oxygen rocket ejector experiments.

  17. The UK sounding rocket and balloon programme

    International Nuclear Information System (INIS)

    Delury, J.T.

    1980-01-01

    The UK civil science balloon and rocket programmes for 1979/80/81 are summarised and the areas of scientific interest for the period 1981/85 mentioned. In the main the facilities available are 10 in number balloons up to 40 m cu ft launched from USA or Australia and up to 10 in number 7 1/2'' diameter Petrel rockets. This paper outlines the 1979 and 1980 programmes and explains the longer term plans covering the next 5 years. (Auth.)

  18. Ramjet Application Possibilities for Increasing Fire Range of the Multiple Launch Rocket Systems Ammunition

    Directory of Open Access Journals (Sweden)

    V. N. Zubov

    2015-01-01

    Full Text Available The article considers a possibility to increase a flying range of the perspective rockets equipped with the control unit with aerodynamic controllers for the multiple launch rocket systems “Smerch”.To increase a flying range and reduce a starting mass of the rocket, the paper studies a possibility to replace the single-mode rocket engine used in the solid-fuel rocket motor for the direct-flow propulsion jet engine (DFPJE with not head sector air intakes. The DFPJE is implemented according to the classical scheme with a fuel charged in the combustion chamber. A separated solid propellant starting accelerator provides the rocket acceleration to reach a speed necessary for the DFPJE to run.When designing the DFPJE a proper choice of not head air intake parameters is one of the most difficult points. For this purpose a COSMOS Flow Simulation software package and analytical dependences were used to define the following: a boundary layer thickness where an air intake is set, maximum permissible and appropriate angles of attack and deviation angles of controllers at the section where the DFPJE works, and some other parameters as well.Calculation of DFPJE characteristics consisted in determining parameters of an air-gas path of the propulsion system, geometrical sizes of the pipeline flow area, sizes of a fuel charge, and dependence of the propulsion system impulse on the flight height and speed. Calculations were performed both in thermodynamic statement of problem and in using software package of COSMOS Flow Simulation.As a result of calculations and design engineering activities the air intake profile is created and mass-dimensional characteristics of DFPJE are defined. Besides, calculations of the starting solid fuel accelerator were carried out. Further design allowed us to create the rocket shape, estimate its mass-dimensional characteristics, and perform ballistic calculations, which proved that achieving a range of 120 km for the rocket is

  19. Nitrous Oxide/Paraffin Hybrid Rocket Engines

    Science.gov (United States)

    Zubrin, Robert; Snyder, Gary

    2010-01-01

    Nitrous oxide/paraffin (N2OP) hybrid rocket engines have been invented as alternatives to other rocket engines especially those that burn granular, rubbery solid fuels consisting largely of hydroxyl- terminated polybutadiene (HTPB). Originally intended for use in launching spacecraft, these engines would also be suitable for terrestrial use in rocket-assisted takeoff of small airplanes. The main novel features of these engines are (1) the use of reinforced paraffin as the fuel and (2) the use of nitrous oxide as the oxidizer. Hybrid (solid-fuel/fluid-oxidizer) rocket engines offer advantages of safety and simplicity over fluid-bipropellant (fluid-fuel/fluid-oxidizer) rocket en - gines, but the thrusts of HTPB-based hybrid rocket engines are limited by the low regression rates of the fuel grains. Paraffin used as a solid fuel has a regression rate about 4 times that of HTPB, but pure paraffin fuel grains soften when heated; hence, paraffin fuel grains can, potentially, slump during firing. In a hybrid engine of the present type, the paraffin is molded into a 3-volume-percent graphite sponge or similar carbon matrix, which supports the paraffin against slumping during firing. In addition, because the carbon matrix material burns along with the paraffin, engine performance is not appreciably degraded by use of the matrix.

  20. Internal Flow Simulation of Enhanced Performance Solid Rocket Booster for the Space Transportation System

    Science.gov (United States)

    Ahmad, Rashid A.; McCool, Alex (Technical Monitor)

    2001-01-01

    An enhanced performance solid rocket booster concept for the space shuttle system has been proposed. The concept booster will have strong commonality with the existing, proven, reliable four-segment Space Shuttle Reusable Solid Rocket Motors (RSRM) with individual component design (nozzle, insulator, etc.) optimized for a five-segment configuration. Increased performance is desirable to further enhance safety/reliability and/or increase payload capability. Performance increase will be achieved by adding a fifth propellant segment to the current four-segment booster and opening the throat to accommodate the increased mass flow while maintaining current pressure levels. One development concept under consideration is the static test of a "standard" RSRM with a fifth propellant segment inserted and appropriate minimum motor modifications. Feasibility studies are being conducted to assess the potential for any significant departure in component performance/loading from the well-characterized RSRM. An area of concern is the aft motor (submerged nozzle inlet, aft dome, etc.) where the altered internal flow resulting from the performance enhancing features (25% increase in mass flow rate, higher Mach numbers, modified subsonic nozzle contour) may result in increased component erosion and char. To assess this issue and to define the minimum design changes required to successfully static test a fifth segment RSRM engineering test motor, internal flow studies have been initiated. Internal aero-thermal environments were quantified in terms of conventional convective heating and discrete phase alumina particle impact/concentration and accretion calculations via Computational Fluid Dynamics (CFD) simulation. Two sets of comparative CFD simulations of the RSRM and the five-segment (IBM) concept motor were conducted with CFD commercial code FLUENT. The first simulation involved a two-dimensional axi-symmetric model of the full motor, initial grain RSRM. The second set of analyses

  1. 12-inch x-ray image intensifier with thin metal input window

    Energy Technology Data Exchange (ETDEWEB)

    Obata, Yoshiharu; Anno, Hidero; Harao, Norio [Toshiba Corp., Kawasaki, Kanagawa (Japan)

    1982-08-01

    Borosilicate glass has been used for X-ray input window of image intensifiers (I.I.) up to now. Now two new types of 12-inch metal I.I., RT12301C and RT12302C, have been developed. They use convex 1-mm aluminum (instead of 5-mm borosilicate glass) for the input window. Adopting a high-performance penta-electronic lens and a new type of light guide CsI film, these intensifiers have greatly improved contrast, quantum detection efficiency (QDE) and resolution capability. In spite of low dosage, image quality equivalent to that in the conventional direct radiograph is obtained through combined use of the new-type 12-inch metal I.I. with 0.3-mm small-focal-spot X-ray tube. Great contribution to digital radiography is expected of this I.I.

  2. Health physics challenges involved with opening a "seventeen-inch" concrete waste vault.

    Science.gov (United States)

    Sullivan, Patrick T; Pizzulli, Michelle

    2005-05-01

    This paper describes the various activities involved with opening a sealed legacy "Seventeen-inch" concrete vault and the health physics challenges and solutions employed. As part of a legacy waste stream that was removed from the former Hazardous Waste Management Facility at Brookhaven National Laboratory, the "Seventeen-inch" concrete vault labeled 1-95 was moved to the new Waste Management Facility for ultimate disposal. Because the vault contained 239Pu foils with a total activity in excess of the transuranic waste limits, the foils needed to be removed and repackaged for disposal. Conventional diamond wire saws could not be used because of facility constraints, so this project relied mainly on manual techniques. The planning and engineering controls put in place enabled personnel to open the vault and remove the waste while keeping dose as low as reasonably achievable.

  3. Fabrication of 0.5-inch diameter FBR mixed oxide fuel pellets

    International Nuclear Information System (INIS)

    Rasmussen, D.E.; Benecke, M.W.; McCord, R.B.

    1979-01-01

    Large diameter (0.535 inch) mixed oxide fuel pellets for Fast Breeder Reactor application were successfully fabricated by the cold-press-and-sinter technique. Enriched UO 2 , PuO 2 -UO 2 , and PuO 2 -ThO 2 compositions were fabricated into nominally 90% theoretical density pellets for the UO 2 and PuO 2 -UO 2 compositions, and 88% and 93% T.D. for the PuO 2 -ThO 2 compositions. Some processing adjustments were required to achieve satisfactory pellet quality and density. Furnace heating rate was reduced from 200 to 50 0 C/h for the organic binder burnout cycle for the large, 0.535-inch diameter pellets to eliminate pellet cracking during sintering. Additional preslugging steps and die wall lubrication during pressing were used to eliminate pressing cracks in the PuO 2 -ThO 2 pellets

  4. Design of a 18F production system at ORNL 86-inch cyclotron

    International Nuclear Information System (INIS)

    Shaeffer, M.C.; Barreto, F.; Datesh, J.R.; Goldstein, B.R.

    1977-01-01

    A target system for the production of 18 F by proton bombardment of H 2 18 O was designed for the ORNL 86-inch cyclotron facility. The system consists of concentric titanium and aluminum cylinders. Oxygen-18-enriched H 2 O circulates through the inner titanium cylinder and through an external heat exchanger with cooling water flowing in the annulus. Yields of 5.0 curies are expected for a 250-μA proton beam current and 24-min irradiation time

  5. IMPACT TESTING OF MATERIALS USING AN EIGHT-INCH AIR GUN AND COMPUTER REDUCTION OF DATA

    Energy Technology Data Exchange (ETDEWEB)

    Thorne, L. F.

    1973-10-01

    A mechanical shock actuator has been converted into an air gun capable of firing 8-inch-·diameter (20.32 cm) projectiles to velocities exceeding 1000 fps (304.8 m/ s). This new capability has been used to study the effect of impact velocity upon the energy.absorbed by crushable materials. Shockpulse data is reduced by computer techniques and test results are displayed in either tabular or graphic format by use of the C DC 6600 Calcomp plotter.

  6. Child seat belt guidelines: Examining the 4 feet 9 inches rule as the standard.

    Science.gov (United States)

    Morse, Amber M; Aitken, Mary E; Mullins, Samantha H; Miller, Beverly K; Pomtree, Mindy M; Ulloa, Erin M; Montgomery, Jeffrey S; Saylors, Marie E

    2017-11-01

    Current American Academy of Pediatrics recommendations regarding transition from child safety/booster seat to adult safety belt use indicate that children should be at least 4 feet 9 inches, 8 years old, or 80 pounds. Proper fit in the vehicle seat, assessed with a five-point fit test, should also be met. Although most children reach 4 feet 9 inches around age 8 years, each child and vehicle presents a unique combination; thus a child may not fit appropriately in all vehicle types using only the 4 feet 9 inches requirement. We enrolled children, aged 7 years to 12 years, into our study. Height, weight, and demographic data were obtained. A Child Passenger Safety Technician then performed the five-point fit test in each of a uniform lineup of five vehicles. Data were collected on fit in the standard vehicle seat and also in a booster seat. We set 90% as the threshold proportion of children who meet all criteria for proper fit to validate current recommendations of a height of 4 feet 9 inches. Data were collected on 388 children. The percentage of 90% proper fit was met in the compact car and small sport-utility vehicle (SUV). However, only 80 (77%) of 104 students (p guidelines for an adult seat belt do not meet safety requirements for fit, especially in larger, commonly used vehicles (large SUVs and trucks). This emphasizes the need for evaluation of fit by a trained personnel and/or development of standard back seat dimensions in all vehicles for maximum safety. Epidemiologic level 1.

  7. Heat Exchanger Design and Testing for a 6-Inch Rotating Detonation Engine

    Science.gov (United States)

    2013-03-01

    water-cooled rotating detonation engine ( RDE ) run on hydrogen and air. The change in water temperature as it cooled the engine was used to find the...a quick-response resistance temperature detector (RTD) was used in an uncooled RDE of similar dimension to the cooled RDE to estimate the transient...double-checking my design calculations, providing his experience with cooling the 3-inch RDE , and for providing technical expertise in regard to the

  8. Worcester 1 Inch Solenoid-Actuated Gas Operated SCHe System Valves

    International Nuclear Information System (INIS)

    VAN KATWIJK, C.

    2000-01-01

    1 inch Gas-operated full-port ball valves incorporate a solenoid and limit switches as integral parts of the actuator. These valves are normally open and fail safe to the open position (GOV-1*02 and 1*06 fail closed) to provide a flow path of helium gas to the MCO under helium purge and off-normal conditions when the MCO is isolated

  9. Plasma waves observed by sounding rockets

    International Nuclear Information System (INIS)

    Kimura, I.

    1977-01-01

    Observations of plasma wave phenomena have been conducted with several rockets launched at Kagoshima Space Center, Kyushu, Japan, and at Showa Base, Antarctica. This report presents some results of the observations in anticipation of having valuable comments from other plasma physicists, especially from those who are concerned with laboratory plasma. In the K-9M-41 rocket experiment, VLF plasma waves were observed. In this experiment, the electron beam of several tens of uA was emitted from a hot cathode when a positive dc bias changing from 0 to 10V at 1V interval each second was applied to a receiving dipole antenna. The discrete emissions with 'U' shaped frequency spectrum were observed for the dc bias over 3 volts. The U emissions appeared twice per spin period of the rocket. Similar rocket experiment was performed at Showa Base using a loop and dipole antenna and without hot cathode. Emissions were observed with varying conditions. At present, the authors postulate that such emissions may be produced just in the vicinity of a rocket due to a kind of wake effect. (Aoki, K.)

  10. Laser-fusion rocket for interplanetary propulsion

    International Nuclear Information System (INIS)

    Hyde, R.A.

    1983-01-01

    A rocket powered by fusion microexplosions is well suited for quick interplanetary travel. Fusion pellets are sequentially injected into a magnetic thrust chamber. There, focused energy from a fusion Driver is used to implode and ignite them. Upon exploding, the plasma debris expands into the surrounding magnetic field and is redirected by it, producing thrust. This paper discusses the desired features and operation of the fusion pellet, its Driver, and magnetic thrust chamber. A rocket design is presented which uses slightly tritium-enriched deuterium as the fusion fuel, a high temperature KrF laser as the Driver, and a thrust chamber consisting of a single superconducting current loop protected from the pellet by a radiation shield. This rocket can be operated with a power-to-mass ratio of 110 W gm -1 , which permits missions ranging from occasional 9 day VIP service to Mars, to routine 1 year, 1500 ton, Plutonian cargo runs

  11. Development of nuclear rocket engine technology

    International Nuclear Information System (INIS)

    Gunn, S.V.

    1989-01-01

    Research sponsored by the Atomic Energy Commission, the USAF, and NASA (later on) in the area of nuclear rocket propulsion is discussed. It was found that a graphite reactor, loaded with highly concentrated Uranium 235, can be used to heat high pressure liquid hydrogen to temperatures of about 4500 R, and to expand the hydrogen through a high expansion ratio rocket nozzle assembly. The results of 20 reactor tests conducted at the Nevada Test Site between July 1959 and June 1969 are analyzed. On the basis of these results, the feasibility of solid graphite reactor/nuclear rocket engines is revealed. It is maintained that this technology will support future space propulsion requirements, using liquid hydrogen as the propellant, for thrust requirements ranging from 25,000 lbs to 250,000 lbs, with vacuum specific impulses of at least 850 sec and with full engine throttle capability. 12 refs

  12. Metallic Hydrogen: A Game Changing Rocket Propellant

    Science.gov (United States)

    Silvera, Isaac F.

    2016-01-01

    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.

  13. Structural and mechanical design challenges of space shuttle solid rocket boosters separation and recovery subsystems

    Science.gov (United States)

    Woodis, W. R.; Runkle, R. E.

    1985-01-01

    The design of the space shuttle solid rocket booster (SRB) subsystems for reuse posed some unique and challenging design considerations. The separation of the SRBs from the cluster (orbiter and external tank) at 150,000 ft when the orbiter engines are running at full thrust meant the two SRBs had to have positive separation forces pushing them away. At the same instant, the large attachments that had reacted launch loads of 7.5 million pounds thrust had to be servered. These design considerations dictated the design requirements for the pyrotechnics and separation rocket motors. The recovery and reuse of the two SRBs meant they had to be safely lowered to the ocean, remain afloat, and be owed back to shore. In general, both the pyrotechnic and recovery subsystems have met or exceeded design requirements. In twelve vehicles, there has only been one instance where the pyrotechnic system has failed to function properly.

  14. Additive Manufacturing for Affordable Rocket Engines

    Science.gov (United States)

    West, Brian; Robertson, Elizabeth; Osborne, Robin; Calvert, Marty

    2016-01-01

    Additive manufacturing (also known as 3D printing) technology has the potential to drastically reduce costs and lead times associated with the development of complex liquid rocket engine systems. NASA is using 3D printing to manufacture rocket engine components including augmented spark igniters, injectors, turbopumps, and valves. NASA is advancing the process to certify these components for flight. Success Story: MSFC has been developing rocket 3D-printing technology using the Selective Laser Melting (SLM) process. Over the last several years, NASA has built and tested several injectors and combustion chambers. Recently, MSFC has 3D printed an augmented spark igniter for potential use the RS-25 engines that will be used on the Space Launch System. The new design is expected to reduce the cost of the igniter by a factor of four. MSFC has also 3D printed and tested a liquid hydrogen turbopump for potential use on an Upper Stage Engine. Additive manufacturing of the turbopump resulted in a 45% part count reduction. To understanding how the 3D printed parts perform and to certify them for flight, MSFC built a breadboard liquid rocket engine using additive manufactured components including injectors, turbomachinery, and valves. The liquid rocket engine was tested seven times in 2016 using liquid oxygen and liquid hydrogen. In addition to exposing the hardware to harsh environments, engineers learned to design for the new manufacturing technique, taking advantage of its capabilities and gaining awareness of its limitations. Benefit: The 3D-printing technology promises reduced cost and schedule for rocket engines. Cost is a function of complexity, and the most complicated features provide the largest opportunities for cost reductions. This is especially true where brazes or welds can be eliminated. The drastic reduction in part count achievable with 3D printing creates a waterfall effect that reduces the number of processes and drawings, decreases the amount of touch

  15. Annular Internal-External-Expansion Rocket Nozzles for Large Booster Applications

    Science.gov (United States)

    Connors, James F.; Cubbison, Robert W.; Mitchell, Glenn A.

    1961-01-01

    For large-thrust booster applications, annular rocket nozzles employing both internal and external expansion are investigated. In these nozzles, free-stream air flows through the center as well as around the outside of the exiting jet. Flaps for deflecting the rocket exhaust are incorporated on the external-expansion surface for thrust-vector control. In order to define nozzle off-design performance, thrust vectoring effectiveness, and external stream effects, an experimental investigation was conducted on two annular nozzles with area ratios of 15 and 25 at Mach 0, 2, and 3 in the Lewis 10- by 10-foot wind tunnel. Air, pressurized to 600 pounds per square inch absolute, was used to simulate the exhaust flow. For a nozzle-pressure-ratio range of 40 to 1000, the ratio of actual to ideal thrust was essentially constant at 0.98 for both nozzles. Compared with conventional convergent-divergent configurations on hypothetical boost missions, the performance gains of the annular nozzle could yield significant orbital payload increases (possibly 8 to 17 percent). A single flap on the external-expansion surface of the area-ratio-25 annular nozzle produced a side force equal to 4 percent of the axial force with no measurable loss in axial thrust.

  16. Energy production using fission fragment rockets

    International Nuclear Information System (INIS)

    Chapline, G.; Matsuda, Y.

    1991-08-01

    Fission fragment rockets are nuclear reactors with a core consisting of thin fibers in a vacuum, and which use magnetic fields to extract the fission fragments from the reactor core. As an alternative to ordinary nuclear reactors, fission fragment rockets would have the following advantages: Approximately twice as efficient if one can directly convert the fission fragment energy into electricity; by reducing the buildup of a fission fragment inventory in the reactor one could avoid a Chernobyl type disaster; and collecting the fission fragments outside the reactor could simplify the waste disposal problem. 6 refs., 4 figs., 2 tabs

  17. Large Liquid Rocket Testing: Strategies and Challenges

    Science.gov (United States)

    Rahman, Shamim A.; Hebert, Bartt J.

    2005-01-01

    Rocket propulsion development is enabled by rigorous ground testing in order to mitigate the propulsion systems risks that are inherent in space flight. This is true for virtually all propulsive devices of a space vehicle including liquid and solid rocket propulsion, chemical and non-chemical propulsion, boost stage and in-space propulsion and so forth. In particular, large liquid rocket propulsion development and testing over the past five decades of human and robotic space flight has involved a combination of component-level testing and engine-level testing to first demonstrate that the propulsion devices were designed to meet the specified requirements for the Earth to Orbit launchers that they powered. This was followed by a vigorous test campaign to demonstrate the designed propulsion articles over the required operational envelope, and over robust margins, such that a sufficiently reliable propulsion system is delivered prior to first flight. It is possible that hundreds of tests, and on the order of a hundred thousand test seconds, are needed to achieve a high-reliability, flight-ready, liquid rocket engine system. This paper overviews aspects of earlier and recent experience of liquid rocket propulsion testing at NASA Stennis Space Center, where full scale flight engines and flight stages, as well as a significant amount of development testing has taken place in the past decade. The liquid rocket testing experience discussed includes testing of engine components (gas generators, preburners, thrust chambers, pumps, powerheads), as well as engine systems and complete stages. The number of tests, accumulated test seconds, and years of test stand occupancy needed to meet varying test objectives, will be selectively discussed and compared for the wide variety of ground test work that has been conducted at Stennis for subscale and full scale liquid rocket devices. Since rocket propulsion is a crucial long-lead element of any space system acquisition or

  18. [Psychomotor reaction in primates placed in ballistic flight in rockets].

    Science.gov (United States)

    Grandpierre, R; Chatelier, G

    1968-01-01

    Psychomotor reactions have been studied in Nemestrina Macaques during flights in rockets in which there were periods of weightlessness lasting about 10 minutes. The animals had undergone motor conditioning and responded to a light signal by pressing on a button placed on a panel where there were four other undifferentiated buttons. The lighting of the panel and pressing on the different buttons in reply was recorded with respect to time. Movement was observed by film recording, and electrical activity of the brain was kept under observation by six different recordings of the cortex. Activity of stretching and bending muscles of the legs was recorded, as well as breathing, and heart activity was recorded on an electrocardiogram. These psychological parameters were transmitted to the earth by long-distance measurement during the entire flight. The recordings made it possible to confirm that the absence of the sensation of weight seems to reduce the alertness of the animals. This fact had already been observed in Wistar's rat and cat. The sensor-psychomotor responses undergo important individual variations and depend on the motivation of the subjects. A very well motivated one reacted after a short period of adaption, as it did in the laboratory: one that was much less motivated would not work until after weightlessness has ended, and showed signs of drowsiness. These observations made during the flights in rockets confirm those of Chambers et al., Harris et al., Lilly et al. made in immersion studies; they show the important role of weight in maintaining alertness, but that changes of psychomotor activity are particularly related to motivation.

  19. Vibration control of uncertain multiple launch rocket system using radial basis function neural network

    Science.gov (United States)

    Li, Bo; Rui, Xiaoting

    2018-01-01

    Poor dispersion characteristics of rockets due to the vibration of Multiple Launch Rocket System (MLRS) have always restricted the MLRS development for several decades. Vibration control is a key technique to improve the dispersion characteristics of rockets. For a mechanical system such as MLRS, the major difficulty in designing an appropriate control strategy that can achieve the desired vibration control performance is to guarantee the robustness and stability of the control system under the occurrence of uncertainties and nonlinearities. To approach this problem, a computed torque controller integrated with a radial basis function neural network is proposed to achieve the high-precision vibration control for MLRS. In this paper, the vibration response of a computed torque controlled MLRS is described. The azimuth and elevation mechanisms of the MLRS are driven by permanent magnet synchronous motors and supposed to be rigid. First, the dynamic model of motor-mechanism coupling system is established using Lagrange method and field-oriented control theory. Then, in order to deal with the nonlinearities, a computed torque controller is designed to control the vibration of the MLRS when it is firing a salvo of rockets. Furthermore, to compensate for the lumped uncertainty due to parametric variations and un-modeled dynamics in the design of the computed torque controller, a radial basis function neural network estimator is developed to adapt the uncertainty based on Lyapunov stability theory. Finally, the simulated results demonstrate the effectiveness of the proposed control system and show that the proposed controller is robust with regard to the uncertainty.

  20. Nanomanipulation of 2 inch wafer fabrication of vertically aligned carbon nanotube arrays by nanoimprint lithography

    DEFF Research Database (Denmark)

    Bu, Ian Y. Y.; Eichhorn, Volkmar; Carlson, Kenneth

    2011-01-01

    Carbon nanotube (CNT) arrays are typically defined by electron beam lithography (EBL), and hence limited to small areas due to the low throughput. To obtain wafer‐scale fabrication we propose large area thermal nanoimprint lithography (NIL). A 2‐inch stamp master is defined using EBL for subsequent......, efficient production of wafer‐scale/larger arrays of CNTs has been achieved. The CNTs have been deposited by wafer‐scale plasma enhanced chemical vapour deposition (PECVD) of C2H2/NH3. Substrates containing such nanotubes have been used to automate nanorobotic manipulation sequences of individual CNTs...

  1. First thin AC-coupled silicon strip sensors on 8-inch wafers

    Energy Technology Data Exchange (ETDEWEB)

    Bergauer, T., E-mail: thomas.bergauer@oeaw.ac.at [Institute of High Energy Physics of the Austrian Academy of Sciences, Nikolsdorfer Gasse 18, 1050 Wien (Vienna) (Austria); Dragicevic, M.; König, A. [Institute of High Energy Physics of the Austrian Academy of Sciences, Nikolsdorfer Gasse 18, 1050 Wien (Vienna) (Austria); Hacker, J.; Bartl, U. [Infineon Technologies Austria AG, Siemensstrasse 2, 9500 Villach (Austria)

    2016-09-11

    The Institute of High Energy Physics (HEPHY) in Vienna and the semiconductor manufacturer Infineon Technologies Austria AG developed a production process for planar AC-coupled silicon strip sensors manufactured on 200 μm thick 8-inch p-type wafers. In late 2015, the first wafers were delivered featuring the world's largest AC-coupled silicon strip sensors. Detailed electrical measurements were carried out at HEPHY, where single strip and global parameters were measured. Mechanical studies were conducted and the long-term behavior was investigated using a climate chamber. Furthermore, the electrical properties of various test structures were investigated to validate the quality of the manufacturing process.

  2. High spatial resolution gamma imaging detector based on a 5 inch diameter R3292 Hamamatsu PSPMT

    International Nuclear Information System (INIS)

    Wojcik, R.; Majewski, S.; Kross, B.; Weisenberger, A.G.; Steinbach, D.

    1998-01-01

    High resolution imaging gamma-ray detectors were developed using Hamamatsu's 5 inch diameter R3292 position sensitive PMT (PSPMT) and a variety of crystal scintillator arrays. Special readout techniques were used to maximize the active imaging area while reducing the number of readout channels. Spatial resolutions approaching 1 mm were obtained in a broad energy range from 20 to 511 keV. Results are also presented of coupling the scintillator arrays to the PMT via imaging light guides consisting of acrylic optical fibers

  3. The 88-Inch Cyclotron: A One-Stop Facility for Electronics Radiation and Detector Testing

    Energy Technology Data Exchange (ETDEWEB)

    Kireeff Covo, M.; Albright, R. A.; Ninemire, B. F.; Johnson, M. B.; Hodgkinson, A.; Loew, T.; Benitez, J. Y.; Todd, D. S.; Xie, D. Z.; Perry, T.; Phair, L.; Bernsteiny, L. A.; Bevins, J.; Brown, J. A.; Goldblum, B. L.; Harasty, M.; Harrig, K. P.; Laplace, T. A.; Matthews, E. F.; Bushmaker, A.; Walker, D.; Oklejas, V.; Hopkins, A. R.; Bleuel, D. L.; Chen, J.; Cronin, S. B.

    2017-10-01

    In outer space down to the altitudes routinely flown by larger aircrafts, radiation can pose serious issues for microelectronics circuits. The 88-Inch Cyclotron at Lawrence Berkeley National Laboratory is a sector-focused cyclotron and home of the Berkeley Accelerator Space Effects Facility, where the effects of energetic particles on sensitive microelectronics are studied with the goal of designing electronic systems for the space community. This paper describes the flexibility of the facility and its capabilities for testing the bombardment of electronics by heavy ions, light ions, and neutrons. Experimental capabilities for the generation of neutron beams from deuteron breakups and radiation testing of carbon nanotube field effect transistor will be discussed.

  4. Elemental Water Impact Test: Phase 2 36-Inch Aluminum Tank Head

    Science.gov (United States)

    Vassilakos, Gregory J.

    2014-01-01

    Spacecraft are being designed based on LS-DYNA simulations of water landing impacts. The Elemental Water Impact Test (EWIT) series was undertaken to assess the accuracy of LS-DYNA water impact simulations. EWIT Phase 2 featured a 36-inch aluminum tank head. The tank head was outfitted with one accelerometer, twelve pressure transducers, three string potentiometers, and four strain gages. The tank head was dropped from heights of 1 foot and 2 feet. The focus of this report is the correlation of analytical models against test data. As a measure of prediction accuracy, peak responses from the baseline LS-DYNA model were compared to peak responses from the tests.

  5. Project Waiver of American Iron and Steel Requirements to the Napa Sanitation District for 24-Inch Diameter Butterfly Valves

    Science.gov (United States)

    Waiver approval by EPA pursuant to the American Iron and Steel Requirements of the Clean Water Act Section 608 to the Napa Sanitation District in California for the purchase of 24-inch butterfly valves.

  6. Comparison of Engine Cycle Codes for Rocket-Based Combined Cycle Engines

    Science.gov (United States)

    Waltrup, Paul J.; Auslender, Aaron H.; Bradford, John E.; Carreiro, Louis R.; Gettinger, Christopher; Komar, D. R.; McDonald, J.; Snyder, Christopher A.

    2002-01-01

    This paper summarizes the results from a one day workshop on Rocket-Based Combined Cycle (RBCC) Engine Cycle Codes held in Monterey CA in November of 2000 at the 2000 JANNAF JPM with the authors as primary participants. The objectives of the workshop were to discuss and compare the merits of existing Rocket-Based Combined Cycle (RBCC) engine cycle codes being used by government and industry to predict RBCC engine performance and interpret experimental results. These merits included physical and chemical modeling, accuracy and user friendliness. The ultimate purpose of the workshop was to identify the best codes for analyzing RBCC engines and to document any potential shortcomings, not to demonstrate the merits or deficiencies of any particular engine design. Five cases representative of the operating regimes of typical RBCC engines were used as the basis of these comparisons. These included Mach 0 sea level static and Mach 1.0 and Mach 2.5 Air-Augmented-Rocket (AAR), Mach 4 subsonic combustion ramjet or dual-mode scramjet, and Mach 8 scramjet operating modes. Specification of a generic RBCC engine geometry and concomitant component operating efficiencies, bypass ratios, fuel/oxidizer/air equivalence ratios and flight dynamic pressures were provided. The engine included an air inlet, isolator duct, axial rocket motor/injector, axial wall fuel injectors, diverging combustor, and exit nozzle. Gaseous hydrogen was used as the fuel with the rocket portion of the system using a gaseous H2/O2 propellant system to avoid cryogenic issues. The results of the workshop, even after post-workshop adjudication of differences, were surprising. They showed that the codes predicted essentially the same performance at the Mach 0 and I conditions, but progressively diverged from a common value (for example, for fuel specific impulse, Isp) as the flight Mach number increased, with the largest differences at Mach 8. The example cases and results are compared and discussed in this paper.

  7. Jidosha's Motors

    OpenAIRE

    Shirakawa Okuma, Rosely; Calderón Orejuela, Javier

    2016-01-01

    La tesis narra la situación de una empresa concesionaria de vehículos nuevos, Jidosha's Motors, perteneciente a una corporación japonesa que cuenta con una cultura muy arraigada de ética y de cumplimiento. Se plantean respuestas, se identifican problemas y sus alternativas de solución para una toma adecuada de decisiones por parte de los directivos, siguiendo una estructura de análisis de situaciones de negocios (ASN). Tesis

  8. NASA rocket launches student project into space

    OpenAIRE

    Crumbley, Liz

    2005-01-01

    A project that began in 2002 will culminate at sunrise on Tuesday, March 15, when a team of Virginia Tech engineering students watch a payload section they designed lift off aboard a sounding rocket from a launch pad at NASA's Wallops Island Flight Facility and travel 59 miles into space.

  9. Straw Rockets Are out of This World

    Science.gov (United States)

    Gillman, Joan

    2013-01-01

    To capture students' excitement and engage their interest in rocketships and visiting planets in the solar system, the author designed lessons that give students the opportunity to experience the joys and challenges of developing straw rockets, and then observing which design can travel the longest distance. The lessons are appropriate for…

  10. Government Relations: It's Not Rocket Science

    Science.gov (United States)

    Radway, Mike

    2007-01-01

    Many people in the early childhood education field are afraid of government relations work, intimidated by politicians, and believe the whole process is unseemly. The author asserts that they should not be afraid nor be intimidated because government relations is not rocket science and fundamentally officeholders are no different from the rest of…

  11. Rocket Based Combined Cycle (RBCC) engine inlet

    Science.gov (United States)

    2004-01-01

    Pictured is a component of the Rocket Based Combined Cycle (RBCC) engine. This engine was designed to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsion systems and ultimately a Single Stage to Orbit (SSTO) air breathing propulsion system.

  12. Microcomputers, Model Rockets, and Race Cars.

    Science.gov (United States)

    Mirus, Edward A., Jr.

    1985-01-01

    The industrial education orientation program at Wisconsin School for the Deaf (WSD) presents problem-solving situations to all seventh- and eighth-grade hearing-impaired students. WSD developed user-friendly microcomputer software to guide students individually through complex computations involving model race cars and rockets while freeing…

  13. An Analysis of Rocket Propulsion Testing Costs

    Science.gov (United States)

    Ramirez, Carmen; Rahman, Shamim

    2010-01-01

    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.

  14. A new alignment procedure for the South African Astronomical Observatory's 74-inch telescope

    Science.gov (United States)

    Crause, Lisa A.; Booth, John A.; Doss, David; Loubser, Egan; O'Connor, James E.; Sass, Craig; Sickafoose, Amanda A.; Worters, Hannah L.

    2016-07-01

    Considerable effort has gone into improving the performance and reliability of the SAAO's 74-inch telescope. This included replacing the telescope encoders, refining the pointing model and increasing the telescope throughput. The latter involved re-aluminising the primary and formulating a procedure to ensure optimal alignment of the telescope mirrors. To this end, we developed the necessary hardware and techniques to ensure that such alignment is achieved and maintained, particularly following re-aluminising of the mirrors. In essence, the procedure involves: placing a Taylor Hobson Alignment Telescope on the mechanical rotation axis of the 74-inch (which we define to be the optical axis, since the Cassegrain instruments attach to the associated turntable), then adjusting the tip/tilt of the secondary mirror to get it onto that axis and, lastly, adjusting the tip/tilt of the primary mirror to eliminate coma. An eyepiece (or wavefront camera) is installed at the Cassegrain port for this final step since comatic star images indicate the need to tip/tilt the primary mirror to align it to the secondary. Tuning out any brightness gradients seen in an out-of-focus image of a bright star may also be used for feedback when adjusting the tip/tilt of the primary mirror to null coma.

  15. Structural reinforcement of a {theta}16 inches tee made during operation with composite material

    Energy Technology Data Exchange (ETDEWEB)

    Beim, Andre [Tresca Engenharia Ltda., Sao Paulo, SP (Brazil); Vilani, Eduardo Cesar [Rust Engenharia Ltda., Diadema, SP (Brazil)

    2009-07-01

    An industrial installation went through a turnaround to increase capacity. A tie-in line had to be made for this, and a 16 inches diameter branch was required to be made on an existing 16 inches pipe. The joining of these elements resulted in a 'tee' that was supposed to receive a reinforcement plate. This reinforcement plate was not installed before the plant start-up. Code calculations and a finite element stress analysis showed that reinforcement was necessary. The only viable solution was the application of a composite material reinforcement, designed to substitute the missing reinforcement plate, and reduce the stresses to allowable levels. A new finite element stress analysis was made to determine the required thickness of this reinforcement. The first part of this work shows the results of the finite element stress analysis. Figures with stress contours of the analyses show the results. The second part shows the details of the installation of the reinforcement, which was made during normal operation of the plant. Pictures illustrate the main steps of the installation procedure. (author)

  16. Production of exotic beams at the LBL 88-Inch Cyclotron by the ISOL method

    International Nuclear Information System (INIS)

    1990-04-01

    The Users of the LBL 88-Inch Cyclotron are preparing a proposal to produce exotic, i.e., radioactive beams. The facility will consist of a high-current 30 MeV cyclotron to generate the radioactive nuclei, an ECR source that can be coupled to different production targets, and the 88-Inch Cyclotron to accelerate the radioactive ions. Thus, the basic concept is that of the double cyclotron system pioneered at Louvain-la-Neuve, although the initial emphasis will be on producing a variety of light proton-rich beams at energies up to 10 MeV/A. At this workshop we wish to outline what is being planned, to invite comments and suggestions, and, especially, to encourage participation. We believe that this facility will be an important step toward establishing the scientific and technical basis for a National High Intensity Facility. This can be achieved through active participation by members of the radioactive beam (RB) community in (1) experiments with high quality radioactive beams of moderate intensity and, (2) R ampersand D on high beam-power targets and highly efficient ion sources. 5 refs., 4 figs

  17. Study on water leak-tightness of small leaks on a 1 inch cylinder valve

    International Nuclear Information System (INIS)

    Miyazawa, T.; Kasai, Y.; Inabe, N.; Aritomi, M.

    2002-01-01

    Practical thresholds for water leak-tightness of small leaks were determined by experimentation. Measurements for small leak samples were taken of air leakage rates and water leakage rates for identical leak samples in order to identify parameters that influence water leak-tightness threshold. Four types of leaks were evaluated: a fine wire inserted in an O-ring seal, a glass capillary tube, a stainless steel orifice, and a scratched valve stem on a 1 inch UF 6 cylinder valve. Experimental results demonstrated that the key parameter for water leak-tightness is the opening size of the leak hole. The maximum allowable hole size to achieve water leak-tightness ranged from 10 to 20 μm in diameter in this study. Experimental results with 1 inch UF 6 cylinder valve samples demonstrated that the acceptance criteria for preshipment leakage test, 1x10 -3 ref-cm 3 .s -1 , as prescribed in ANSI N14.5 is an appropriate value from the point of view of water leak-tightness for enriched UF 6 packages. The mechanism of water leak-tightness is plugging by tiny particles existing in water. The water used in experiments in this study contained far fewer particles than in water assumed to be encountered under accident conditions of transport. Therefore, the water leak-tightness threshold determined in this study is a conservative value in a practical evaluation. (author)

  18. Artist autonomy in a digital era: The case of Nine Inch Nails

    Directory of Open Access Journals (Sweden)

    Steven C. Brown

    2012-09-01

    Full Text Available A 2009 presentation by Michael Masnick (CEO and founder of insight company Floor64 entitled ‘How Trent Reznor and Nine Inch Nails represent the Future of the Music Business’ brought the success of the business models employed by Reznor in distributing Nine Inch Nails’ music into the spotlight. The present review provides a comprehensive timeline of the band circa 2005-2010, evaluating the success of the distribution methods employed in accordance with Masnick’s (2009 proposed business model of connecting with fans and providing them with a reason to buy. The model is conceptualised in the wider context in which Reznor’s distribution methods take place (including a brief consideration of Radiohead’s much cited pay-what-you- want model, addressing the perceived gaps in the model by exploring the involvement of musical preferences; age and consumer purchasing behavior and fan worship. Implications are discussed concerning the applicability of the model for new and emerging bands.

  19. Rocketing into the future the history and technology of rocket planes

    CERN Document Server

    van Pelt, Michel

    2012-01-01

    Rocketing into the Future journeys into the exciting world of rocket planes, examining the exotic concepts and actual flying vehicles that have been devised over the last one hundred years. Lavishly illustrated with over 150 photographs, it recounts the history of rocket planes from the early pioneers who attached simple rockets on to their wooden glider airplanes to the modern world of high-tech research vehicles. The book then looks at the possibilities for the future. The technological and economic challenges of the Space Shuttle proved insurmountable, and thus the program was unable to fulfill its promise of low-cost access to space. However, the burgeoning market of suborbital space tourism may yet give the necessary boost to the development of a truly reusable spaceplane.

  20. The Alabama Space and Rocket Center: The Second Decade.

    Science.gov (United States)

    Buckbee, Edward O.

    1983-01-01

    The Alabama Space and Rocket Center in Huntsville, the world's largest rocket and space museum, includes displays illustrating American rocket history, exhibits and demonstrations on rocketry principles and experiences, and simulations of space travel. A new project includes an integrated recreational-educational complex, described in the three…

  1. 14 CFR 437.67 - Tracking a reusable suborbital rocket.

    Science.gov (United States)

    2010-01-01

    ... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Tracking a reusable suborbital rocket. 437... a reusable suborbital rocket. A permittee must— (a) During permitted flight, measure in real time the position and velocity of its reusable suborbital rocket; and (b) Provide position and velocity...

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

    Directory of Open Access Journals (Sweden)

    Kalomba Mboyi

    2015-04-01

    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.

  3. Measurements of temperature profiles at the exit of small rockets.

    Science.gov (United States)

    Griggs, M; Harshbarger, F C

    1966-02-01

    The sodium line reversal technique was used to determine the reversal temperature profile across the exit of small rockets. Measurements were made on one 73-kg thrust rocket, and two 23-kg thrust rockets with different injectors. The large rocket showed little variation of reversal temperature across the plume. However, the 23-kg rockets both showed a large decrease of reversal temperature from the axis to the edge of the plume. In addition, the sodium line reversal technique of temperature measurement was compared with an infrared technique developed in these laboratories.

  4. Analysis of rocket flight stability based on optical image measurement

    Science.gov (United States)

    Cui, Shuhua; Liu, Junhu; Shen, Si; Wang, Min; Liu, Jun

    2018-02-01

    Based on the abundant optical image measurement data from the optical measurement information, this paper puts forward the method of evaluating the rocket flight stability performance by using the measurement data of the characteristics of the carrier rocket in imaging. On the basis of the method of measuring the characteristics of the carrier rocket, the attitude parameters of the rocket body in the coordinate system are calculated by using the measurements data of multiple high-speed television sets, and then the parameters are transferred to the rocket body attack angle and it is assessed whether the rocket has a good flight stability flying with a small attack angle. The measurement method and the mathematical algorithm steps through the data processing test, where you can intuitively observe the rocket flight stability state, and also can visually identify the guidance system or failure analysis.

  5. Pressure-Equalizing Cradle for Booster Rocket Mounting

    Science.gov (United States)

    Rutan, Elbert L. (Inventor)

    2015-01-01

    A launch system and method improve the launch efficiency of a booster rocket and payload. A launch aircraft atop which the booster rocket is mounted in a cradle, is flown or towed to an elevation at which the booster rocket is released. The cradle provides for reduced structural requirements for the booster rocket by including a compressible layer, that may be provided by a plurality of gas or liquid-filled flexible chambers. The compressible layer contacts the booster rocket along most of the length of the booster rocket to distribute applied pressure, nearly eliminating bending loads. Distributing the pressure eliminates point loading conditions and bending moments that would otherwise be generated in the booster rocket structure during carrying. The chambers may be balloons distributed in rows and columns within the cradle or cylindrical chambers extending along a length of the cradle. The cradle may include a manifold communicating gas between chambers.

  6. Rocket propellants with reduced smoke and high burning rates

    Energy Technology Data Exchange (ETDEWEB)

    Menke, K.; Eisele, S. [Fraunhofer-Institut fuer Chemische Technologie (ICT), Pfinztal-Berghausen (Germany)

    1997-07-01

    Rocket propellants with reduced smoke and high burning rates recommend themselves for use in a rocket motor for high accelerating tactical missiles. They serve for an improved camouflage on the battle field and may enable guidance control due to the higher transmission of their rocket plume compared to traditional aluminized composite propellants. In this contribution the material based ranges of performance and properties of three non aluminized rocket propellants will be introduced and compared to each other. The selected formulations based on AP/HTPB; AP/PU/TMETN and AP/HMX/GAP/TMETN have roughly the same specific impulse of I{sub SP}=2430 Ns/kg at 70:1 expansion ratio. The burning rates in the pressure range from 10-18 MPa vary from to 26-33 mm/s for the AP/HTPB propellant, 52-68 mm/s for the formulation based on AP/PU/TMETN and 28-39 mm/s for the propellant based on AP/HMX/GAP. With 58% and 20% AP-contents the propellants with nitrate ester plasticizers create a much smaller secondary signature than the AP/HTPB representative containing 85% AP. Their disadvantage, however, is the connection of high performance to a high level of energetic plasticizer. For this reason, the very fast burning propellant based on AP/PU/TMETN is endowed with a low elastic modulus and is limited to a grain configuration which isn`t exposed too much to the fast and turbulent airstream. The mechanical properties of the AP/HMX/GAP-propellant are as good or better as those of the AP/HTPB propellant. The first one exhibits the same performance and burn rates as the composite representative but produces only one fifth of HCl exhaust. For this reason it is recommended for missile applications, which must have high accelerating power together with a significantly reduced plume signature and smoke production. (orig.) [Deutsch] Rauchreduzierte Festtreibstoffe mit hohen Abbrandgeschwindigkeiten bieten sich fuer den Antrieb hochbeschleunigender taktischer Flugkoerper an, da sie gegenueber

  7. DARK BURSTS IN THE SWIFT ERA: THE PALOMAR 60 INCH-SWIFT EARLY OPTICAL AFTERGLOW CATALOG

    International Nuclear Information System (INIS)

    Cenko, S. B.; Harrison, F. A.; Kelemen, J.; Fox, D. B.; Kulkarni, S. R.; Kasliwal, M. M.; Ofek, E. O.; Rau, A.; Gal-Yam, A.; Frail, D. A.; Moon, D.-S.

    2009-01-01

    We present multicolor optical observations of long-duration γ-ray bursts (GRBs) made over a three-year period with the robotic Palomar 60 inch telescope (P60). Our sample consists of all 29 events discovered by Swift for which P60 began observations less than 1 hr after the burst trigger. We were able to recover 80% of the optical afterglows from this prompt sample, and we attribute this high efficiency to our red coverage. Like Melandri et al. (2008), we find that a significant fraction (∼50%) of Swift events show a suppression of the optical flux with regard to the X-ray emission (the so-called 'dark' bursts). Our multicolor photometry demonstrates this is likely due in large part to extinction in the host galaxy. We argue that previous studies, by selecting only the brightest and best-sampled optical afterglows, have significantly underestimated the amount of dust present in typical GRB environments.

  8. Development of laser cladding system to repair wall thinning of 1-inch heat exchanger tube

    International Nuclear Information System (INIS)

    Terada, Takaya

    2013-01-01

    We developed a laser cladding system to repair the inner wall wastage of heat exchanger tubes. Our system, which is designed to repair thinning tube walls within 100 mm from the edge of a heat exchanger tube, consists of a fiber laser, a composite-type optical fiberscope, a coupling device, a laser processing head, and a wire-feeding device. All of these components were reconfigured from the technologies of FBR maintenance. The laser processing head, which has a 15-mm outer diameter, was designed to be inserted into a 1-inch heat exchanger tube. We mounted a heatproof broadband mirror for laser cladding and fiberscope observation with visible light inside the laser processing head. The wire-feeding device continuously supplied 0.4-mm wire to the laser irradiation spot with variable feeding speeds from 0.5 to 20 mm/s. We are planning to apply our proposed system to the maintenance of aging industrial plants. (author)

  9. Analysis of ATLAS 6-inch cold leg break simulation with MARS-KS code

    Energy Technology Data Exchange (ETDEWEB)

    Kim, Se Yun; Jun, Hwang Yong; Ha, Sang Jun [Korea Electric Power Company, Daejeon (Korea, Republic of)

    2011-05-15

    A Domestic Standard Problem (DSP) exercise using ATLAS facility has been organized by KAERI. As the second DSP exercise, the 6-inch cold leg bottom break was determined. This experiment is the counterpart test to the DVI line break to verify the safety performance of the DVI method over the traditional CLI method. Compared with the large break LOCA, the phases of the small break LOCA prior to core recovery occur over a long period. The blowdown, natural circulation, loop seal clearance, boil-off, and core recovery phase should be investigated minutely with relevant models of safety analysis codes in order to predict these thermal hydraulic phenomena correctly. To investigate the ECC bypass phenomena, a finer study on the thermalhydraulic behavior in upper annulus downcomer was carried out

  10. Testing of one-inch UF{sub 6} cylinder valves under simulated fire conditions

    Energy Technology Data Exchange (ETDEWEB)

    Elliott, P.G. [Martin Marietta Energy Systems, Inc., Paducah, KY (United States)

    1991-12-31

    Accurate computational models which predict the behavior of UF{sub 6} cylinders exposed to fires are required to validate existing firefighting and emergency response procedures. Since the cylinder valve is a factor in the containment provided by the UF{sub 6} cylinder, its behavior under fire conditions has been a necessary assumption in the development of such models. Consequently, test data is needed to substantiate these assumptions. Several studies cited in this document provide data related to the behavior of a 1-inch UF{sub 6} cylinder valve in fire situations. To acquire additional data, a series of tests were conducted at the Paducah Gaseous Diffusion Plant (PGDP) under a unique set of test conditions. This document describes this testing and the resulting data.

  11. ECR heavy-ion source for the LBL 88-inch cyclotron

    International Nuclear Information System (INIS)

    Clark, D.J.; Kalnins, J.G.; Lyneis, C.M.

    1983-03-01

    An Electron Cyclotron Resonance (ECR) heavy-ion source is under construction at the LBL 88-Inch Cyclotron. This source will produce very-high-charge-state heavy ions, such as 0 8 + and Ar 12 + , which will increase cyclotron energies by a factor of 2-4, up to A = 80. It is a two-stage source using room-temperature coils, a permanent-magnet sextupole, and a 6-9 GHz microwave system. Design features include adjustable first-to-second-stage plasma coupling, a variable second-stage mirror ratio, high-conductance radial pumping of the second stage, and a beam-diagnostic system. A remotely movable extraction electrode will optimize extraction efficiency. The project includes construction of a transport line and improvements to the cyclotron axial-injection system. The construction period is expected to be two years

  12. Recent developments in high charge state heavy ion beams at the LBL 88-inch Cyclotron

    International Nuclear Information System (INIS)

    Gough, R.A.; Clark, D.J.; Glasgow, L.R.

    1978-01-01

    Recent advances in design and operation of the internal PIG sources at the LBL 88-Inch Cyclotron have led to the development of high charge state (0.4 16 O 8+ . Total external intensities of these beams range from 10 12 particles/s for 6 Li 3+ to 0.1 particles/s for 16 O 8+ . Techniques have been developed for routine tune-out of the low intensity beams. These include use of model beams and reliance on the large systematic data base of cyclotron parameters which has been developed over many years of operation. Techniques for delivery of these weak beams to the experimental target areas are presented. Source design and operation, including special problems associated with Li, Be, and B beams are discussed

  13. Evaluation of new 5 inch photomultiplier for use in threshold Cherenkov detectors with aerogel radiator

    International Nuclear Information System (INIS)

    Wojtsekhowski, B.; Zorn, C.; Flyckt, S.O.

    2000-01-01

    A cost effective alternative to UV-sensitive 5 inch PMTs often used with threshold Aerogel Cherenkov detectors has been developed and tested. The photomultiplier -XP4572-is a variation of the Photonis XP4512 glass window tube with improved electron collection efficiency. Fast timing and high gain were only moderately compromised. The effective quantum efficiency has been measured as twice that of a Burle 8854 Quantacon when exposed to a Cherenkov spectrum generated by Ru-106 electrons (les;3.54 MeV) through 1 cm of high index, high transparency Matsushita Electric aerogel (n=1.05). This new phototube is being installed in an aerogel-based Cherenkov detector for Hall A at Jefferson Lab

  14. Development and characterization of a CCD camera system for use on six-inch manipulator systems

    International Nuclear Information System (INIS)

    Logory, L.M.; Bell, P.M.; Conder, A.D.; Lee, F.D.

    1996-01-01

    The Lawrence Livermore National Laboratory has designed, constructed, and fielded a compact CCD camera system for use on the Six Inch Manipulator (SIM) at the Nova laser facility. The camera system has been designed to directly replace the 35 mm film packages on all active SIM-based diagnostics. The unit's electronic package is constructed for small size and high thermal conductivity using proprietary printed circuit board technology, thus reducing the size of the overall camera and improving its performance when operated within the vacuum environment of the Nova laser target chamber. The camera has been calibrated and found to yield a linear response, with superior dynamic range and signal-to-noise levels as compared to T-Max 3200 optic film, while providing real-time access to the data. Limiting factors related to fielding such devices on Nova will be discussed, in addition to planned improvements of the current design

  15. LOFT transient thermal analysis for 10 inch primary coolant blowdown piping weld

    International Nuclear Information System (INIS)

    Howell, S.K.

    1978-01-01

    A flaw in a weld in the 10 inch primary coolant blowdown piping was discovered by LOFT personnel. As a result of this, a thermal analysis and fracture mechanics analysis was requested by LOFT personnel. The weld and pipe section were analyzed for a complete thermal cycle, heatup and Loss of Coolant Experiment (LOCE), using COUPLE/MOD2, a two-dimensional finite element heat conduction code. The finite element representation used in this analysis was generated by the Applied Mechanics Branch. The record of nodal temperatures for the entire transient was written on tape VSN=T9N054, and has been forwarded to the Applied Mechanics Branch for use in their mechanical analysis. Specific details and assumptions used in this analysis are found in appropriate sections of this report

  16. Current status of rocket developments in universities -development of a small hybrid rocket with a swirling oxidizer flow type engine

    OpenAIRE

    Yuasa, Saburo; Kitagawa, Koki

    2005-01-01

    To develop an experimental small hybrid rocket with a swirling gaseous oxygen flow type engine, we made a flight model engine. Burning tests of the engine showed that a maximum thrust of 692 N and a specific impulse of 263 s (at sea level) were achieved. We designed a small hybrid rocket with this engine. The rocket measured 1.8 m in length and 15.4 kg in mass. To confirm the flight stability of the rocket, wind tunnel tests using a 112-scale model of the rocket and simulations of the flight ...

  17. Rocket center Peenemünde — Personal memories

    Science.gov (United States)

    Dannenberg, Konrad; Stuhlinger, Ernst

    Von Braun built his first rockets as a young teenager. At 14, he started making plans for rockets for human travel to the Moon and Mars. The German Army began a rocket program in 1929. Two years later, Colonel (later General) Becker contacted von Braun who experimented with rockets in Berlin, gave him a contract in 1932, and, jointly with the Air Force, in 1936 built the rocket center Peenemünde where von Braun and his team developed the A-4 (V-2) rocket under Army auspices, while the Air Force developed the V-1 (buzz bomb), wire-guided bombs, and rocket planes. Albert Speer, impressed by the work of the rocketeers, allowed a modest growth of the Peenemünde project; this brought Dannenberg to the von Braun team in 1940. Hitler did not believe in rockets; he ignored the A-4 project until 1942 when he began to support it, expecting that it could turn the fortunes of war for him. He drastically increased the Peenemünde work force and allowed the transfer of soldiers from the front to Peenemünde; that was when Stuhlinger, in 1943, came to Peenemünde as a Pfc.-Ph.D. Later that year, Himmler wrenched the authority over A-4 production out of the Army's hands, put it under his command, and forced production of the immature rocket at Mittelwerk, and its military deployment against targets in France, Belgium, and England. Throughout the development of the A-4 rocket, von Braun was the undisputed leader of the project. Although still immature by the end of the war, the A-4 had proceeded to a status which made it the first successful long-range precision rocket, the prototype for a large number of military rockets built by numerous nations after the war, and for space rockets that launched satellites and traveled to the Moon and the planets.

  18. AJ26 rocket engine testing news briefing

    Science.gov (United States)

    2010-01-01

    NASA's John C. Stennis Space Center Director Gene Goldman (center) stands in front of a 'pathfinder' rocket engine with Orbital Sciences Corp. President and Chief Operating Officer J.R. Thompson (left) and Aerojet President Scott Seymour during a Feb. 24 news briefing at the south Mississippi facility. The leaders appeared together to announce a partnership for testing Aerojet AJ26 rocket engines at Stennis. The engines will be used to power Orbital's Taurus II space vehicles to provide commercial cargo transportation missions to the International Space Station for NASA. During the event, the Stennis partnership with Orbital was cited as an example of the new direction of NASA to work with commercial interests for space travel and transport.

  19. Nuclear thermal rockets using indigenous Martian propellants

    International Nuclear Information System (INIS)

    Zubrin, R.M.

    1989-01-01

    This paper considers a novel concept for a Martian descent and ascent vehicle, called NIMF (for nuclear rocket using indigenous Martian fuel), the propulsion for which will be provided by a nuclear thermal reactor which will heat an indigenous Martian propellant gas to form a high-thrust rocket exhaust. The performance of each of the candidate Martian propellants, which include CO2, H2O, CH4, N2, CO, and Ar, is assessed, and the methods of propellant acquisition are examined. Attention is also given to the issues of chemical compatibility between candidate propellants and reactor fuel and cladding materials, and the potential of winged Mars supersonic aircraft driven by this type of engine. It is shown that, by utilizing the nuclear landing craft in combination with a hydrogen-fueled nuclear thermal interplanetary vehicle and a heavy lift booster, it is possible to achieve a manned Mars mission in one launch. 6 refs

  20. Multisized Inert Particle Loading for Solid Rocket Axial Combustion Instability Suppression

    Directory of Open Access Journals (Sweden)

    David R. Greatrix

    2012-01-01

    Full Text Available In the present investigation, various factors and trends, related to the usage of two or more sets of inert particles comprised of the same material (nominally aluminum but at different diameters for the suppression of axial shock wave development, are numerically predicted for a composite-propellant cylindrical-grain solid rocket motor. The limit pressure wave magnitudes at a later reference time in a given pulsed firing simulation run are collected for a series of runs at different particle sizes and loading distributions and mapped onto corresponding attenuation trend charts. The inert particles’ presence in the central core flow is demonstrated to be an effective means of instability symptom suppression, in correlating with past experimental successes in the usage of particles. However, the predicted results of this study suggest that one needs to be careful when selecting more than one size of particle for a given motor application.

  1. Design and qualification of an UHV system for operation on sounding rockets

    Energy Technology Data Exchange (ETDEWEB)

    Grosse, Jens, E-mail: jens.grosse@dlr.de; Braxmaier, Claus [Center of Applied Space Technology and Microgravity (ZARM), University of Bremen, Bremen, 28359, Germany and German Aerospace Center (DLR) Bremen, Bremen, 28359 (Germany); Seidel, Stephan Tobias; Becker, Dennis; Lachmann, Maike Diana [Institute of Quantum Optics, Leibniz University Hanover, Hanover, 30167 (Germany); Scharringhausen, Marco [German Aerospace Center (DLR) Bremen, Bremen, 28359 (Germany); Rasel, Ernst Maria [Institute of Quantum Optics, Leibniz University Hanover, Hanover, 30167, Bremen (Germany)

    2016-05-15

    The sounding rocket mission MAIUS-1 has the objective to create the first Bose–Einstein condensate in space; therefore, its scientific payload is a complete cold atom experiment built to be launched on a VSB-30 sounding rocket. An essential part of the setup is an ultrahigh vacuum system needed in order to sufficiently suppress interactions of the cooled atoms with the residual background gas. Contrary to vacuum systems on missions aboard satellites or the international space station, the required vacuum environment has to be reached within 47 s after motor burn-out. This paper contains a detailed description of the MAIUS-1 vacuum system, as well as a description of its qualification process for the operation under vibrational loads of up to 8.1 g{sub RMS} (where RMS is root mean square). Even though a pressure rise dependent on the level of vibration was observed, the design presented herein is capable of regaining a pressure of below 5 × 10{sup −10} mbar in less than 40 s when tested at 5.4 g{sub RMS}. To the authors' best knowledge, it is the first UHV system qualified for operation on a sounding rocket.

  2. Rocket Based Combined Cycle (RBCC) Engine

    Science.gov (United States)

    2004-01-01

    Pictured is an artist's concept of the Rocket Based Combined Cycle (RBCC) launch. The RBCC's overall objective is to provide a technology test bed to investigate critical technologies associated with opperational usage of these engines. The program will focus on near term technologies that can be leveraged to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsions systems and ultimately a Single Stage To Orbit (SSTO) air breathing propulsion system.

  3. Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

    International Nuclear Information System (INIS)

    Emrich, William J. Jr.

    2008-01-01

    To support a potential future development of a nuclear thermal rocket engine, a state-of-the-art non nuclear experimental test setup has been constructed to evaluate the performance characteristics of candidate fuel element materials and geometries in representative environments. The test device simulates the environmental conditions (minus the radiation) to which nuclear rocket fuel components could be subjected during reactor operation. Test articles mounted in the simulator are inductively heated in such a manner as to accurately reproduce the temperatures and heat fluxes normally expected to occur as a result of nuclear fission while at the same time being exposed to flowing hydrogen. This project is referred to as the Nuclear Thermal Rocket Element Environment Simulator or NTREES. The NTREES device is located at the Marshall Space flight Center in a laboratory which has been modified to accommodate the high powers required to heat the test articles to the required temperatures and to handle the gaseous hydrogen flow required for the tests. Other modifications to the laboratory include the installation of a nitrogen gas supply system and a cooling water supply system. During the design and construction of the facility, every effort was made to comply with all pertinent regulations to provide assurance that the facility could be operated in a safe and efficient manner. The NTREES system can currently supply up to 50 kW of inductive heating to the fuel test articles, although the facility has been sized to eventually allow test article heating levels of up to several megawatts

  4. Sounding rocket study of auroral electron precipitation

    International Nuclear Information System (INIS)

    McFadden, J.P.

    1985-01-01

    Measurement of energetic electrons in the auroral zone have proved to be one of the most useful tools in investigating the phenomena of auroral arc formation. This dissertation presents a detailed analysis of the electron data from two sounding rocket campaigns and interprets the measurements in terms of existing auroral models. The Polar Cusp campaign consisted of a single rocket launched from Cape Parry, Canada into the afternoon auroral zone at 1:31:13 UT on January 21, 1982. The results include the measurement of a narrow, magnetic field aligned electron flux at the edge of an arc. This electron precipitation was found to have a remarkably constant 1.2 eV temperature perpendicular to the magnetic field over a 200 to 900 eV energy range. The payload also made simultaneous measurements of both energetic electrons and 3-MHz plasma waves in an auroral arc. Analysis has shown that the waves are propagating in the upper hybrid band and should be generated by a positive slope in the parallel electron distribution. A correlation was found between the 3-MHz waves and small positive slopes in the parallel electron distribution but experimental uncertainties in the electron measurement were large enough to influence the analysis. The BIDARCA campaign consisted of two sounding rockets launched from Poker Flat and Fort Yukon, Alaska at 9:09:00 UT and 9:10:40 UT on February 7, 1984

  5. Fine motor control

    Science.gov (United States)

    ... gross (large, general) motor control. An example of gross motor control is waving an arm in greeting. Problems ... out the child's developmental age. Children develop fine motor skills over time, by practicing and being taught. To ...

  6. Ultrasonographic median nerve cross-section areas measured by 8-point 'inching test' for idiopathic carpal tunnel syndrome: a correlation of nerve conduction study severity and duration of clinical symptoms

    International Nuclear Information System (INIS)

    Chen, Shu-Fang; Lu, Cheng-Hsien; Huang, Chi-Ren; Chuang, Yao-Chung; Tsai, Nai-Wen; Chang, Chiung-Chih; Chang, Wen-Neng

    2011-01-01

    Incremental palmar stimulation of the median nerve sensory conduction at the wrist, the 'inching test', provides an assessment with reference to segments proximal and distal to the entrapment. This study used high-resolution ultrasonography (US) to measure the median nerve's cross-section areas (CSAs) like the 'inching test' and to correlate with the nerve conduction study (NCS) severity and duration of carpal tunnel syndrome (CTS). Two hundred and twelve (212) 'CTS-hands' from 135 CTS patients and 50 asymptomatic hands ('A-hands') from 25 control individuals were enrolled. The median nerve CSAs were measured at the 8-point marked as i4, i3, i2, i1, w, o1, o2, and 03 in inching test. The NCS severities were classified into six groups based on motor and sensory responses (i.e., negative, minimal, mild, moderate, severe, and extreme). Results of US studies were compared in terms of NCS severity and duration of clinical CTS symptoms. There was significantly larger CSA of the NCS negative group of 'CTS-hands' than of 'A-hands'. The cut-off values of the CSAs of the NCS negative CTS group were 12.5 mm 2 , 11.5 mm 2 and 10.1 mm 2 at the inlet, wrist crease, and outlet, respectively. Of the 212 'CTS-hands', 32 were NCS negative while 40 had minimal, 43 mild, 85 moderate, 10 severe, and two extreme NCS severities. The CSAs of 'CTS-hands' positively correlated with different NCS severities and with the duration of CTS symptoms. By duration of clinical symptoms, 12 of the 212 'CTS-hands' were in the 1 month group; 82 in >1 month and ≤12 months group, and 118 in >12 months group. In 'inching test', segments i4-i3 and i3-i2 were the most common 'positive-site'. The corresponding CSAs measured at i4 and i3, but not at i2, were significantly larger than those measured at points that were not 'positive-site'. Using the 8-point measurement of the median

  7. Special-Purpose High-Torque Permanent-Magnet Motors

    Science.gov (United States)

    Doane, George B., III

    1995-01-01

    Permanent-magnet brushless motors that must provide high commanded torques and satisfy unusual heat-removal requirement are developed. Intended for use as thrust-vector-control actuators in large rocket engines. Techniques and concepts used to design improved motors for special terrestrial applications. Conceptual motor design calls for use of rotor containing latest high-energy-product rare-earth permanent magnets so that motor produces required torque while drawing smallest possible currents from power supply. Torque generated by electromagnetic interaction between stator and permanent magnets in rotor when associated electronic circuits applied appropriately temporally and spatially phased currents to stator windings. Phase relationships needed to produce commanded torque computed in response to torque command and to electronically sensed angular position of rotor relative to stator.

  8. A simplified computational fluid-dynamic approach to the oxidizer injector design in hybrid rockets

    Science.gov (United States)

    Di Martino, Giuseppe D.; Malgieri, Paolo; Carmicino, Carmine; Savino, Raffaele

    2016-12-01

    Fuel regression rate in hybrid rockets is non-negligibly affected by the oxidizer injection pattern. In this paper a simplified computational approach developed in an attempt to optimize the oxidizer injector design is discussed. Numerical simulations of the thermo-fluid-dynamic field in a hybrid rocket are carried out, with a commercial solver, to investigate into several injection configurations with the aim of increasing the fuel regression rate and minimizing the consumption unevenness, but still favoring the establishment of flow recirculation at the motor head end, which is generated with an axial nozzle injector and has been demonstrated to promote combustion stability, and both larger efficiency and regression rate. All the computations have been performed on the configuration of a lab-scale hybrid rocket motor available at the propulsion laboratory of the University of Naples with typical operating conditions. After a preliminary comparison between the two baseline limiting cases of an axial subsonic nozzle injector and a uniform injection through the prechamber, a parametric analysis has been carried out by varying the oxidizer jet flow divergence angle, as well as the grain port diameter and the oxidizer mass flux to study the effect of the flow divergence on heat transfer distribution over the fuel surface. Some experimental firing test data are presented, and, under the hypothesis that fuel regression rate and surface heat flux are proportional, the measured fuel consumption axial profiles are compared with the predicted surface heat flux showing fairly good agreement, which allowed validating the employed design approach. Finally an optimized injector design is proposed.

  9. Motor control for a brushless DC motor

    Science.gov (United States)

    Peterson, William J. (Inventor); Faulkner, Dennis T. (Inventor)

    1985-01-01

    This invention relates to a motor control system for a brushless DC motor having an inverter responsively coupled to the motor control system and in power transmitting relationship to the motor. The motor control system includes a motor rotor speed detecting unit that provides a pulsed waveform signal proportional to rotor speed. This pulsed waveform signal is delivered to the inverter to thereby cause an inverter fundamental current waveform output to the motor to be switched at a rate proportional to said rotor speed. In addition, the fundamental current waveform is also pulse width modulated at a rate proportional to the rotor speed. A fundamental current waveform phase advance circuit is controllingly coupled to the inverter. The phase advance circuit is coupled to receive the pulsed waveform signal from the motor rotor speed detecting unit and phase advance the pulsed waveform signal as a predetermined function of motor speed to thereby cause the fundamental current waveform to be advanced and thereby compensate for fundamental current waveform lag due to motor winding reactance which allows the motor to operate at higher speeds than the motor is rated while providing optimal torque and therefore increased efficiency.

  10. LOX/hydrocarbon rocket engine analytical design methodology development and validation. Volume 2: Appendices

    Science.gov (United States)

    Niiya, Karen E.; Walker, Richard E.; Pieper, Jerry L.; Nguyen, Thong V.

    1993-05-01

    This final report includes a discussion of the work accomplished during the period from Dec. 1988 through Nov. 1991. The objective of the program was to assemble existing performance and combustion stability models into a usable design methodology capable of designing and analyzing high-performance and stable LOX/hydrocarbon booster engines. The methodology was then used to design a validation engine. The capabilities and validity of the methodology were demonstrated using this engine in an extensive hot fire test program. The engine used LOX/RP-1 propellants and was tested over a range of mixture ratios, chamber pressures, and acoustic damping device configurations. This volume contains time domain and frequency domain stability plots which indicate the pressure perturbation amplitudes and frequencies from approximately 30 tests of a 50K thrust rocket engine using LOX/RP-1 propellants over a range of chamber pressures from 240 to 1750 psia with mixture ratios of from 1.2 to 7.5. The data is from test configurations which used both bitune and monotune acoustic cavities and from tests with no acoustic cavities. The engine had a length of 14 inches and a contraction ratio of 2.0 using a 7.68 inch diameter injector. The data was taken from both stable and unstable tests. All combustion instabilities were spontaneous in the first tangential mode. Although stability bombs were used and generated overpressures of approximately 20 percent, no tests were driven unstable by the bombs. The stability instrumentation included six high-frequency Kistler transducers in the combustion chamber, a high-frequency Kistler transducer in each propellant manifold, and tri-axial accelerometers. Performance data is presented, both characteristic velocity efficiencies and energy release efficiencies, for those tests of sufficient duration to record steady state values.

  11. Assessment of the MARS-KS Code Using Atlas 6-inch cold leg Break Test

    Energy Technology Data Exchange (ETDEWEB)

    Kang, D. G.; Kim, J. S.; Ahn, S. H.; Seul, K. W. [Korea Institute of Nuclear Safety, Daejeon (Korea, Republic of)

    2012-03-15

    An integral effect test on the SBLOCA (Small-Break Loss of Coolant Accident) aiming at 6-inch cold leg bottom break, SB-CL-09, was conducted with the Atlas on November, 13, 2009, by KAERI. In this study, the calculations using MARS-KS Vt1.2 code were conducted for 6-inch cold leg break test of Atlas (SB-CL-09) which is the second domestic standard problem (Dsp-02) to assess MARS-KS code capability to simulate the transient thermal-hydraulic behavior for SBLOCA. The steady state was determined by conducting a null transient calculation and the errors between the calculated and measured values are acceptable for almost primary/secondary system parameters. The predicted pressurizer pressure agrees relatively well with the experimental data and the predicted break flow and mass are in good agreement with experiment. In MARS-KS calculation, the decrease of core collapsed water level is predicted well in blowdown phase, but just before LSC, water level is higher than experiment. However, the sudden decrease and increase of water level is higher than experiment. However, the sudden decrease and increase of water level at the LSC are predicted qualitatively. After LSC, there is another water level dip at Sit injection time which is not in experiment. It is considered that this phenomenon is caused by rapid depressurization of downcomer due to significant condensation rate of vapor in downcomer when Sit water flows in it. For the downcomer water level is predicted well, however, it is significantly over-predicted at SIT injection time, water level is predicted well, however, it is significantly over-predicted at SIT injection time after SIT water flows in downcomer. Predicted cladding temperature generally agrees well with the experiment, while there is peak at SIT injection time in calculation which is not in experiment. The loop seals of 1A, 2B intermediate leg are cleared around 400 seconds in experiment, while only that of 1A is cleared in MARS-KS calculation at the

  12. Rocket measurements of electron density irregularities during MAC/SINE

    Science.gov (United States)

    Ulwick, J. C.

    1989-01-01

    Four Super Arcas rockets were launched at the Andoya Rocket Range, Norway, as part of the MAC/SINE campaign to measure electron density irregularities with high spatial resolution in the cold summer polar mesosphere. They were launched as part of two salvos: the turbulent/gravity wave salvo (3 rockets) and the EISCAT/SOUSY radar salvo (one rocket). In both salvos meteorological rockets, measuring temperature and winds, were also launched and the SOUSY radar, located near the launch site, measured mesospheric turbulence. Electron density irregularities and strong gradients were measured by the rocket probes in the region of most intense backscatter observed by the radar. The electron density profiles (8 to 4 on ascent and 4 on descent) show very different characteristics in the peak scattering region and show marked spatial and temporal variability. These data are intercompared and discussed.

  13. Introduction to the Special Issue on Sounding Rockets and Instrumentation

    OpenAIRE

    Christe, Steven; Zeiger, Ben; Pfaff, Rob; Garcia, Michael

    2016-01-01

    Rocket technology, originally developed for military applications, has provided a low-cost observing platform to carry critical and rapid-response scientific investigations for over 70 years. Even with the development of launch vehicles that could put satellites into orbit, high altitude sounding rockets have remained relevant. In addition to science observations, sounding rockets provide a unique technology test platform and a valuable training ground for scientists and engineers. Most impor...

  14. The Norwegian sounding rocket programme 1978-81

    International Nuclear Information System (INIS)

    Landmark, B.

    1978-01-01

    The Norwegian sounding rocket programme is reasonably well defined up to and including the winter of 1981/82. All the projects have been planned and will be carried out in international cooperation. Norwegian scientists so far plan to participate in a number of 24 rocket payloads over the period. Out of these 18 will be launched from the Andoya rocket range, 3 from Esrange and 3 from the siple station in the antarctic. (author)

  15. Design evaluation of the 20-cm (8-inch) secondary burner system

    International Nuclear Information System (INIS)

    Rode, J.S.

    1977-08-01

    This report describes an evaluation of the design of the existing 20-cm (8-inch) engineering-scale secondary burner system in the HTGR reprocessing cold pilot plant at General Atomic Co. The purpose of this evaluation is to assess the suitability of the existing design as a prototype of the HTGR Recycle Demonstration Facility (HRDF) secondary burner system and to recommend alternatives where the existing design is thought to be unsuitable as a prototype. This evaluation has led to recommendations for the parallel development of two integrated design concepts for a prototype secondary burner system. One concept utilizes the existing burner heating and cooling subsystems in order to minimize development risk, but simplifies a number of other features associated with remote maintenance and burner operation. The other concept, which offers maximum cost reduction, utilizes internal gas cooling of the burner, retains the existing heating subsystem for design compatibility, but requires considerable development to reduce the risk to acceptable limits. These concepts, as well as other design alternatives, are described and evaluated

  16. Isolation and identification of culturable extremely halophilic archaea of Inche-Boroun wetland

    Directory of Open Access Journals (Sweden)

    Mehrnoosh Rasooli

    2013-11-01

    Full Text Available Haloarchaeal diversity of Inche-Boroun wetland in north of Iran in Golestan province was investigated by using culture-dependent methods. Sampling was carried out in May and September 2010. In each sampling, 4 distinct region of wetland were analyzed by using complex media like MGM, JCM168, MH1 and an alkaliphilic medium containing 23% salts. After incubation at 40ºC, a total of 406 isolates were prepared and 2.1×106 CFU/ml were obtained in culture media. Among all isolates, 361 isolates were obtained from MGM and 39 isolates from JCM 168, 3 isolates from MH1 and 3 isolate from alkaliphilic media. Initial morphological, biochemical and physiological tests were performed. According to the results, 45 isolates were selected and phylogenetic analysis of 16S rRNA was performed for them. Among selected strains, 40 isolates belonged to Halobacteriacaea and were related to Haloarcula, Halorubrum, Haloferax, Halobellus, Halogeometricum, Halobacterium, Halolamina, Halorhabdus and Halostagnicola (respectively 30, 27.5, 17.5, 10, 5.2, 2.6, 2.6, 2.6 and 2.6 percent of Haloarchaeal strains. A total of 5 strains belonged to the kingdom of Bacteria and were related to Rhodovibrio, Pseudomonas and Salicola (respectively 40, 40 and 20 percent of bacterial strains. According to our results and the limited numbers of haloarchaeal genera that having been discovered until now, it seemed that the culturable prokaryotic populations in this hypersaline environment was diverse.

  17. Analysis of improved and original designs of a 16 inch long penultimate stage turbine blade

    International Nuclear Information System (INIS)

    Carnero, A.; Kubiak, J.A.; Mendez, R.

    1994-01-01

    A finite element analysis of 16 inch long penultimate stage (L-1) blade was carried out to evaluate the improved and the original designs. The original design of the blade involved the ''blade-tenon-shroud'' system to make blade groups (6 blades per group). The improved design applied the concept of Integral Shroud Blade (ISB). Thus all the blades made a 360 degree group. The paper presents an application of the finite element analysis method to compute the natural frequencies, steady-state and alternating stresses, deformation due to forces acting on the blades and modal shapes of the blade group. In the case of the improved design it was also necessary to carry out computation of the dynamic response of a 360 degree blade-disk arc. This was to include the effect of the flexible disk fastening where blade and disk interaction were important to identify certain resonant conditions. It was concluded from the finite element results, that the steady-state stresses in the improved blade were lower, and the tangential mode shapes were eliminated. This was a great advantage since in the original design the first tangential mode shape and the higher steady-state stresses in the tenon contributed to the frequent failure of the ''blade-tenon-shroud'' system

  18. Test results of a jet impingement from a 4 inch pipe under BWR LOCA conditions

    International Nuclear Information System (INIS)

    Isozaki, Toshikuni; Yano, Toshikazu; Miyazaki, Noriyuki; Kato, Rokuro; Kurihara, Ryoichi; Ueda, Shuzo; Miyazono, Shohachiro

    1982-09-01

    Hypothetical instantaneous pipe rupture is now considered to be one of the design basis accidents during the operation of the light water reactor. If a pipe rupture accidnet occurs, the pipe will start moving with the sudden discharge of internal fluid. So, the various apparatus such as pipe whip restraints and jet deflectors are being installed near the postulated break location to protect the nuclear power plants against the effect of postulated pipe rupture. Pipe whipping test and jet discharge test are now being conducted at the Division of Reactor Safety of the Japan Atomic Energy Research Institute. This report describes the test results of the jet discharge from a 4 inch pipe under BWR LOCA condition. In front of the pipe exit the target disk of 1000 mm in diameter was installed. The distance between the pipe exit and the target was 500 mm. 13 pressure transducers and 13 thermocouples were mounted on the target disk to measure the pressure and temperature increase due to jet impingement on the target. (author)

  19. Pipe rupture test results: 4-inch pipe whip tests under PWR LOCA conditions

    International Nuclear Information System (INIS)

    Miyazaki, Noriyuki; Ueda, Shuzo; Isozaki, Toshikuni; Kato, Rokuro; Kurihara, Ryoichi; Yano, Toshikazu; Miyazono, Shohachiro

    1982-09-01

    This report summarizes the results of 4-inch pipe whip tests (RUN No. 5506, 5507, 5508 and 5604) under the PWR LOCA conditions. The dynamic behaviors of the test pipe and restraints were studied in the tests. In the tests, the gap between the test pipe and the restraints was kept at the constant value of 8.85 mm and the overhang length was varied from 250 mm to 650 mm. The dynamic behaviors of the test pipe and the restraint were made clear by the outputs of strain gages and the measurements of residual deformations. The data of water hammer in subcooled water were also obtained by the pressure transducers mounted on the test pipe. The main conclusions obtained from the tests are as follows. (1) The whipping of pipe can be prevented more effectively as the overhang length becomes shorter. (2) The load acting on the restraint-support structure becomes larger as the overhang length becomes shorter. (3) The restraint farther from the break location does not limit the pipe movement except for the first impact when the overhang length is long. (4) The ultimate moment M sub(u) of the pipe at the restraint location can be used to predict the plastic collapse of the whipping pipe. (5) The restraints slide along the pipe axis and are subjected to bending moment, when the overhang length is long. (author)

  20. Summary of LaRC 2-inch Erectable Joint Hardware Heritage Test Data

    Science.gov (United States)

    Dorsey, John T.; Watson, Judith J.

    2016-01-01

    As the National Space Transportation System (STS, also known as the Space Shuttle) went into service during the early 1980's, NASA envisioned many missions of exploration and discovery that could take advantage of the STS capabilities. These missions included: large orbiting space stations, large space science telescopes and large spacecraft for manned missions to the Moon and Mars. The missions required structures that were significantly larger than the payload volume available on the STS. NASA Langley Research Center (LaRC) conducted studies to design and develop the technology needed to assemble the large space structures in orbit. LaRC focused on technology for erectable truss structures, in particular, the joint that connects the truss struts at the truss nodes. When the NASA research in large erectable space structures ended in the early 1990's, a significant amount of structural testing had been performed on the LaRC 2-inch erectable joint that was never published. An extensive set of historical information and data has been reviewed and the joint structural testing results from this historical data are compiled and summarized in this report.

  1. A new facility for advanced rocket propulsion research

    Science.gov (United States)

    Zoeckler, Joseph G.; Green, James M.; Raitano, Paul

    1993-06-01

    A new test facility was constructed at the NASA Lewis Research Center Rocket Laboratory for the purpose of conducting rocket propulsion research at up to 8.9 kN (2000 lbf) thrust, using liquid oxygen and gaseous hydrogen propellants. A laser room adjacent to the test cell provides access to the rocket engine for advanced laser diagnostic systems. The size and location of the test cell provide the ability to conduct large amounts of testing in short time periods, with rapid turnover between programs. These capabilities make the new test facility an important asset for basic and applied rocket propulsion research.

  2. Wave-particle interaction phenomena observed by antarctic rockets

    International Nuclear Information System (INIS)

    Kimura, I.; Hirasawa, T.

    1979-01-01

    Rocket measurements of wave and particles activities made at Syowa Station in Antarctica during IMS period are reviewed. Nine rockets were used for such observations, out of which 6 rockets were launched in the auroral sky. In the VLF frequency range, 0 - 10 KHz, wideband spectra of wave electric and magnetic fields, Poynting flux and the direction of propagation vector were measured for chorus, ELF and VLF hiss, and for electrostatic noises. In the MF and HF range, the dynamic frequency spectra of 0.1 - 10 MHz were measured. The relationship of these wave phenomena with energetic particle activities measured by the same rockets are discussed. (author)

  3. Impact and mitigation of stratospheric ozone depletion by chemical rockets

    International Nuclear Information System (INIS)

    Mcdonald, A.J.

    1992-03-01

    The American Institute of Aeronautics and Astronautics (AIAA) conducted a workshop in conjunction with the 1991 AIAA Joint Propulsion Conference in Sacramento, California, to assess the impact of chemical rocket propulsion on the environment. The workshop included recognized experts from the fields of atmospheric physics and chemistry, solid rocket propulsion, liquid rocket propulsion, government, and environmental agencies, and representatives from several responsible environmental organizations. The conclusion from this workshop relative to stratospheric ozone depletion was that neither solid nor liquid rocket launchers have a significant impact on stratospheric ozone depletion, and that there is no real significant difference between the two

  4. Unsupervised Anomaly Detection for Liquid-Fueled Rocket Prop...

    Data.gov (United States)

    National Aeronautics and Space Administration — Title: Unsupervised Anomaly Detection for Liquid-Fueled Rocket Propulsion Health Monitoring. Abstract: This article describes the results of applying four...

  5. High efficiency motors; Motores de alta eficiencia

    Energy Technology Data Exchange (ETDEWEB)

    Uranga Favela, Ivan Jaime [Energia Controlada de Mexico, S. A. de C. V., Mexico, D. F. (Mexico)

    1993-12-31

    This paper is a technical-financial study of the high efficiency and super-premium motors. As it is widely known, more than 60% of the electrical energy generated in the country is used for the operation of motors, in industry as well as in commerce. Therefore the importance that the motors have in the efficient energy use. [Espanol] El presente trabajo es un estudio tecnico-financiero de los motores de alta eficiencia y los motores super premium. Como es ampliamente conocido, mas del 60% de la energia electrica generada en el pais, es utilizada para accionar motores, dentro de la industria y el comercio. De alli la importancia que los motores tienen en el uso eficiente de la energia.

  6. High efficiency motors; Motores de alta eficiencia

    Energy Technology Data Exchange (ETDEWEB)

    Uranga Favela, Ivan Jaime [Energia Controlada de Mexico, S. A. de C. V., Mexico, D. F. (Mexico)

    1992-12-31

    This paper is a technical-financial study of the high efficiency and super-premium motors. As it is widely known, more than 60% of the electrical energy generated in the country is used for the operation of motors, in industry as well as in commerce. Therefore the importance that the motors have in the efficient energy use. [Espanol] El presente trabajo es un estudio tecnico-financiero de los motores de alta eficiencia y los motores super premium. Como es ampliamente conocido, mas del 60% de la energia electrica generada en el pais, es utilizada para accionar motores, dentro de la industria y el comercio. De alli la importancia que los motores tienen en el uso eficiente de la energia.

  7. Advanced electric motor technology flux mapping

    Science.gov (United States)

    Doane, George B., III; Campbell, Warren; Dean, Garvin

    1993-01-01

    Design of electric motors which fulfill the needs of Thrust Vector Control (TVC) actuators used in large rocket propelled launch vehicles is covered. To accomplish this end the methodology of design is laid out in some detail. In addition a point design of a motor to fulfill the requirements of a certain actuator specified by MSFC is accomplished and reported upon. In the course of this design great stress has been placed on ridding the actuator of internally generated heat. To conduct the heat out of the motor use is made of the unique properties of the in house MSFC designed driving electronics. This property is that as along as they are operated in a quasi-linear mode the electronics nullify the effects of armature inductance as far as the phase of the armature current versus the rotor position is concerned. Actually the additional inductance due to the extended end turns in this design is of benefit because in the shorted armature failure mode the armature current in the fault (caused by the rotor flux sweeping past the armature) is diminished at a given rotor speed when compared to a more conventional motor with lower inductance. The magnetic circuit is analyzed using electromagnetic finite element methods.

  8. Ablative Material Testing at Lewis Rocket Lab

    Science.gov (United States)

    1997-01-01

    The increasing demand for a low-cost, reliable way to launch commercial payloads to low- Earth orbit has led to the need for inexpensive, expendable propulsion systems for new launch vehicles. This, in turn, has renewed interest in less complex, uncooled rocket engines that have combustion chambers and exhaust nozzles fabricated from ablative materials. A number of aerospace propulsion system manufacturers have utilized NASA Lewis Research Center's test facilities with a high degree of success to evaluate candidate materials for application to new propulsion devices.

  9. Water Rockets. Get Funny With Newton's Laws

    Directory of Open Access Journals (Sweden)

    Manuel Roca Vicent

    2017-01-01

    Full Text Available The study of the movement of the rocket has been used for decades to encourage students in the study of physics. This system has an undeniable interest to introduce concepts such as properties of gases, laws of Newton,  exchange  between  different  types  of  energy  and  its  conservation  or fluid  mechanics.  Our  works has  been  to  build  and  launch  these  rockets  in  different  educational  levels  and  in  each  of  these  ones  have introduced  the  part  of  Physics  more  suited  to  the  knowledge  of  our  students.  The  aim  of  the  learning experience  is  to  launch  the  rocket  as  far  as  possible  and  learn  to  predict  the  travelled  distance,  using Newton's  laws  and fluid  mechanics.  After  experimentation  we  demonstrated  to  be  able  to  control  the parameters that improve the performance of our rocket, such as the  fill factor, the volume and mass of the empty  bottle,  liquid  density,  launch  angle,  pressure  prior  air  release.  In addition, it is a fun experience can be attached to all levels of education in primary and high school.

  10. Numerical investigations of hybrid rocket engines

    Science.gov (United States)

    Betelin, V. B.; Kushnirenko, A. G.; Smirnov, N. N.; Nikitin, V. F.; Tyurenkova, V. V.; Stamov, L. I.

    2018-03-01

    Paper presents the results of numerical studies of hybrid rocket engines operating cycle including unsteady-state transition stage. A mathematical model is developed accounting for the peculiarities of diffusion combustion of fuel in the flow of oxidant, which is composed of oxygen-nitrogen mixture. Three dimensional unsteady-state simulations of chemically reacting gas mixture above thermochemically destructing surface are performed. The results show that the diffusion combustion brings to strongly non-uniform fuel mass regression rate in the flow direction. Diffusive deceleration of chemical reaction brings to the decrease of fuel regression rate in the longitudinal direction.

  11. Contamination-free sounding rocket Langmuir probe

    Science.gov (United States)

    Amatucci, W. E.; Schuck, P. W.; Walker, D. N.; Kintner, P. M.; Powell, S.; Holback, B.; Leonhardt, D.

    2001-04-01

    A technique for removing surface contaminants from a sounding rocket spherical Langmuir probe is presented. Contamination layers present on probe surfaces can skew the collected data, resulting in the incorrect determination of plasma parameters. Despite following the usual probe cleaning techniques that are used prior to a launch, the probe surface can become coated with layers of adsorbed neutral gas in less than a second when exposed to atmosphere. The laboratory tests reported here show that by heating the probe from the interior using a small halogen lamp, adsorbed neutral particles can be removed from the probe surface, allowing accurate plasma parameter measurements to be made.

  12. Nuclear thermal rocket engine operation and control

    International Nuclear Information System (INIS)

    Gunn, S.V.; Savoie, M.T.; Hundal, R.

    1993-06-01

    The operation of a typical Rover/Nerva-derived nuclear thermal rocket (NTR) engine is characterized and the control requirements of the NTR are defined. A rationale for the selection of a candidate diverse redundant NTR engine control system is presented and the projected component operating requirements are related to the state of the art of candidate components and subsystems. The projected operational capabilities of the candidate system are delineated for the startup, full-thrust, shutdown, and decay heat removal phases of the engine operation. 9 refs

  13. Numerical simulations of a sounding rocket in ionospheric plasma: Effects of magnetic field on the wake formation and rocket potential

    Science.gov (United States)

    Darian, D.; Marholm, S.; Paulsson, J. J. P.; Miyake, Y.; Usui, H.; Mortensen, M.; Miloch, W. J.

    2017-09-01

    The charging of a sounding rocket in subsonic and supersonic plasma flows with external magnetic field is studied with numerical particle-in-cell (PIC) simulations. A weakly magnetized plasma regime is considered that corresponds to the ionospheric F2 layer, with electrons being strongly magnetized, while the magnetization of ions is weak. It is demonstrated that the magnetic field orientation influences the floating potential of the rocket and that with increasing angle between the rocket axis and the magnetic field direction the rocket potential becomes less negative. External magnetic field gives rise to asymmetric wake downstream of the rocket. The simulated wake in the potential and density may extend as far as 30 electron Debye lengths; thus, it is important to account for these plasma perturbations when analyzing in situ measurements. A qualitative agreement between simulation results and the actual measurements with a sounding rocket is also shown.

  14. The dimensional changes of a 24 inch diameter vessel during a campaign of successive glass making runs

    International Nuclear Information System (INIS)

    Chidley, B.E.

    1983-01-01

    Throughout the life of a 24 inch diameter Glass Making Vessel used on the inactive Harvest Vitrification Plant dimensional measurements have been recorded. When glass is present the diameter increases due to thermal stresses. Some permanent longitudinal extension (about 1%) occurred over the 20 runs comprising the glass-making campaign. (author)

  15. Physico-Chemical Research on the Sounding Rocket Maser 13

    Science.gov (United States)

    Lockowandt, Christian; Kemi, Stig; Abrahamsson, Mattias; Florin, Gunnar

    MASER is a sounding rocket platform for short-duration microgravity experiments, providing the scientific community with an excellent microgravity tool. The MASER programme has been running by SSC from 1987 and has up to 2012 provided twelve successful flights for microgravity missions with 6-7 minutes of microgravity, the g-level is normally below 1x10-5 g. The MASER 13 is planned to be launched in spring 2015 from Esrange Space Center in Northern Sweden. The rocket will carry four ESA financed experiment modules. The MASER 13 vehicle will be propelled by the 2-stage solid fuel VSB-30 rocket motor, which provided the 390 kg payload with an apogee of 260 km and 6 and a half minutes of microgravity. Swedish Space Corporation carries out the MASER missions for ESA and the program is also available for other customers. The payload comprise four different experiment modules of which three could be defined as physic-chemical research; XRMON-SOL, CDIC-3, MEDI. It also comprises the Maser Service Module and the recovery system. The Service Module provided real-time 5 Mbps down-link of compressed experiment digital video data from the on-board cameras, as well as high-speed housekeeping telemetry data. XRMON-SOL In this experiment the influence of gravity on the formation of an equiaxed microstructure will be investigated. Special attention will be put on the aspect of nucleation, segregation and impingement. The experiment scope is to melt and solidify an AlCu-alloy sample in microgravity. The solidification will be performed in an isothermal environment. The solidification process will be monitored and recorded with X-ray image during the whole flight, images will also be down-linked to ground for real-time monitoring and possible interaction. CDIC-3 The goal is to study in migrogravity the spatio-temporal dynamics of a chemical front travelling in a thin solution layer open to the air and specifically the respective role of Marangoni and density-related hydrodynamic

  16. Software for Collaborative Engineering of Launch Rockets

    Science.gov (United States)

    Stanley, Thomas Troy

    2003-01-01

    The Rocket Evaluation and Cost Integration for Propulsion and Engineering software enables collaborative computing with automated exchange of information in the design and analysis of launch rockets and other complex systems. RECIPE can interact with and incorporate a variety of programs, including legacy codes, that model aspects of a system from the perspectives of different technological disciplines (e.g., aerodynamics, structures, propulsion, trajectory, aeroheating, controls, and operations) and that are used by different engineers on different computers running different operating systems. RECIPE consists mainly of (1) ISCRM a file-transfer subprogram that makes it possible for legacy codes executed in their original operating systems on their original computers to exchange data and (2) CONES an easy-to-use filewrapper subprogram that enables the integration of legacy codes. RECIPE provides a tightly integrated conceptual framework that emphasizes connectivity among the programs used by the collaborators, linking these programs in a manner that provides some configuration control while facilitating collaborative engineering tradeoff studies, including design to cost studies. In comparison with prior collaborative-engineering schemes, one based on the use of RECIPE enables fewer engineers to do more in less time.

  17. A Flight Demonstration of Plasma Rocket Propulsion

    Science.gov (United States)

    Petro, Andrew; Chang-Diaz, Franklin; Schwenterly, WIlliam; Hitt, Michael; Lepore, Joseph

    2000-01-01

    The Advanced Space Propulsion Laboratory at the NASA Johnson Space Center has been engaged in the development of a variable specific impulse magnetoplasma rocket (V ASIMR) for several years. This type of rocket could be used in the future to propel interplanetary spacecraft and has the potential to open the entire solar system to human exploration. One feature of this propulsion technology is the ability to vary its specific impulse so that it can be operated in a mode that maximizes propellant efficiency or a mode that maximizes thrust. Variation of specific impulse and thrust enhances the ability to optimize interplanetary trajectories and results in shorter trip times and lower propellant requirements than with a fixed specific impulse. In its ultimate application for interplanetary travel, the VASIMR would be a multi-megawatt device. A much lower power system is being designed for demonstration in the 2004 timeframe. This first space demonstration would employ a lO-kilowatt thruster aboard a solar powered spacecraft in Earth orbit. The 1O-kilowatt V ASIMR demonstration unit would operate for a period of several months with hydrogen or deuterium propellant with a specific impulse of 10,000 seconds.

  18. The Off-plane Grating Rocket Experiment

    Science.gov (United States)

    Donovan, Benjamin

    2018-01-01

    The next generation of X-ray spectrometers necessitate significant increases in both resolution and effective area to achieve the science goals set forth in the 2010 Decadal Survey and the 2013 Astrophysics Roadmap. The Off-plane Grating Rocket Experiment (OGRE), an X-ray spectroscopy suborbital rocket payload currently scheduled for launch in Q3 2020, will serve as a testbed for several key technologies which can help achieve the desired performance increases of future spectrometers. OGRE will be the first instrument to fly mono-crystalline silicon X-ray mirrors developed at NASA Goddard Space Flight Center. The payload will also utilize an array of off-plane gratings manufactured at The Pennsylvania State University. Additionally, the focal plane will be populated with an array of four electron-multiplying CCDs developed by the Open University and XCAM Ltd. With these key technologies, OGRE hopes to achieve the highest resolution on-sky soft X-ray spectrum to date. We discuss the optical design, expected performance, and the current status of the payload.

  19. Small Rocket/Spacecraft Technology (SMART) Platform

    Science.gov (United States)

    Esper, Jaime; Flatley, Thomas P.; Bull, James B.; Buckley, Steven J.

    2011-01-01

    The NASA Goddard Space Flight Center (GSFC) and the Department of Defense Operationally Responsive Space (ORS) Office are exercising a multi-year collaborative agreement focused on a redefinition of the way space missions are designed and implemented. A much faster, leaner and effective approach to space flight requires the concerted effort of a multi-agency team tasked with developing the building blocks, both programmatically and technologically, to ultimately achieve flights within 7-days from mission call-up. For NASA, rapid mission implementations represent an opportunity to find creative ways for reducing mission life-cycle times with the resulting savings in cost. This in tum enables a class of missions catering to a broader audience of science participants, from universities to private and national laboratory researchers. To that end, the SMART (Small Rocket/Spacecraft Technology) micro-spacecraft prototype demonstrates an advanced avionics system with integrated GPS capability, high-speed plug-and-playable interfaces, legacy interfaces, inertial navigation, a modular reconfigurable structure, tunable thermal technology, and a number of instruments for environmental and optical sensing. Although SMART was first launched inside a sounding rocket, it is designed as a free-flyer.

  20. Gas core nuclear rocket feasibility project

    International Nuclear Information System (INIS)

    Howe, S.D.; DeVolder, B.; Thode, L.; Zerkle, D.

    1997-09-01

    The next giant leap for mankind will be the human exploration of Mars. Almost certainly within the next thirty years, a human crew will brave the isolation, the radiation, and the lack of gravity to walk on and explore the Red planet. However, because the mission distances and duration will be hundreds of times greater than the lunar missions, a human crew will face much greater obstacles and a higher risk than those experienced during the Apollo program. A single solution to many of these obstacles is to dramatically decrease the mission duration by developing a high performance propulsion system. The gas core nuclear rocket (GCNR) has the potential to be such a system. The gas core concept relies on the use of fluid dynamic forces to create and maintain a vortex. The vortex is composed of a fissile material which will achieve criticality and produce high power levels. By radiatively coupling to the surrounding fluids, extremely high temperatures in the propellant and, thus, high specific impulses can be generated. The ship velocities enabled by such performance may allow a 9 month round trip, manned Mars mission to be considered. Alternatively, one might consider slightly longer missions in ships that are heavily shielded against the intense Galactic Cosmic Ray flux to further reduce the radiation dose to the crew. The current status of the research program at the Los Alamos National Laboratory into the gas core nuclear rocket feasibility will be discussed

  1. Nuclear Thermal Rocket Simulation in NPSS

    Science.gov (United States)

    Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas M.

    2013-01-01

    Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic-metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

  2. SCORE - Sounding-rocket Coronagraphic Experiment

    Science.gov (United States)

    Fineschi, Silvano; Moses, Dan; Romoli, Marco

    The Sounding-rocket Coronagraphic Experiment - SCORE - is a The Sounding-rocket Coronagraphic Experiment - SCORE - is a coronagraph for multi-wavelength imaging of the coronal Lyman-alpha lines, HeII 30.4 nm and HI 121.6 nm, and for the broad.band visible-light emission of the polarized K-corona. SCORE has flown successfully in 2009 acquiring the first images of the HeII line-emission from the extended corona. The simultaneous observation of the coronal Lyman-alpha HI 121.6 nm, has allowed the first determination of the absolute helium abundance in the extended corona. This presentation will describe the lesson learned from the first flight and will illustrate the preparations and the science perspectives for the second re-flight approved by NASA and scheduled for 2016. The SCORE optical design is flexible enough to be able to accommodate different experimental configurations with minor modifications. This presentation will describe one of such configurations that could include a polarimeter for the observation the expected Hanle effect in the coronal Lyman-alpha HI line. The linear polarization by resonance scattering of coronal permitted line-emission in the ultraviolet (UV) can be modified by magnetic fields through the Hanle effect. Thus, space-based UV spectro-polarimetry would provide an additional new tool for the diagnostics of coronal magnetism.

  3. Study of flow induce vibration inside 3.5 inch hard disk drives

    Directory of Open Access Journals (Sweden)

    Wichitpon Seepangmon

    2014-06-01

    Full Text Available This study focused on flow induced vibration of head stack assembly (HSA in a 3.5 inch hard disk drive with 5 disks and 10 read/write heads. We studied the effects of air flow on gimbal flex and resonance on arm. The comparison of vibrations on slider between the normal model and the experiment has been done for verifying the model. The peaks of frequency in experiment match the normal model at 1,040 1,320 and 1,400 Hz respectively. After that, the RNG K-ε turbulence model was used to determine the turbulent air flow of 7,200 rpm hard disk drive. The comparison between the normal model and the model with spoiler was investigated by using, computational fluid dynamics software (ANSYS and FLUENT. The results shown velocity magnitudes at the arm were decreased by 0.725 - 57.689 % and pressure dropped by 74.028 - 87.222 %. The velocity magnitudes at the gimbal flex were decreased by 5.522 - 14.291 % and pressure dropped by 48.440 - 82.947 %. The peak of vibrations on arm and gimbal flex was occurred at the frequency 1200 Hz. The model with spoiler could reduce vibration at arm by 2.56 - 95.601 % and reduce vibration at gimbal flex by 4.065 - 95.503 %. In the conclusion, the model with a spoiler could decrease the vibration at all surface of the arm and gimbal flex due to the velocity and pressure reduction[1][4].

  4. Neutron flux density and secondary-particle energy spectra at the 184-inch synchrocyclotron medical facility

    International Nuclear Information System (INIS)

    Smith, A.R.; Schimmerling, W.; Henson, A.M.; Kanstein, L.L.; McCaslin, J.B.; Stephens, L.D.; Thomas, R.H.; Ozawa, J.; Yeater, F.W.

    1978-07-01

    Helium ions, with an energy of 920 MeV, produced by the 184-inch synchrocyclotron of the Lawrence Berkeley Laboratory are now being used in a pilot series to determine their efficacy in the treatment of tumors of large volume. The techniques for production of the large uniform radiation fields required for these treatments involve the use of beam-limiting collimators and energy degraders. Interaction of the primary beam with these beam components produces secondary charged particles and neutrons. The sources of neutron production in the beam transport system of the alpha-particle beam have been identified and their magnitudes have been determined. Measurements with activation detectors and pulse counters of differing energy responses have been used to determine secondary particle spectra at various locations on the patient table. These spectra are compared to a calculation of neutron production based on best estimates derived from published cross sections. Agreement between the calculated spectra and those derived from experimental measurements is obtained (at the 10 to 20% level) when the presence of charged particles is taken into account. The adsorbed dose in soft tissue is not very sensitive to the shape of the incident neutron energy spectrum, and the values obtained from unfolding the experimental measurements agree with the values obtained from the calculated spectra within the estimated uncertainty of +-25%. These values are about 3 x 10 -3 rad on the beam axis and about 1 x 10 -3 rad at 20 cm or more from the beam axis, per rad deposited by the incident alpha-particle beam. Estimates of upper limit dose to the lens of the eye and red bone marrow are approximately 10 rad and approximately 1 rad, respectively, for a typical treatment plan. The absorbed dose to the lens of the eye is thus well below the threshold value for cataractogenesis estimated for fission neutrons. An upper limit for the risk of leukemia is estimated to be approximately 0.04%

  5. Nonlinear Longitudinal Mode Instability in Liquid Propellant Rocket Engine Preburners

    Science.gov (United States)

    Sims, J. D. (Technical Monitor); Flandro, Gary A.; Majdalani, Joseph; Sims, Joseph D.

    2004-01-01

    Nonlinear pressure oscillations have been observed in liquid propellant rocket instability preburner devices. Unlike the familiar transverse mode instabilities that characterize primary combustion chambers, these oscillations appear as longitudinal gas motions with frequencies that are typical of the chamber axial acoustic modes. In several respects, the phenomenon is similar to longitudinal mode combustion instability appearing in low-smoke solid propellant motors. An important feature is evidence of steep-fronted wave motions with very high amplitude. Clearly, gas motions of this type threaten the mechanical integrity of associated engine components and create unacceptably high vibration levels. This paper focuses on development of the analytical tools needed to predict, diagnose, and correct instabilities of this type. For this purpose, mechanisms that lead to steep-fronted, high-amplitude pressure waves are described in detail. It is shown that such gas motions are the outcome of the natural steepening process in which initially low amplitude standing acoustic waves grow into shock-like disturbances. The energy source that promotes this behavior is a combination of unsteady combustion energy release and interactions with the quasi-steady mean chamber flow. Since shock waves characterize the gas motions, detonation-like mechanisms may well control the unsteady combustion processes. When the energy gains exceed the losses (represented mainly by nozzle and viscous damping), the waves can rapidly grow to a finite amplitude limit cycle. Analytical tools are described that allow the prediction of the limit cycle amplitude and show the dependence of this wave amplitude on the system geometry and other design parameters. This information can be used to guide corrective procedures that mitigate or eliminate the oscillations.

  6. Assessment of exposure-response functions for rocket-emission toxicants

    National Research Council Canada - National Science Library

    Subcommittee on Rocket-Emission Toxicants, National Research Council

    ... aborted launch that results in a rocket being destroyed near the ground. Assessment of Exposure-Response Functions for Rocket-Emmission Toxicants evaluates the model and the data used for three rocket emission toxicants...

  7. Evaluation of the Effect of Exhausts from Liquid and Solid Rockets on Ozone Layer

    Science.gov (United States)

    Yamagiwa, Yoshiki; Ishimaki, Tetsuya

    This paper reports the analytical results of the influences of solid rocket and liquid rocket exhausts on ozone layer. It is worried about that the exhausts from solid propellant rockets cause the ozone depletion in the ozone layer. Some researchers try to develop the analytical model of ozone depletion by rocket exhausts to understand its physical phenomena and to find the effective design of rocket to minimize its effect. However, these models do not include the exhausts from liquid rocket although there are many cases to use solid rocket boosters with a liquid rocket at the same time in practical situations. We constructed combined analytical model include the solid rocket exhausts and liquid rocket exhausts to analyze their effects. From the analytical results, we find that the exhausts from liquid rocket suppress the ozone depletion by solid rocket exhausts.

  8. Theoretical and Experimental Analysis of the Physics of Water Rockets

    Science.gov (United States)

    Barrio-Perotti, R.; Blanco-Marigorta, E.; Fernandez-Francos, J.; Galdo-Vega, M.

    2010-01-01

    A simple rocket can be made using a plastic bottle filled with a volume of water and pressurized air. When opened, the air pressure pushes the water out of the bottle. This causes an increase in the bottle momentum so that it can be propelled to fairly long distances or heights. Water rockets are widely used as an educational activity, and several…

  9. Project Stratos; reaching space with a student-built rocket

    NARCIS (Netherlands)

    Haneveer, M.

    2013-01-01

    In the spring of 2009 a team of 15 TU Delft students travelled to Kiruna, Sweden with only one goal: to launch the rocket Stratos I they had been working on for 2 years to an altitude of over 12km, thereby claiming the European Amateur Rocket Altitude record. These students were part of Delft

  10. Nuclear Rocket Facility Decommissioning Project: Controlled Explosive Demolition of Neutron-Activated Shield Wall

    International Nuclear Information System (INIS)

    Michael R, Kruzic

    2008-01-01

    Located in Area 25 of the Nevada Test Site (NTS), the Test Cell A (TCA) Facility (Figure 1) was used in the early to mid-1960s for testing of nuclear rocket engines, as part of the Nuclear Rocket Development Program, to further space travel. Nuclear rocket testing resulted in the activation of materials around the reactors and the release of fission products and fuel particles. The TCA facility, known as Corrective Action Unit 115, was decontaminated and decommissioned (D and D) from December 2004 to July 2005 using the Streamlined Approach for Environmental Restoration (SAFER) process, under the Federal Facility Agreement and Consent Order. The SAFER process allows environmental remediation and facility closure activities (i.e., decommissioning) to occur simultaneously, provided technical decisions are made by an experienced decision maker within the site conceptual site model. Facility closure involved a seven-step decommissioning strategy. First, preliminary investigation activities were performed, including review of process knowledge documentation, targeted facility radiological and hazardous material surveys, concrete core drilling and analysis, shield wall radiological characterization, and discrete sampling, which proved to be very useful and cost-effective in subsequent decommissioning planning and execution and worker safety. Second, site setup and mobilization of equipment and personnel were completed. Third, early removal of hazardous materials, including asbestos, lead, cadmium, and oil, was performed ensuring worker safety during more invasive demolition activities. Process piping was to be verified void of contents. Electrical systems were de-energized and other systems were rendered free of residual energy. Fourth, areas of high radiological contamination were decontaminated using multiple methods. Contamination levels varied across the facility. Fixed beta/gamma contamination levels ranged up to 2 million disintegrations per minute (dpm)/100

  11. Nuclear Rocket Facility Decommissioning Project: Controlled Explosive Demolition of Neutron-Activated Shield Wall

    Energy Technology Data Exchange (ETDEWEB)

    Michael R. Kruzic

    2008-06-01

    Located in Area 25 of the Nevada Test Site (NTS), the Test Cell A (TCA) Facility (Figure 1) was used in the early to mid-1960s for testing of nuclear rocket engines, as part of the Nuclear Rocket Development Program, to further space travel. Nuclear rocket testing resulted in the activation of materials around the reactors and the release of fission products and fuel particles. The TCA facility, known as Corrective Action Unit 115, was decontaminated and decommissioned (D&D) from December 2004 to July 2005 using the Streamlined Approach for Environmental Restoration (SAFER) process, under the Federal Facility Agreement and Consent Order. The SAFER process allows environmental remediation and facility closure activities (i.e., decommissioning) to occur simultaneously, provided technical decisions are made by an experienced decision maker within the site conceptual site model. Facility closure involved a seven-step decommissioning strategy. First, preliminary investigation activities were performed, including review of process knowledge documentation, targeted facility radiological and hazardous material surveys, concrete core drilling and analysis, shield wall radiological characterization, and discrete sampling, which proved to be very useful and cost-effective in subsequent decommissioning planning and execution and worker safety. Second, site setup and mobilization of equipment and personnel were completed. Third, early removal of hazardous materials, including asbestos, lead, cadmium, and oil, was performed ensuring worker safety during more invasive demolition activities. Process piping was to be verified void of contents. Electrical systems were de-energized and other systems were rendered free of residual energy. Fourth, areas of high radiological contamination were decontaminated using multiple methods. Contamination levels varied across the facility. Fixed beta/gamma contamination levels ranged up to 2 million disintegrations per minute (dpm)/100

  12. Computational fluid dynamics and frequency-dependent finite-difference time-domain method coupling for the interaction between microwaves and plasma in rocket plumes

    International Nuclear Information System (INIS)

    Kinefuchi, K.; Funaki, I.; Shimada, T.; Abe, T.

    2012-01-01

    Under certain conditions during rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency transmissions. To understand the relevant physical processes involved in this phenomenon and establish a prediction process for in-flight attenuation levels, we attempted to measure microwave attenuation caused by rocket exhaust plumes in a sea-level static firing test for a full-scale solid propellant rocket motor. The microwave attenuation level was calculated by a coupling simulation of the inviscid-frozen-flow computational fluid dynamics of an exhaust plume and detailed analysis of microwave transmissions by applying a frequency-dependent finite-difference time-domain method with the Drude dispersion model. The calculated microwave attenuation level agreed well with the experimental results, except in the case of interference downstream the Mach disk in the exhaust plume. It was concluded that the coupling estimation method based on the physics of the frozen plasma flow with Drude dispersion would be suitable for actual flight conditions, although the mixing and afterburning in the plume should be considered depending on the flow condition.

  13. Computational fluid dynamics and frequency-dependent finite-difference time-domain method coupling for the interaction between microwaves and plasma in rocket plumes

    Energy Technology Data Exchange (ETDEWEB)

    Kinefuchi, K. [Department of Aeronautics and Astronautics, University of Tokyo, 7-3-1, Hongo, Bunkyo-ku, Tokyo 113-8656 (Japan); Funaki, I.; Shimada, T.; Abe, T. [Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency, 3-1-1, Yoshinodai, Chuo-ku, Sagamihara, Kanagawa 252-5210 (Japan)

    2012-10-15

    Under certain conditions during rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency transmissions. To understand the relevant physical processes involved in this phenomenon and establish a prediction process for in-flight attenuation levels, we attempted to measure microwave attenuation caused by rocket exhaust plumes in a sea-level static firing test for a full-scale solid propellant rocket motor. The microwave attenuation level was calculated by a coupling simulation of the inviscid-frozen-flow computational fluid dynamics of an exhaust plume and detailed analysis of microwave transmissions by applying a frequency-dependent finite-difference time-domain method with the Drude dispersion model. The calculated microwave attenuation level agreed well with the experimental results, except in the case of interference downstream the Mach disk in the exhaust plume. It was concluded that the coupling estimation method based on the physics of the frozen plasma flow with Drude dispersion would be suitable for actual flight conditions, although the mixing and afterburning in the plume should be considered depending on the flow condition.

  14. NASTRAN buckling study of a linear induction motor reaction rail

    Science.gov (United States)

    Williams, J. G.

    1973-01-01

    NASTRAN was used to study problems associated with the installation of a linear induction motor reaction rail test track. Specific problems studied include determination of the critical axial compressive buckling stress and establishment of the lateral stiffness of the reaction rail under combined loads. NASTRAN results were compared with experimentally obtained values and satisfactory agreement was obtained. The reaction rail was found to buckle at an axial compressive stress of 11,400 pounds per square inch. The results of this investigation were used to select procedures for installation of the reaction rail.

  15. Rocket experiment METS - Microwave Energy Transmission in Space

    Science.gov (United States)

    Kaya, N.; Matsumoto, H.; Akiba, R.

    A Microwave Energy Transmission in Space (METS) rocket experiment is being planned by the Solar Power Satellite Working Group at the Institute of Space and Astronautical Science in Japan for the forthcoming International Space Year, 1992. The METS experiment is an advanced version of the previous MINIX rocket experiment (Matsumoto et al., 1990). This paper describes a conceptual design of the METS rocket experiment. It aims at verifying a newly developed microwave energy transmission system for space use and to study nonlinear effects of the microwave energy beam in the space plasma environment. A high power microwave of 936 W will be transmitted by the new phased-array antenna from a mother rocket to a separated target (daughter rocket) through the ionospheric plasma. The active phased-array system has a capability of focusing the microwave energy around any spatial point by controlling the digital phase shifters individually.

  16. Rocket experiment METS Microwave Energy Transmission in Space

    Science.gov (United States)

    Kaya, N.; Matsumoto, H.; Akiba, R.

    A METS (Microwave Energy Transmission in Space) rocket experiment is being planned by the SPS (Solar Power Satellite) Working Group at the Institute of Space and Astronautical Science (ISAS) in Japan for the forthcoming International Space Year (ISY), 1992. The METS experiment is an advanced version of our MINIX rocket experiment. This paper describes the conceptual design for the METS rocket experiment. Aims are to verify the feasibility of a newly developed microwave energy transmission system designed for use in space and to study nonlinear effects of the microwave energy beam on space plasma. A high power microwave (936 W) will be transmitted by a new phase-array antenna from a mother rocket to a separate target (daughter rocket) through the Earth's ionospheric plasma. The active phased-array system has the capability of being able to focus the microwave energy at any spatial point by individually controlling the digital phase shifters.

  17. Photometric observations of local rocket-atmosphere interactions

    Science.gov (United States)

    Greer, R. G. H.; Murtagh, D. P.; Witt, G.; Stegman, J.

    1983-06-01

    Photometric measurements from rocket flights which recorded a strong foreign luminance in the altitude region between 90 and 130 km are reported. From one Nike-Orion rocket the luminance appeared on both up-leg and down-leg; from a series of Petrel rockets the luminance was apparent only on the down-leg. The data suggest that the luminance may be distributed mainly in the wake region along the rocket trajectory. The luminance is believed to be due to a local interaction between the rocket and the atmosphere although the precise nature of the interaction is unknown. It was measured at wavelengths ranging from 275 nm to 1.61 microns and may be caused by a combination of reactions.

  18. Developments in REDES: The Rocket Engine Design Expert System

    Science.gov (United States)

    Davidian, Kenneth O.

    1990-01-01

    The Rocket Engine Design Expert System (REDES) was developed at NASA-Lewis to collect, automate, and perpetuate the existing expertise of performing a comprehensive rocket engine analysis and design. Currently, REDES uses the rigorous JANNAF methodology to analyze the performance of the thrust chamber and perform computational studies of liquid rocket engine problems. The following computer codes were included in REDES: a gas properties program named GASP; a nozzle design program named RAO; a regenerative cooling channel performance evaluation code named RTE; and the JANNAF standard liquid rocket engine performance prediction code TDK (including performance evaluation modules ODE, ODK, TDE, TDK, and BLM). Computational analyses are being conducted by REDES to provide solutions to liquid rocket engine thrust chamber problems. REDES was built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

  19. Electric motor handbook

    CERN Document Server

    Chalmers, B J

    2013-01-01

    Electric Motor Handbook aims to give practical knowledge in a wide range of capacities such as plant design, equipment specification, commissioning, operation and maintenance. The book covers topics such as the modeling of steady-state motor performance; polyphase induction, synchronous, and a.c. commutator motors; ambient conditions, enclosures, cooling and loss dissipation; and electrical supply systems and motor drives. Also covered are topics such as variable-speed drives and motor control; materials and motor components; insulation types, systems, and techniques; and the installation, sit

  20. Lunar mission design using nuclear thermal rockets

    International Nuclear Information System (INIS)

    Stancati, M.L.; Collins, J.T.; Borowski, S.K.

    1991-01-01

    The NERVA-class Nuclear Thermal Rocket (NTR), with performance nearly double that of advanced chemical engines, has long been considered an enabling technology for human missions to Mars. NTR engines address the demanding trip time and payload delivery needs of both cargo-only and piloted flights. But NTR can also reduce the Earth launch requirements for manned lunar missions. First use of NTR for the Moon would be less demanding and would provide a test-bed for early operations experience with this powerful technology. Study of application and design options indicates that NTR propulsion can be integrated with the Space Exploration Initiative scenarios to deliver performance gains while managing controlled, long-term disposal of spent reactors to highly stable orbits

  1. Rocket Engine Innovations Advance Clean Energy

    Science.gov (United States)

    2012-01-01

    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.

  2. Additive Manufacturing a Liquid Hydrogen Rocket Engine

    Science.gov (United States)

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

    2016-01-01

    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. Rockets: Physical science teacher's guide with activities

    Science.gov (United States)

    Vogt, Gregory L.; Rosenberg, Carla R. (Editor)

    1993-01-01

    This guide begins with background information sections on the history of rocketry, scientific principles, and practical rocketry. The sections on scientific principles and practical rocketry are based on Isaac Newton's three laws of motion. These laws explain why rockets work and how to make them more efficient. The background sections are followed with a series of physical science activities that demonstrate the basic science of rocketry. Each activity is designed to be simple and take advantage of inexpensive materials. Construction diagrams, materials and tools lists, and instructions are included. A brief discussion elaborates on the concepts covered in the activities and is followed with teaching notes and discussion questions. The guide concludes with a glossary of terms, suggested reading list, NASA educational resources, and an evaluation questionnaire with a mailer.

  4. Tidal analysis of Met rocket wind data

    Science.gov (United States)

    Bedinger, J. F.; Constantinides, E.

    1976-01-01

    A method of analyzing Met Rocket wind data is described. Modern tidal theory and specialized analytical techniques were used to resolve specific tidal modes and prevailing components in observed wind data. A representation of the wind which is continuous in both space and time was formulated. Such a representation allows direct comparison with theory, allows the derivation of other quantities such as temperature and pressure which in turn may be compared with observed values, and allows the formation of a wind model which extends over a broader range of space and time. Significant diurnal tidal modes with wavelengths of 10 and 7 km were present in the data and were resolved by the analytical technique.

  5. Optical measurements in rocket engine liquid sprays

    Science.gov (United States)

    Feikema, Douglas A.

    1994-01-01

    The performance of liquid propellant rocket engines is dependent upon many elements of the entire system. One of the most fundamental and most critical is the performance of the injector elements. Their characterization is an important part of the development of combustion devices. Optical measurements within these environments have proven to be invaluable tools in quantifying the physical environment of two phase flows. The effort reported herein involves the measurement of drop velocity, drop size, and most importantly mass flux using Phase-Doppler Particle Anemometry within a spray generated by a single swirl injector element operating in atmospheric pressure conditions. The mass flux has been determined and validated by mechanical patternation methods and by profile integration of the mass flux.

  6. Rocket Testing and Integrated System Health Management

    Science.gov (United States)

    Figueroa, Fernando; Schmalzel, John

    2005-01-01

    Integrated System Health Management (ISHM) describes a set of system capabilities that in aggregate perform: determination of condition for each system element, detection of anomalies, diagnosis of causes for anomalies, and prognostics for future anomalies and system behavior. The ISHM should also provide operators with situational awareness of the system by integrating contextual and timely data, information, and knowledge (DIaK) as needed. ISHM capabilities can be implemented using a variety of technologies and tools. This chapter provides an overview of ISHM contributing technologies and describes in further detail a novel implementation architecture along with associated taxonomy, ontology, and standards. The operational ISHM testbed is based on a subsystem of a rocket engine test stand. Such test stands contain many elements that are common to manufacturing systems, and thereby serve to illustrate the potential benefits and methodologies of the ISHM approach for intelligent manufacturing.

  7. Reusable Rocket Engine Turbopump Health Management System

    Science.gov (United States)

    Surko, Pamela

    1994-01-01

    A health monitoring expert system software architecture has been developed to support condition-based health monitoring of rocket engines. Its first application is in the diagnosis decisions relating to the health of the high pressure oxidizer turbopump (HPOTP) of Space Shuttle Main Engine (SSME). The post test diagnostic system runs off-line, using as input the data recorded from hundreds of sensors, each running typically at rates of 25, 50, or .1 Hz. The system is invoked after a test has been completed, and produces an analysis and an organized graphical presentation of the data with important effects highlighted. The overall expert system architecture has been developed and documented so that expert modules analyzing other line replaceable units may easily be added. The architecture emphasizes modularity, reusability, and open system interfaces so that it may be used to analyze other engines as well.

  8. Particle bed reactor nuclear rocket concept

    International Nuclear Information System (INIS)

    Ludewig, H.

    1991-01-01

    The particle bed reactor nuclear rocket concept consists of fuel particles (in this case (U,Zr)C with an outer coat of zirconium carbide). These particles are packed in an annular bed surrounded by two frits (porous tubes) forming a fuel element; the outer one being a cold frit, the inner one being a hot frit. The fuel element are cooled by hydrogen passing in through the moderator. These elements are assembled in a reactor assembly in a hexagonal pattern. The reactor can be either reflected or not, depending on the design, and either 19 or 37 elements, are used. Propellant enters in the top, passes through the moderator fuel element and out through the nozzle. Beryllium used for the moderator in this particular design to withstand the high radiation exposure implied by the long run times

  9. Formulation, Casting, and Evaluation of Paraffin-Based Solid Fuels Containing Energetic and Novel Additives for Hybrid Rockets

    Science.gov (United States)

    Larson, Daniel B.; Desain, John D.; Boyer, Eric; Wachs, Trevor; Kuo, Kenneth K.; Borduin, Russell; Koo, Joseph H.; Brady, Brian B.; Curtiss, Thomas J.; Story, George

    2012-01-01

    This investigation studied the inclusion of various additives to paraffin wax for use in a hybrid rocket motor. Some of the paraffin-based fuels were doped with various percentages of LiAlH4 (up to 10%). Addition of LiAlH4 at 10% was found to increase regression rates between 7 - 10% over baseline paraffin through tests in a gaseous oxygen hybrid rocket motor. Mass burn rates for paraffin grains with 10% LiAlH4 were also higher than those of the baseline paraffin. RDX was also cast into a paraffin sample via a novel casting process which involved dissolving RDX into dimethylformamide (DMF) solvent and then drawing a vacuum on the mixture of paraffin and RDX/DMF in order to evaporate out the DMF. It was found that although all DMF was removed, the process was not conducive to generating small RDX particles. The slow boiling generated an inhomogeneous mixture of paraffin and RDX. It is likely that superheating the DMF to cause rapid boiling would likely reduce RDX particle sizes. In addition to paraffin/LiAlH4 grains, multi-walled carbon nanotubes (MWNT) were cast in paraffin for testing in a hybrid rocket motor, and assorted samples containing a range of MWNT percentages in paraffin were imaged using SEM. The fuel samples showed good distribution of MWNT in the paraffin matrix, but the MWNT were often agglomerated, indicating that a change to the sonication and mixing processes were required to achieve better uniformity and debundled MWNT. Fuel grains with MWNT fuel grains had slightly lower regression rate, likely due to the increased thermal conductivity to the fuel subsurface, reducing the burning surface temperature.

  10. MLEP-Fail calibration for 1/8 inch thick cast plate of 17-4 steel.

    Energy Technology Data Exchange (ETDEWEB)

    Corona, Edmundo [Sandia National Lab. (SNL-NM), Albuquerque, NM (United States)

    2018-02-01

    The purpose of the work presented in this memo was to calibrate the Sierra material model Multilinear Elastic-Plastic Hardening Model with Failure (MLEP-Fail) for 1/8 inch thick cast plate of 17-4 steel. The calibration approach is essentially the same as that recently used in a previous memo using data from smooth and notched tensile specimens. The notched specimens were manufactured with three notch radii R = 1=8, 1/32 and 1/64 inches. The dimensions of the smooth and notched specimens are given in the prints in Appendix A. Two cast plates, Plate 3 and Plate 4, with nominally identical properties were considered.

  11. Handbook on linear motor application

    International Nuclear Information System (INIS)

    1988-10-01

    This book guides the application for Linear motor. It lists classification and speciality of Linear Motor, terms of linear-induction motor, principle of the Motor, types on one-side linear-induction motor, bilateral linear-induction motor, linear-DC Motor on basic of the motor, linear-DC Motor for moving-coil type, linear-DC motor for permanent-magnet moving type, linear-DC motor for electricity non-utility type, linear-pulse motor for variable motor, linear-pulse motor for permanent magneto type, linear-vibration actuator, linear-vibration actuator for moving-coil type, linear synchronous motor, linear electromagnetic motor, linear electromagnetic solenoid, technical organization and magnetic levitation and linear motor and sensor.

  12. Interacting adiabatic quantum motor

    Science.gov (United States)

    Bruch, Anton; Kusminskiy, Silvia Viola; Refael, Gil; von Oppen, Felix

    2018-05-01

    We present a field-theoretic treatment of an adiabatic quantum motor. We explicitly discuss a motor called the Thouless motor which is based on a Thouless pump operating in reverse. When a sliding periodic potential is considered to be the motor degree of freedom, a bias voltage applied to the electron channel sets the motor in motion. We investigate a Thouless motor whose electron channel is modeled as a Luttinger liquid. Interactions increase the gap opened by the periodic potential. For an infinite Luttinger liquid the coupling-induced friction is enhanced by electron-electron interactions. When the Luttinger liquid is ultimately coupled to Fermi liquid reservoirs, the dissipation reduces to its value for a noninteracting electron system for a constant motor velocity. Our results can also be applied to a motor based on a nanomagnet coupled to a quantum spin Hall edge.

  13. Chronic motor tic disorder

    Science.gov (United States)

    Chronic vocal tic disorder; Tic - chronic motor tic disorder ... Chronic motor tic disorder is more common than Tourette syndrome . Chronic tics may be forms of Tourette syndrome. Tics usually start ...

  14. Holographic elements and curved slit used to enlarge field of view in rocket detection system

    Science.gov (United States)

    Breton, Mélanie; Fortin, Jean; Lessard, Roger A.; Châteauneuf, Marc

    2006-09-01

    Rocket detection over a wide field of view is an important issue in the protection of light armored vehicle. Traditionally, the detection occurs in UV band, but recent studies have shown the existence of significant emission peaks in the visible and near infrared at rocket launch time. The use of the visible region is interesting in order to reduce the weight and cost of systems. Current methods to detect those specific peaks involve use of interferometric filters. However, they fail to combine wide angle with wavelength selectivity. A linear array of volume holographic elements combined with a curved exit slit is proposed for the development of a wide field of view sensor for the detection of solid propellant motor launch flash. The sensor is envisaged to trigger an active protection system. On the basis of geometric theory, a system has been designed. It consists of a collector, a linear array of holographic elements, a curved slit and a detector. The collector is an off-axis parabolic mirror. Holographic elements are recorded subdividing a hologram film in regions, each individually exposed with a different incidence angle. All regions have a common diffraction angle. The incident angle determines the instantaneous field of view of the elements. The volume hologram performs the function of separating and focusing the diffracted beam on an image plane to achieve wavelength filtering. Conical diffraction property is used to enlarge the field of view in elevation. A curved slit was designed to correspond to oblique incidence of the holographic linear array. It is situated at the image plane and filters the diffracted spectrum toward the sensor. The field of view of the design was calculated to be 34 degrees. This was validated by a prototype tested during a field trial. Results are presented and analyzed. The system succeeded in detecting the rocket launch flash at desired fields of view.

  15. Piezoelectric Motors, an Overview

    OpenAIRE

    Karl Spanner; Burhanettin Koc

    2016-01-01

    Piezoelectric motors are used in many industrial and commercial applications. Various piezoelectric motors are available in the market. All of the piezoelectric motors use the inverse piezoelectric effect, where microscopically small oscillatory motions are converted into continuous or stepping rotary or linear motions. Methods of obtaining long moving distance have various drive and functional principles that make these motors categorized into three groups: resonance-drive (piezoelectric ult...

  16. Electric Motor Thermal Management

    Energy Technology Data Exchange (ETDEWEB)

    Bennion, Kevin S [National Renewable Energy Laboratory (NREL), Golden, CO (United States)

    2017-09-01

    Thermal management enables more efficient and cost-effective motors. This Annual Merit Review presentation describes the technical accomplishments and progress in electric motor thermal management R&D over the last year. This project supports a broad industry demand for data, analysis methods, and experimental techniques to improve and better understand motor thermal management.

  17. Programmable dc motor controller

    Science.gov (United States)

    Hopwood, J. E.

    1982-11-01

    A portable programmable dc motor controller, with features not available on commercial instruments was developed for controlling fixtures during welding processes. The controller can be used to drive any dc motor having tachometer feedback and motor requirements not exceeding 30 volts, 3 amperes. Among the controller's features are delayed start time, upslope time, speed, and downslope time.

  18. THE ADIABATIC DEMAGNETIZATION REFRIGERATOR FOR THE MICRO-X SOUNDING ROCKET TELESCOPE

    International Nuclear Information System (INIS)

    Wikus, P.; Bagdasarova, Y.; Figueroa-Feliciano, E.; Leman, S. W.; Rutherford, J. M.; Trowbridge, S. N.; Adams, J. S.; Bandler, S. R.; Eckart, M. E.; Kelley, R. L.; Kilbourne, C. A.; Porter, F. S.; Doriese, W. B.; McCammon, D.

    2010-01-01

    The Micro-X Imaging X-ray Spectrometer is a sounding rocket payload slated for launch in 2011. An array of Transition Edge Sensors, which is operated at a bath temperature of 50 mK, will be used to obtain a high resolution spectrum of the Puppis-A supernova remnant. An Adiabatic Demagnetization Refrigerator (ADR) with a 75 gram Ferric Ammonium Alum (FAA) salt pill in the bore of a 4 T superconducting magnet provides a stable heat sink for the detector array only a few seconds after burnout of the rocket motors. This requires a cold stage design with very short thermal time constants. A suspension made from Kevlar strings holds the 255 gram cold stage in place. It is capable of withstanding loads in excess of 200 g. Stable operation of the TES array in proximity to the ADR magnet is ensured by a three-stage magnetic shielding system which consists of a superconducting can, a high-permeability shield and a bucking coil. The development and testing of the Micro-X payload is well underway.

  19. Investigation of Post-Flight Solid Rocket Booster Thermal Protection System

    Science.gov (United States)

    Nelson, Linda A.

    2006-01-01

    After every Shuttle mission, the Solid Rocket Boosters (SRBs) are recovered and observed for missing material. Most of the SRB is covered with a cork-based thermal protection material (MCC-l). After the most recent shuttle mission, STS-114, the forward section of the booster appeared to have been impacted during flight. The darkened fracture surfaces indicated that this might have occurred early in flight. The scope of the analysis included microscopic observations to assess the degree of heat effects and locate evidence of the impact source as well as chemical analysis of the fracture surfaces and recovered foreign material using Fourier Transform Infrared Spectroscopy and Scanning Electron Microscopy/Energy Dispersive Spectroscopy. The amount of heat effects and presence of soot products on the fracture surface indicated that the material was impacted prior to SRB re-entry into the atmosphere. Fragments of graphite fibers found on these fracture surfaces were traced to slag inside the Solid Rocket Motor (SRM) that forms during flight as the propellant is spent and is ejected throughout the descent of the SRB after separation. The direction of the impact mark matches with the likely trajectory of SRBs tumbling prior to re-entry.

  20. Thrust Augmented Nozzle for a Hybrid Rocket with a Helical Fuel Port

    Science.gov (United States)

    Marshall, Joel H.

    A thrust augmented nozzle for hybrid rocket systems is investigated. The design lever-ages 3-D additive manufacturing to embed a helical fuel port into the thrust chamber of a hybrid rocket burning gaseous oxygen and ABS plastic as propellants. The helical port significantly increases how quickly the fuel burns, resulting in a fuel-rich exhaust exiting the nozzle. When a secondary gaseous oxygen flow is injected into the nozzle downstream of the throat, all of the remaining unburned fuel in the plume spontaneously ignites. This secondary reaction produces additional high pressure gases that are captured by the nozzle and significantly increases the motor's performance. Secondary injection and combustion allows a high expansion ratio (area of the nozzle exit divided by area of the throat) to be effective at low altitudes where there would normally be significantly flow separation and possibly an embedded shock wave due. The result is a 15 percent increase in produced thrust level with no loss in engine efficiency due to secondary injection. Core flow efficiency was increased significantly. Control tests performed using cylindrical fuel ports with secondary injection, and helical fuel ports without secondary injection did not exhibit this performance increase. Clearly, both the fuel-rich plume and secondary injection are essential features allowing the hybrid thrust augmentation to occur. Techniques for better design optimization are discussed.

  1. Estimates of the radiation environment for a nuclear rocket engine

    International Nuclear Information System (INIS)

    Courtney, J.C.; Manohara, H.M.; Williams, M.L.

    1992-01-01

    Ambitious missions in deep space, such as manned expeditions to Mars, require nuclear propulsion if they are to be accomplished in a reasonable length of time. Current technology is adequate to support the use of nuclear fission as a source of energy for propulsion; however, problems associated with neutrons and gammas leaking from the rocket engine must be addressed. Before manned or unmanned space flights are attempted, an extensive ground test program on the rocket engine must be completed. This paper compares estimated radiation levels and nuclear heating rates in and around the rocket engine for both a ground test and space environments

  2. Water Impact Prediction Tool for Recoverable Rockets

    Science.gov (United States)

    Rooker, William; Glaese, John; Clayton, Joe

    2011-01-01

    Reusing components from a rocket launch can be cost saving. NASA's space shuttle system has reusable components that return to the Earth and impact the ocean. A primary example is the Space Shuttle Solid Rocket Booster (SRB) that descends on parachutes to the Earth after separation and impacts the ocean. Water impact generates significant structural loads that can damage the booster, so it is important to study this event in detail in the design of the recovery system. Some recent examples of damage due to water impact include the Ares I-X First Stage deformation as seen in Figure 1 and the loss of the SpaceX Falcon 9 First Stage.To ensure that a component can be recovered or that the design of the recovery system is adequate, an adequate set of structural loads is necessary for use in failure assessments. However, this task is difficult since there are many conditions that affect how a component impacts the water and the resulting structural loading that a component sees. These conditions include the angle of impact with respect to the water, the horizontal and vertical velocities, the rotation rate, the wave height and speed, and many others. There have been attempts to simulate water impact. One approach is to analyze water impact using explicit finite element techniques such as those employed by the LS-Dyna tool [1]. Though very detailed, this approach is time consuming and would not be suitable for running Monte Carlo or optimization analyses. The purpose of this paper is to describe a multi-body simulation tool that runs quickly and that captures the environments a component might see. The simulation incorporates the air and water interaction with the component, the component dynamics (i.e. modes and mode shapes), any applicable parachutes and lines, the interaction of winds and gusts, and the wave height and speed. It is capable of quickly conducting Monte Carlo studies to better capture the environments and genetic algorithm optimizations to reproduce a

  3. Motor/generator

    Science.gov (United States)

    Hickam, Christopher Dale [Glasford, IL

    2008-05-13

    A motor/generator is provided for connecting between a transmission input shaft and an output shaft of a prime mover. The motor/generator may include a motor/generator housing, a stator mounted to the motor/generator housing, a rotor mounted at least partially within the motor/generator housing and rotatable about a rotor rotation axis, and a transmission-shaft coupler drivingly coupled to the rotor. The transmission-shaft coupler may include a clamp, which may include a base attached to the rotor and a plurality of adjustable jaws.

  4. Piezoelectric Motors, an Overview

    Directory of Open Access Journals (Sweden)

    Karl Spanner

    2016-02-01

    Full Text Available Piezoelectric motors are used in many industrial and commercial applications. Various piezoelectric motors are available in the market. All of the piezoelectric motors use the inverse piezoelectric effect, where microscopically small oscillatory motions are converted into continuous or stepping rotary or linear motions. Methods of obtaining long moving distance have various drive and functional principles that make these motors categorized into three groups: resonance-drive (piezoelectric ultrasonic motors, inertia-drive, and piezo-walk-drive. In this review, a comprehensive summary of piezoelectric motors, with their classification from initial idea to recent progress, is presented. This review also includes some of the industrial and commercial applications of piezoelectric motors that are presently available in the market as actuators.

  5. Motor degradation prediction methods

    Energy Technology Data Exchange (ETDEWEB)

    Arnold, J.R.; Kelly, J.F.; Delzingaro, M.J.

    1996-12-01

    Motor Operated Valve (MOV) squirrel cage AC motor rotors are susceptible to degradation under certain conditions. Premature failure can result due to high humidity/temperature environments, high running load conditions, extended periods at locked rotor conditions (i.e. > 15 seconds) or exceeding the motor`s duty cycle by frequent starts or multiple valve stroking. Exposure to high heat and moisture due to packing leaks, pressure seal ring leakage or other causes can significantly accelerate the degradation. ComEd and Liberty Technologies have worked together to provide and validate a non-intrusive method using motor power diagnostics to evaluate MOV rotor condition and predict failure. These techniques have provided a quick, low radiation dose method to evaluate inaccessible motors, identify degradation and allow scheduled replacement of motors prior to catastrophic failures.

  6. Neuroplasticity & Motor Learning

    DEFF Research Database (Denmark)

    Jensen, Jesper Lundbye

    Practice of a new motor task is usually associated with an improvement in performance. Indeed, if we stop practicing and return the next day to the same task, we find that our performance has been maintained and may even be better than it was at the start of the first day. This improvement...... is a measure of our ability to form and store a motor memory of the task. However, the initial memory of the task is labile and may be subject to interference. During and following motor learning plastic changes occur within the central nervous system. On one hand these changes are driven by motor practice......, on the other hand the changes underlie the formation of motor memory and the retention of improved motor performance. During motor learning changes may occur at many different levels within the central nervous system dependent on the type of task and training. Here, we demonstrate different studies from our...

  7. Computational and Experimental Investigation of Liquid Propellant Rocket Combustion Instability

    Data.gov (United States)

    National Aeronautics and Space Administration — Combustion instability has been a problem faced by rocket engine developers since the 1940s. The complicated phenomena has been highly unpredictable, causing engine...

  8. Simulation and experimental research on line throwing rocket with flight

    Directory of Open Access Journals (Sweden)

    Wen-bin Gu

    2014-06-01

    Full Text Available The finite segment method is used to model the line throwing rocket system. A dynamic model of line throwing rocket with flight motion based on Kane's method is presented by the kinematics description of the system and the consideration of the forces acting on the system. The experiment designed according to the parameters of the dynamic model is made. The simulation and experiment results, such as range, velocity and flight time, are compared and analyzed. The simulation results are basically agreed with the test data, which shows that the flight motion of the line throwing rocket can be predicted by the dynamic model. A theoretical model and guide for the further research on the disturbance of rope and the guidance, flight control of line throwing rocket are provided by the dynamic modeling.

  9. Propellant Flow Actuated Piezoelectric Rocket Engine Igniter, Phase I

    Data.gov (United States)

    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,...

  10. Dynamical Model of Rocket Propellant Loading with Liquid Hydrogen

    Data.gov (United States)

    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...

  11. Magnesium Based Rockets for Martian Exploration, Phase I

    Data.gov (United States)

    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)...

  12. Distributed Rocket Engine Testing Health Monitoring System, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — The on-ground and Distributed Rocket Engine Testing Health Monitoring System (DiRETHMS) provides a system architecture and software tools for performing diagnostics...

  13. Distributed Rocket Engine Testing Health Monitoring System, Phase II

    Data.gov (United States)

    National Aeronautics and Space Administration — Leveraging the Phase I achievements of the Distributed Rocket Engine Testing Health Monitoring System (DiRETHMS) including its software toolsets and system building...

  14. Infrasound from the 2009 and 2017 DPRK rocket launches

    Science.gov (United States)

    Evers, L. G.; Assink, J. D.; Smets, P. SM

    2018-06-01

    Supersonic rockets generate low-frequency acoustic waves, that is, infrasound, during the launch and re-entry. Infrasound is routinely observed at infrasound arrays from the International Monitoring System, in place for the verification of the Comprehensive Nuclear-Test-Ban Treaty. Association and source identification are key elements of the verification system. The moving nature of a rocket is a defining criterion in order to distinguish it from an isolated explosion. Here, it is shown how infrasound recordings can be associated, which leads to identification of the rocket. Propagation modelling is included to further constrain the source identification. Four rocket launches by the Democratic People's Republic of Korea in 2009 and 2017 are analysed in which multiple arrays detected the infrasound. Source identification in this region is important for verification purposes. It is concluded that with a passive monitoring technique such as infrasound, characteristics can be remotely obtained on sources of interest, that is, infrasonic intelligence, over 4500+ km.

  15. Manufacturing Advanced Channel Wall Rocket Liners, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — This SBIR will adapt and demonstrate a low cost flexible method of manufacturing channel wall liquid rocket nozzles and combustors, while providing developers a...

  16. Defense Against Rocket Attacks in the Presence of False Cues

    National Research Council Canada - National Science Library

    Harari, Lior

    2008-01-01

    Rocket attacks on civilian and military targets, from both Hezbollah (South Lebanon) and Hamas (Gaza strip) have been causing a major operational problem for the Israeli Defense Force for over two decades...

  17. LOX/Methane Regeneratively-Cooled Rocket Engine Development

    Data.gov (United States)

    National Aeronautics and Space Administration — The purpose of this project is to advance the technologies required to build a subcritical regeneratively cooled liquid oxygen/methane rocket combustion chamber for...

  18. Advanced Vortex Hybrid Rocket Engine (AVHRE), Phase II

    Data.gov (United States)

    National Aeronautics and Space Administration — ORBITEC proposes to develop a unique Advanced Vortex Hybrid Rocket Engine (AVHRE) to achieve a safe, highly-reliable, low-cost and uniquely versatile propulsion...

  19. Taming Liquid Hydrogen: The Centaur Upper Stage Rocket

    Science.gov (United States)

    Dawson, Virginia P.; Bowles, Mark D.

    2004-01-01

    The Centaur is one of the most powerful rockets in the world. As an upper-stage rocket for the Atlas and Titan boosters it has been a reliable workhorse for NASA for over forty years and has played an essential role in many of NASA's adventures into space. In this CD-ROM you will be able to explore the Centaur's history in various rooms to this virtual museum. Visit the "Movie Theater" to enjoy several video documentaries on the Centaur. Enter the "Interview Booth" to hear and read interviews with scientists and engineers closely responsible for building and operating the rocket. Go to the "Photo Gallery" to look at numerous photos of the rocket throughout its history. Wander into the "Centaur Library" to read various primary documents of the Centaur program. Finally, stop by the "Observation Deck" to watch a virtual Centaur in flight.

  20. Advanced Vortex Hybrid Rocket Engine (AVHRE), Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — Orbital Technologies Corporation (ORBITEC) proposes to develop a unique Advanced Vortex Hybrid Rocket Engine (AVHRE) to achieve a highly-reliable, low-cost and...