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Sample records for rocket systems volume

  1. Space shuttle solid rocket booster recovery system definition, volume 1

    NASA Technical Reports Server (NTRS)

    1973-01-01

    The performance requirements, preliminary designs, and development program plans for an airborne recovery system for the space shuttle solid rocket booster are discussed. The analyses performed during the study phase of the program are presented. The basic considerations which established the system configuration are defined. A Monte Carlo statistical technique using random sampling of the probability distribution for the critical water impact parameters was used to determine the failure probability of each solid rocket booster component as functions of impact velocity and component strength capability.

  2. Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study, volume 2

    NASA Technical Reports Server (NTRS)

    1989-01-01

    The Liquid Rocket Booster (LRB) Systems Definition Handbook presents the analyses and design data developed during the study. The Systems Definition Handbook (SDH) contains three major parts: the LRB vehicles definition; the Pressure-Fed Booster Test Bed (PFBTB) study results; and the ALS/LRB study results. Included in this volume are the results of all trade studies; final configurations with supporting rationale and analyses; technology assessments; long lead requirements for facilities, materials, components, and subsystems; operational requirements and scenarios; and safety, reliability, and environmental analyses.

  3. Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Volume 1: Executive summary

    NASA Technical Reports Server (NTRS)

    1989-01-01

    The feasibility of replacing the Space Transportation System (STS) Solid Rocket Boosters (SRBs) with Liquid Rocket Boosters (LRBs) was studied. Feasibility required acceptable technical risk, program costs, and a program plan which supports STS requirements. Three major goals were identified to direct booster design and operation: increased STS safety and reliability; STS/LRB integration with minimum impact; and increased STS performance. Two booster engine designs were studied. The first engine design was a turbopump-fed engine with state-of-the-art design, and the second was a pressure-fed engine which might provide a lower cost alternative to the pump-fed concept. Both booster concepts were carried through to completion of conceptual design and all system impacts and program costs were identified. Applications for LRB use in the Advanced Launch System (ALS) program were studied using the pump-fed LRB baseline concept and variations on the baseline concept. Support for the Pressure-Fed Booster Test Bed (PFBTB) included test program planning and costs and technical support.

  4. Operationally efficient propulsion system study (OEPSS) data book. Volume 10; Air Augmented Rocket Afterburning

    NASA Technical Reports Server (NTRS)

    Farhangi, Shahram; Trent, Donnie (Editor)

    1992-01-01

    A study was directed towards assessing viability and effectiveness of an air augmented ejector/rocket. Successful thrust augmentation could potentially reduce a multi-stage vehicle to a single stage-to-orbit vehicle (SSTO) and, thereby, eliminate the associated ground support facility infrastructure and ground processing required by the eliminated stage. The results of this preliminary study indicate that an air augmented ejector/rocket propulsion system is viable. However, uncertainties resulting from simplified approach and assumptions must be resolved by further investigations.

  5. Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Volume 2: Addendum 1

    NASA Technical Reports Server (NTRS)

    1990-01-01

    The potential of a common Liquid Rocket Booster (LRB) design was evaluated for use with both the Space Transportation System (STS) and the Advanced Launch System (ALS). A goal is to have a common Liquid Oxygen/Liquid Hydrogen (LO2/LH2) engine developed for both the ALS booster and the core stage. The LO2/LH2 option for the STS was evaluated to identify potential LRB program cost reductions. The objective was to identify the structural impacts to the external tank (ET), and to determine if any significant ET re-development costs are required as a result of the larger LO2/LH2 LRB. The potential ET impacts evaluated are presented.

  6. Rocket pollution reduction system

    SciTech Connect

    Geisler, R.L.

    1994-01-04

    A system is provided for reducing the emissions of hydrochloric acid (HCl) from solid fuel rockets, especially during ground disposal. An aqueous solution of an alkali metal hydroxide is injected as a mist into the rocket chamber as the rocket fuel is burned. The reaction of the alkali metal with hydrogen chloride (HCl) produces a salt and thereby minimizes the presence of hydrochloric acid in the rocket exhaust. An injected neutralizing material which reduces hydrochloric acid, but which produces less thrust than an equal weight of rocket fuel, can be injected into an operating rocket which carries a payload high above the earth, with the injected material being injected only while the rocket is at a lower altitude when hydrochloric acid is most undesirable. The injected material can be produced by a small auxiliary rocket device whose exhaust is delivered directly to the main rocket chamber, and with the exhaust of the auxiliary rocket device including a high proportion of magnesium to react with the hydrochloric acid with minimal degradation of rocket performance. 4 figs.

  7. Rockets and People. Volume 1

    NASA Technical Reports Server (NTRS)

    Chertok, Boris E; Siddiqi, Asif A. (Editor)

    2005-01-01

    Much has been written in the West on the history of the Soviet space program but few Westerners have read direct first-hand accounts of the men and women who were behind the many Russian accomplishments in exploring space.The memoirs of Academician Boris Chertok, translated from the original Russian, fills that gap.Chertok began his career as an electrician in 1930 at an aviation factory near Moscow.Twenty-seven years later, he became deputy to the founding figure of the Soviet space program, the mysterious Chief Designer Sergey Korolev. Chertok s sixty-year-long career and the many successes and failures of the Soviet space program constitute the core of his memoirs, Rockets and People. These writings are spread over four volumes. This is volume I. Academician Chertok not only describes and remembers, but also elicits and extracts profound insights from an epic story about a society s quest to explore the cosmos. In Volume 1, Chertok describes his early years as an engineer and ends with the mission to Germany after the end of World War II when the Soviets captured Nazi missile technology and expertise. Volume 2 takes up the story with the development of the world s first intercontinental ballistic missile ICBM) and ends with the launch of Sputnik and the early Moon probes. In Volume 3, Chertok recollects the great successes of the Soviet space program in the 1960s including the launch of the world s first space voyager Yuriy Gagarin as well as many events connected with the Cold War. Finally, in Volume 4, Chertok meditates at length on the massive Soviet lunar project designed to beat the Americans to the Moon in the 1960s, ending with his remembrances of the Energiya-Buran project.

  8. Liquid rocket booster study. Volume 2, book 6, appendix 10: Vehicle systems effects

    NASA Technical Reports Server (NTRS)

    1989-01-01

    Three tasks were undertaken by Eagle Engineering as a part of the Liquid Rocket Booster (LRB) study. Task 1 required Eagle to supply current data relative to the Space Shuttle vehicle and systems affected by an LRB substitution. Tables listing data provided are presented. Task 2 was to evaluate and compare shuttle impacts of candidate LRB configuration in concert with overall trades of analysis activity. Three selected configurations with emphasis on flight loads, separation dynamics, and cost comparison are presented. Task 3 required the development of design guidelines and requirements to minimize impacts to the Space Shuttle system from all LRB substitution. Results are presented for progress to date.

  9. Laser rocket system analysis

    NASA Technical Reports Server (NTRS)

    Jones, W. S.; Forsyth, J. B.; Skratt, J. P.

    1979-01-01

    The laser rocket systems investigated in this study were for orbital transportation using space-based, ground-based and airborne laser transmitters. The propulsion unit of these systems utilizes a continuous wave (CW) laser beam focused into a thrust chamber which initiates a plasma in the hydrogen propellant, thus heating the propellant and providing thrust through a suitably designed nozzle and expansion skirt. The specific impulse is limited only by the ability to adequately cool the thruster and the amount of laser energy entering the engine. The results of the study showed that, with advanced technology, laser rocket systems with either a space- or ground-based laser transmitter could reduce the national budget allocated to space transportation by 10 to 345 billion dollars over a 10-year life cycle when compared to advanced chemical propulsion systems (LO2-LH2) of equal capability. The variation in savings depends upon the projected mission model.

  10. Mars Rocket Propulsion System

    NASA Technical Reports Server (NTRS)

    Zubrin, Robert; Harber, Dan; Nabors, Sammy

    2008-01-01

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

  11. Exergy Analysis of Rocket Systems

    NASA Technical Reports Server (NTRS)

    Gilbert, Andrew; Mesmer, Bryan; Watson, Michael D.

    2015-01-01

    Exergy is defined as the useful work available from a system in a specified environment. Exergy analysis allows for comparison between different system designs, and allows for comparison of subsystem efficiencies within system designs. The proposed paper explores the relationship between the fundamental rocket equation and an exergy balance equation. A previously derived exergy equation related to rocket systems is investigated, and a higher fidelity analysis will be derived. The exergy assessments will enable informed, value-based decision making when comparing alternative rocket system designs, and will allow the most efficient configuration among candidate configurations to be determined.

  12. Space shuttle solid rocket booster recovery system definition. Volume 2: SRB water impact Monte Carlo computer program, user's manual

    NASA Technical Reports Server (NTRS)

    1973-01-01

    The HD 220 program was created as part of the space shuttle solid rocket booster recovery system definition. The model was generated to investigate the damage to SRB components under water impact loads. The random nature of environmental parameters, such as ocean waves and wind conditions, necessitates estimation of the relative frequency of occurrence for these parameters. The nondeterministic nature of component strengths also lends itself to probabilistic simulation. The Monte Carlo technique allows the simultaneous perturbation of multiple independent parameters and provides outputs describing the probability distribution functions of the dependent parameters. This allows the user to determine the required statistics for each output parameter.

  13. Liquid Rocket Booster Integration Study. Volume 2: Study synopsis

    NASA Technical Reports Server (NTRS)

    1988-01-01

    The impacts of introducing liquid rocket booster engines (LRB) into the Space Transportation System (STS)/Kennedy Space Center (KSC) launch environment are identified and evaluated. Proposed ground systems configurations are presented along with a launch site requirements summary. Prelaunch processing scenarios are described and the required facility modifications and new facility requirements are analyzed. Flight vehicle design recommendations to enhance launch processing are discussed. Processing approaches to integrate LRB with existing STS launch operations are evaluated. The key features and significance of launch site transition to a new STS configuration in parallel with ongoing launch activities are enumerated. This volume is the study summary of the five volume series.

  14. Liquid rocket booster integration study. Volume 1: Executive summary

    NASA Technical Reports Server (NTRS)

    1988-01-01

    The impacts of introducing liquid rocket booster engines (LRB) into the Space Transportation System (STS)/Kennedy Space Center (KSC) launch environment are identified and evaluated. Proposed ground systems configurations are presented along with a launch site requirements summary. Prelaunch processing scenarios are described and the required facility modifications and new facility requirements are analyzed. Flight vehicle design recommendations to enhance launch processing are discussed. Processing approaches to integrate LRB with existing STS launch operations are evaluated. The key features and significance of launch site transition to a new STS configuration in parallel with ongoing launch activities are enumerated. This volume is the executive summary of the five volume series.

  15. Liquid rocket booster integration study. Volume 5, part 1: Appendices

    NASA Technical Reports Server (NTRS)

    1988-01-01

    The impacts of introducing liquid rocket booster engines (LRB) into the Space Transportation System (STS)/Kennedy Space Center (KSC) launch environment are identified and evaluated. Proposed ground systems configurations are presented along with a launch site requirements summary. Prelaunch processing scenarios are described and the required facility modifications and new facility requirements are analyzed. Flight vehicle design recommendations to enhance launch processing are discussed. Processing approaches to integrate LRB with existing STS launch operations are evaluated. The key features and significance of launch site transition to a new STS configuration in parallel with ongoing launch activities are enumerated. This volume is the appendices of the five volume series.

  16. Rocket Based Combined Cycle (RBCC) Propulsion Workshop, volume 2

    NASA Technical Reports Server (NTRS)

    Chojnacki, Kent T.

    1992-01-01

    The goal of the Rocket Based Combined Cycle (RBCC) Propulsion Technology Workshop, was to impart technology information to the propulsion community with respect to hypersonic combined cycle propulsion capabilities. The major recommendation resulting from this technology workshop was as follows: conduct a systems-level applications study to define the desired propulsion system and vehicle technology requirements for LEO launch vehicles. All SSTO and TSTO options using the various propulsion systems (airbreathing combined cycle, rocket-based combined cycle, and all rocket) must be considered. Such a study should be accomplished as soon as possible. It must be conducted with a consistent set of ground rules and assumptions. Additionally, the study should be conducted before any major expenditures on a RBCC technology development program occur.

  17. Mean Flow Augmented Acoustics in Rocket Systems

    NASA Technical Reports Server (NTRS)

    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

  18. World Data Center A (rockets and satellites) catalogue of data. Volume 1, part A: Sounding rockets

    NASA Technical Reports Server (NTRS)

    1972-01-01

    A cumulative listing of all scientifically successful rockets that have been identified from various sources is presented. The listing starts with the V-2 rocket launched on 7 March 1947 and contains all rockets identified up to 31 December 1971.

  19. Mean Flow Augmented Acoustics in Rocket Systems

    NASA Technical Reports Server (NTRS)

    Fischbach, Sean R.

    2015-01-01

    Combustion instability in solid rocket motors and liquid engines is a complication that continues to plague designers and engineers. Many rocket systems experience violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process and gas dynamics. During sever cases of combustion instability fluctuation amplitudes can reach values equal to or greater than the average chamber pressure. Large amplitude oscillations lead to damaged injectors, loss of rocket performance, damaged payloads, and in some cases breach of case/loss of mission. Historic difficulties in modeling and predicting combustion instability has reduced most rocket systems experiencing instability into a costly fix through testing paradigm or to scrap the system entirely.

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

    NASA Technical Reports Server (NTRS)

    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.

  1. PC programs for the prediction of the linear stability behavior of liquid propellant propulsion systems and application to current MSFC rocket engine test programs, volume 1

    NASA Technical Reports Server (NTRS)

    Doane, George B., III; Armstrong, W. C.

    1990-01-01

    Research on propulsion stability (chugging and acoustic modes), and propellant valve control was investigated. As part of the activation of the new liquid propulsion test facilities, it is necessary to analyze total propulsion system stability. To accomplish this, several codes were built to run on desktop 386 machines. These codes enable one to analyze the stability question associated with the propellant feed systems. In addition, further work was adapted to this computing environment and furnished along with other codes. This latter inclusion furnishes those interested in high frequency oscillatory combustion behavior (that does not couple to the feed system) a set of codes for study of proposed liquid rocket engines.

  2. Integrated System Test of an Airbreathing Rocket

    NASA Technical Reports Server (NTRS)

    Mack, Gregory; Beaudry, Charles; Ketchum, Andrew; McArthur, J. Craig (Technical Monitor)

    2002-01-01

    This viewgraph presentation provides information on NASA's attempts to develop an air-breathing propulsion in an effort to make future space transportation safer, more reliable and significantly less expensive than today's missions. Spacecraft powered by air-breathing rocket engines would be completely reusable, able to take off and land at airport runways and ready to fly again within days. A radical new engine project is called the Integrated System Tests of an Air-breathing Rocket, or ISTAR.

  3. SRB/SLEEC (Solid Rocket Booster/Shingle Lap Extendible Exit Cone) feasibility study, volume 2. Appendix A: Design study for a SLEEC actuation system

    NASA Technical Reports Server (NTRS)

    Thompson, D. S.

    1986-01-01

    The results are presented of a design feasibility study of a self-contained (powered) actuation system for a Shingle Lap Extendible Exit Cone (SLEEC) for Transportation System (STS). The evolution of the SLEEC actuation system design is reviewed, the final design concept is summarized, and the results of the detailed study of the final concept of the actuation system are treated. A conservative design using proven mechanical components was established as a major program priority. The final mechanical design has a very low development risk since the components, which consist of ballscrews, gearing, flexible shaft drives, and aircraft cables, have extensive aerospace applications and a history of proven reliability. The mathematical model studies have shown that little or no power is required to deploy the SLEEC actuation system because acceleration forces and internal pressure from the rocket plume provide the required energies. A speed control brake is incorporated in the design in order to control the rate of deployment.

  4. Dual mode nuclear rocket system applications.

    NASA Technical Reports Server (NTRS)

    Boretz, J. E.; Bell, J. M.; Plebuch, R. K.; Priest, C. C.

    1972-01-01

    Mission areas where the dual-mode nuclear rocket system is superior to nondual-mode systems are demonstrated. It is shown that the dual-mode system is competitive with the nondual-mode system even for those specific missions and particular payload configurations where it does not have a clear-cut advantage.

  5. Liquid Rocket Booster Study. Volume 2, Book 1

    NASA Technical Reports Server (NTRS)

    1989-01-01

    The recommended Liquid Rocket Booster (LRB) concept is shown which uses a common main engine with the Advanced Launch System (ALS) which burns LO2 and LH2. The central rationale is based on the belief that the U.S. can only afford one big new rocket engine development in the 1990's. A LO2/LH2 engine in the half million pound thrust class could satisfy STS LRB, ALS, and Shuttle C (instead of SSMEs). Development costs and higher production rates can be shared by NASA and USAF. If the ALS program does not occur, the LO2/RP-1 propellants would produce slight lower costs for and STS LRB. When the planned Booster Engine portion of the Civil Space Transportation Initiatives has provided data on large pressure fed LO2/RP-1 engines, then the choice should be reevaluated.

  6. Mean Flow Augmented Acoustics in Rocket Systems

    NASA Technical Reports Server (NTRS)

    Fischbach, Sean

    2014-01-01

    Combustion instability 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. Recent advances in energy based modeling of combustion instabilities require accurate determination of acoustic frequencies and mode shapes. Of particular interest is the acoustic mean flow interactions within the converging section of a rocket nozzle, where gradients of pressure, density, and velocity become large. The expulsion of unsteady energy through the nozzle of a rocket is identified as the predominate source of acoustic damping for most rocket systems. 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 present study aims to implement the French model within the COMSOL Multiphysiscs framework and analyzes one of the author's presented test cases.

  7. National Institute for Rocket Propulsion Systems (NIRPS): Solutions Facilitator

    NASA Technical Reports Server (NTRS)

    Brown, Tom

    2011-01-01

    National Institute for Rocket Propulsion Systems (NIRPS) "Solutions" plans to enable our nation's future in rocket propulsion systems by leveraging existing skills and capabilities to support industry's future needs

  8. Demand-type gas supply system for rocket borne thin-window proportional counters

    NASA Technical Reports Server (NTRS)

    Acton, L. W.; Caravalho, R.; Catura, R. C.; Joki, E. G.

    1977-01-01

    A simple closed loop control system has been developed to maintain the gas pressure in thin-window proportional counters during rocket flights. This system permits convenient external control of detector pressure and system flushing rate. The control system is activated at launch with the sealing of a reference volume at the existing system pressure. Inflight control to plus or minus 2 torr at a working pressure of 760 torr has been achieved on six rocket flights.

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

    NASA Technical Reports Server (NTRS)

    1972-01-01

    A technical analysis of the solid propellant rocket engines for use with the space shuttle is presented. The subjects discussed are: (1) solid rocket motor stage recovery, (2) environmental effects, (3) man rating of the solid propellant rocket engines, (4) system safety analysis, (5) ground support equipment, and (6) transportation, assembly, and checkout.

  10. A fluidic sounding rocket motor ignition system.

    NASA Technical Reports Server (NTRS)

    Marchese, V. P.; Rakowsky, E. L.; Bement, L. J.

    1972-01-01

    Fluidic sounding rocket motor ignition has been found to be feasible using a system without stored energy and with the complete absence of electrical energy and wiring. The fluidic ignitor is based on a two component aerodynamic resonance heating device called the pneumatic match. Temperatures in excess of 600 C were generated in closed resonance tubes which were excited by a free air jet from a simple convergent nozzle. Using nitrocellulose interface material, ignition of boron potassium nitrate was accomplished with air supply pressures as low as 45 psi. This paper describes an analytical and experimental program which established a fluidic rocket motor ignition system concept incorporating a pneumatic match with a simple hand pump as the only energy source.

  11. Space Shuttle solid rocket booster dewatering system

    NASA Technical Reports Server (NTRS)

    Fishel, K. R.

    1982-01-01

    After the launch of the Space Shuttle, the two solid rocket boosters (SRB's) are jettisoned into the ocean where they float in a spar (vertical) mode. It is cost effective to recover the SRB's. A remote controlled submersible vehicle has been developed to aid in their recovery. The vehicle is launched from a support ship, maneuvered to the SRB, then taken to depth and guided into the rocket nozzle. It then dewaters the SRB, using compressed air from the ship, and seals the nozzle. When dewatered, the SRB floats in a log (horizontal) mode and can be towed to port for reuse. The design of the remote controlled vehicle and its propulsion system is presented.

  12. Solid rocket booster performance evaluation model. Volume 4: Program listing

    NASA Technical Reports Server (NTRS)

    1974-01-01

    All subprograms or routines associated with the solid rocket booster performance evaluation model are indexed in this computer listing. An alphanumeric list of each routine in the index is provided in a table of contents.

  13. Liquid rocket booster integration study. Volume 3: Study products. Part 2: Sections 8-19

    NASA Technical Reports Server (NTRS)

    1988-01-01

    The impacts of introducing liquid rocket booster engines (LRB) into the Space Transportation System (STS)/Kennedy Space Center (KSC) launch environment are identified and evaluated. Proposed ground systems configurations are presented along with a launch site requirements summary. Prelaunch processing scenarios are described and the required facility modifications and new facility requirements are analyzed. Flight vehicle design recommendations to enhance launch processing are discussed. Processing approaches to integrate LRB with existing STS launch operations are evaluated. The key features and significance of launch site transition to a new STS configuration in parallel with ongoing launch activities are enumerated. This volume is part two of the study products section of the five volume series.

  14. Liquid rocket booster integration study. Volume 3, part 1: Study products

    NASA Technical Reports Server (NTRS)

    1988-01-01

    The impacts of introducing liquid rocket booster engines (LRB) into the Space Transportation System (STS)/Kennedy Space Center (KSC) launch environment are identified and evaluated. Proposed ground systems configurations are presented along with a launch site requirements summary. Prelaunch processing scenarios are described and the required facility modifications and new facility requirements are analyzed. Flight vehicle design recommendations to enhance launch processing are discussed. Processing approaches to integrate LRB with existing STS launch operations are evaluated. The key features and significance of launch site transition to a new STS configuration in parallel with ongoing launch activities are enumerated. This volume is part one of the study products section of the five volume series.

  15. Rocket Testing and Integrated System Health Management

    NASA Technical Reports Server (NTRS)

    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.

  16. A rocket engine design expert system

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth J.

    1989-01-01

    The overall structure and capabilities of an expert system designed to evaluate rocket engine performance are described. The expert system incorporates a JANNAF standard reference computer code to determine rocket engine performance and a state-of-the-art finite element computer code to calculate the interactions between propellant injection, energy release in the combustion chamber, and regenerative cooling heat transfer. Rule-of-thumb heuristics were incorporated for the hydrogen-oxygen coaxial injector design, including a minimum gap size constraint on the total number of injector elements. One-dimensional equilibrium chemistry was employed in the energy release analysis of the combustion chamber and three-dimensional finite-difference analysis of the regenerative cooling channels was used to calculate the pressure drop along the channels and the coolant temperature as it exits the coolant circuit. Inputting values to describe the geometry and state properties of the entire system is done directly from the computer keyboard. Graphical display of all output results from the computer code analyses is facilitated by menu selection of up to five dependent variables per plot.

  17. A Rocket Engine Design Expert System

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth J.

    1989-01-01

    The overall structure and capabilities of an expert system designed to evaluate rocket engine performance are described. The expert system incorporates a JANNAF standard reference computer code to determine rocket engine performance and a state of the art finite element computer code to calculate the interactions between propellant injection, energy release in the combustion chamber, and regenerative cooling heat transfer. Rule-of-thumb heuristics were incorporated for the H2-O2 coaxial injector design, including a minimum gap size constraint on the total number of injector elements. One dimensional equilibrium chemistry was used in the energy release analysis of the combustion chamber. A 3-D conduction and/or 1-D advection analysis is used to predict heat transfer and coolant channel wall temperature distributions, in addition to coolant temperature and pressure drop. Inputting values to describe the geometry and state properties of the entire system is done directly from the computer keyboard. Graphical display of all output results from the computer code analyses is facilitated by menu selection of up to five dependent variables per plot.

  18. Reusable Rocket Engine Turbopump Health Management System

    NASA Technical Reports Server (NTRS)

    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.

  19. Parallelization of Rocket Engine System Software (Press)

    NASA Technical Reports Server (NTRS)

    Cezzar, Ruknet

    1996-01-01

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

  20. Engine protection system for recoverable rocket booster

    NASA Technical Reports Server (NTRS)

    Shelby, Jr., Jerry A. (Inventor)

    1994-01-01

    A rocket engine protection system for a recoverable rocket booster which is arranged to land in a salt water body in substantially a nose down attitude. The system includes an inflatable bag which is stowed on a portion of a flat annular rim of the aft skirt of the booster. The bag is hinged at opposing sides and is provided with springs that urge the bag open. The bag is latched in a stowed position during launch and prior to landing for recovery is unlatched to permit the bag to be urged open and into sealing engagement with the rim. A source of pressurized gas further inflates the bag and urges it into sealing engagement with the rim of the skirt where it is locked into position. The gas provides a positive pressure upon the interior of the bag to preclude entry of salt water into the skirt and into contact with the engine. A flotation arrangement may assist in precluding the skirt of the booster from becoming submerged.

  1. Solid rocket booster thermal radiation model. Volume 2: User's manual

    NASA Technical Reports Server (NTRS)

    Lee, A. L.

    1976-01-01

    A user's manual was prepared for the computer program of a solid rocket booster (SRB) thermal radiation model. The following information was included: (1) structure of the program, (2) input information required, (3) examples of input cards and output printout, (4) program characteristics, and (5) program listing.

  2. Solid rocket booster performance evaluation model. Volume 2: Users manual

    NASA Technical Reports Server (NTRS)

    1974-01-01

    This users manual for the solid rocket booster performance evaluation model (SRB-II) contains descriptions of the model, the program options, the required program inputs, the program output format and the program error messages. SRB-II is written in FORTRAN and is operational on both the IBM 370/155 and the MSFC UNIVAC 1108 computers.

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

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

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

  4. Workshop on the Suborbital Science Sounding Rocket Program, Volume 1

    NASA Technical Reports Server (NTRS)

    1991-01-01

    The unique characteristics of the sounding rocket program is described, with its importance to space science stressed, especially in providing UARS correlative measurements. The program provided opportunities to do innovative scientific studies in regions not other wise accessible; it was a testbed for developing new technologies; and its key attributes were flexibility, reliability, and economy. The proceedings of the workshop are presented in viewgraph form, including the objectives of the workshop and the workshop agenda.

  5. Developments in REDES: The rocket engine design expert system

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth O.

    1990-01-01

    The Rocket Engine Design Expert System (REDES) is being developed at the 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 is built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

  6. Developments in REDES: The Rocket Engine Design Expert System

    NASA Technical Reports Server (NTRS)

    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.

  7. Solid rocket booster thermal radiation model, volume 1

    NASA Technical Reports Server (NTRS)

    Watson, G. H.; Lee, A. L.

    1976-01-01

    A solid rocket booster (SRB) thermal radiation model, capable of defining the influence of the plume flowfield structure on the magnitude and distribution of thermal radiation leaving the plume, was prepared and documented. Radiant heating rates may be calculated for a single SRB plume or for the dual SRB plumes astride the space shuttle. The plumes may be gimbaled in the yaw and pitch planes. Space shuttle surface geometries are simulated with combinations of quadric surfaces. The effect of surface shading is included. The computer program also has the capability to calculate view factors between the SRB plumes and space shuttle surfaces as well as surface-to-surface view factors.

  8. Solid rocket booster performance evaluation model. Volume 1: Engineering description

    NASA Technical Reports Server (NTRS)

    1974-01-01

    The space shuttle solid rocket booster performance evaluation model (SRB-II) is made up of analytical and functional simulation techniques linked together so that a single pass through the model will predict the performance of the propulsion elements of a space shuttle solid rocket booster. The available options allow the user to predict static test performance, predict nominal and off nominal flight performance, and reconstruct actual flight and static test performance. Options selected by the user are dependent on the data available. These can include data derived from theoretical analysis, small scale motor test data, large motor test data and motor configuration data. The user has several options for output format that include print, cards, tape and plots. Output includes all major performance parameters (Isp, thrust, flowrate, mass accounting and operating pressures) as a function of time as well as calculated single point performance data. The engineering description of SRB-II discusses the engineering and programming fundamentals used, the function of each module, and the limitations of each module.

  9. Space transportation system solid rocket booster thrust vector control system

    NASA Technical Reports Server (NTRS)

    Verble, A. J., Jr.; Mccool, A. A.; Potter, J. H.

    1979-01-01

    The Solid Rocket Booster, Thrust Vector Control (TVC) system was designed in accordance with the following requirements: self-contained power supply, fail-safe operation, 20 flight uses after exposure to seawater landings, optimized cost, and component interchangeability. Trade studies were performed which led to the selection of a recirculating hydraulic system powered by Auxiliary Power Units (APU) which drive the hydraulic actuators and gimbal the solid rocket motor nozzle. Other approaches for the system design were studied in arriving at the recirculating hydraulic system powered by an APU. These systems must withstand the imposed environment and be usable for a minimum of 20 Space Transportation System flights with a minimum of refurbishment. The TVC system has completed the major portion of qualification and verification tests and is prepared to be cleared for the first Shuttle flight (STS-1). Substantiation data will include analytical and test data.

  10. Space Transportation System solid rocket booster thrust vector control system

    NASA Technical Reports Server (NTRS)

    Verble, A. J., Jr.; Mccool, A. A.; Potter, J. H.

    1980-01-01

    The Solid Rocket Booster, Thrust Vector Control (TVC) system was designed in accordance with the following requirements: self-contained power supply, failsafe operation, 20 flight uses after exposure to seawater landings, optimized cost, and component interchangeability. Trade studies were performed which led to the selection of a recirculating hydraulic system powered by Auxiliary Power Units (APU) which drive the hydraulic actuators and gimbal the solid rocket motor nozzle. Other approaches for the system design were studied in arriving at the recirculating hydraulic system powered by an APU. These systems must withstand the imposed environment and be usable for a minimum of 20 Space Transportation System flights with a minimum of refurbishment. The TVC system completed the required qualification and verification tests and is certified for the intended application. Substantiation data include analytical and test data.

  11. SRB-3D Solid Rocket Booster performance prediction program. Volume 2: Sample case

    NASA Technical Reports Server (NTRS)

    Winkler, J. C.

    1976-01-01

    The sample case presented in this volume is an asymmetrical eight sector thermal gradient performance prediction for the solid rocket motor. This motor is the TC-227A-75 grain design and the initial grain geometry is assumed to be symmetrical about the motors longitudinal axis.

  12. Advanced small rocket chambers. Option 3: 110 1bf Ir-Re rocket, volume 1

    NASA Astrophysics Data System (ADS)

    Jassowski, Donald M.; Schoenman, Leonard

    1995-02-01

    This report describes the AJ10-221, a high performance Iridium-coated Rhenium (Ir-Re) 110 lbf (490N) welded rocket chamber with 286:1 area ratio nozzle. This engine was designed, built, and hot fired for over 6 hours on this program. It demonstrated an I(s) of 321.8 sec, which is 10 sec higher than conventional 110 lbf silicide coated Cb chambers now in use. The approach used in this portion of the program was to demonstrate the performance improvement that can be made by the elimination of fuel film cooling and the use of a high temperature (4000F) (2200C) iridium-coated rhenium (Ir-Re) rocket chamber. Detailed thermal, performance, mechanical, and dynamic design analyses of the full engine were conducted by Aerojet. Two Ir-Re chambers were built to the Aerojet design by Ultramet, using the chemical vapor deposition (CVD) process. Incorporation of a secondary mixing device or Boundary Layer Trip (BLT) within the combustion chamber (Aerojet Patents 4882904 and 4936091) results in improvement in flow uniformity, and a significant life and performance increase. The 110 lbf engine design was verified in bolt-up hardware tests at sea level and altitude. The effects of injector design on performance were studied. Two duplicate injectors were fabricated matching the preferred design and were demonstrated to be interchangeable in operation. One of these units was fabricated matching the preferred design and was demonstrated to be interchangeable in operation. One of these units was welded into a flight type thruster which was tested for an accumulated duration of 22,590 sec in 93 firings, one of which was a continuous burn of two hours. A design deficiency in the C-103 nozzle near the Re-Cb transition joint was discovered, studied and corrected design has been prepared. Workhardening studies have been conducted to investigate methods for increasing the low yield strength of the Re in the annealed conditions. An advanced 490N high performance engine has been demonstrated

  13. Advanced small rocket chambers. Option 3: 110 1bf Ir-Re rocket, volume 2

    NASA Technical Reports Server (NTRS)

    Jassowski, Donald M.; Schoenman, Leonard

    1995-01-01

    This is the second part of a two-part report that describes the AJ10-221, a high performance iridium-coated rhenium (Ir-Re) 110 lbf (490N) welded rocket chamber with 286:1 area ratio nozzle. This engine was designed, built, and hot fired for over 6 hours on this program. It demonstrated an I(s) of 321.8 sec, which is 10 sec higher than conventional 110 lbf silicide coated Cb chambers now in use. The approach used in this portion of the program was to demonstrate the performance improvement that can be made by the elimination of fuel film cooling and the use of a high temperature (4000 F) (2200 C) iridium-coated rhenium (Ir-Re) rocket chamber. Detailed thermal, performance, mechanical, and dynamic design analyses of the full engine were conducted by Aerojet. Two Ir-Re chambers were built to the Aerojet design by Ultramet, using the chemical vapor deposition (CVD) process. Incorporation of a secondary mixing device or Boundary Layer Trip (BLT) within the combustion chamber (Aerojet Patents 4882904 and 4936091) results in improvement in flow uniformity, and a significant life and performance increase. The 110 lbf engine design was verified in bolt-up hardware tests at sea level and altitude. The effects of injector design on performance were studied. Two duplicate injectors were fabricated matching the preferred design and were demonstrated to be interchangeable in operation. One of these units were welded into a flight type thruster which was tested for an accumulated duration of 22,590 sec in 93 firings, one of which was a continuous burn of two hours. A design deficiency in the C-103 nozzle near the Re-Cb transition joint was discovered, studied and corrected design has been prepared. Workhardening studies have been conducted to investigate methods for increasing the low yield strength of the Re in the annealed conditions. An advanced 490N high performance engine has been demonstrated which, when proven to be capable of withstanding launch vibration, is ready for

  14. Advanced Small Rocket Chambers. Option 3: 110 1Bf Ir-Re Rocket, Volume 1

    NASA Technical Reports Server (NTRS)

    Jassowski, Donald M.; Schoenman, Leonard

    1995-01-01

    This report describes the AJ10-221, a high performance Iridium-coated Rhenium (Ir-Re) 110 lbf (490N) welded rocket chamber with 286:1 area ratio nozzle. This engine was designed, built, and hot fired for over 6 hours on this program. It demonstrated an I(s) of 321.8 sec, which is 10 sec higher than conventional 110 lbf silicide coated Cb chambers now in use. The approach used in this portion of the program was to demonstrate the performance improvement that can be made by the elimination of fuel film cooling and the use of a high temperature (4000F) (2200C) iridium-coated rhenium (Ir-Re) rocket chamber. Detailed thermal, performance, mechanical, and dynamic design analyses of the full engine were conducted by Aerojet. Two Ir-Re chambers were built to the Aerojet design by Ultramet, using the chemical vapor deposition (CVD) process. Incorporation of a secondary mixing device or Boundary Layer Trip (BLT) within the combustion chamber (Aerojet Patents 4882904 and 4936091) results in improvement in flow uniformity, and a significant life and performance increase. The 110 lbf engine design was verified in bolt-up hardware tests at sea level and altitude. The effects of injector design on performance were studied. Two duplicate injectors were fabricated matching the preferred design and were demonstrated to be interchangeable in operation. One of these units was fabricated matching the preferred design and was demonstrated to be interchangeable in operation. One of these units was welded into a flight type thruster which was tested for an accumulated duration of 22,590 sec in 93 firings, one of which was a continuous burn of two hours. A design deficiency in the C-103 nozzle near the Re-Cb transition joint was discovered, studied and corrected design has been prepared. Workhardening studies have been conducted to investigate methods for increasing the low yield strength of the Re in the annealed conditions. An advanced 490N high performance engine has been demonstrated

  15. Advanced small rocket chambers. Option 3: 110 1bf Ir-Re rocket, volume 2

    NASA Astrophysics Data System (ADS)

    Jassowski, Donald M.; Schoenman, Leonard

    1995-02-01

    This is the second part of a two-part report that describes the AJ10-221, a high performance iridium-coated rhenium (Ir-Re) 110 lbf (490N) welded rocket chamber with 286:1 area ratio nozzle. This engine was designed, built, and hot fired for over 6 hours on this program. It demonstrated an I(s) of 321.8 sec, which is 10 sec higher than conventional 110 lbf silicide coated Cb chambers now in use. The approach used in this portion of the program was to demonstrate the performance improvement that can be made by the elimination of fuel film cooling and the use of a high temperature (4000 F) (2200 C) iridium-coated rhenium (Ir-Re) rocket chamber. Detailed thermal, performance, mechanical, and dynamic design analyses of the full engine were conducted by Aerojet. Two Ir-Re chambers were built to the Aerojet design by Ultramet, using the chemical vapor deposition (CVD) process. Incorporation of a secondary mixing device or Boundary Layer Trip (BLT) within the combustion chamber (Aerojet Patents 4882904 and 4936091) results in improvement in flow uniformity, and a significant life and performance increase. The 110 lbf engine design was verified in bolt-up hardware tests at sea level and altitude. The effects of injector design on performance were studied. Two duplicate injectors were fabricated matching the preferred design and were demonstrated to be interchangeable in operation. One of these units were welded into a flight type thruster which was tested for an accumulated duration of 22,590 sec in 93 firings, one of which was a continuous burn of two hours. A design deficiency in the C-103 nozzle near the Re-Cb transition joint was discovered, studied and corrected design has been prepared. Workhardening studies have been conducted to investigate methods for increasing the low yield strength of the Re in the annealed conditions. An advanced 490N high performance engine has been demonstrated which, when proven to be capable of withstanding launch vibration, is ready for

  16. RocketCam systems for providing situational awareness on rockets, spacecraft, and other remote platforms

    NASA Astrophysics Data System (ADS)

    Ridenoure, Rex

    2004-09-01

    Space-borne imaging systems derived from commercial technology have been successfully employed on launch vehicles for several years. Since 1997, over sixty such imagers - all in the product family called RocketCamTM - have operated successfully on 29 launches involving most U.S. launch systems. During this time, these inexpensive systems have demonstrated their utility in engineering analysis of liftoff and ascent events, booster performance, separation events and payload separation operations, and have also been employed to support and document related ground-based engineering tests. Such views from various vantage points provide not only visualization of key events but stunning and extremely positive public relations video content. Near-term applications include capturing key events on Earth-orbiting spacecraft and related proximity operations. This paper examines the history to date of RocketCams on expendable and manned launch vehicles, assesses their current utility on rockets, spacecraft and other aerospace vehicles (e.g., UAVs), and provides guidance for their use in selected defense and security applications. Broad use of RocketCams on defense and security projects will provide critical engineering data for developmental efforts, a large database of in-situ measurements onboard and around aerospace vehicles and platforms, compelling public relations content, and new diagnostic information for systems designers and failure-review panels alike.

  17. Dynamic Analysis of Sounding Rocket Pneumatic System Revision

    NASA Technical Reports Server (NTRS)

    Armen, Jerald

    2010-01-01

    The recent fusion of decades of advancements in mathematical models, numerical algorithms and curve fitting techniques marked the beginning of a new era in the science of simulation. It is becoming indispensable to the study of rockets and aerospace analysis. In pneumatic system, which is the main focus of this paper, particular emphasis will be placed on the efforts of compressible flow in Attitude Control System of sounding rocket.

  18. Computational Analysis for Rocket-Based Combined-Cycle Systems During Rocket-Only Operation

    NASA Technical Reports Server (NTRS)

    Steffen, C. J., Jr.; Smith, T. D.; Yungster, S.; Keller, D. J.

    2000-01-01

    A series of Reynolds-averaged Navier-Stokes calculations were employed to study the performance of rocket-based combined-cycle systems operating in an all-rocket mode. This parametric series of calculations were executed within a statistical framework, commonly known as design of experiments. The parametric design space included four geometric and two flowfield variables set at three levels each, for a total of 729 possible combinations. A D-optimal design strategy was selected. It required that only 36 separate computational fluid dynamics (CFD) solutions be performed to develop a full response surface model, which quantified the linear, bilinear, and curvilinear effects of the six experimental variables. The axisymmetric, Reynolds-averaged Navier-Stokes simulations were executed with the NPARC v3.0 code. The response used in the statistical analysis was created from Isp efficiency data integrated from the 36 CFD simulations. The influence of turbulence modeling was analyzed by using both one- and two-equation models. Careful attention was also given to quantify the influence of mesh dependence, iterative convergence, and artificial viscosity upon the resulting statistical model. Thirteen statistically significant effects were observed to have an influence on rocket-based combined-cycle nozzle performance. It was apparent that the free-expansion process, directly downstream of the rocket nozzle, can influence the Isp efficiency. Numerical schlieren images and particle traces have been used to further understand the physical phenomena behind several of the statistically significant results.

  19. Rocket noise filtering system using digital filters

    NASA Technical Reports Server (NTRS)

    Mauritzen, David

    1990-01-01

    A set of digital filters is designed to filter rocket noise to various bandwidths. The filters are designed to have constant group delay and are implemented in software on a general purpose computer. The Parks-McClellan algorithm is used. Preliminary tests are performed to verify the design and implementation. An analog filter which was previously employed is also simulated.

  20. Knowledge Preservation for Design of Rocket Systems

    NASA Technical Reports Server (NTRS)

    Moreman, Douglas

    2002-01-01

    An engineer at NASA Lewis RC presented a challenge to us at Southern University. Our response to that challenge, stated circa 1993, has evolved into the Knowledge Preservation Project which is here reported. The stated problem was to capture some of the knowledge of retiring NASA engineers and make it useful to younger engineers via computers. We evolved that initial challenge to this - design a system of tools such that, with this system, people might efficiently capture and make available via commonplace computers, deep knowledge of retiring NASA engineers. In the process of proving some of the concepts of this system, we would (and did) capture knowledge from some specific engineers and, so, meet the original challenge along the way to meeting the new. Some of the specific knowledge acquired, particularly that on the RL- 10 engine, was directly relevant to design of rocket engines. We considered and rejected some of the techniques popular in the days we began - specifically "expert systems" and "oral histories". We judged that these old methods had too high a cost per sentence preserved. That cost could be measured in hours of labor of a "knowledge professional". We did spend, particularly in the grant preceding this one, some time creating a couple of "concept maps", one of the latest ideas of the day, but judged this also to be costly in time of a specially trained knowledge-professional. We reasoned that the cost in specialized labor could be lowered if less time were spent being selective about sentences from the engineers and in crafting replacements for those sentences. The trade-off would seem to be that our set of sentences would be less dense in information, but we found a computer-based way around this seeming defect. Our plan, details of which we have been carrying out, was to find methods of extracting information from experts which would be capable of gaining cooperation, and interest, of senior engineers and using their time in a way they would

  1. Study of solid rocket motor for space shuttle booster, volume 2, book 3, appendix A

    NASA Technical Reports Server (NTRS)

    1972-01-01

    A systems requirements analysis for the solid propellant rocket engine to be used with the space shuttle was conducted. The systems analysis was developed to define the physical and functional requirements for the systems and subsystems. The operations analysis was performed to identify the requirements of the various launch operations, mission operations, ground operations, and logistic and flight support concepts.

  2. Liquid rocket booster integration study. Volume 4: Reviews and presentation material

    NASA Technical Reports Server (NTRS)

    1988-01-01

    Liquid rocket booster integration study is presented. Volume 4 contains materials presented at the MSFC/JSC/KSC Integrated Reviews and Working Group Sessions, and the Progress Reviews presented to the KSC Study Manager. The following subject areas are covered: initial impact assessment; conflicts with the on-going STS mission; access to the LRB at the PAD; the activation schedule; transition requirements; cost methodology; cost modelling approach; and initial life cycle cost.

  3. Flip-Flop Recovery System for sounding rocket payloads

    NASA Technical Reports Server (NTRS)

    Flores, A., Jr.

    1986-01-01

    The design, development, and testing of the Flip-Flop Recovery System, which protects sensitive forward-mounted instruments from ground impact during sounding rocket payload recovery operations, are discussed. The system was originally developed to reduce the impact damage to the expensive gold-plated forward-mounted spectrometers in two existing Taurus-Orion rocket payloads. The concept of the recovery system is simple: the payload is flipped over end-for-end at a predetermined time just after parachute deployment, thus minimizing the risk of damage to the sensitive forward portion of the payload from ground impact.

  4. Influence of curvature in regenerative cooling system of rocket engine

    NASA Astrophysics Data System (ADS)

    Torres, Y.; Stefanini, L.; Suslov, D.

    2009-09-01

    Thermomechanical loads in rocket engines can be drastically reduced by a reliable cooling system. The regenerative cooling system uses propellants as coolant which flows through milled cooling channels in the chamber walls. Due to centrifugal forces, dynamic secondary motions appear in cooling-channel curvatures, which strongly modify heat transfer. Three-dimensional (3D) numerical calculations have been performed in order to compare this heat flux modification with empirical correlations. Different turbulence models and wall treatments have been tested to develop a complete numerical data base about asymmetrical (concave side) heat transfer in curved cooling channels of rocket engine.

  5. Reusable rocket engine intelligent control system framework design, phase 2

    NASA Technical Reports Server (NTRS)

    Nemeth, ED; Anderson, Ron; Ols, Joe; Olsasky, Mark

    1991-01-01

    Elements of an advanced functional framework for reusable rocket engine propulsion system control are presented for the Space Shuttle Main Engine (SSME) demonstration case. Functional elements of the baseline functional framework are defined in detail. The SSME failure modes are evaluated and specific failure modes identified for inclusion in the advanced functional framework diagnostic system. Active control of the SSME start transient is investigated, leading to the identification of a promising approach to mitigating start transient excursions. Key elements of the functional framework are simulated and demonstration cases are provided. Finally, the advanced function framework for control of reusable rocket engines is presented.

  6. National Institute for Rocket Propulsion Systems 1st Annual Workshop

    NASA Technical Reports Server (NTRS)

    Doreswamy, Rajiv; Fry, Emma; Swindell, Tina

    2012-01-01

    The National Institute for Rocket Propulsion Systems (NIRPS) is a Government -wide initiative that seeks to ensure the resiliency of the Nation fs rocket propulsion community in order for the enterprise to remain vibrant and capable of providing reliable and affordable propulsion systems for the nation fs defense, civil and commercial needs. Recognizing that rocket propulsion is a multi-use technology that ensures the nation fs leadership in aerospace, the Government has a vested interest in maintaining this strategic capability through coordinated and synchronized acquisition programs and continual investments in research and development. NIRPS is a resource for collaboration and integration between all sectors of the U.S. propulsion enterprise, supporting policy development options, identifying technology requirements, and offering solutions that maximize national resources while ensuring that capability exists to meet future demand. NIRPS functions as a multi ]agency organization that our nation fs decision makers can look to for comprehensive information regarding all issues concerning the propulsion enterprise.

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

    NASA Technical Reports Server (NTRS)

    1972-01-01

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

  8. Effect of Correlation on Multi-Engine Rocket Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Baker, J. R.; Breneman, J. E.

    1989-01-01

    A matter of great concern in the design and operation of multi-engine rocket propulsion systems is the effect of the premature shutdown of one engine on the vehicle. This probability that a premature shutdown will cause a vehicle loss is termed correlation. Based on airbreathing experiences as well as rocket engine data the best estimate of this correlation is made and then applied to the overall multi-engine reliability problem to demonstrate its potential effect. At this point, follow-on analyses are pointed out that illustrate how any potential failures that may cause a correlatable event can be eliminated; thus bringing this correlation to almost 0.

  9. Rocket Based Combined Cycle (RBCC) Propulsion Technology Workshop. Volume 1: Executive summary

    NASA Technical Reports Server (NTRS)

    Chojnacki, Kent T.

    1992-01-01

    The goal of the Rocket-Based Combined Cycle (RBCC) Propulsion Technology Workshop was to assess the RBCC propulsion system's viability for Earth-to-Orbit (ETO) transportation systems. This was accomplished by creating a forum (workshop) in which past work in the field of RBCC propulsion systems was reviewed, current technology status was evaluated, and future technology programs in the field of RBCC propulsion systems were postulated, discussed, and recommended.

  10. Network Flow Simulation of Fluid Transients in Rocket Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Bandyopadhyay, Alak; Hamill, Brian; Ramachandran, Narayanan; Majumdar, Alok

    2011-01-01

    Fluid transients, also known as water hammer, can have a significant impact on the design and operation of both spacecraft and launch vehicle propulsion systems. These transients often occur at system activation and shutdown. The pressure rise due to sudden opening and closing of valves of propulsion feed lines can cause serious damage during activation and shutdown of propulsion systems. During activation (valve opening) and shutdown (valve closing), pressure surges must be predicted accurately to ensure structural integrity of the propulsion system fluid network. In the current work, a network flow simulation software (Generalized Fluid System Simulation Program) based on Finite Volume Method has been used to predict the pressure surges in the feed line due to both valve closing and valve opening using two separate geometrical configurations. The valve opening pressure surge results are compared with experimental data available in the literature and the numerical results compared very well within reasonable accuracy (< 5%) for a wide range of inlet-to-initial pressure ratios. A Fast Fourier Transform is preformed on the pressure oscillations to predict the various modal frequencies of the pressure wave. The shutdown problem, i.e. valve closing problem, the simulation results are compared with the results of Method of Characteristics. Most rocket engines experience a longitudinal acceleration, known as "pogo" during the later stage of engine burn. In the shutdown example problem, an accumulator has been used in the feed system to demonstrate the "pogo" mitigation effects in the feed system of propellant. The simulation results using GFSSP compared very well with the results of Method of Characteristics.

  11. Integrated model development for liquid fueled rocket propulsion systems

    NASA Technical Reports Server (NTRS)

    Santi, L. Michael

    1993-01-01

    As detailed in the original statement of work, the objective of phase two of this research effort was to develop a general framework for rocket engine performance prediction that integrates physical principles, a rigorous mathematical formalism, component level test data, system level test data, and theory-observation reconciliation. Specific phase two development tasks are defined.

  12. Universal Data Handling System for Sounding Rockets and Balloons

    NASA Astrophysics Data System (ADS)

    Andersson, G.

    2015-09-01

    Data handling systems (DHS) used in service systems and experiment modules on sounding rockets and balloons have traditionally been different in design. A study was performed in 2012 at SSC to evaluate the feasibility of a common system usable across different platforms. The outcome was the “Unified DHS system”. The new DHS is very modular in design and can easily be adapted to different mission scenarios.

  13. Development of a 12-Thrust Chamber Kerosene /Oxygen Primary Rocket Sub-System for an Early (1964) Air-Augmented Rocket Ground-Test System

    NASA Technical Reports Server (NTRS)

    Pryor, D.; Hyde, E. H.; Escher, W. J. D.

    1999-01-01

    Airbreathing/Rocket combined-cycle, and specifically rocket-based combined- cycle (RBCC), propulsion systems, typically employ an internal engine flow-path installed primary rocket subsystem. To achieve acceptably short mixing lengths in effecting the "air augmentation" process, a large rocket-exhaust/air interfacial mixing surface is needed. This leads, in some engine design concepts, to a "cluster" of small rocket units, suitably arrayed in the flowpath. To support an early (1964) subscale ground-test of a specific RBCC concept, such a 12-rocket cluster was developed by NASA's Marshall Space Flight Center (MSFC). The small primary rockets used in the cluster assembly were modified versions of an existing small kerosene/oxygen water-cooled rocket engine unit routinely tested at MSFC. Following individual thrust-chamber tests and overall subsystem qualification testing, the cluster assembly was installed at the U. S. Air Force's Arnold Engineering Development Center (AEDC) for RBCC systems testing. (The results of the special air-augmented rocket testing are not covered here.) While this project was eventually successfully completed, a number of hardware integration problems were met, leading to catastrophic thrust chamber failures. The principal "lessons learned" in conducting this early primary rocket subsystem experimental effort are documented here as a basic knowledge-base contribution for the benefit of today's RBCC research and development community.

  14. Expert System Architecture for Rocket Engine Numerical Simulators: A Vision

    NASA Technical Reports Server (NTRS)

    Mitra, D.; Babu, U.; Earla, A. K.; Hemminger, Joseph A.

    1998-01-01

    Simulation of any complex physical system like rocket engines involves modeling the behavior of their different components using mostly numerical equations. Typically a simulation package would contain a set of subroutines for these modeling purposes and some other ones for supporting jobs. A user would create an input file configuring a system (part or whole of a rocket engine to be simulated) in appropriate format understandable by the package and run it to create an executable module corresponding to the simulated system. This module would then be run on a given set of input parameters in another file. Simulation jobs are mostly done for performance measurements of a designed system, but could be utilized for failure analysis or a design job such as inverse problems. In order to use any such package the user needs to understand and learn a lot about the software architecture of the package, apart from being knowledgeable in the target domain. We are currently involved in a project in designing an intelligent executive module for the rocket engine simulation packages, which would free any user from this burden of acquiring knowledge on a particular software system. The extended abstract presented here will describe the vision, methodology and the problems encountered in the project. We are employing object-oriented technology in designing the executive module. The problem is connected to the areas like the reverse engineering of any simulation software, and the intelligent systems for simulation.

  15. Cycle Trades for Nuclear Thermal Rocket Propulsion Systems

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

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

  16. A compact and robust diode laser system for atom interferometry on a sounding rocket

    NASA Astrophysics Data System (ADS)

    Schkolnik, V.; Hellmig, O.; Wenzlawski, A.; Grosse, J.; Kohfeldt, A.; Döringshoff, K.; Wicht, A.; Windpassinger, P.; Sengstock, K.; Braxmaier, C.; Krutzik, M.; Peters, A.

    2016-08-01

    We present a diode laser system optimized for laser cooling and atom interferometry with ultra-cold rubidium atoms aboard sounding rockets as an important milestone toward space-borne quantum sensors. Design, assembly and qualification of the system, combing micro-integrated distributed feedback (DFB) diode laser modules and free space optical bench technology, is presented in the context of the MAIUS (Matter-wave Interferometry in Microgravity) mission. This laser system, with a volume of 21 l and total mass of 27 kg, passed all qualification tests for operation on sounding rockets and is currently used in the integrated MAIUS flight system producing Bose-Einstein condensates and performing atom interferometry based on Bragg diffraction. The MAIUS payload is being prepared for launch in fall 2016. We further report on a reference laser system, comprising a rubidium stabilized DFB laser, which was operated successfully on the TEXUS 51 mission in April 2015. The system demonstrated a high level of technological maturity by remaining frequency stabilized throughout the mission including the rocket's boost phase.

  17. Space Shuttle Solid Rocket Booster Lightweight Recovery System

    NASA Technical Reports Server (NTRS)

    Wolf, Dean; Runkle, Roy E.

    1995-01-01

    The cancellation of the Advanced Solid Rocket Booster Project and the earth-to-orbit payload requirements for the Space Station dictated that the National Aeronautics and Space Administration (NASA) look at performance enhancements from all Space Transportation System (STS) elements (Orbiter Project, Space Shuttle Main Engine Project, External Tank Project, Solid Rocket Motor Project, & Solid Rocket Booster Project). The manifest for launching of Space Station components indicated that an additional 12-13000 pound lift capability was required on 10 missions and 15-20,000 pound additional lift capability is required on two missions. Trade studies conducted by all STS elements indicate that by deleting the parachute Recovery System (and associated hardware) from the Solid Rocket Boosters (SRBS) and going to a lightweight External Tank (ET) the 20,000 pound additional lift capability can be realized for the two missions. The deletion of the parachute Recovery System means the loss of four SRBs and this option is two expensive (loss of reusable hardware) to be used on the other 10 Space Station missions. Accordingly, each STS element looked at potential methods of weight savings, increased performance, etc. As the SRB and ET projects are non-propulsive (i.e. does not have launch thrust elements) their only contribution to overall payload enhancement can be achieved by the saving of weight while maintaining adequate safety factors and margins. The enhancement factor for the SRB project is 1:10. That is for each 10 pounds saved on the two SRBS; approximately 1 additional pound of payload in the orbiter bay can be placed into orbit. The SRB project decided early that the SRB recovery system was a prime candidate for weight reduction as it was designed in the early 1970s and weight optimization had never been a primary criteria.

  18. Liquid Rocket Booster (LRB) for the Space Transportion System (STS) systems study. Appendix D: Trade study summary for the liquid rocket booster

    NASA Technical Reports Server (NTRS)

    1989-01-01

    Trade studies plans for a number of elements in the Liquid Rocket Booster (LRB) component of the Space Transportation System (STS) are given in viewgraph form. Some of the elements covered include: avionics/flight control; avionics architecture; thrust vector control studies; engine control electronics; liquid rocket propellants; propellant pressurization systems; recoverable spacecraft; cryogenic tanks; and spacecraft construction materials.

  19. Precision volume measurement system.

    SciTech Connect

    Fischer, Erin E.; Shugard, Andrew D.

    2004-11-01

    A new precision volume measurement system based on a Kansas City Plant (KCP) design was built to support the volume measurement needs of the Gas Transfer Systems (GTS) department at Sandia National Labs (SNL) in California. An engineering study was undertaken to verify or refute KCP's claims of 0.5% accuracy. The study assesses the accuracy and precision of the system. The system uses the ideal gas law and precise pressure measurements (of low-pressure helium) in a temperature and computer controlled environment to ratio a known volume to an unknown volume.

  20. Carrier rockets

    NASA Astrophysics Data System (ADS)

    Aleksandrov, V. A.; Vladimirov, V. V.; Dmitriev, R. D.; Osipov, S. O.

    This book takes into consideration domestic and foreign developments related to launch vehicles. General information concerning launch vehicle systems is presented, taking into account details of rocket structure, basic design considerations, and a number of specific Soviet and American launch vehicles. The basic theory of reaction propulsion is discussed, giving attention to physical foundations, the various types of forces acting on a rocket in flight, basic parameters characterizing rocket motion, the effectiveness of various approaches to obtain the desired velocity, and rocket propellants. Basic questions concerning the classification of launch vehicles are considered along with construction and design considerations, aspects of vehicle control, reliability, construction technology, and details of structural design. Attention is also given to details of rocket motor design, the basic systems of the carrier rocket, and questions of carrier rocket development.

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

    NASA Technical Reports Server (NTRS)

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

    2010-01-01

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

  2. 15 CFR 744.3 - Restrictions on Certain Rocket Systems (including ballistic missile systems and space launch...

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 15 Commerce and Foreign Trade 2 2011-01-01 2011-01-01 false Restrictions on Certain Rocket Systems (including ballistic missile systems and space launch vehicles and sounding rockets) and Unmanned Air...: END-USER AND END-USE BASED § 744.3 Restrictions on Certain Rocket Systems (including ballistic...

  3. 15 CFR 744.3 - Restrictions on Certain Rocket Systems (including ballistic missile systems and space launch...

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 15 Commerce and Foreign Trade 2 2013-01-01 2013-01-01 false Restrictions on Certain Rocket Systems (including ballistic missile systems and space launch vehicles and sounding rockets) and Unmanned Air...: END-USER AND END-USE BASED § 744.3 Restrictions on Certain Rocket Systems (including ballistic...

  4. 15 CFR 744.3 - Restrictions on Certain Rocket Systems (including ballistic missile systems and space launch...

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 15 Commerce and Foreign Trade 2 2012-01-01 2012-01-01 false Restrictions on Certain Rocket Systems (including ballistic missile systems and space launch vehicles and sounding rockets) and Unmanned Air...: END-USER AND END-USE BASED § 744.3 Restrictions on Certain Rocket Systems (including ballistic...

  5. 15 CFR 744.3 - Restrictions on Certain Rocket Systems (including ballistic missile systems and space launch...

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 15 Commerce and Foreign Trade 2 2014-01-01 2014-01-01 false Restrictions on Certain Rocket Systems (including ballistic missile systems and space launch vehicles and sounding rockets) and Unmanned Air...: END-USER AND END-USE BASED § 744.3 Restrictions on Certain Rocket Systems (including ballistic...

  6. 15 CFR 744.3 - Restrictions on Certain Rocket Systems (including ballistic missile systems and space launch...

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 15 Commerce and Foreign Trade 2 2010-01-01 2010-01-01 false Restrictions on Certain Rocket Systems (including ballistic missile systems and space launch vehicles and sounding rockets) and Unmanned Air...: END-USER AND END-USE BASED § 744.3 Restrictions on Certain Rocket Systems (including ballistic...

  7. Automatic volume calibration system

    SciTech Connect

    Gates, A.J.; Aaron, C.C.

    1985-05-06

    The Automatic Volume Calibration System presently consists of three independent volume-measurement subsystems and can possibly be expanded to five subsystems. When completed, the system will manually or automatically perform the sequence of valve-control and data-acquisition operations required to measure given volumes. An LSI-11 minicomputer controls the vacuum and pressure sources and controls solenoid control valves to open and close various volumes. The input data are obtained from numerous displacement, temperature, and pressure sensors read by the LSI-11. The LSI-11 calculates the unknown volume from the data acquired during the sequence of valve operations. The results, based on the Ideal Gas Law, also provide information for feedback and control. This paper describes the volume calibration system, its subsystems, and the integration of the various instrumentation used in the system's design and development. 11 refs., 13 figs., 4 tabs.

  8. Liquid rocket booster study. Volume 2, book 5, appendix 9: LRB alternate applications and evolutionary growth

    NASA Technical Reports Server (NTRS)

    1989-01-01

    The analyses performed in assessing the merit of the Liquid Rocket Booster concept for use in alternate applications such as for Shuttle C, for Standalone Expendable Launch Vehicles, and possibly for use with the Air Force's Advanced Launch System are presented. A comparison is also presented of the three LRB candidate designs, namely: (1) the LO2/LH2 pump fed, (2) the LO2/RP-1 pump fed, and (3) the LO2/RP-1 pressure fed propellant systems in terms of evolution along with design and cost factors, and other qualitative considerations. A further description is also presented of the recommended LRB standalone, core-to-orbit launch vehicle concept.

  9. Feasibility study of superconducting LSM rocket launcher system

    NASA Technical Reports Server (NTRS)

    Yoshida, Kinjiro; Ohashi, Takaaki; Shiraishi, Katsuto; Takami, Hiroshi

    1994-01-01

    A feasibility study is presented concerning an application of a superconducting linear synchronous motor (LSM) to a large-scale rocket launcher, whose acceleration guide tube of LSM armature windings is constructed 1,500 meters under the ground. The rocket is released from the linear launcher just after it gets to a peak speed of about 900 kilometers per hour, and it flies out of the guide tube to obtain the speed of 700 kilometers per hour at the height of 100 meters above ground. The linear launcher is brought to a stop at the ground surface for a very short time of 5 seconds by a quick control of deceleration. Very large current variations in the single-layer windings of the LSM armature, which are produced at the higher speed region of 600 to 900 kilometers per hour, are controlled successfully by adopting the double-layer windings. The proposed control method makes the rocket launcher ascend stably in the superconducting LSM system, controlling the Coriolis force.

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

    NASA Technical Reports Server (NTRS)

    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.

  11. Intelligent control system for rocket engines

    NASA Astrophysics Data System (ADS)

    1993-02-01

    An intelligent control system (ICS) for a reusable space propulsion system for future launch vehicles is considered which is being developed in the NASA Lewis Research Center. A functional framework within which new engine-control functionalities are organized is developed for an SSME-like engine with expanded actuation capability. Control and diagnostic functions of this framework include primary engine control, real-time engine diagnostics, component condition monitoring, and sensor/actuator fault tolerance. It is noted that the controller should intelligently manage engine operation to achieve mission objectives while minimizing between-flight maintenance and maximizing engine life and performance.

  12. Large liquid rocket engine transient performance simulation system

    NASA Technical Reports Server (NTRS)

    Mason, J. R.; Southwick, R. D.

    1991-01-01

    A simulation system, ROCETS, was designed and developed to allow cost-effective computer predictions of liquid rocket engine transient performance. The system allows a user to generate a simulation of any rocket engine configuration using component modules stored in a library through high-level input commands. The system library currently contains 24 component modules, 57 sub-modules and maps, and 33 system routines and utilities. FORTRAN models from other sources can be operated in the system upon inclusion of interface information on comment cards. Operation of the simulation is simplified for the user by run, execution, and output processors. The simulation system makes available steady-state trim balance, transient operation, and linear partial generation. The system utilizes a modern equation solver for efficient operation of the simulations. Transient integration methods include integral and differential forms for the trapezoidal, first order Gear, and second order Gear corrector equations. A detailed technology test bed engine (TTBE) model was generated to be used as the acceptance test of the simulation system. The general level of model detail was that reflected in the Space Shuttle Main Engine DTM. The model successfully obtained steady-state balance in main stage operation and simulated throttle transients, including engine starts and shutdown. A NASA FORTRAN control model was obtained, ROCETS interface installed in comment cards, and operated with the TTBE model in closed-loop transient mode.

  13. Development of miniaturised low cost attitude determination system for sounding rockets

    NASA Astrophysics Data System (ADS)

    Bekkeng, Jan Kenneth; Booij, Wilfred; Moen, J.

    2005-08-01

    Spacecraft attitude (orientation) information is needed in order to transform scientific vector measurements in the reference frame of the rocket into a more meaningful Earth-fixed reference frame. By fusing data from a 3-axial magnetometer, a sun sensor and three rate gyros the rockets attitude can be determined (reconstructed). Since the system does not need to determine the attitude in real time (the attitude data is not used to control the rocket orientation), all data from the attitude sensors can be transmitted back to ground, where they are fused to estimate an absolute orientation of the rocket. A prototype inertial measurement unit and a miniature high accuracy lens-less sun sensor for spinning rocket is under development. A test version of both instruments will be launched on a single stage Hotel Payload sounding rocket from Andøya Rocket Range in July 2005.

  14. Nuclear thermal rocket workshop reference system Rover/NERVA

    NASA Technical Reports Server (NTRS)

    Borowski, Stanley K.

    1991-01-01

    The Rover/NERVA engine system is to be used as a reference, against which each of the other concepts presented in the workshop will be compared. The following topics are reviewed: the operational characteristics of the nuclear thermal rocket (NTR); the accomplishments of the Rover/NERVA programs; and performance characteristics of the NERVA-type systems for both Mars and lunar mission applications. Also, the issues of ground testing, NTR safety, NASA's nuclear propulsion project plans, and NTR development cost estimates are briefly discussed.

  15. Design of Electrical Systems for Rocket Propulsion Test Facilities at the John C. Stennis Space Center

    NASA Technical Reports Server (NTRS)

    Hughes, Mark S.; Davis, Dawn M.; Bakker, Henry J.; Jensen, Scott L.

    2007-01-01

    This viewgraph presentation reviews the design of the electrical systems that are required for the testing of rockets at the Rocket Propulsion Facility at NASA Stennis Space Center (NASA SSC). NASA/SSC s Mission in Rocket Propulsion Testing Is to Acquire Test Performance Data for Verification, Validation and Qualification of Propulsion Systems Hardware. These must be accurate reliable comprehensive and timely. Data acquisition in a rocket propulsion test environment is challenging: severe temporal transient dynamic environments, large thermal gradients, vacuum to 15 ksi pressure regimes SSC has developed and employs DAS, control systems and control systems and robust instrumentation that effectively satisfies these challenges.

  16. Unsteady Analyses of Valve Systems in Rocket Engine Testing Environments

    NASA Technical Reports Server (NTRS)

    Shipman, Jeremy; Hosangadi, Ashvin; Ahuja, Vineet

    2004-01-01

    This paper discusses simulation technology used to support the testing of rocket propulsion systems by performing high fidelity analyses of feed system components. A generalized multi-element framework has been used to perform simulations of control valve systems. This framework provides the flexibility to resolve the structural and functional complexities typically associated with valve-based high pressure feed systems that are difficult to deal with using traditional Computational Fluid Dynamics (CFD) methods. In order to validate this framework for control valve systems, results are presented for simulations of a cryogenic control valve at various plug settings and compared to both experimental data and simulation results obtained at NASA Stennis Space Center. A detailed unsteady analysis has also been performed for a pressure regulator type control valve used to support rocket engine and component testing at Stennis Space Center. The transient simulation captures the onset of a modal instability that has been observed in the operation of the valve. A discussion of the flow physics responsible for the instability and a prediction of the dominant modes associated with the fluctuations is presented.

  17. Distributed Health Monitoring System for Reusable Liquid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Lin, C. F.; Figueroa, F.; Politopoulos, T.; Oonk, S.

    2009-01-01

    The ability to correctly detect and identify any possible failure in the systems, subsystems, or sensors within a reusable liquid rocket engine is a major goal at NASA John C. Stennis Space Center (SSC). A health management (HM) system is required to provide an on-ground operation crew with an integrated awareness of the condition of every element of interest by determining anomalies, examining their causes, and making predictive statements. However, the complexity associated with relevant systems, and the large amount of data typically necessary for proper interpretation and analysis, presents difficulties in implementing complete failure detection, identification, and prognostics (FDI&P). As such, this paper presents a Distributed Health Monitoring System for Reusable Liquid Rocket Engines as a solution to these problems through the use of highly intelligent algorithms for real-time FDI&P, and efficient and embedded processing at multiple levels. The end result is the ability to successfully incorporate a comprehensive HM platform despite the complexity of the systems under consideration.

  18. Hybrid rocket propulsion systems for outer planet exploration missions

    NASA Astrophysics Data System (ADS)

    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.

  19. Precision volume measuring system

    SciTech Connect

    Klevgard, P.A.

    1984-11-01

    An engineering study was undertaken to calibrate and certify a precision volume measurement system that uses the ideal gas law and precise pressure measurements (of low-pressure helium) to ratio a known to an unknown volume. The constant-temperature, computer-controlled system was tested for thermodynamic instabilities, for precision (0.01%), and for bias (0.01%). Ratio scaling was used to optimize the quartz crystal pressure transducer calibration.

  20. The 2003 Goddard Rocket Replica Project: A Reconstruction of the World's First Functional Liquid Rocket System

    NASA Technical Reports Server (NTRS)

    Farr, R. A.; Elam, S. K.; Hicks, G. D.; Sanders, T. M.; London, J. R.; Mayne, A. W.; Christensen, D. L.

    2003-01-01

    As a part of NASA s 2003 Centennial of Flight celebration, engineers and technicians at Marshall Space Flight Center (MSFC), Huntsville, Alabama, in cooperation with the Alabama-Mississippi AIAA Section, have reconstructed historically accurate, functional replicas of Dr. Robert H. Goddard s 1926 first liquid- fuel rocket. The purposes of this project were to clearly understand, recreate, and document the mechanisms and workings of the 1926 rocket for exhibit and educational use, creating a vital resource for researchers studying the evolution of liquid rocketry for years to come. The MSFC team s reverse engineering activity has created detailed engineering-quality drawings and specifications describing the original rocket and how it was built, tested, and operated. Static hot-fire tests, as well as flight demonstrations, have further defined and quantified the actual performance and engineering actual performance and engineering challenges of this major segment in early aerospace history.

  1. Development of a hybrid rocket engine for the laboratory rocket system

    NASA Astrophysics Data System (ADS)

    Robinson, Joel W.

    1993-10-01

    The design, production, and testing of a Plexiglas-fueled, gaseous oxygen hybrid rocket engine is described. Hybrid fundamentals are discussed and previous testing is examined. Questions concerning the use of Plexiglas as a fuel and its performance are addressed. The experiment was divided into analyzing the ignition of the Plexiglas in ambient conditions, testing the engine with the verification of nitrogen and oxygen flow, the engine without the nozzle assembly and the engine with the nozzle assembly attached. The resulting graphs of data taken are explained. The small scale engine has answered valuable questions concerning the principles of hybrid rocket engine testing.

  2. Shuttle crew escape systems (CES) rocket test at Hurricane Mesa, Utah

    NASA Technical Reports Server (NTRS)

    1987-01-01

    Shuttle crew escape systems (CES) tractor rocket tests conducted at Hurricane Mesa, Utah. This preliminary ground test of the tractor rocket will lead up to in-air evaluations. View shows tractor rocket as it is fired from side hatch mockup. The tractor rocket concept is one of two escape methods being studied to provide crew egress capability during Space Shuttle controlled gliding flight. In-air tests of the system, utilizing a Convair-240 aircraft, will begin 11-19-87 at the Naval Weapons Center in China Lake, California.

  3. From Bridges and Rockets, Lessons for Software Systems

    NASA Technical Reports Server (NTRS)

    Holloway, C. Michael

    2004-01-01

    Although differences exist between building software systems and building physical structures such as bridges and rockets, enough similarities exist that software engineers can learn lessons from failures in traditional engineering disciplines. This paper draws lessons from two well-known failures the collapse of the Tacoma Narrows Bridge in 1940 and the destruction of the space shuttle Challenger in 1986 and applies these lessons to software system development. The following specific applications are made: (1) the verification and validation of a software system should not be based on a single method, or a single style of methods; (2) the tendency to embrace the latest fad should be overcome; and (3) the introduction of software control into safety-critical systems should be done cautiously.

  4. Wireless Data-Acquisition System for Testing Rocket Engines

    NASA Technical Reports Server (NTRS)

    Lin, Chujen; Lonske, Ben; Hou, Yalin; Xu, Yingjiu; Gang, Mei

    2007-01-01

    A prototype wireless data-acquisition system has been developed as a potential replacement for a wired data-acquisition system heretofore used in testing rocket engines. The traditional use of wires to connect sensors, signal-conditioning circuits, and data acquisition circuitry is time-consuming and prone to error, especially when, as is often the case, many sensors are used in a test. The system includes one master and multiple slave nodes. The master node communicates with a computer via an Ethernet connection. The slave nodes are powered by rechargeable batteries and are packaged in weatherproof enclosures. The master unit and each of the slave units are equipped with a time-modulated ultra-wide-band (TMUWB) radio transceiver, which spreads its RF energy over several gigahertz by transmitting extremely low-power and super-narrow pulses. In this prototype system, each slave node can be connected to as many as six sensors: two sensors can be connected directly to analog-to-digital converters (ADCs) in the slave node and four sensors can be connected indirectly to the ADCs via signal conditioners. The maximum sampling rate for streaming data from any given sensor is about 5 kHz. The bandwidth of one channel of the TM-UWB radio communication system is sufficient to accommodate streaming of data from five slave nodes when they are fully loaded with data collected through all possible sensor connections. TM-UWB radios have a much higher spatial capacity than traditional sinusoidal wave-based radios. Hence, this TM-UWB wireless data-acquisition can be scaled to cover denser sensor setups for rocket engine test stands. Another advantage of TM-UWB radios is that it will not interfere with existing wireless transmission. The maximum radio-communication range between the master node and a slave node for this prototype system is about 50 ft (15 m) when the master and slave transceivers are equipped with small dipole antennas. The range can be increased by changing to

  5. Integrated System Health Management (ISHM) Implementation in Rocket Engine Testing

    NASA Technical Reports Server (NTRS)

    Figueroa, Fernando; Morris, Jon; Turowski, Mark; Franzl, Richard; Walker, Mark; Kapadia, Ravi; Venkatesh, Meera

    2010-01-01

    A pilot operational ISHM capability has been implemented for the E-2 Rocket Engine Test Stand (RETS) and a Chemical Steam Generator (CSG) test article at NASA Stennis Space Center. The implementation currently includes an ISHM computer and a large display in the control room. The paper will address the overall approach, tools, and requirements. It will also address the infrastructure and architecture. Specific anomaly detection algorithms will be discussed regarding leak detection and diagnostics, valve validation, and sensor validation. It will also describe development and use of a Health Assessment Database System (HADS) as a repository for measurements, health, configuration, and knowledge related to a system with ISHM capability. It will conclude with a discussion of user interfaces, and a description of the operation of the ISHM system prior, during, and after testing.

  6. SRB/SLEEC (Solid Rocket Booster/Shingle Lap Extendible Exit Cone) feasibility study, volume 1

    NASA Technical Reports Server (NTRS)

    Baker, William H., Jr.

    1986-01-01

    A preliminary design and analysis was completed for a SLEEC (Shingle Lap Extendible Exit Cone) which could be incorporated on the Space Transportation System (STS) Solid Rocket Booster (SRB). Studies were completed which predicted weights and performance increases and development plans were prepared for the full-scale bench and static test of SLEEC. In conjunction with the design studies, a series of supporting analyses were performed to assure the validity and feasibility of performance, fabrication, cost, and reliability for the selected design. The feasibility and required amounts of bench, static firing, and flight tests considered necessary for the successful incorporation of SLEEC on the Shuttle SRBs were determined. Preliminary plans were completed which define both a follow on study effort and a development program.

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

    NASA Technical Reports Server (NTRS)

    1972-01-01

    An analysis of the solid propellant rocket engines for use with the space shuttle booster was conducted. A definition of the specific solid propellant rocket engine stage designs, development program requirements, production requirements, launch requirements, and cost data for each program phase were developed.

  8. Thirteenth Workshop for Computational Fluid Dynamic Applications in Rocket Propulsion and Launch Vehicle Technology. Volume 1

    NASA Technical Reports Server (NTRS)

    Williams, R. W. (Compiler)

    1996-01-01

    The purpose of the workshop was to discuss experimental and computational fluid dynamic activities in rocket propulsion and launch vehicles. The workshop was an open meeting for government, industry, and academia. A broad number of topics were discussed including computational fluid dynamic methodology, liquid and solid rocket propulsion, turbomachinery, combustion, heat transfer, and grid generation.

  9. Demonstration of a sterilizable solid rocket motor system

    NASA Technical Reports Server (NTRS)

    Mastrolia, E. J.; Santerre, G. M.; Lambert, W. L.

    1975-01-01

    A solid propellant rocket motor containing 60.9 Kg (134-lb) of propellant was successfully static fired after being subjected to eight heat sterilization cycles (three 54-hour cycles plus five 40-hour cycles) at 125 C (257 F). The test motor, a modified SVM-3 chamber, incorporated a flexible grain retention system of EPR rubber to relieve thermal shrinkage stresses. The propellant used in the motor was ANB-3438, and 84 wt% solids system (18 wt% aluminum) containing 66 wt% stabilized ammonium perchlorate oxidizer and a saturated hydroxylterminated polybutadiene binder. Bonding of the propellant to the EPR insulation (GenGard V-4030) was provided by the use of SD-886, an epoxy urethane restriction.

  10. Materials Problems in Chemical Liquid-Propellant Rocket Systems

    NASA Technical Reports Server (NTRS)

    Gilbert, L. L.

    1959-01-01

    With the advent of the space age, new adjustments in technical thinking and engineering experience are necessary. There is an increasing and extensive interest in the utilization of materials for components to be used at temperatures ranging from -423 to over 3500 deg F. This paper presents a description of the materials problems associated with the various components of chemical liquid rocket systems. These components include cooled and uncooled thrust chambers, injectors, turbine drive systems, propellant tanks, and cryogenic propellant containers. In addition to materials limitations associated with these components, suggested research approaches for improving materials properties are made. Materials such as high-temperature alloys, cermets, carbides, nonferrous alloys, plastics, refractory metals, and porous materials are considered.

  11. Non-Rocket Earth-Moon Transport System

    NASA Technical Reports Server (NTRS)

    Bolonkin, Alexander

    2002-01-01

    This paper proposes a new method and transportation system to travel to the Moon. This transportation system uses a mechanical energy transfer and requires only minimal energy so that it provides a 'Free Trip' into space. The method uses the rotary and kinetic energy of the Moon. This paper presents the theory and results of computations for the project provided Free Trips (without rockets and spend a big energy) to the Moon for six thousand people annually. The project uses artificial materials like nanotubes and whiskers that have a ratio of tensile strength to density equal 4 million meters. In the future, nanotubes will be produced that can reach a specific stress up 100 millions meter and will significantly improve the parameters of suggested project. The author is prepared to discuss the problems with serious organizations that want to research and develop these innovations.

  12. Qualitative model-based diagnostics for rocket systems

    NASA Technical Reports Server (NTRS)

    Maul, William; Meyer, Claudia; Jankovsky, Amy; Fulton, Christopher

    1993-01-01

    A diagnostic software package is currently being developed at NASA LeRC that utilizes qualitative model-based reasoning techniques. These techniques can provide diagnostic information about the operational condition of the modeled rocket engine system or subsystem. The diagnostic package combines a qualitative model solver with a constraint suspension algorithm. The constraint suspension algorithm directs the solver's operation to provide valuable fault isolation information about the modeled system. A qualitative model of the Space Shuttle Main Engine's oxidizer supply components was generated. A diagnostic application based on this qualitative model was constructed to process four test cases: three numerical simulations and one actual test firing. The diagnostic tool's fault isolation output compared favorably with the input fault condition.

  13. Recovery system of sounding rocket S-520 type of ISAS

    NASA Astrophysics Data System (ADS)

    Hinada, M.; Hayashi, T.; Matsuo, H.; Tsukamoto, S.; Ohshima, T.; Maeda, Y.; Kamata, Y.

    The design, performance, equipment, and initial flight test results for the S-520 solid fueled rocket are described. The S-520 carries a 408 kg payload, delivers 28,000 lb thrust, has a specific impulse of 29 sec, attains a velocity of 2.13 km/sec, and can reach 232 km altitude. The first S-520 was 9 m long and 0.524 m in diameter and had a three-stage configuration: engine, recovery test unit (RCU), and scientific instruments. The RTU contained parachute and flotation systems and a Loran-C locator beacon. The initial flight in September 1981 resulted in a splashdown 320 km downrange, with recovery accomplished within 4 hr of launch. Further flight tests and design upgrades are planned.

  14. Solid Rocket Booster (SRB) Flight System Integration at Its Best

    NASA Technical Reports Server (NTRS)

    Wood, T. David; Kanner, Howard S.; Freeland, Donna M.; Olson, Derek T.

    2011-01-01

    The Solid Rocket Booster (SRB) element integrates all the subsystems needed for ascent flight, entry, and recovery of the combined Booster and Motor system. These include the structures, avionics, thrust vector control, pyrotechnic, range safety, deceleration, thermal protection, and retrieval systems. This represents the only human-rated, recoverable and refurbishable solid rocket ever developed and flown. Challenges included subsystem integration, thermal environments and severe loads (including water impact), sometimes resulting in hardware attrition. Several of the subsystems evolved during the program through design changes. These included the thermal protection system, range safety system, parachute/recovery system, and others. Because the system was recovered, the SRB was ideal for data and imagery acquisition, which proved essential for understanding loads, environments and system response. The three main parachutes that lower the SRBs to the ocean are the largest parachutes ever designed, and the SRBs are the largest structures ever to be lowered by parachutes. SRB recovery from the ocean was a unique process and represented a significant operational challenge; requiring personnel, facilities, transportation, and ground support equipment. The SRB element achieved reliability via extensive system testing and checkout, redundancy management, and a thorough postflight assessment process. However, the in-flight data and postflight assessment process revealed the hardware was affected much more strongly than originally anticipated. Assembly and integration of the booster subsystems required acceptance testing of reused hardware components for each build. Extensive testing was done to assure hardware functionality at each level of stage integration. Because the booster element is recoverable, subsystems were available for inspection and testing postflight, unique to the Shuttle launch vehicle. Problems were noted and corrective actions were implemented as needed

  15. Thirteenth Workshop for Computational Fluid Dynamic Applications in Rocket Propulsion and Launch Vehicle Technology. Volume 2

    NASA Technical Reports Server (NTRS)

    Williams, R. W. (Compiler)

    1996-01-01

    This conference publication includes various abstracts and presentations given at the 13th Workshop for Computational Fluid Dynamic Applications in Rocket Propulsion and Launch Vehicle Technology held at the George C. Marshall Space Flight Center April 25-27 1995. The purpose of the workshop was to discuss experimental and computational fluid dynamic activities in rocket propulsion and launch vehicles. The workshop was an open meeting for government, industry, and academia. A broad number of topics were discussed including computational fluid dynamic methodology, liquid and solid rocket propulsion, turbomachinery, combustion, heat transfer, and grid generation.

  16. Liquid rocket booster study. Volume 2, book 3, appendices 2-5: PPIP, transition plan, AMOS plan, and environmental analysis

    NASA Technical Reports Server (NTRS)

    1988-01-01

    This Preliminary Project Implementation Plan (PPIP) was used to examine the feasibility of replacing the current Solid Rocket Boosters on the Space Shuttle with Liquid Rocket Boosters (LRBs). The need has determined the implications of integrating the LRB with the Space Transportation System as the earliest practical date. The purpose was to identify and define all elements required in a full scale development program for the LRB. This will be a reference guide for management of the LRB program, addressing such requirement as design and development, configuration management, performance measurement, manufacturing, product assurance and verification, launch operations, and mission operations support.

  17. Large liquid rocket engine transient performance simulation system

    NASA Technical Reports Server (NTRS)

    Mason, J. R.; Southwick, R. D.

    1989-01-01

    Phase 1 of the Rocket Engine Transient Simulation (ROCETS) program consists of seven technical tasks: architecture; system requirements; component and submodel requirements; submodel implementation; component implementation; submodel testing and verification; and subsystem testing and verification. These tasks were completed. Phase 2 of ROCETS consists of two technical tasks: Technology Test Bed Engine (TTBE) model data generation; and system testing verification. During this period specific coding of the system processors was begun and the engineering representations of Phase 1 were expanded to produce a simple model of the TTBE. As the code was completed, some minor modifications to the system architecture centering on the global variable common, GLOBVAR, were necessary to increase processor efficiency. The engineering modules completed during Phase 2 are listed: INJTOO - main injector; MCHBOO - main chamber; NOZLOO - nozzle thrust calculations; PBRNOO - preburner; PIPE02 - compressible flow without inertia; PUMPOO - polytropic pump; ROTROO - rotor torque balance/speed derivative; and TURBOO - turbine. Detailed documentation of these modules is in the Appendix. In addition to the engineering modules, several submodules were also completed. These submodules include combustion properties, component performance characteristics (maps), and specific utilities. Specific coding was begun on the system configuration processor. All functions necessary for multiple module operation were completed but the SOLVER implementation is still under development. This system, the Verification Checkout Facility (VCF) allows interactive comparison of module results to store data as well as provides an intermediate checkout of the processor code. After validation using the VCF, the engineering modules and submodules were used to build a simple TTBE.

  18. Block 2 Solid Rocket Motor (SRM) conceptual design study, volume 1

    NASA Technical Reports Server (NTRS)

    1986-01-01

    Segmented and monolithic Solid Rocket Motor (SRM) design concepts were evaluated with emphasis on joints and seals. Particular attention was directed to eliminating deficiencies in the SRM High Performance Motor (HPM). The selected conceptual design is described and discussed.

  19. Block 2 Solid Rocket Motor (SRM) conceptual design study. Volume 1: Appendices

    NASA Technical Reports Server (NTRS)

    1986-01-01

    The design studies task implements the primary objective of developing a Block II Solid Rocket Motor (SRM) design offering improved flight safety and reliability. The SRM literature was reviewed. The Preliminary Development and Validation Plan is presented.

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

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

    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.

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

    NASA Astrophysics Data System (ADS)

    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.

  2. Prediction of pressure fluctuation in sounding rockets and manifolded recovery systems

    NASA Technical Reports Server (NTRS)

    Laudadio, J. F.

    1972-01-01

    The determination of altitude by means of barometric sensors in sounding rocket applications is discussed. A method for predicting the performance of such sensing systems is needed. A method is developed for predicting the pressure-time response of a volume subjected to subsonic air flow through from one to four passages. The pressure calculation is based on one-dimensional gas flow with friction. A computed program has been developed which solves the differential equations using a self-starting predictor-corrector integration technique. The input data required are the pressure sensing system dimensions, pressure forcing function(s) at the inlet port(s), and a trajectory over the time of analysis (altitude-velocity-time), if the forcing function is trajectory dependent. The program then computes the pressure-temperature history of the gas in the manifold over the time interval specified.

  3. Modified modular imaging system designed for a sounding rocket experiment

    NASA Astrophysics Data System (ADS)

    Veach, Todd J.; Scowen, Paul A.; Beasley, Matthew; Nikzad, Shouleh

    2012-09-01

    We present the design and system calibration results from the fabrication of a charge-coupled device (CCD) based imaging system designed using a modified modular imager cell (MIC) used in an ultraviolet sounding rocket mission. The heart of the imaging system is the MIC, which provides the video pre-amplifier circuitry and CCD clock level filtering. The MIC is designed with standard four-layer FR4 printed circuit board (PCB) with surface mount and through-hole components for ease of testing and lower fabrication cost. The imager is a 3.5k by 3.5k LBNL p-channel CCD with enhanced quantum efficiency response in the UV using delta-doping technology at JPL. The recently released PCIe/104 Small-Cam CCD controller from Astronomical Research Cameras, Inc (ARC) performs readout of the detector. The PCIe/104 Small-Cam system has the same capabilities as its larger PCI brethren, but in a smaller form factor, which makes it ideally suited for sub-orbital ballistic missions. The overall control is then accomplished using a PCIe/104 computer from RTD Embedded Technologies, Inc. The design, fabrication, and testing was done at the Laboratory for Astronomical and Space Instrumentation (LASI) at Arizona State University. Integration and flight calibration are to be completed at the University of Colorado Boulder before integration into CHESS.

  4. Simulation of an advanced techniques of ion propulsion Rocket system

    NASA Astrophysics Data System (ADS)

    Bakkiyaraj, R.

    2016-07-01

    The ion propulsion rocket system is expected to become popular with the development of Deuterium,Argon gas and Hexagonal shape Magneto hydrodynamic(MHD) techniques because of the stimulation indirectly generated the power from ionization chamber,design of thrust range is 1.2 N with 40 KW of electric power and high efficiency.The proposed work is the study of MHD power generation through ionization level of Deuterium gas and combination of two gaseous ions(Deuterium gas ions + Argon gas ions) at acceleration stage.IPR consists of three parts 1.Hexagonal shape MHD based power generator through ionization chamber 2.ion accelerator 3.Exhaust of Nozzle.Initially the required energy around 1312 KJ/mol is carrying out the purpose of deuterium gas which is changed to ionization level.The ionized Deuterium gas comes out from RF ionization chamber to nozzle through MHD generator with enhanced velocity then after voltage is generated across the two pairs of electrode in MHD.it will produce thrust value with the help of mixing of Deuterium ion and Argon ion at acceleration position.The simulation of the IPR system has been carried out by MATLAB.By comparing the simulation results with the theoretical and previous results,if reaches that the proposed method is achieved of thrust value with 40KW power for simulating the IPR system.

  5. Study of solid rocket motor for space shuttle booster. Volume 4: Cost

    NASA Technical Reports Server (NTRS)

    1972-01-01

    The cost data for solid propellant rocket engines for use with the space shuttle are presented. The data are based on the selected 156 inch parallel and series burn configurations. Summary cost data are provided for the production of the 120 inch and 260 inch configurations. Graphs depicting parametric cost estimating relationships are included.

  6. SRB-3D Solid Rocket Booster performance prediction program. Volume 3: Programmer's manual

    NASA Technical Reports Server (NTRS)

    Winkler, J. C.

    1976-01-01

    The programmer's manual for the Modified Solid Rocket Booster Performance Prediction Program (SRB-3D) describes the major control routines of SRB-3D, followed by a super index listing of the program and a cross-reference of the program variables.

  7. Study of solid rocket motors for a space shuttle booster. Volume 3: Program acquisition planning

    NASA Technical Reports Server (NTRS)

    Vonderesch, A. H.

    1972-01-01

    Plans for conducting Phase C/D for a solid rocket motor booster vehicle are presented. Methods for conducting this program with details of scheduling, testing, and program management and control are included. The requirements of the space shuttle program to deliver a minimum cost/maximum reliability booster vehicle are examined.

  8. Test of Re-Entry Systems at Estrange Using Sounding Rockets and Stratospheric Balloons

    NASA Astrophysics Data System (ADS)

    Lockowandt, C.; Abrahamsson, M.; Florin, G.

    2015-09-01

    Stratospheric balloons and sounding rockets can provide an ideal in-flight platform for performing re-entry and other high speed tests off different types of vehicles and techniques. They are also ideal platforms for testing different types of recovery systems such as airbrakes and parachutes. This paper expands on some examples of platforms and missions for drop tests from balloons as well as sounding rockets launched from Esrange Space Center, a facility run by Swedish Space Corporation SSC in northern Sweden.

  9. A Low Cost GPS System for Real-Time Tracking of Sounding Rockets

    NASA Technical Reports Server (NTRS)

    Markgraf, M.; Montenbruck, O.; Hassenpflug, F.; Turner, P.; Bull, B.; Bauer, Frank (Technical Monitor)

    2001-01-01

    This paper describes the development as well as the on-ground and the in-flight evaluation of a low cost Global Positioning System (GPS) system for real-time tracking of sounding rockets. The flight unit comprises a modified ORION GPS receiver and a newly designed switchable antenna system composed of a helical antenna in the rocket tip and a dual-blade antenna combination attached to the body of the service module. Aside from the flight hardware a PC based terminal program has been developed to monitor the GPS data and graphically displays the rocket's path during the flight. In addition an Instantaneous Impact Point (IIP) prediction is performed based on the received position and velocity information. In preparation for ESA's Maxus-4 mission, a sounding rocket test flight was carried out at Esrange, Kiruna, on 19 Feb. 2001 to validate existing ground facilities and range safety installations. Due to the absence of a dedicated scientific payload, the flight offered the opportunity to test multiple GPS receivers and assess their performance for the tracking of sounding rockets. In addition to the ORION receiver, an Ashtech G12 HDMA receiver and a BAE (Canadian Marconi) Allstar receiver, both connected to a wrap-around antenna, have been flown on the same rocket as part of an independent experiment provided by the Goddard Space Flight Center. This allows an in-depth verification and trade-off of different receiver and antenna concepts.

  10. Reusable Rocket Engine Advanced Health Management System. Architecture and Technology Evaluation: Summary

    NASA Technical Reports Server (NTRS)

    Pettit, C. D.; Barkhoudarian, S.; Daumann, A. G., Jr.; Provan, G. M.; ElFattah, Y. M.; Glover, D. E.

    1999-01-01

    In this study, we proposed an Advanced Health Management System (AHMS) functional architecture and conducted a technology assessment for liquid propellant rocket engine lifecycle health management. The purpose of the AHMS is to improve reusable rocket engine safety and to reduce between-flight maintenance. During the study, past and current reusable rocket engine health management-related projects were reviewed, data structures and health management processes of current rocket engine programs were assessed, and in-depth interviews with rocket engine lifecycle and system experts were conducted. A generic AHMS functional architecture, with primary focus on real-time health monitoring, was developed. Fourteen categories of technology tasks and development needs for implementation of the AHMS were identified, based on the functional architecture and our assessment of current rocket engine programs. Five key technology areas were recommended for immediate development, which (1) would provide immediate benefits to current engine programs, and (2) could be implemented with minimal impact on the current Space Shuttle Main Engine (SSME) and Reusable Launch Vehicle (RLV) engine controllers.

  11. The development of a post-test diagnostic system for rocket engines

    NASA Technical Reports Server (NTRS)

    Zakrajsek, June F.

    1991-01-01

    An effort was undertaken by NASA to develop an automated post-test, post-flight diagnostic system for rocket engines. The automated system is designed to be generic and to automate the rocket engine data review process. A modular, distributed architecture with a generic software core was chosen to meet the design requirements. The diagnostic system is initially being applied to the Space Shuttle Main Engine data review process. The system modules currently under development are the session/message manager, and portions of the applications section, the component analysis section, and the intelligent knowledge server. An overview is presented of a rocket engine data review process, the design requirements and guidelines, the architecture and modules, and the projected benefits of the automated diagnostic system.

  12. Stage separation study of Nike-Black Brant V Sounding Rocket System

    NASA Technical Reports Server (NTRS)

    Ferragut, N. J.

    1976-01-01

    A new Sounding Rocket System has been developed. It consists of a Nike Booster and a Black Brant V Sustainer with slanted fins which extend beyond its nozzle exit plane. A cursory look was taken at different factors which must be considered when studying a passive separation system. That is, one separation system without mechanical constraints in the axial direction and which will allow separation due to drag differential accelerations between the Booster and the Sustainer. The equations of motion were derived for rigid body motions and exact solutions were obtained. The analysis developed could be applied to any other staging problem of a Sounding Rocket System.

  13. Rocket injector anomalies study. Volume 1: Description of the mathematical model and solution procedure

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

    The capability of simulating three dimensional two phase reactive flows with combustion in the liquid fuelled rocket engines is demonstrated. This was accomplished by modifying an existing three dimensional computer program (REFLAN3D) with Eulerian Lagrangian approach to simulate two phase spray flow, evaporation and combustion. The modified code is referred as REFLAN3D-SPRAY. The mathematical formulation of the fluid flow, heat transfer, combustion and two phase flow interaction of the numerical solution procedure, boundary conditions and their treatment are described.

  14. Dynamics of variable mass systems with application to the star 48 solid rocket motor

    NASA Technical Reports Server (NTRS)

    Eke, F. O.

    1984-01-01

    Existing methods for the derivation of equations of motion of variable mass systems are reviewed and compared, the end product being a system of general dynamical equations for variable mass systems. These equations are used to study the lateral stability problem associated with the Star 48 solid rocket engine. It is shown that the shape of the combustion chamber could have a significant effect on the lateral stability of the rocket; specifically, a short and wide combustion chamber is destabilizing, while a long and narrow chamber is stabilizing.

  15. User's manual for rocket combustor interactive design (ROCCID) and analysis computer program. Volume 1: User's manual

    NASA Technical Reports Server (NTRS)

    Muss, J. A.; Nguyen, T. V.; Johnson, C. W.

    1991-01-01

    The user's manual for the rocket combustor interactive design (ROCCID) computer program is presented. The program, written in Fortran 77, provides a standardized methodology using state of the art codes and procedures for the analysis of a liquid rocket engine combustor's steady state combustion performance and combustion stability. The ROCCID is currently capable of analyzing mixed element injector patterns containing impinging like doublet or unlike triplet, showerhead, shear coaxial, and swirl coaxial elements as long as only one element type exists in each injector core, baffle, or barrier zone. Real propellant properties of oxygen, hydrogen, methane, propane, and RP-1 are included in ROCCID. The properties of other propellants can easily be added. The analysis model in ROCCID can account for the influence of acoustic cavities, helmholtz resonators, and radial thrust chamber baffles on combustion stability. ROCCID also contains the logic to interactively create a combustor design which meets input performance and stability goals. A preliminary design results from the application of historical correlations to the input design requirements. The steady state performance and combustion stability of this design is evaluated using the analysis models, and ROCCID guides the user as to the design changes required to satisfy the user's performance and stability goals, including the design of stability aids. Output from ROCCID includes a formatted input file for the standardized JANNAF engine performance prediction procedure.

  16. Rockets for spin recovery

    NASA Technical Reports Server (NTRS)

    Whipple, R. D.

    1980-01-01

    The potential effectiveness of rockets as an auxiliary means for an aircraft to effect recovery from spins was investigated. The advances in rocket technology produced by the space effort suggested that currently available systems might obviate many of the problems encountered in earlier rocket systems. A modern fighter configuration known to exhibit a flat spin mode was selected. An analytical study was made of the thrust requirements for a rocket spin recovery system for the subject configuration. These results were then applied to a preliminary systems study of rocket components appropriate to the problem. Subsequent spin tunnel tests were run to evaluate the analytical results.

  17. The Development of a Fiber Optic Raman Temperature Measurement System for Rocket Flows

    NASA Technical Reports Server (NTRS)

    Degroot, Wim A.

    1992-01-01

    A fiberoptic Raman diagnostic system for H2/O2 rocket flows is currently under development. This system is designed for measurement of temperature and major species concentration in the combustion chamber and part of the nozzle of a 100 Newton thrust rocket currently undergoing testing. This paper describes a measurement system based on the spontaneous Raman scattering phenomenon. An analysis of the principles behind the technique is given. Software is developed to measure temperature and major species concentrations by comparing theoretical Raman scattering spectra with experimentally obtained spectra. Equipment selection and experimental approach are summarized. This experimental program is part of a program, which is in progress, to evaluate Navier-Stokes based analyses for this class of rocket.

  18. Current and Future Critical Issues in Rocket Propulsion Systems

    NASA Technical Reports Server (NTRS)

    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).

  19. Disturbance Rejection Based Test Rocket Control System Design and Validation

    NASA Astrophysics Data System (ADS)

    Yang, H.; Zhang, S.; Li, T.; Zhang, Y.

    2015-09-01

    This paper presents a novel design and validation for the three-channel attitude controller of a STT test rocket based on the extended state observer approach. The uniform second order integral-chain state space model is firstly established for the control variable of the angle of attack, angle of sideslip and roll angle. Combined with the pole placement, the extended state observer is applied to the disturbance rejection design of the attitude controller. Through numerical and hardware-in-the-loop simulation with uncertainties considered, the effectiveness and robustness of the controller are illustrated and verified. Finally, the performance of the controller is validated by flight-test with satisfactory results.

  20. NASA's Hydrogen Outpost: The Rocket Systems Area at Plum Brook Station

    NASA Technical Reports Server (NTRS)

    Arrighi, Robert S.

    2016-01-01

    "There was pretty much a general knowledge about hydrogen and its capabilities," recalled former researcher Robert Graham. "The question was, could you use it in a rocket engine? Do we have the technology to handle it? How will it cool? Will it produce so much heat release that we can't cool the engine? These were the questions that we had to address." The National Aeronautics and Space Administration's (NASA) Glenn Research Center, referred to historically as the Lewis Research Center, made a concerted effort to answer these and related questions in the 1950s and 1960s. The center played a critical role transforming hydrogen's theoretical potential into a flight-ready propellant. Since then NASA has utilized liquid hydrogen to send humans and robots to the Moon, propel dozens of spacecraft across the universe, orbit scores of satellite systems, and power 135 space shuttle flights. Rocket pioneers had recognized hydrogen's potential early on, but its extremely low boiling temperature and low density made it impracticable as a fuel. The Lewis laboratory first demonstrated that liquid hydrogen could be safely utilized in rocket and aircraft propulsion systems, then perfected techniques to store, pump, and cleanly burn the fuel, as well as use it to cool the engine. The Rocket Systems Area at Lewis's remote testing area, Plum Brook Station, played a little known, but important role in the center's hydrogen research efforts. This publication focuses on the activities at the Rocket Systems Area, but it also discusses hydrogen's role in NASA's space program and Lewis's overall hydrogen work. The Rocket Systems Area included nine physically modest test sites and three test stands dedicated to liquid-hydrogen-related research. In 1962 Cleveland Plain Dealer reporter Karl Abram claimed, "The rocket facility looks more like a petroleum refinery. Its test rigs sprout pipes, valves and tanks. During the night test runs, excess hydrogen is burned from special stacks in the best

  1. A three body dynamic simulation of a seated tractor rocket escape system for the Space Shuttle

    NASA Technical Reports Server (NTRS)

    Ondler, R. M.

    1989-01-01

    In the tractor-rocket seated-extraction candidate system for Space Shuttle Orbiter crew escape, the crewmember is pulled from his seat and away from the Orbiter via an elastic pendant, using a system of rails to guide the extraction trajectory through an opening on the window frame for flight deck crew and through the side hatch for the middeck crew. A three-body simulation has been developed to model the flight-mechanics aspects of the concept, where the three bodies are the astronaut (six DOF), the tractor rocket (six DOF), and the Shuttle Orbiter (three DOF); attention is given to crewmembers' clearance of the Orbiter structure and engine plumes.

  2. Volume measuring system

    NASA Technical Reports Server (NTRS)

    Oele, J. S.

    1975-01-01

    Chamber is designed to be airtight; it includes face mask for person to breathe outside air so that he does not disturb chamber environment. Chamber includes piston to vary air volume inside. Also included are two microphone transducers which record pressure information inside chamber.

  3. Study of solid rocket motors for a space shuttle booster. Volume 2, book 3, addendum 1: Cost estimating data

    NASA Technical Reports Server (NTRS)

    Vonderesch, A. H.

    1972-01-01

    A second iteration of the program baseline configuration and cost for the solid propellant rocket engines used with the space shuttle booster system is presented. The purpose of the study was to ensure that total program costs were complete and to review areas where costs might be overly conservative and could be reduced. Labor and material were analyzed in more depth, more definition was prepared to separate recurring from nonrecurring costs, and the operations portions of the engine and stage were separated into more identifiable activities.

  4. Study of solid rocket motor for space shuttle booster, volume 2, book 5, appendices E thru H

    NASA Technical Reports Server (NTRS)

    1972-01-01

    Preliminary parametric studies were performed to establish size, weight and packaging arrangements for aerodynamic decelerator devices that could be used for recovery of the expended solid propellant rocket motors used in the launch phase of the Space Shuttle System. Computations were made using standard engineering analysis techniques. Terminal stage parachutes were sized to provide equilibrium descent velocities for water entry that are presently thought to be acceptable without developing loads that could exceed the boosters structural integrity. The performance characteristics of the aerodynamic parachute decelerator devices considered are based on analysis and prior test results for similar configurations and are assumed to be maintained at the scale requirements of the present problem.

  5. An intelligent control system for rocket engines - Need, vision, and issues

    NASA Technical Reports Server (NTRS)

    Lorenzo, Carl F.; Merrill, Walter C.

    1991-01-01

    Several components of intelligence are defined. Within the context of these definitions an intelligent control system for rocket engines is described. The description includes a framework for development of an intelligent control system, including diagnostics, coordination, and direct control. Some current results and issues are presented.

  6. Cyclic fatigue analysis of rocket thrust chambers. Volume 1: OFHC copper chamber low cycle fatigue

    NASA Technical Reports Server (NTRS)

    Miller, R. W.

    1974-01-01

    A three-dimensional finite element elasto-plastic strain analysis was performed for the throat section of a regeneratively cooled rocket combustion chamber. The analysis employed the RETSCP finite element computer program. The analysis included thermal and pressure loads, and the effects of temperature dependent material properties, to determine the strain range corresponding to the chamber operating cycle. The analysis was performed for chamber configuration and operating conditions corresponding to a hydrogen-oxygen combustion chamber which was fatigue tested to failure. The computed strain range at typical chamber operating conditions was used in conjunction with oxygen-free, high-conductivity (OHFC) copper isothermal fatigue test data to predict chamber low-cycle fatigue life.

  7. Thrust chamber life prediction. Volume 1: Mechanical and physical properties of high performance rocket nozzle materials

    NASA Technical Reports Server (NTRS)

    Esposito, J. J.; Zabora, R. F.

    1975-01-01

    Pertinent mechanical and physical properties of six high conductivity metals were determined. The metals included Amzirc, NARloy Z, oxygen free pure copper, electroformed copper, fine silver, and electroformed nickel. Selection of these materials was based on their possible use in high performance reusable rocket nozzles. The typical room temperature properties determined for each material included tensile ultimate strength, tensile yield strength, elongation, reduction of area, modulus of elasticity, Poisson's ratio, density, specific heat, thermal conductivity, and coefficient of thermal expansion. Typical static tensile stress-strain curves, cyclic stress-strain curves, and low-cycle fatigue life curves are shown. Properties versus temperature are presented in graphical form for temperatures from 27.6K (-410 F) to 810.9K (1000 F).

  8. Solid rocket booster performance evaluation model. Volume 3: Sample case. [propellant combustion simulation/internal ballistics

    NASA Technical Reports Server (NTRS)

    1974-01-01

    The solid rocket booster performance evaluation model (SRB-11) is used to predict internal ballistics in a sample motor. This motor contains a five segmented grain. The first segment has a 14 pointed star configuration with a web which wraps partially around the forward dome. The other segments are circular in cross-section and are tapered along the interior burning surface. Two of the segments are inhibited on the forward face. The nozzle is not assumed to be submerged. The performance prediction is broken into two simulation parts: the delivered end item specific impulse and the propellant properties which are required as inputs for the internal ballistics module are determined; and the internal ballistics for the entire burn duration of the motor are simulated.

  9. Non-rocket Earth-Moon transportation system

    NASA Astrophysics Data System (ADS)

    Bolonkin, A.

    Author suggests and researches one of his methods of flights to outer Space, described in book "Non Rocket Flights in Space", which is prepared and offered for publication. In given report the method and facilities named "Bolonkin Transport System" (BTS) for delivering of payload and people to Moon and back is presented. BTS can be used also for free trip to outer Space up at altitude 60,000 km and more. BTS can be applying as a trust system for atmospheric supersonic aircrafts, and as a free energy source. This method uses, in general, the rotary and kinetic energy of the Moon. The manuscript contains the theory and results of computation of special Project. This project uses three cables (main and two for driving of loads) from artificial material: fiber, whiskers, nanotubes, with the specific tensile strength (ratio the tensile stress to density) k=/=4*10^7 or more. The nanotubes with same and better parameters are received in scientific laboratories. Theoretical limit of nanotubes SWNT is about k=100*10^7. The upper end of the cable is connected to the Moon. The lower end of the cable is connected to an aircraft (or buoy), which flies (i.e. glides or slides) in Earth atmosphere along the planet's surface. The aircraft (and Moon) has devices, which allows the length of cables to be changed. The device would consists of a spool, motor, brake, transmission, and controller. The facility could have devices for delivering people and payloads t o the Moon and back using the suggested Transport System. The delivery devices include: containers, cables, motors, brakes, and controllers. If the aircraft is small and the cable is strong the motion of the Moon can be used to move the airplane. For example (see enclosed project), if the airplane weighs 15 tons and has an aerodynamic ratio (the lift force to the drag force) equal 5, a thrust of 3000 kg would be enough for the aircraft to fly for infinity without requiring any fuel. The aircraft could use a small turbine engine

  10. Dynamics of a variable mass system applied to spacecraft rocket attitude theory

    NASA Astrophysics Data System (ADS)

    Mudge, Jason Dominic

    This research project is a study of the dynamics of a variable mass system. The scope of this research project is to gain understanding as to how a variable mass system will behave. The intent is to bring the level of understanding of variable mass dynamics higher and closer to the level of constant mass dynamics in the area of spacecrafts in particular. A main contribution is the finding of a set of criteria to minimize or eliminate the deviation of the nutation angle (or cone angle or angle of attack) of spacecraft rockets passively, i.e. without active control. The motivation for this research project is the Star 48 anomaly. The Star 48 is a solid rocket motor which has propelled (boosted) communication satellites from lower earth orbit to a higher one during the 1980's. The anomaly is that when the spacecraft rocket is being propelled, the nutation angle may deviate excessively which is considered undesirable. In the first part of this research project, a variable mass system is described and defined and the governing equations are derived. The type of governing equations derived are those that are most useful for analyzing the motion of a spacecraft rocket. The method of derivation makes use of Leibnitz Theorem, Divergence Theorem and Newton's Second Law of Motion. Next, the governing equations are specialized with several assumptions which are generally accepted assumptions applied in the analysis of spacecraft rockets. With these assumptions, the form governing equations is discussed and then the equations are solved analytically for the system's angular velocity. Having solved for the angular velocity of the system, the attitude of the system is obtained using a unique method which circumvents the nonlinearities that exist using Euler Angles and their kinematical equations. The attitude is approximately found analytically and a set of criteria is discussed which will minimize or eliminate the deviation of the nutation angle of a spacecraft rocket. Finally

  11. National Institute for Rocket Propulsion Systems 2012 Annual Report: A Year of Progress and Challenge

    NASA Technical Reports Server (NTRS)

    Thomas, L. Dale; Doreswamy, Rajiv; Fry, Emma Kiele

    2013-01-01

    The National Institute for Rocket Propulsion Systems (NIRPS) maintains and advances U.S. leadership in all aspects of rocket propulsion for defense, civil, and commercial uses. The Institute's creation is in response to widely acknowledged concerns about the U.S. rocket propulsion base dating back more than a decade. U.S. leadership in rocket and missile propulsion is threatened by long-term industry downsizing, a shortage of new solid and liquid propulsion programs, limited ability to attract and retain fresh talent, and discretionary federal budget pressures. Numerous trade and independent studies cite erosion of this capability as a threat to national security and the U.S. economy resulting in a loss of global competitiveness for the U.S. propulsion industry. This report covers the period between May 2011 and December 2012, which includes the creation and transition to operations of NIRPS. All subsequent reports will be annual. The year 2012 has been an eventful one for NIRPS. In its first full year, the new team overcame many obstacles and explored opportunities to ensure the institute has a firm foundation for the future. NIRPS is now an active organization making contributions to the development, sustainment, and strategy of the rocket propulsion industry in the United States. This report describes the actions taken by the NIRPS team to determine the strategy, organizational structure, and goals of the Institute. It also highlights key accomplishments, collaborations with other organizations, and the strategic framework for the Institute.

  12. NERVA-derived rocket module for solar system exploration

    SciTech Connect

    Zweig, H.R.; Cooper, M.H.

    1993-06-01

    A 50,000 pound thrust nuclear thermal rocket engine module concept based on Rover/NERVA technology is presented. Key engine features selected for reliability and safety have been integrated into this concept to provide 4.5 hours of full-thrust operation at a specific impulse of over 850 seconds. Those features include a single turbopump with an expander turbine, tank-head start in space, open-cycle decay heat removal with minimal loss of propellant, reactor cold end axial reflector, and tie tubes which are insulated end which are not used to provide turbine power. The tie tube configuration complements the open cycle decay heat removal concept. Retractable safety rods for water immersion subcriticality were considered and a design concept was developed. Other important safety issues were identified, and their method for accommodation in the design were considered for future implementation. 5 refs.

  13. Another Look at Rocket Thrust

    ERIC Educational Resources Information Center

    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…

  14. State Machine Modeling of the Space Launch System Solid Rocket Boosters

    NASA Technical Reports Server (NTRS)

    Harris, Joshua A.; Patterson-Hine, Ann

    2013-01-01

    The Space Launch System is a Shuttle-derived heavy-lift vehicle currently in development to serve as NASA's premiere launch vehicle for space exploration. The Space Launch System is a multistage rocket with two Solid Rocket Boosters and multiple payloads, including the Multi-Purpose Crew Vehicle. Planned Space Launch System destinations include near-Earth asteroids, the Moon, Mars, and Lagrange points. The Space Launch System is a complex system with many subsystems, requiring considerable systems engineering and integration. To this end, state machine analysis offers a method to support engineering and operational e orts, identify and avert undesirable or potentially hazardous system states, and evaluate system requirements. Finite State Machines model a system as a finite number of states, with transitions between states controlled by state-based and event-based logic. State machines are a useful tool for understanding complex system behaviors and evaluating "what-if" scenarios. This work contributes to a state machine model of the Space Launch System developed at NASA Ames Research Center. The Space Launch System Solid Rocket Booster avionics and ignition subsystems are modeled using MATLAB/Stateflow software. This model is integrated into a larger model of Space Launch System avionics used for verification and validation of Space Launch System operating procedures and design requirements. This includes testing both nominal and o -nominal system states and command sequences.

  15. Sounding rockets in Antarctica

    NASA Technical Reports Server (NTRS)

    Alford, G. C.; Cooper, G. W.; Peterson, N. E.

    1982-01-01

    Sounding rockets are versatile tools for scientists studying the atmospheric region which is located above balloon altitudes but below orbital satellite altitudes. Three NASA Nike-Tomahawk sounding rockets were launched from Siple Station in Antarctica in an upper atmosphere physics experiment in the austral summer of 1980-81. The 110 kg payloads were carried to 200 km apogee altitudes in a coordinated project with Arcas rocket payloads and instrumented balloons. This Siple Station Expedition demonstrated the feasibility of launching large, near 1,000 kg, rocket systems from research stations in Antarctica. The remoteness of research stations in Antarctica and the severe environment are major considerations in planning rocket launching expeditions.

  16. A urine volume measurement system

    NASA Technical Reports Server (NTRS)

    Poppendiek, H. F.; Mouritzen, G.; Sabin, C. M.

    1972-01-01

    An improved urine volume measurement system for use in the unusual environment of manned space flight is reported. The system utilizes a low time-constant thermal flowmeter. The time integral of the transient response of the flowmeter gives the urine volume during a void as it occurs. In addition, the two phase flows through the flowmeter present no problem. Developments of the thermal flowmeter and a verification of the predicted performance characteristics are summarized.

  17. Rocket engine system reliability analyses using probabilistic and fuzzy logic techniques

    NASA Technical Reports Server (NTRS)

    Hardy, Terry L.; Rapp, Douglas C.

    1994-01-01

    The reliability of rocket engine systems was analyzed by using probabilistic and fuzzy logic techniques. Fault trees were developed for integrated modular engine (IME) and discrete engine systems, and then were used with the two techniques to quantify reliability. The IRRAS (Integrated Reliability and Risk Analysis System) computer code, developed for the U.S. Nuclear Regulatory Commission, was used for the probabilistic analyses, and FUZZYFTA (Fuzzy Fault Tree Analysis), a code developed at NASA Lewis Research Center, was used for the fuzzy logic analyses. Although both techniques provided estimates of the reliability of the IME and discrete systems, probabilistic techniques emphasized uncertainty resulting from randomness in the system whereas fuzzy logic techniques emphasized uncertainty resulting from vagueness in the system. Because uncertainty can have both random and vague components, both techniques were found to be useful tools in the analysis of rocket engine system reliability.

  18. Dumbo: A pachydermal rocket motor

    NASA Technical Reports Server (NTRS)

    Kirk, Bill

    1991-01-01

    A brief historical account is given of the Dumbo nuclear reactor, a type of folded flow reactor that could be used for rocket propulsion. Much of the information is given in viewgraph form. Viewgraphs show details of the reactor system, fuel geometry, and key characteristics of the system (folded flow, use of fuel washers, large flow area, small fuel volume, hybrid modulator, and cermet fuel).

  19. Probabilistic risk assessment of the Space Shuttle. Phase 3: A study of the potential of losing the vehicle during nominal operation. Volume 4: System models and data analysis

    NASA Astrophysics Data System (ADS)

    Fragola, Joseph R.; Maggio, Gaspare; Frank, Michael V.; Gerez, Luis; McFadden, Richard H.; Collins, Erin P.; Ballesio, Jorge; Appignani, Peter L.; Karns, James J.

    1995-02-01

    In this volume, volume 4 (of five volumes), the discussion is focussed on the system models and related data references and has the following subsections: space shuttle main engine, integrated solid rocket booster, orbiter auxiliary power units/hydraulics, and electrical power system.

  20. Probabilistic risk assessment of the Space Shuttle. Phase 3: A study of the potential of losing the vehicle during nominal operation. Volume 4: System models and data analysis

    NASA Technical Reports Server (NTRS)

    Fragola, Joseph R.; Maggio, Gaspare; Frank, Michael V.; Gerez, Luis; Mcfadden, Richard H.; Collins, Erin P.; Ballesio, Jorge; Appignani, Peter L.; Karns, James J.

    1995-01-01

    In this volume, volume 4 (of five volumes), the discussion is focussed on the system models and related data references and has the following subsections: space shuttle main engine, integrated solid rocket booster, orbiter auxiliary power units/hydraulics, and electrical power system.

  1. Study of solid rocket motor for space, shuttle booster, volume 2, book 4 appendices B thru D

    NASA Technical Reports Server (NTRS)

    1972-01-01

    The mass properties and related data for the solid propellant rocket engine for use with the space shuttle are presented. Data for three solid propellant rocket engines are provided. The three designs considered are: (1) baseline parallel burn, (2) optional parallel burn, and (3) baseline series burn. Layouts of the respective designs to show design and dimensional data are included.

  2. Feasibility Investigation on the Development of a Structural Damage Diagnostic and Monitoring System for Rocket Engines

    NASA Technical Reports Server (NTRS)

    Shen, Ji Y.; Sharpe, Lonnie, Jr.

    1998-01-01

    The research activity for this project is mainly to investigate the necessity and feasibility to develop a structural health monitoring system for rocket engines, and to carry out a research plan for further development of the system. More than one hundred technical papers have been searched and reviewed during the period. We concluded after this investigation that adding a new module in NASA's existing automated diagnostic system to monitor the healthy condition of rocket engine structures is a crucial task, and it's possible to develop such a system based upon the vibrational-based nondestructive damage assessment techniques. A number of such techniques have been introduced. Their advantages and disadvantages are also discussed. A global research plan has been figured out. As the first step of the overall research plan, a proposal for the next fiscal year has been submitted.

  3. Evaluation and Characterization Study of Dual Pulse Laser-Induced Spark (DPLIS) For Rocket Engine Ignition System Application

    NASA Technical Reports Server (NTRS)

    Trinh, Huu P.; Early, Jim; Osborne, Robin

    2002-01-01

    This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Space Flight Center (MSFC). Initial hot-fire tests in a small-scale rocket chamber at MSFC have demonstrated the DPLIS concept having two main advantages over existing laser ignition concepts. First, the DPLIS can be potentially optimized its laser pulse format to maximize the initial plasma volume, the plasma lifetime, as well as the flame kernel growth rate. Characterization studies of the laser pulse format are now underway with experiments of igniting gaseous hydrogen/air in a Hencken burner. Once ignition is achieved, the flame is open to the atmosphere. This open environment allows easy access for diagnostics of the ignition phenomenon. The quick turn-around time of conducting experiments on this burner make it more amenable for conducting a large number of experiments for statistical analysis of the sensitivity of the test parameters. The results from these experiments will help optimize the laser format for future testing in an H2/O2 subscale rocket chamber.

  4. A demonstration of an intelligent control system for a reusable rocket engine

    NASA Technical Reports Server (NTRS)

    Musgrave, Jeffrey L.; Paxson, Daniel E.; Litt, Jonathan S.; Merrill, Walter C.

    1992-01-01

    An Intelligent Control System for reusable rocket engines is under development at NASA Lewis Research Center. The primary objective is to extend the useful life of a reusable rocket propulsion system while minimizing between flight maintenance and maximizing engine life and performance through improved control and monitoring algorithms and additional sensing and actuation. This paper describes current progress towards proof-of-concept of an Intelligent Control System for the Space Shuttle Main Engine. A subset of identifiable and accommodatable engine failure modes is selected for preliminary demonstration. Failure models are developed retaining only first order effects and included in a simplified nonlinear simulation of the rocket engine for analysis under closed loop control. The engine level coordinator acts as an interface between the diagnostic and control systems, and translates thrust and mixture ratio commands dictated by mission requirements, and engine status (health) into engine operational strategies carried out by a multivariable control. Control reconfiguration achieves fault tolerance if the nominal (healthy engine) control cannot. Each of the aforementioned functionalities is discussed in the context of an example to illustrate the operation of the system in the context of a representative failure. A graphical user interface allows the researcher to monitor the Intelligent Control System and engine performance under various failure modes selected for demonstration.

  5. Development of eddy current testing system for inspection of combustion chambers of liquid rocket engines.

    PubMed

    He, D F; Zhang, Y Z; Shiwa, M; Moriya, S

    2013-01-01

    An eddy current testing (ECT) system using a high sensitive anisotropic magnetoresistive (AMR) sensor was developed. In this system, a 20 turn circular coil with a diameter of 3 mm was used to produce the excitation field. A high sensitivity AMR sensor was used to measure the magnetic field produced by the induced eddy currents. A specimen made of copper alloy was prepared to simulate the combustion chamber of liquid rocket. Scanning was realized by rotating the chamber with a motor. To reduce the influence of liftoff variance during scanning, a dual frequency excitation method was used. The experimental results proved that ECT system with an AMR sensor could be used to check liquid rocket combustion chamber.

  6. A Real Time Differential GPS Tracking System for NASA Sounding Rockets

    NASA Technical Reports Server (NTRS)

    Bull, Barton; Bauer, Frank (Technical Monitor)

    2000-01-01

    Sounding rockets are suborbital launch vehicles capable of carrying scientific payloads to several hundred miles in altitude. These missions return a variety of scientific data including: chemical makeup and physical processes taking place in the atmosphere, natural radiation surrounding the Earth, data on the Sun, stars, galaxies and many other phenomena. In addition, sounding rockets provide a reasonably economical means of conducting engineering tests for instruments and devices to be used on satellites and other spacecraft prior to their use in these more expensive missions. Typically around thirty of these rockets are launched each year, from established ranges at Wallops Island, Virginia; Poker Flat Research Range, Alaska; White Sands Missile Range, New Mexico and from a number of ranges outside the United States. Many times launches are conducted from temporary launch ranges in remote parts of the world requiring considerable expense to transport and operate tracking radars. In order to support these missions, an inverse differential GPS system has been developed. The flight system consists of a small, inexpensive receiver, a preamplifier and a wrap-around antenna. A rugged, compact, portable ground station extracts GPS data from the raw payload telemetry stream, performs a real time differential solution and graphically displays the rocket's path relative to a predicted trajectory plot. In addition to generating a real time navigation solution, the system has been used for payload recovery, timing, data timetagging, precise tracking of multiple payloads and slaving of optical tracking systems for over the horizon acquisition. This paper discusses, in detail, the flight and ground hardware, as well as data processing and operational aspects of the system, and provides evidence of the system accuracy.

  7. Numerical simulations of the flowfields of industrial ventilation systems and solar rocket plume

    SciTech Connect

    Yu, Shengtao.

    1989-01-01

    The motivation for this research is to incorporate modern numerical methods in modeling the flowfields of two systems: (1) industrial ventilation systems and (2) solar rocket plume. For both systems, calculations of the velocity, temperature, turbulence properties, and species concentration of flowfields were performed. Brief discussions of the two topics follow: (1) Industrial ventilation systems. An open vessel equipped with a push-pull ventilation system to control toxic vapor and a flanged suction inlet to control grinding particles and welding fumes has been analyzed. The computational method involves solving the two-dimensional turbulent flow equations for the conservation of mass, momentum, energy, turbulence properties, and chemical species in finite form. The method provides information needed by engineers to assess the effectiveness of their designs. In order to verify the accuracy of the theoretical analysis, a two-dimensional push-pull system prototype was set up and color schlieren photography and hot wire anemometry were performed. Favorable agreement was found between the experimental data and calculated results. (2) Solar rocket plume. The interaction of the solar rocket plume and the solar concentrator is studied by flow-field analysis. Such interaction can adversely affect the collector performance through fouling, excessive heat, or pressure loading. The geometrical shape of the concentrator is such that only the flow from the nozzle boundary layer can reach it, but the thrust levels of interest lead to very thick boundary layers. A time-marching Parabolized Navier-Stokes (PNS) scheme is developed to calculate the flowfields inside nozzles. The Method of Characteristics (MOC) is used to simulate the flow of rocket plume. Results show that both pressure and heat transfer effects are low, but that they increase as the chamber pressure or the thrust level size is reduced.

  8. Rocket system for development testing of a retardation parachute for a supersonic store

    SciTech Connect

    Rollstin, L.R.

    1986-01-01

    A solid-propellant rocket booster system has been developed to support the development testing of a parachute system for the supersonic retardation of an 800-lb store. The parachute deployment flight condition requirements ranged from a dynamic pressure of 1800 psf to 4400 psf with a corresponding Mach number of 1.3 to 2.3. Also, this development testing was supported by the design and development of a small ''tractor'' (pulling type) rocket motor which affected the required rapid and symmetrical deployment of the parachute in the supersonic flight environment. A data reduction procedure was developed to combine payload accelerometer data with the optical or radar track to enhance the accuracy of the flight environment parameters during parachute deployment and the extreme deceleration phase.

  9. Test of a life support system with Hirudo medicinalis in a sounding rocket.

    PubMed

    Lotz, R G; Baum, P; Bowman, G H; Klein, K D; von Lohr, R; Schrotter, L

    1972-01-01

    Two Nike-Tomahawk rockets each carrying two Biosondes were launched from Wallops Island, Virginia, the first on 10 December 1970 and the second on 16 December 1970. The primary objective of both flights was to test the Biosonde life support system under a near weightless environment and secondarily to subject the Hirudo medicinalis to the combined stresses of a rocket flight. The duration of the weightless environment was approximately 6.5 minutes. Data obtained during the flight by telemetry was used to ascertain the operation of the system and the movements of the leeches during flight. Based on the information obtained, it has been concluded that the operation of the Biosondes during the flight was similar to that observed in the laboratory. The experiment and equipment are described briefly and the flight results presented.

  10. Analysis of Flow-System Starting Dynamics of Turbopump-Fed Liquid-Propellant Rocket

    NASA Technical Reports Server (NTRS)

    Krebs, Richard P.; Hart, Clint E.

    1959-01-01

    Two rocket configurations with turbopump drive were investigated analytically. In one configuration the inlet pressure to the turbine was fixed at the design value. The second configuration employed a "bootstrap" technique for supplying energy to the turbine. An injector was the chief resistance between the pump and the rocket combustion chamber. From the analysis two parameters were developed from which the speed response time of the turbopump, the flow response time, and the maximum dynamic line loss could be evaluated. These parameters were functions of turbopump moment of inertia, design performance of the turbine, and flow-system geometry. The moment of inertia of the turbopump and the ratio of turbine torque at zero speed to design torque had the most influence on the starting dynamics of the flow system. These parameters were also applicable to the bootstrap configuration as long as the inlet pressure to the turbine exceeded half the design value.

  11. Simulation Based on Ion Propulsion Rocket System with Using Negative ion - Negative Ion Pair Techniques

    NASA Astrophysics Data System (ADS)

    Sathiyavel, C.

    2016-07-01

    Ion propulsion rocket system is expected to become popular with the development of ion-ion pair techniques because of their stimulated of low propellant, Design of Thrust range is 1N with low electric power and high efficiency. A Negative ion-Negative ion pair of ion propulsion rocket system is proposed in this work .Negative Ion Based Rocket system consists of three parts 1.ionization chamber 2. Repulsion force and ion accelerator 3. Exhaust of Nozzle. The Negative ions from electro negatively gas are produced by attachment of the gas ,such as chlorine with electron emitted from a Electron gun ionization chamber. The formulate of large stable negative ion is achievable in chlorine gas with respect to electron affinity (∆E). The electron affinity is a measure of the energy change when an electron is added to a neutral atom to form a negative ion. When a neutral chlorine atom in the gaseous form picks up an electron to form a Cl- ion, it releases energy of 349 kJ/mol or 3.6 ev/atom. It is said to have an electron affinity of -349 kJ/mol ,the negative sign indicating that energy is released during this process .The mechanisms of attachment involve the formation of intermediate states. In that reason for , the highly repulsive force created between the same negative ions. The distance between same negative ions is important for the evaluate of the rocket thrust and is also determined by the exhaust velocity of the propellant. The mass flow rate of propellant is achieved by the ratio of total mass of the propellant (Kg) needed for operation to time period(s). Accelerate the Negative ions to a high velocity in the thrust vector direction with a significantly intense Magnetic field and the exhaust of negative ions through Nozzle. The simulation of the ion propulsion system has been carried out by MATLAB. By comparing the simulation results with the theoretical and previous results, we have found that the proposed method is achieved of thrust value with estimated

  12. Congreve Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

    The British fired Congreve rockets against the United States in the War of 1812. As a result Francis Scott Key coined the phrase the 'rocket's red glare.' Congreve had used a 16-foot guide stick to help stabilize his rocket. William Hale, another British inventor, invented the stickless rocket in 1846. The U.S. Army used the Hale rocket more than 100 years ago in the war with Mexico. Rockets were also used to a limited extent by both sides in the American Civil War.

  13. Preliminary design of a pressurization system for small bipropellant rocket engines

    NASA Astrophysics Data System (ADS)

    Stanley, Steven

    A study was conducted on the feasibility of developing a device or system that would improve the performance of small, bipropellant rockets through pressurization of the propellants. Due to the limitations in the space industry, namely high development costs and resistance to change, the new approach needed to be as simple and robust as possible. After reviewing several different potential methodologies, a concept was developed from first principles based on small gas turbine engine fuel injection approaches. The concept is simple and has heritage in the field of gas turbine engines, but it is new for the field of rocket propulsion. Using the basic physics of the proposed baseline concept, a simulation was developed to optimize the design parameters and to explore the trade space. Exercising the resulting simulation led to the identification of the critical design parameters and key performance metrics. During the iteration process, the design was updated and finalized. The resulting configuration appears to be feasible and has the potential of providing a new capability for small bipropellant rockets. Based upon the results of the study, recommendations were developed and a plan was created to further the development of the pump.

  14. Direct measurement of the impulse in a magnetic thrust chamber system for laser fusion rocket

    SciTech Connect

    Maeno, Akihiro; Yamamoto, Naoji; Nakashima, Hideki; Fujioka, Shinsuke; Johzaki, Tomoyuki; Mori, Yoshitaka; Sunahara, Atsushi

    2011-08-15

    An experiment is conducted to measure an impulse for demonstrating a magnetic thrust chamber system for laser fusion rocket. The impulse is produced by the interaction between plasma and magnetic field. In the experiment, the system consists of plasma and neodymium permanent magnets. The plasma is created by a single-beam laser aiming at a polystyrene spherical target. The impulse is 1.5 to 2.2 {mu}Ns by means of a pendulum thrust stand, when the laser energy is 0.7 J. Without magnetic field, the measured impulse is found to be zero. These results indicate that the system for generating impulse is working.

  15. System Engineering and Technical Challenges Overcome in the J-2X Rocket Engine Development Project

    NASA Technical Reports Server (NTRS)

    Ballard, Richard O.

    2012-01-01

    Beginning in 2006, NASA initiated the J-2X engine development effort to develop an upper stage propulsion system to enable the achievement of the primary objectives of the Constellation program (CxP): provide continued access to the International Space Station following the retirement of the Space Station and return humans to the moon. The J-2X system requirements identified to accomplish this were very challenging and the time expended over the five years following the beginning of the J- 2X effort have been noteworthy in the development of innovations in both the fields for liquid rocket propulsion and system engineering.

  16. System Modeling and Diagnostics for Liquefying-Fuel Hybrid Rockets

    NASA Technical Reports Server (NTRS)

    Poll, Scott; Iverson, David; Ou, Jeremy; Sanderfer, Dwight; Patterson-Hine, Ann

    2003-01-01

    A Hybrid Combustion Facility (HCF) was recently built at NASA Ames Research Center to study the combustion properties of a new fuel formulation that burns approximately three times faster than conventional hybrid fuels. Researchers at Ames working in the area of Integrated Vehicle Health Management recognized a good opportunity to apply IVHM techniques to a candidate technology for next generation launch systems. Five tools were selected to examine various IVHM techniques for the HCF. Three of the tools, TEAMS (Testability Engineering and Maintenance System), L2 (Livingstone2), and RODON, are model-based reasoning (or diagnostic) systems. Two other tools in this study, ICS (Interval Constraint Simulator) and IMS (Inductive Monitoring System) do not attempt to isolate the cause of the failure but may be used for fault detection. Models of varying scope and completeness were created, both qualitative and quantitative. In each of the models, the structure and behavior of the physical system are captured. In the qualitative models, the temporal aspects of the system behavior and the abstraction of sensor data are handled outside of the model and require the development of additional code. In the quantitative model, less extensive processing code is also necessary. Examples of fault diagnoses are given.

  17. Analysis and Results from a Flush Airdata Sensing System in Close Proximity to Firing Rocket Nozzles

    NASA Technical Reports Server (NTRS)

    Ali, Aliyah N.; Borrer, Jerry L.

    2013-01-01

    This paper presents information regarding the nosecap Flush Airdata Sensing (FADS) system on Orion’s Pad Abort 1 (PA-1) vehicle. The purpose of the nosecap FADS system was to test whether or not useful data could be obtained from a FADS system if it was placed in close proximity to firing rocket nozzles like the Attitude Control Motor (ACM) nozzles on the PA-1 Launch Abort System. The nosecap FADS system used pressure measurements from a series of pressure ports which were arranged in a cruciform pattern and flush with the surface of the vehicle to estimate values of angle of attack, angle of sideslip, Mach number, impact pressure, and freestream static pressure. This paper will present the algorithms employed by the FADS system along with the development of the calibration datasets and a comparison of the final results to the Best Estimated Trajectory (BET) data for PA-1. Also presented in this paper is a Computational Fluid Dynamics (CFD) study to explore the impact of the ACM on the nosecap FADS system. The comparison of the nosecap FADS system results to the BET and the CFD study showed that more investigation is needed to quantify the impact of the firing rocket motors on the FADS system.

  18. Nuclear Thermal Rocket - Arc Jet Integrated System Model

    NASA Technical Reports Server (NTRS)

    Taylor, Brian D.; Emrich, William

    2016-01-01

    In the post-shuttle era, space exploration is moving into a new regime. Commercial space flight is in development and is planned to take on much of the low earth orbit space flight missions. With the development of a heavy lift launch vehicle, the Space Launch, System, NASA has become focused on deep space exploration. Exploration into deep space has traditionally been done with robotic probes. More ambitious missions such as manned missions to asteroids and Mars will require significant technology development. Propulsion system performance is tied to the achievability of these missions and the requirements of other developing technologies that will be required. Nuclear thermal propulsion offers a significant improvement over chemical propulsion while still achieving high levels of thrust. Opportunities exist; however, to build upon what would be considered a standard nuclear thermal engine to attain improved performance, thus further enabling deep space missions. This paper discuss the modeling of a nuclear thermal system integrated with an arc jet to further augment performance. The performance predictions and systems impacts are discussed.

  19. Rocket Flight.

    ERIC Educational Resources Information Center

    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)

  20. Torpedo Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

    All through the 13th to the 15th Centuries there were reports of many rocket experiments. For example, Joanes de Fontana of Italy designed a surface-rurning, rocket-powered torpedo for setting enemy ships on fire

  1. General view of the Solid Rocket Booster's (SRB) Solid Rocket ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    General view of the Solid Rocket Booster's (SRB) Solid Rocket Motor Segments in the Surge Building of the Rotation Processing and Surge Facility at Kennedy Space Center awaiting transfer to the Vehicle Assembly Building and subsequent mounting and assembly on the Mobile Launch Platform. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  2. Liquid rocket actuators and operators. [in spacecraft control systems

    NASA Technical Reports Server (NTRS)

    1973-01-01

    All the types of actuators and associated operators used in booster, upper stage, and spacecraft propulsion and reaction-control systems except for chemical-explosive actuators and turbine actuators are discussed. Discussion of static and dynamic seals, mechanical transmission of motion, and instrumentation is included to the extent that actuator or operator design is affected. Selection of the optimum actuator configuration is discussed for specific application which require a tradeoff study that considers all the relevant factors: available energy sources, load capacity, stroke, speed of response, leakage limitations, environmental conditions, chemical compatibility, storage life and conditions, size, weight, and cost. These factors are interrelated with overall control-system design evaluations that are beyond the scope of this monograph; however, literature references are cited for a detailed review of the general considerations. Perinent advanced-state-of-the-art design concepts are surveyed briefly.

  3. Progress toward an advanced condition monitoring system for reusable rocket engines

    NASA Technical Reports Server (NTRS)

    Maram, J.; Barkhoudarian, S.

    1987-01-01

    A new generation of advanced sensor technologies will allow the direct measurement of critical/degradable rocket engine components' health and the detection of degraded conditions before component deterioration affects engine performance, leading to substantial improvements in reusable engines' operation and maintenance. When combined with a computer-based engine condition-monitoring system, these sensors can furnish a continuously updated data base for the prediction of engine availability and advanced warning of emergent maintenance requirements. Attention is given to the case of a practical turbopump and combustion device diagnostic/prognostic health-monitoring system.

  4. Technical report analysis and design: Study of solid rocket motors for a space shuttle booster, volume 2, book 1, supplement 1

    NASA Technical Reports Server (NTRS)

    1972-01-01

    An analysis and design effort was conducted as part of the study of solid rocket motor for a space shuttle booster. The 156-inch-diameter, parallel burn solid rocket motor was selected as its baseline because it is transportable and is the most cost-effective, reliable system that has been developed and demonstrated. The basic approach was to concentrate on the selected baseline design, and to draw from the baseline sufficient data to describe the alternate approaches also studied. The following conclusions were reached with respect to technical feasibility of the use of solid rocket booster motors for the space shuttle vehicle: (1) The 156-inch, parallel-burn baseline SRM design meets NASA's study requirements while incorporating conservative safety factors. (2) The solid rocket motor booster represents a cost-effective approach. (3) Baseline costs are conservative and are based on a demonstrated design. (4) Recovery and reuse are feasible and offer substantial cost savings. (5) Abort can be accomplished successfully. (6) Ecological effects are acceptable.

  5. Reusable rocket engine turbopump health monitoring system, part 3

    NASA Technical Reports Server (NTRS)

    Perry, John G.

    1989-01-01

    Degradation mechanisms and sensor identification/selection resulted in a list of degradation modes and a list of sensors that are utilized in the diagnosis of these degradation modes. The sensor list is divided into primary and secondary indicators of the corresponding degradation modes. The signal conditioning requirements are discussed, describing the methods of producing the Space Shuttle Main Engine (SSME) post-hot-fire test data to be utilized by the Health Monitoring System. Development of the diagnostic logic and algorithms is also presented. The knowledge engineering approach, as utilized, includes the knowledge acquisition effort, characterization of the expert's problem solving strategy, conceptually defining the form of the applicable knowledge base, and rule base, and identifying an appropriate inferencing mechanism for the problem domain. The resulting logic flow graphs detail the diagnosis/prognosis procedure as followed by the experts. The nature and content of required support data and databases is also presented. The distinction between deep and shallow types of knowledge is identified. Computer coding of the Health Monitoring System is shown to follow the logical inferencing of the logic flow graphs/algorithms.

  6. The Mars Exploration Rover (MER) Transverse Impulse Rocket System (TIRS)

    NASA Technical Reports Server (NTRS)

    SanMartin, Alejandro Miguel; Bailey, Erik

    2005-01-01

    In a very short period of time the MER project successfully developed and tested a system, TIRS/DIMES, to improve the probability of success in the presence of large Martian winds. The successful development of TIRS/DIMES played a big role in the landing site selection process by enabling the landing of Spirit on Gusev crater, a site of very high scientific interest but with known high wind conditions. The performance of TIRS by Spirit at Gusev Crater was excellent. The velocity prediction error was small and Big TIRS was fired reducing the impact horizontal velocity from approximately 23 meters per second to approximately 11 meters per second, well within the airbag capabilities. The performance of TIRS by Opportunity at Meridiani was good. The velocity prediction error was rather large (approximately 6 meters per second, a less than 2 sigma value, but TIRS did not fire which was the correct action.

  7. Overview of rocket engine control

    NASA Technical Reports Server (NTRS)

    Lorenzo, Carl F.; Musgrave, Jeffrey L.

    1991-01-01

    The issues of Chemical Rocket Engine Control are broadly covered. The basic feedback information and control variables used in expendable and reusable rocket engines, such as Space Shuttle Main Engine, are discussed. The deficiencies of current approaches are considered and a brief introduction to Intelligent Control Systems for rocket engines (and vehicles) is presented.

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

    NASA Technical Reports Server (NTRS)

    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.

  9. Rocket engine failure detection using system identification techiques

    NASA Technical Reports Server (NTRS)

    Meyer, Claudia M.; Zakrajsek, June F.

    1990-01-01

    The theoretical foundation and application of two univariate failure detection algorithms to Space Shuttle Main Engine (SSME) test firing data is presented. Both algorithms were applied to data collected during steady state operation of the engine. One algorithm, the time series algorithm, is based on time series techniques and involves the computation of autoregressive models. Times series techniques have been previously applied to SSME data. The second algorithm is based on standard signal processing techniques. It consists of tracking the variations in the average signal power with time. The average signal power algorithm is a newly proposed SSME failure detection algorithm. Seven nominal test firings were used to develop failure indication thresholds for each algorithm. These thresholds were tested using four anomalous firings and one additional nominal firing. Both algorithms provided significantly earlier failure indication times than did the current redline limit system. Neither algorithm gave false failure indications for the nominal firing. The strengths and weaknesses of the two algorithms are discussed and compared. The average signal algorithm was found to have several advantages over the time series algorithm.

  10. Aluminum-fueled rockets for the space transportation system

    NASA Technical Reports Server (NTRS)

    Cutler, Andrew Hall

    1992-01-01

    Aluminum-fueled engines, used to propel orbital transfer vehicles (OTV's), offer benefits to the Space Transportation System (STS) if scrap aluminum can be scavenged at a reasonable cost. Aluminum scavenged from Space Shuttle external tanks could replace propellants hauled from Earth, thus allowing more payloads to be sent to their final destinations at the same Shuttle launch rate. To allow OTV use of aluminum fuel, two new items would be required: a facility to reprocess aluminum from external tanks and an engine for the OTV which could burn aluminum. Design of the orbital transfer vehicle would have to differ substantially from current concepts for it to carry and use the aluminum fuel. The aluminum reprocessing facility would probably have a mass of under 15 metric tons and would probably cost less that $200,000,000. Development of an aluminum-burning engine would no doubt be extremely expensive (1 to 2 billion dollars), but this amount would be adequately repaid by increased STS throughput. Engine production cost is difficult to estimate, but even an extremely high cost (e.g., $250,000,000 per engine) would not significantly increase orbit-raising expenses.

  11. Recession Curve Generation for the Space Shuttle Solid Rocket Booster Thermal Protection System Coatings

    NASA Technical Reports Server (NTRS)

    Kanner, Howard S.; Stuckey, C. Irvin; Davis, Darrell W.; Davis, Darrell (Technical Monitor)

    2002-01-01

    Ablatable Thermal Protection System (TPS) coatings are used on the Space Shuttle Vehicle Solid Rocket Boosters in order to protect the aluminum structure from experiencing excessive temperatures. The methodology used to characterize the recession of such materials is outlined. Details of the tests, including the facility, test articles and test article processing are also presented. The recession rates are collapsed into an empirical power-law relation. A design curve is defined using a 95-percentile student-t distribution. based on the nominal results. Actual test results are presented for the current acreage TPS material used.

  12. SRB-3D Solid Rocket Booster performance prediction program. Volume 1: Engineering description/users information manual

    NASA Technical Reports Server (NTRS)

    Winkler, J. C.

    1976-01-01

    The modified Solid Rocket Booster Performance Evaluation Model (SRB-3D) was developed as an extension to the internal ballistics module of the SRB-2 performance program. This manual contains the engineering description of SRB-3D which describes the approach used to develop the 3D concept and an explanation of the modifications which were necessary to implement these concepts.

  13. Study of solid rocket motors for a space shuttle booster. Volume 2, book 3: Cost estimating data

    NASA Technical Reports Server (NTRS)

    Vanderesch, A. H.

    1972-01-01

    Cost estimating data for the 156 inch diameter, parallel burn solid rocket propellant engine selected for the space shuttle booster are presented. The costing aspects on the baseline motor are initially considered. From the baseline, sufficient data is obtained to provide cost estimates of alternate approaches.

  14. Contact diagnostics of combustion products of rocket engines, their units, and systems

    NASA Astrophysics Data System (ADS)

    Ivanov, N. N.; Ivanov, A. N.

    2013-12-01

    This article is devoted to a new block-module device used in the diagnostics of condensed combustion products of rocket engines during research and development with liquid-propellant rocket engines (Glushko NPO Energomash; engines RD-171, RD-180, and RD-191) and solid-propellant rocket motors. Soot samplings from the supersonic high-temperature jet of a high-power liquid-propellant rocket engine were taken by the given device for the first time in practice for closed-exhaust lines. A large quantity of significant results was also obtained during a combustion investigation of solid propellants within solid-propellant rocket motors.

  15. Solid rocket motor conceptual design - The development of a design optimization expert system with a hypertext user interface

    NASA Astrophysics Data System (ADS)

    Clegern, James B.

    1993-06-01

    Solid rocket motor (SRM) design prototypes can be rapidly formulated and evaluated by the use of advanced computer-based methodologies that apply expert system and artificial intelligence software to the SRM design optimization processes. The research program that was carried out, and is reported in this paper, was to formulate a computer-based SRM expert system for motor design and optimization, with the assistance of a hypertext software algorithm that provides a user-friendly interface. With this interface for parameter input, the design engineer can quickly obtain rocket motor designs that satisfy the performance mission of the SRM, as well as meet criteria for optimized (minimum) motor mass. The computer-based software has been designated as the Solid Rocket Motor Conceptual Design Optimization System (SRMCDOS). The main purpose of this SRM design system is to aid the SRM design engineer in making the best initial design selections and thereby reducing the overall 'design cycle time' of a project.

  16. The rationale/benefits of nuclear thermal rocket propulsion for NASA's lunar space transportation system

    NASA Astrophysics Data System (ADS)

    Borowski, Stanley K.

    1994-09-01

    The solid core nuclear thermal rocket (NTR) represents the next major evolutionary step in propulsion technology. With its attractive operating characteristics, which include high specific impulse (approximately 850-1000 s) and engine thrust-to-weight (approximately 4-20), the NTR can form the basis for an efficient lunar space transportation system (LTS) capable of supporting both piloted and cargo missions. Studies conducted at the NASA Lewis Research Center indicate that an NTR-based LTS could transport a fully-fueled, cargo-laden, lunar excursion vehicle to the Moon, and return it to low Earth orbit (LEO) after mission completion, for less initial mass in LEO than an aerobraked chemical system of the type studied by NASA during its '90-Day Study.' The all-propulsive NTR-powered LTS would also be 'fully reusable' and would have a 'return payload' mass fraction of approximately 23 percent--twice that of the 'partially reusable' aerobraked chemical system. Two NTR technology options are examined--one derived from the graphite-moderated reactor concept developed by NASA and the AEC under the Rover/NERVA (Nuclear Engine for Rocket Vehicle Application) programs, and a second concept, the Particle Bed Reactor (PBR). The paper also summarizes NASA's lunar outpost scenario, compares relative performance provided by different LTS concepts, and discusses important operational issues (e.g., reusability, engine 'end-of life' disposal, etc.) associated with using this important propulsion technology.

  17. Dynamic certification of a thrust-measuring system for large solid rocket motors

    NASA Technical Reports Server (NTRS)

    Lemaster, R. A.; Runyan, R. B.

    1983-01-01

    The J-4 Rocket Test Cell and its Thrust Measuring System (TMS) at Arnold Engineering Development Center were modified to provide multicomponent force measurement of large solid rocket motors having nozzle gimbaling capability. To verify the structural integrity of a combined TMS and motor system, a large finite element model of the TMS and motor was developed using the NASTRAN computer program. Due to the importance of obtaining accurate estimates for the dynamic force levels, it was necessary to certify that the model adequately simulated the physical system. This was accomplished by performing a modal analysis test on the TMS and motor combination. The objectives were to discuss the physical characteristics of the TMS and motor that influence the NASTRAN model; to compare the frequency response characteristics computed using the NASTRAN model to those obtained from the modal analysis test; to discuss the experiences gained in modal analysis testing; and to demonstrate how state-of-the-art experimental and analytical methods are used in the design of ground test facilities.

  18. The Rationale/Benefits of Nuclear Thermal Rocket Propulsion for NASA's Lunar Space Transportation System

    NASA Technical Reports Server (NTRS)

    Borowski, Stanley K.

    1994-01-01

    The solid core nuclear thermal rocket (NTR) represents the next major evolutionary step in propulsion technology. With its attractive operating characteristics, which include high specific impulse (approximately 850-1000 s) and engine thrust-to-weight (approximately 4-20), the NTR can form the basis for an efficient lunar space transportation system (LTS) capable of supporting both piloted and cargo missions. Studies conducted at the NASA Lewis Research Center indicate that an NTR-based LTS could transport a fully-fueled, cargo-laden, lunar excursion vehicle to the Moon, and return it to low Earth orbit (LEO) after mission completion, for less initial mass in LEO than an aerobraked chemical system of the type studied by NASA during its '90-Day Study.' The all-propulsive NTR-powered LTS would also be 'fully reusable' and would have a 'return payload' mass fraction of approximately 23 percent--twice that of the 'partially reusable' aerobraked chemical system. Two NTR technology options are examined--one derived from the graphite-moderated reactor concept developed by NASA and the AEC under the Rover/NERVA (Nuclear Engine for Rocket Vehicle Application) programs, and a second concept, the Particle Bed Reactor (PBR). The paper also summarizes NASA's lunar outpost scenario, compares relative performance provided by different LTS concepts, and discusses important operational issues (e.g., reusability, engine 'end-of life' disposal, etc.) associated with using this important propulsion technology.

  19. Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Appendix B: Liquid rocket booster acoustic and thermal environments

    NASA Technical Reports Server (NTRS)

    1989-01-01

    The ascent thermal environment and propulsion acoustic sources for the Martin-Marietta Corporation designed Liquid Rocket Boosters (LRB) to be used with the Space Shuttle Orbiter and External Tank are described. Two designs were proposed: one using a pump-fed propulsion system and the other using a pressure-fed propulsion system. Both designs use LOX/RP-1 propellants, but differences in performance of the two propulsion systems produce significant differences in the proposed stage geometries, exhaust plumes, and resulting environments. The general characteristics of the two designs which are significant for environmental predictions are described. The methods of analysis and predictions for environments in acoustics, aerodynamic heating, and base heating (from exhaust plume effects) are also described. The acoustic section will compare the proposed exhaust plumes with the current SRB from the standpoint of acoustics and ignition overpressure. The sections on thermal environments will provide details of the LRB heating rates and indications of possible changes in the Orbiter and ET environments as a result of the change from SRBs to LRBs.

  20. Rediscovering the potential of solid rocket propulsion systems for low cost launch vehicle and upper stage applications

    NASA Astrophysics Data System (ADS)

    Chew, James S. B.

    1992-08-01

    Solid propulsion component technologies for future space systems must not only meet current upper stage and launch vehicle performance requirements, but must have significantly lower system life cycle costs. Innovative system concepts and component technologies must be developed to meet these goals. An in-house Phillips Laboratory, Propulsion Directorate program is working toward this goal. This in-house program is investigating and developing innovative technologies for all the major components in a solid rocket motor. In addition, work is being conducted to better understand the design, development and fabrication processes for conventional solid rocket motor components. Life cycle cost reduction technologies for these components are also being investigated.

  1. Development and Testing of a Methane/Oxygen Catalytic Microtube Ignition System for Rocket Propulsion

    NASA Technical Reports Server (NTRS)

    Deans, Matthew

    2012-01-01

    This study sought to develop a catalytic ignition advanced torch system with a unique catalyst microtube design that could serve as a low energy alternative or redundant system for the ignition of methane and oxygen rockets. Development and testing of iterations of hardware was carried out to create a system that could operate at altitude and produce a torch. A unique design was created that initiated ignition via the catalyst and then propagated into external staged ignition. This system was able to meet the goals of operating across a range of atmospheric and altitude conditions with power inputs on the order of 20 to 30 watts with chamber pressures and mass flow rates typical of comparable ignition systems for a 100 lbf engine.

  2. Development and Testing of a Methane/Oxygen Catalytic Microtube Ignition System for Rocket Propulsion

    NASA Technical Reports Server (NTRS)

    Deans, Matthew C.; Schneider, Steven J.

    2012-01-01

    This study sought to develop a catalytic ignition advanced torch system with a unique catalyst microtube design that could serve as a low energy alternative or redundant system for the ignition of methane and oxygen rockets. Development and testing of iterations of hardware was carried out to create a system that could operate at altitude and produce a torch. A unique design was created that initiated ignition via the catalyst and then propagated into external staged ignition. This system was able to meet the goals of operating across a range of atmospheric and altitude conditions with power inputs on the order of 20 to 30 watts with chamber pressures and mass flow rates typical of comparable ignition systems for a 100 Ibf engine.

  3. Combustion Instability in an Acid-Heptane Rocket with a Pressurized-Gas Propellant Pumping System

    NASA Technical Reports Server (NTRS)

    Tischler, Adelbert O.; Bellman, Donald R.

    1951-01-01

    Results of experimental measurements of low-frequency combustion instability of a 300-pound thrust acid-heptane rocket engine were compared to the trends predicted by an analysis of combustion instability in a rocket engine with a pressurized-gas propellant pumping system. The simplified analysis, which assumes a monopropellant model, was based on the concept of a combustion the delay occurring from the moment of propellant injection to the moment of propellant combustion. This combustion time delay was experimentally measured; the experimental values were of approximately half the magnitude predicted by the analysis. The pressure-fluctuation frequency for a rocket engine with a characteristic length of 100 inches and operated at a combustion-chamber pressure of 280 pounds per square inch absolute was 38 cycles per second; the analysis indicated. a frequency of 37 cycles per second. Increasing combustion-chamber characteristic length decreased the pressure-fluctuation frequency, in conformity to the analysis. Increasing the chamber operating pressure or increasing the injector pressure drop increased the frequency. These latter two effects are contrary to the analysis; the discrepancies are attributed to the conflict between the assumptions made to simplify the analysis and the experimental conditions. Oxidant-fuel ratio had no apparent effect on the experimentally measured pressure-fluctuation frequency for acid-heptane ratios from 3.0 to 7.0. The frequencies decreased with increased amplitude of the combustion-chamber pressure variations. The analysis indicated that if the combustion time delay were sufficiently short, low-frequency combustion instability would be eliminated.

  4. An Object-Oriented Graphical User Interface for a Reusable Rocket Engine Intelligent Control System

    NASA Technical Reports Server (NTRS)

    Litt, Jonathan S.; Musgrave, Jeffrey L.; Guo, Ten-Huei; Paxson, Daniel E.; Wong, Edmond; Saus, Joseph R.; Merrill, Walter C.

    1994-01-01

    An intelligent control system for reusable rocket engines under development at NASA Lewis Research Center requires a graphical user interface to allow observation of the closed-loop system in operation. The simulation testbed consists of a real-time engine simulation computer, a controls computer, and several auxiliary computers for diagnostics and coordination. The system is set up so that the simulation computer could be replaced by the real engine and the change would be transparent to the control system. Because of the hard real-time requirement of the control computer, putting a graphical user interface on it was not an option. Thus, a separate computer used strictly for the graphical user interface was warranted. An object-oriented LISP-based graphical user interface has been developed on a Texas Instruments Explorer 2+ to indicate the condition of the engine to the observer through plots, animation, interactive graphics, and text.

  5. An object-oriented graphical user interface for a reusable rocket engine intelligent control system

    NASA Astrophysics Data System (ADS)

    Litt, Jonathan S.; Musgrave, Jeffrey L.; Guo, Ten-Huei; Paxson, Daniel E.; Wong, Edmond; Saus, Joseph R.; Merrill, Walter C.

    1994-10-01

    An intelligent control system for reusable rocket engines under development at NASA Lewis Research Center requires a graphical user interface to allow observation of the closed-loop system in operation. The simulation testbed consists of a real-time engine simulation computer, a controls computer, and several auxiliary computers for diagnostics and coordination. The system is set up so that the simulation computer could be replaced by the real engine and the change would be transparent to the control system. Because of the hard real-time requirement of the control computer, putting a graphical user interface on it was not an option. Thus, a separate computer used strictly for the graphical user interface was warranted. An object-oriented LISP-based graphical user interface has been developed on a Texas Instruments Explorer 2+ to indicate the condition of the engine to the observer through plots, animation, interactive graphics, and text.

  6. Uranium droplet core nuclear rocket

    NASA Technical Reports Server (NTRS)

    Anghaie, Samim

    1991-01-01

    Uranium droplet nuclear rocket is conceptually designed to utilize the broad temperature range ofthe liquid phase of metallic uranium in droplet configuration which maximizes the energy transfer area per unit fuel volume. In a baseline system dissociated hydrogen at 100 bar is heated to 6000 K, providing 2000 second of Isp. Fission fragments and intense radian field enhance the dissociation of molecular hydrogen beyond the equilibrium thermodynamic level. Uranium droplets in the core are confined and separated by an axisymmetric vortex flow generated by high velocity tangential injection of hydrogen in the mid-core regions. Droplet uranium flow to the core is controlled and adjusted by a twin flow nozzle injection system.

  7. Study of solid rocket motors for a space shuttle booster. Volume 2, book 1: Analysis and design

    NASA Technical Reports Server (NTRS)

    1972-01-01

    An analysis of the factors which determined the selection of the solid rocket propellant engines for the space shuttle booster is presented. The 156 inch diameter, parallel burn engine was selected because of its transportability, cost effectiveness, and reliability. Other factors which caused favorable consideration are: (1) recovery and reuse are feasible and offer substantial cost savings, (2) abort can be easily accomplished. and (3) ecological effects are acceptable.

  8. Convective Heat Transfer in the Reusable Solid Rocket Motor of the Space Transportation System

    NASA Technical Reports Server (NTRS)

    Ahmad, Rashid A.; Cash, Stephen F. (Technical Monitor)

    2002-01-01

    This simulation involved a two-dimensional axisymmetric model of a full motor initial grain of the Reusable Solid Rocket Motor (RSRM) of the Space Transportation System (STS). It was conducted with CFD (computational fluid dynamics) commercial code FLUENT. This analysis was performed to: a) maintain continuity with most related previous analyses, b) serve as a non-vectored baseline for any three-dimensional vectored nozzles, c) provide a relatively simple application and checkout for various CFD solution schemes, grid sensitivity studies, turbulence modeling and heat transfer, and d) calculate nozzle convective heat transfer coefficients. The accuracy of the present results and the selection of the numerical schemes and turbulence models were based on matching the rocket ballistic predictions of mass flow rate, head end pressure, vacuum thrust and specific impulse, and measured chamber pressure drop. Matching these ballistic predictions was found to be good. This study was limited to convective heat transfer and the results compared favorably with existing theory. On the other hand, qualitative comparison with backed-out data of the ratio of the convective heat transfer coefficient to the specific heat at constant pressure was made in a relative manner. This backed-out data was devised to match nozzle erosion that was a result of heat transfer (convective, radiative and conductive), chemical (transpirating), and mechanical (shear and particle impingement forces) effects combined.

  9. Use of System Safety Risk Assessments for the Space Shuttle Reusable Solid Rocket Motor (RSRM)

    NASA Technical Reports Server (NTRS)

    Greenhalgh, Phillip O.; McCool, Alex (Technical Monitor)

    2001-01-01

    This paper discusses the System Safety approach used to assess risk for the Space Shuttle Reusable Solid Rocket Motor (RSRM). Previous to the first RSRM flight in the fall of 1988, all systems were analyzed extensively to assure that hazards were identified, assessed and that the baseline risk was understood and appropriately communicated. Since the original RSRM baseline was established, Thiokol and NASA have implemented a number of initiatives that have further improved the RSRM. The robust design, completion of rigorous testing and flight success of the RSRM has resulted in a wise reluctance to make changes. One of the primary assessments required to accompany the documentation of each proposed change and aid in the decision making process is a risk assessment. Documentation supporting proposed changes, including the risk assessments from System Safety, are reviewed and assessed by Thiokol and NASA technical management. After thorough consideration, approved changes are implemented adding improvements to and reducing risk of the Space Shuttle RSRM.

  10. Zero Boil-Off System Design and Thermal Analysis of the Bimodal Thermal Nuclear Rocket

    SciTech Connect

    Christie, Robert J.; Plachta, David W.

    2006-01-20

    Mars exploration studies at NASA are evaluating vehicles that incorporate Bimodal Nuclear Thermal Rocket (BNTR) propulsion which use a high temperature nuclear fission reactor and hydrogen to produce thermal propulsion. The hydrogen propellant is to be stored in liquid state for periods up to 18 months. To prevent boil-off of the liquid hydrogen, a system of passive and active components are needed to prevent heat from entering the tanks and to remove any heat that does. This report describes the design of the system components used for the BNTR Crew Transfer Vehicle and the thermal analysis performed. The results show that Zero Boil-Off (ZBO) can be achieved with the electrical power allocated for the ZBO system.

  11. GPS Sounding Rocket Developments

    NASA Technical Reports Server (NTRS)

    Bull, Barton

    1999-01-01

    Sounding rockets are suborbital launch vehicles capable of carrying scientific payloads several hundred miles in altitude. These missions return a variety of scientific data including; chemical makeup and physical processes taking place in the atmosphere, natural radiation surrounding the Earth, data on the Sun, stars, galaxies and many other phenomena. In addition, sounding rockets provide a reasonably economical means of conducting engineering tests for instruments and devices used on satellites and other spacecraft prior to their use in more expensive activities. This paper addresses the NASA Wallops Island history of GPS Sounding Rocket experience since 1994 and the development of highly accurate and useful system.

  12. Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Appendix A: Stress analysis report for the pump-fed and pressure-fed liquid rocket booster

    NASA Technical Reports Server (NTRS)

    1989-01-01

    Pressure effects on the pump-fed Liquid Rocket Booster (LRB) of the Space Transportation System are examined. Results from the buckling tests; bending moments tests; barrel, propellant tanks, frame XB1513, nose cone, and intertank tests; and finite element examination of forward and aft skirts are presented.

  13. Rocket University at KSC

    NASA Technical Reports Server (NTRS)

    Sullivan, Steven J.

    2014-01-01

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

  14. Method for providing real-time control of a gaseous propellant rocket propulsion system

    NASA Technical Reports Server (NTRS)

    Morris, Brian G. (Inventor)

    1991-01-01

    The new and improved methods and apparatus disclosed provide effective real-time management of a spacecraft rocket engine powered by gaseous propellants. Real-time measurements representative of the engine performance are compared with predetermined standards to selectively control the supply of propellants to the engine for optimizing its performance as well as efficiently managing the consumption of propellants. A priority system is provided for achieving effective real-time management of the propulsion system by first regulating the propellants to keep the engine operating at an efficient level and thereafter regulating the consumption ratio of the propellants. A lower priority level is provided to balance the consumption of the propellants so significant quantities of unexpended propellants will not be left over at the end of the scheduled mission of the engine.

  15. A study of the durability of beryllium rocket engines. [space shuttle reaction control system

    NASA Technical Reports Server (NTRS)

    Paster, R. D.; French, G. C.

    1974-01-01

    An experimental test program was performed to demonstrate the durability of a beryllium INTEREGEN rocket engine when operating under conditions simulating the space shuttle reaction control system. A vibration simulator was exposed to the equivalent of 100 missions of X, Y, and Z axes random vibration to demonstrate the integrity of the recently developed injector-to-chamber braze joint. An off-limits engine was hot fired under extreme conditions of mixture ratio, chamber pressure, and orifice plugging. A durability engine was exposed to six environmental cycles interspersed with hot-fire tests without intermediate cleaning, service, or maintenance. Results from this program indicate the ability of the beryllium INTEREGEN engine concept to meet the operational requirements of the space shuttle reaction control system.

  16. Design study of RL10 derivatives. Volume 2: Engine design characteristics. [application of rocket engine to space tug propulsion

    NASA Technical Reports Server (NTRS)

    Adams, A.

    1973-01-01

    The design characteristics of the RL-10 rocket engine are discussed. The results from critical elements evaluation, baseline engine design, parametric and special study tasks are presented. Critical element evaluation established the feasibility of various engine features such as tank head idle, pumped idle, autogenous tank pressurization, and two-phase pumping. Three baseline engines, derived from the RL-10 were conceptually designed. Parametric life and performance data were generated. Special studies were conducted to establish the impact on the engine design of environment, safety, interchangeability, and maintenance.

  17. Rotating and positive-displacement pumps for low-thrust rocket engines. Volume 2: Fabrication and testing

    NASA Technical Reports Server (NTRS)

    Csomor, A.

    1974-01-01

    Rotating and positive displacement pumps of various types were studied for pumping liquid fluorine for low thrust high performance rocket engines. Included in the analysis were: centrifugal, pitot, Barske, Tesla, drag, gear, vane, axial piston, radial piston, diaphragm and helirotor pump concepts. The centrifugal and gear pumps were carried through detail design and fabrication. After preliminary testing in Freon 12, the centrifugal pump was selected for further testing and development. It was tested in Freon 12 to obtain the hydrodynamic performance. Tests were also conducted in liquid fluorine to demonstrate chemical compatibility.

  18. Space transfer concepts and analysis for exploration missions. Implementation plan and element description document (draft final). Volume 3: Nuclear thermal rocket vehicle

    NASA Technical Reports Server (NTRS)

    1991-01-01

    This document presents the nuclear thermal rocket (NTR) concept design developed in support of the Space Transfer Concepts and Analysis for Exploration Missions (STCAEM) study. The evolution of the NTR concept is described along with the requirements, guidelines and assumptions for the design. Operating modes and options are defined and a systems description of the vehicle is presented. Artificial gravity configuration options and space and ground support systems are discussed. Finally, an implementation plan is presented which addresses technology needs, schedules, facilities and costs.

  19. KINETIC: A system code for analyzing Nuclear thermal propulsion rocket engine transients

    NASA Astrophysics Data System (ADS)

    Schmidt, E.; Lazareth, O.; Ludewig, H.

    1993-07-01

    A system code suitable for analyzing Nuclear Thermal Propulsion (NTP) rocket engines is described in this paper. The code consists of a point reactor model and nodes to describe the fluid dynamics and heat transfer mechanism. Feedback from the fuel coolant, moderator and reflector are allowed for, and the control of the reactor is by motion of control elements (drums or rods). The worth of the control clement and feedback coefficients are predetermined. Separate models for the turbo-pump assembly (TPA) and nozzle are also included. The model to be described in this paper is specific for the Particle Bed Reactor (PBR). An illustrative problem is solved. This problem consists of a PBR operating in a blowdown mode.

  20. KINETIC -- a system code for analyzing Nuclear Thermal Propulsion rocket engine transients

    SciTech Connect

    Schmidt, E.; Lazareth, O.; Ludewig, H.

    1993-07-01

    A system code suitable for analyzing Nuclear Thermal Propulsion (NTP) rocket engines is described in this paper. The code consists of a point reactor model and nodes to describe the fluid dynamics and heat transfer mechanism. Feedback from the fuel coolant, moderator and reflector are allowed for, and the control of the reactor is by motion of control elements (drums or rods). The worth of the control clement and feedback coefficients are predetermined. Separate models for the turbo-pump assembly (TPA) and nozzle are also included. The model to be described in this paper is specific for the Particle Bed Reactor (PBR). An illustrative problem is solved. This problem consists of a PBR operating in a blowdown mode.

  1. Kinetic---a system code for analyzing nuclear thermal propulsion rocket engine transients

    SciTech Connect

    Schmidt, E.; Lazareth, O.; Ludewig, H. )

    1993-01-20

    A system code suitable for analyzing Nuclear Thermal Propulsion (NTP) rocket engines is described in this paper. The code consists of a point reactor model and nodes to describe the fluid dynamics and heat transfer mechanism. Feedback from the fuel, coolant, moderator and reflector are allowed for, and the control of the reactor is by motion of controls element (drums or rods). The worth of the control element and feedback coefficients are predetermined. Separate models for the turbo-pump assembly (TPA) and nozzle are also included. The model to be described in this paper is specific for the Particle Bed Reactor (PBR). An illustrative problem is solved. This problem consists of a PBR operating in a blowdown mode.

  2. Status on Technology Development of Optic Fiber-Coupled Laser Ignition System for Rocket Engine Applications

    NASA Technical Reports Server (NTRS)

    Trinh, Huu P.; Early, Jim; Osborne, Robin; Thomas, Matthew; Bossard, John

    2003-01-01

    To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for liquid rocket engine applications. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concept: not only offer system simplicity, but also enhance the combustion performance. Test results have shown that chamber performance is markedly high even at a low chamber length-to-diameter ratio. This incentive can be translated to a convenience in the thrust chamber packaging.

  3. Effects of natural environment on first generation solid rocket booster thermal protection system materials

    NASA Technical Reports Server (NTRS)

    Webb, D. D.

    1988-01-01

    The effort to demonstrate, by real-time exposure, the effects of the natural environment at Kennedy Space Center, Florida, upon the Thermal Protection System (TPS) of the Solid Rocket Booster (SRB) is summarized, and that the overall SRB TPS configuration is verified to meet all requirements for resistance to the conditions associated with outdoor weathering, including: solar radiation; temperature; humidity; precipitation; wind; sand/dust abrasion; static electricity; salt spray; fungus; and atmospheric oxidants. The evaluation criterion for this project was based upon flatwise tensile properties, visual inspection, color change, and thermal performance. Based upon the evaluation of the changes in these properties, it is concluded that properly applied and topcoat-protected TPS can satisfactorily withstand the conditions of the natural environment at KSC for exposures up to six months.

  4. User's manual for rocket combustor interactive design (ROCCID) and analysis computer program. Volume 2: Appendixes A-K

    NASA Technical Reports Server (NTRS)

    Muss, J. A.; Nguyen, T. V.; Johnson, C. W.

    1991-01-01

    The appendices A-K to the user's manual for the rocket combustor interactive design (ROCCID) computer program are presented. This includes installation instructions, flow charts, subroutine model documentation, and sample output files. The ROCCID program, written in Fortran 77, provides a standardized methodology using state of the art codes and procedures for the analysis of a liquid rocket engine combustor's steady state combustion performance and combustion stability. The ROCCID is currently capable of analyzing mixed element injector patterns containing impinging like doublet or unlike triplet, showerhead, shear coaxial and swirl coaxial elements as long as only one element type exists in each injector core, baffle, or barrier zone. Real propellant properties of oxygen, hydrogen, methane, propane, and RP-1 are included in ROCCID. The properties of other propellants can be easily added. The analysis models in ROCCID can account for the influences of acoustic cavities, helmholtz resonators, and radial thrust chamber baffles on combustion stability. ROCCID also contains the logic to interactively create a combustor design which meets input performance and stability goals. A preliminary design results from the application of historical correlations to the input design requirements. The steady state performance and combustion stability of this design is evaluated using the analysis models, and ROCCID guides the user as to the design changes required to satisfy the user's performance and stability goals, including the design of stability aids. Output from ROCCID includes a formatted input file for the standardized JANNAF engine performance prediction procedure.

  5. Evaluation and Characterization Study of Dual Pulse Laser-Induced Spark (DPLIS) for Rocket Engine Ignition System Application

    NASA Technical Reports Server (NTRS)

    Osborne, Robin; Wehrmeyer, Joseph; Trinh, Huu; Early, James

    2003-01-01

    This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Space Flight Center (MSFC). Laser ignition has been used at MSFC in recent test series to successfully ignite RP1/GOX propellants in a subscale rocket chamber, and other past studies by NASA GRC have demonstrated the use of laser ignition for rocket engines. Despite the progress made in the study of this ignition method, the logistics of depositing laser sparks inside a rocket chamber have prohibited its use. However, recent advances in laser designs, the use of fiber optics, and studies of multi-pulse laser formats3 have renewed the interest of rocket designers in this state-of the-art technology which offers the potential elimination of torch igniter systems and their associated mechanical parts, as well as toxic hypergolic ignition systems. In support of this interest to develop an alternative ignition system that meets the risk-reduction demands of Next Generation Launch Technology (NGLT), characterization studies of a dual pulse laser format for laser-induced spark ignition are underway at MSFC. Results obtained at MSFC indicate that a dual pulse format can produce plasmas that absorb the laser energy as efficiently as a single pulse format, yet provide a longer plasma lifetime. In an experiments with lean H2/air propellants, the dual pulse laser format, containing the same total energy of a single laser pulse, produced a spark that was superior in its ability to provide sustained ignition of fuel-lean H2/air propellants. The results from these experiments are being used to optimize a dual pulse laser format for future subscale rocket chamber tests. Besides the ignition enhancement, the dual pulse technique provides a practical way to distribute and deliver laser light to the combustion chamber, an important consideration given the limitation of peak power that can be delivered through optical fibers. With this knowledge, scientists and engineers at Los

  6. Integrated System Health Management: Pilot Operational Implementation in a Rocket Engine Test Stand

    NASA Technical Reports Server (NTRS)

    Figueroa, Fernando; Schmalzel, John L.; Morris, Jonathan A.; Turowski, Mark P.; Franzl, Richard

    2010-01-01

    This paper describes a credible implementation of integrated system health management (ISHM) capability, as a pilot operational system. Important core elements that make possible fielding and evolution of ISHM capability have been validated in a rocket engine test stand, encompassing all phases of operation: stand-by, pre-test, test, and post-test. The core elements include an architecture (hardware/software) for ISHM, gateways for streaming real-time data from the data acquisition system into the ISHM system, automated configuration management employing transducer electronic data sheets (TEDS?s) adhering to the IEEE 1451.4 Standard for Smart Sensors and Actuators, broadcasting and capture of sensor measurements and health information adhering to the IEEE 1451.1 Standard for Smart Sensors and Actuators, user interfaces for management of redlines/bluelines, and establishment of a health assessment database system (HADS) and browser for extensive post-test analysis. The ISHM system was installed in the Test Control Room, where test operators were exposed to the capability. All functionalities of the pilot implementation were validated during testing and in post-test data streaming through the ISHM system. The implementation enabled significant improvements in awareness about the status of the test stand, and events and their causes/consequences. The architecture and software elements embody a systems engineering, knowledge-based approach; in conjunction with object-oriented environments. These qualities are permitting systematic augmentation of the capability and scaling to encompass other subsystems.

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

    NASA Technical Reports Server (NTRS)

    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

  8. GPS Sounding Rocket Developments

    NASA Technical Reports Server (NTRS)

    Bull, Barton

    1999-01-01

    Sounding rockets are suborbital launch vehicles capable of carrying scientific payloads several hundred miles in altitude. These missions return a variety of scientific data including; chemical makeup and physical processes taking place In the atmosphere, natural radiation surrounding the Earth, data on the Sun, stars, galaxies and many other phenomena. In addition, sounding rockets provide a reasonably economical means of conducting engineering tests for instruments and devices used on satellites and other spacecraft prior to their use in more expensive activities. The NASA Sounding Rocket Program is managed by personnel from Goddard Space Flight Center Wallops Flight Facility (GSFC/WFF) in Virginia. Typically around thirty of these rockets are launched each year, either from established ranges at Wallops Island, Virginia, Poker Flat Research Range, Alaska; White Sands Missile Range, New Mexico or from Canada, Norway and Sweden. Many times launches are conducted from temporary launch ranges in remote parts of the world requi6ng considerable expense to transport and operate tracking radars. An inverse differential GPS system has been developed for Sounding Rocket. This paper addresses the NASA Wallops Island history of GPS Sounding Rocket experience since 1994 and the development of a high accurate and useful system.

  9. LOX/hydrocarbon rocket engine analytical design methodology development and validation. Volume 1: Executive summary and technical narrative

    NASA Technical Reports Server (NTRS)

    Pieper, Jerry L.; Walker, Richard E.

    1993-01-01

    During the past three decades, an enormous amount of resources were expended in the design and development of Liquid Oxygen/Hydrocarbon and Hydrogen (LOX/HC and LOX/H2) rocket engines. A significant portion of these resources were used to develop and demonstrate the performance and combustion stability for each new engine. During these efforts, many analytical and empirical models were developed that characterize design parameters and combustion processes that influence performance and stability. Many of these models are suitable as design tools, but they have not been assembled into an industry-wide usable analytical design methodology. The objective of this program was to assemble existing performance and combustion stability models into a usable methodology capable of producing high performing and stable LOX/hydrocarbon and LOX/hydrogen propellant booster engines.

  10. Another Look at Rocket Thrust

    NASA Astrophysics Data System (ADS)

    Hester, Brooke; Burris, Jennifer

    2012-12-01

    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.1,2 Equivalently, in the absence of a net external force, the total momentum of the system, rocket plus ejected gas, remains constant. The law of conservation of momentum is generally used in textbooks3,4 and elsewhere5 in deriving the rocket thrust equation. In this paper we take the somewhat different approach of explicitly applying Newton's second law to the rocket. This method provides a good opportunity to show the importance of choosing carefully the system to which Newton's second law is applied.

  11. A real-time electronic imaging system for solar X-ray observations from sounding rockets

    NASA Technical Reports Server (NTRS)

    Davis, J. M.; Ting, J. W.; Gerassimenko, M.

    1979-01-01

    A real-time imaging system for displaying the solar coronal soft X-ray emission, focussed by a grazing incidence telescope, is described. The design parameters of the system, which is to be used primarily as part of a real-time control system for a sounding rocket experiment, are identified. Their achievement with a system consisting of a microchannel plate, for the conversion of X-rays into visible light, and a slow-scan vidicon, for recording and transmission of the integrated images, is described in detail. The system has a quantum efficiency better than 8 deg above 8 A, a dynamic range of 1000 coupled with a sensitivity to single photoelectrons, and provides a spatial resolution of 15 arc seconds over a field of view of 40 x 40 square arc minutes. The incident radiation is filtered to eliminate wavelengths longer than 100 A. Each image contains 3.93 x 10 to the 5th bits of information and is transmitted to the ground where it is processed by a mini-computer and displayed in real-time on a standard TV monitor.

  12. Rockets Away!

    ERIC Educational Resources Information Center

    Kaahaaina, Nancy

    1997-01-01

    Describes a project that involved a rocket-design competition where students played the roles of McDonnell Douglas employees competing for NASA contracts. Provides a real world experience involving deadlines, design and performance specifications, and budgets. (JRH)

  13. Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Appendix C: Battery report for the liquid rocket booster TVC actuators

    NASA Technical Reports Server (NTRS)

    1989-01-01

    The actuators for control of engine valves and gimbals for a booster require 165 kW or more peak power at 270 volts direct current (VDC) during the 2 or 3 minutes of first stage ascent; other booster devices require much less power at 28 VDC. It is desired that a booster supply its own electrical power and satisfy redundancy requirements of the Solid Rocket Booster Shuttle, when applicable. The power of a Liquid Rocket Booster is therefore provided by two subsystems: Actuator Battery Power (270 VDC) Subsystem for the engine actuators, and Electrical Power and Distribution (28 VDC) Subsystem, to power everything else. Boosters will receive no electrical power from Orbiter, only commands and data, according to current plans. It was concluded that nine 30 volt silver-zinc batteries-in-series be used to provide the 270 volt, 37 kW average (165 kW peak).

  14. System Sensitivity Analysis Applied to the Conceptual Design of a Dual-Fuel Rocket SSTO

    NASA Technical Reports Server (NTRS)

    Olds, John R.

    1994-01-01

    This paper reports the results of initial efforts to apply the System Sensitivity Analysis (SSA) optimization method to the conceptual design of a single-stage-to-orbit (SSTO) launch vehicle. SSA is an efficient, calculus-based MDO technique for generating sensitivity derivatives in a highly multidisciplinary design environment. The method has been successfully applied to conceptual aircraft design and has been proven to have advantages over traditional direct optimization methods. The method is applied to the optimization of an advanced, piloted SSTO design similar to vehicles currently being analyzed by NASA as possible replacements for the Space Shuttle. Powered by a derivative of the Russian RD-701 rocket engine, the vehicle employs a combination of hydrocarbon, hydrogen, and oxygen propellants. Three primary disciplines are included in the design - propulsion, performance, and weights & sizing. A complete, converged vehicle analysis depends on the use of three standalone conceptual analysis computer codes. Efforts to minimize vehicle dry (empty) weight are reported in this paper. The problem consists of six system-level design variables and one system-level constraint. Using SSA in a 'manual' fashion to generate gradient information, six system-level iterations were performed from each of two different starting points. The results showed a good pattern of convergence for both starting points. A discussion of the advantages and disadvantages of the method, possible areas of improvement, and future work is included.

  15. Reactor moderator, pressure vessel, and heat rejection system of an open-cycle gas core nuclear rocket concept

    NASA Technical Reports Server (NTRS)

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

    1973-01-01

    A preliminary design study of a conceptual 6000-megawatt open-cycle gas-core nuclear rocket engine system was made. The engine has a thrust of 196,600 newtons (44,200 lb) and a specific impulse of 4400 seconds. The nuclear fuel is uranium-235 and the propellant is hydrogen. Critical fuel mass was calculated for several reactor configurations. Major components of the reactor (reflector, pressure vessel, and waste heat rejection system) were considered conceptually and were sized.

  16. Longitudinal and lateral-directional static aerodynamic characteristics of an unpowered escape system extraction rocket model with attached launch tubes

    NASA Technical Reports Server (NTRS)

    Huffman, J. K.; Fox, C. H., Jr.; Satterthwaite, R. E.

    1977-01-01

    An escape system extraction rocket proposed for use on the Rotor Systems Research Aircraft was tested at Mach numbers of 0.1 and 0.3 through an angle of attack range from -2 deg to 102 deg and an angle of sideslip range from 0 deg to 15 deg in the Langley 7- by 10-foot high speed tunnel. The data are presented without analysis.

  17. Optical tools and techniques for aligning solar payloads with the SPARCS control system. [Solar Pointing Aerobee Rocket Control System

    NASA Technical Reports Server (NTRS)

    Thomas, N. L.; Chisel, D. M.

    1976-01-01

    The success of a rocket-borne experiment depends not only on the pointing of the attitude control system, but on the alignment of the attitude control system to the payload. To ensure proper alignment, special optical tools and alignment techniques are required. Those that were used in the SPARCS program are described and discussed herein. These tools include theodolites, autocollimators, a 38-cm diameter solar simulator, a high-performance 1-m heliostat to provide a stable solar source during the integration of the rocket payload, a portable 75-cm sun tracker for use at the launch site, and an innovation called the Solar Alignment Prism. Using the real sun as the primary reference under field conditions, the Solar Alignment Prism facilitates the coalignment of the attitude sun sensor with the payload. The alignment techniques were developed to ensure the precise alignment of the solar payloads to the SPARCS attitude sensors during payload integration and to verify the required alignment under field conditions just prior to launch.

  18. Common Data Acquisition Systems (DAS) Software Development for Rocket Propulsion Test (RPT) Test Facilities

    NASA Technical Reports Server (NTRS)

    Hebert, Phillip W., Sr.; Davis, Dawn M.; Turowski, Mark P.; Holladay, Wendy T.; Hughes, Mark S.

    2012-01-01

    The advent of the commercial space launch industry and NASA's more recent resumption of operation of Stennis Space Center's large test facilities after thirty years of contractor control resulted in a need for a non-proprietary data acquisition systems (DAS) software to support government and commercial testing. The software is designed for modularity and adaptability to minimize the software development effort for current and future data systems. An additional benefit of the software's architecture is its ability to easily migrate to other testing facilities thus providing future commonality across Stennis. Adapting the software to other Rocket Propulsion Test (RPT) Centers such as MSFC, White Sands, and Plumbrook Station would provide additional commonality and help reduce testing costs for NASA. Ultimately, the software provides the government with unlimited rights and guarantees privacy of data to commercial entities. The project engaged all RPT Centers and NASA's Independent Verification & Validation facility to enhance product quality. The design consists of a translation layer which provides the transparency of the software application layers to underlying hardware regardless of test facility location and a flexible and easily accessible database. This presentation addresses system technical design, issues encountered, and the status of Stennis development and deployment.

  19. A Low Cost GPS System for Real-Time Tracking of Sounding Rockets

    NASA Technical Reports Server (NTRS)

    Markgraf, M.; Montenbruck, O.; Hassenpflug, F.; Turner, P.; Bull, B.; Bauer, Frank (Technical Monitor)

    2001-01-01

    In an effort to minimize the need for costly, complex, tracking radars, the German Space Operations Center has set up a research project for GPS based tracking of sounding rockets. As part of this project, a GPS receiver based on commercial technology for terrestrial applications has been modified to allow its use under the highly dynamical conditions of a sounding rocket flight. In addition, new antenna concepts are studied as an alternative to proven but costly wrap-around antennas.

  20. Technology Development of a Fiber Optic-Coupled Laser Ignition System for Multi-Combustor Rocket Engines

    NASA Technical Reports Server (NTRS)

    Trinh, Huu P.; Early, Jim; Osborne, Robin; Thomas, Matthew E.; Bossard, John A.

    2002-01-01

    This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Space Flight Center (MSFC). The first two years of the project focus on comprehensive assessments and evaluations of a novel dual-pulse laser concept, flight- qualified laser system, and the technology required to integrate the laser ignition system to a rocket chamber. With collaborations of the Department of Energy/Los Alamos National Laboratory (LANL) and CFD Research Corporation (CFDRC), MSFC has conducted 26 hot fire ignition tests with lab-scale laser systems. These tests demonstrate the concept feasibility of dual-pulse laser ignition to initiate gaseous oxygen (GOX)/liquid kerosene (RP-1) combustion in a rocket chamber. Presently, a fiber optic- coupled miniaturized laser ignition prototype is being implemented at the rocket chamber test rig for future testing. Future work is guided by a technology road map that outlines the work required for maturing a laser ignition system. This road map defines activities for the next six years, with the goal of developing a flight-ready laser ignition system.

  1. Conceptual Engine System Design for NERVA derived 66.7KN and 111.2KN Thrust Nuclear Thermal Rockets

    SciTech Connect

    Fittje, James E.; Buehrle, Robert J.

    2006-01-20

    The Nuclear Thermal Rocket concept is being evaluated as an advanced propulsion concept for missions to the moon and Mars. A tremendous effort was undertaken during the 1960's and 1970's to develop and test NERVA derived Nuclear Thermal Rockets in the 111.2 KN to 1112 KN pound thrust class. NASA GRC is leveraging this past NTR investment in their vehicle concepts and mission analysis studies, and has been evaluating NERVA derived engines in the 66.7 KN to the 111.2 KN thrust range. The liquid hydrogen propellant feed system, including the turbopumps, is an essential component of the overall operation of this system. The NASA GRC team is evaluating numerous propellant feed system designs with both single and twin turbopumps. The Nuclear Engine System Simulation code is being exercised to analyze thermodynamic cycle points for these selected concepts. This paper will present propellant feed system concepts and the corresponding thermodynamic cycle points for 66.7 KN and 111.2 KN thrust NTR engine systems. A pump out condition for a twin turbopump concept will also be evaluated, and the NESS code will be assessed against the Small Nuclear Rocket Engine preliminary thermodynamic data.

  2. Development of limb volume measuring system

    NASA Technical Reports Server (NTRS)

    Bhagat, P. K.; Kadaba, P. K.

    1983-01-01

    The mechanisms underlying the reductions in orthostatic tolerance associated with weightlessness are not well established. Contradictory results from measurements of leg volume changes suggest that altered venomotor tone and reduced blood flow may not be the only contributors to orthostatic intolerance. It is felt that a more accurate limb volume system which is insensitive to environmental factors will aid in better quantification of the hemodynamics of the leg. Of the varous limb volume techniques presently available, the ultrasonic limb volume system has proven to be the best choice. The system as described herein is free from environmental effects, safe, simple to operate and causes negligible radio frequency interference problems. The segmental ultrasonic ultrasonic plethysmograph is expected to provide a better measurement of limb volume change since it is based on cross-sectional area measurements.

  3. Using Monte Carlo techniques and parallel processing for debris hazard analysis of rocket systems

    SciTech Connect

    LaFarge, R.A.

    1994-02-01

    Sandia National Laboratories has been involved with rocket systems for many years. Some of these systems have carried high explosive onboard, while others have had FTS for destruction purposes whenever a potential hazard is detected. Recently, Sandia has also been involved with flight tests in which a target vehicle is intentionally destroyed by a projectile. Such endeavors always raise questions about the safety of personnel and the environment in the event of a premature detonation of the explosive or an activation of the FTS, as well as intentional vehicle destruction. Previous attempts to investigate fragmentation hazards for similar configurations have analyzed fragment size and shape in detail but have computed only a limited number of trajectories to determine the probabilities of impact and casualty expectations. A computer program SAFETIE has been written in support of various SNL flight experiments to compute better approximations of the hazards. SAFETIE uses the AMEER trajectory computer code and the Engineering Sciences Center LAN of Sun workstations to determine more realistically the probability of impact for an arbitrary number of exclusion areas. The various debris generation models are described.

  4. Air-Powered Rockets.

    ERIC Educational Resources Information Center

    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. A dynamic systems analysis of the ''Rocket-Type'' launching of a well-casing tubing string

    SciTech Connect

    McLauchian, R.A.

    1983-11-01

    A dynamic systems analysis was made of the problem of the ''rocket-type'' launching of a well-casing, tubing string from a well at a hole depth of 3300 feet. The objective of this analysis was to assess whether the observed ''rocket-type'' launching of the tubing string could have been caused by nozzle-breakthrough or tubing failure at the SRC No. 1 in Eddy County, New Mexico. The analysis considers tubing weight, fluid buoyancy, external fluid drag and fluid thrust forces on the tubing string. The tubing string and the fluid columns in the moving string and the annular region external to the tubing string are considered as a coupled dynamic system. Several situations of interest are considered relevant to the dynamic behavior of and hence implied failure mode for the tubing string. Summary plots show typical dynamic behavior as a function of internal shut-in pressure and tubing nozzle diameter.

  6. The Space Launch System -The Biggest, Most Capable Rocket Ever Built, for Entirely New Human Exploration Missions Beyond Earth's Orbit

    NASA Technical Reports Server (NTRS)

    Shivers, C. Herb

    2012-01-01

    NASA is developing the Space Launch System -- an advanced heavy-lift launch vehicle that will provide an entirely new capability for human exploration beyond Earth's orbit. The Space Launch System will provide a safe, affordable and sustainable means of reaching beyond our current limits and opening up new discoveries from the unique vantage point of space. The first developmental flight, or mission, is targeted for the end of 2017. The Space Launch System, or SLS, will be designed to carry the Orion Multi-Purpose Crew Vehicle, as well as important cargo, equipment and science experiments to Earth's orbit and destinations beyond. Additionally, the SLS will serve as a backup for commercial and international partner transportation services to the International Space Station. The SLS rocket will incorporate technological investments from the Space Shuttle Program and the Constellation Program in order to take advantage of proven hardware and cutting-edge tooling and manufacturing technology that will significantly reduce development and operations costs. The rocket will use a liquid hydrogen and liquid oxygen propulsion system, which will include the RS-25D/E from the Space Shuttle Program for the core stage and the J-2X engine for the upper stage. SLS will also use solid rocket boosters for the initial development flights, while follow-on boosters will be competed based on performance requirements and affordability considerations.

  7. Description of an experimental (hydrogen peroxide) rocket system and its use in measuring aileron and rudder effectiveness of a light airplane

    NASA Technical Reports Server (NTRS)

    Obryan, T. C.; Goode, M. W.; Gregory, F. D.; Mayo, M. H.

    1980-01-01

    A hydrogen peroxide fueled rocket system, which is to be used as a research tool in flight studies of stall and spin maneuvers, was installed on a light, four place general aviation airplane. The pilot controlled rocket system produces moments about either the roll or the yaw body axis to augment or oppose the aerodynamic forces and inertial moments acting on the airplane during various flight maneuvers, including the spin. These controlled moments of a known magnitude can be used in various ways to help analyze and interpret the importance of the various factors which influence airplane maneuvers. The rocket system and its installation in the airplane are described, and the results of flight rests used to measure rudder and aileron effectiveness at airspeeds above the stall are presented. These tests also serve to demonstrate the operational readiness of the rocket system for future research operations.

  8. NASA Data Acquisition System Software Development for Rocket Propulsion Test Facilities

    NASA Technical Reports Server (NTRS)

    Herbert, Phillip W., Sr.; Elliot, Alex C.; Graves, Andrew R.

    2015-01-01

    Current NASA propulsion test facilities include Stennis Space Center in Mississippi, Marshall Space Flight Center in Alabama, Plum Brook Station in Ohio, and White Sands Test Facility in New Mexico. Within and across these centers, a diverse set of data acquisition systems exist with different hardware and software platforms. The NASA Data Acquisition System (NDAS) is a software suite designed to operate and control many critical aspects of rocket engine testing. The software suite combines real-time data visualization, data recording to a variety formats, short-term and long-term acquisition system calibration capabilities, test stand configuration control, and a variety of data post-processing capabilities. Additionally, data stream conversion functions exist to translate test facility data streams to and from downstream systems, including engine customer systems. The primary design goals for NDAS are flexibility, extensibility, and modularity. Providing a common user interface for a variety of hardware platforms helps drive consistency and error reduction during testing. In addition, with an understanding that test facilities have different requirements and setups, the software is designed to be modular. One engine program may require real-time displays and data recording; others may require more complex data stream conversion, measurement filtering, or test stand configuration management. The NDAS suite allows test facilities to choose which components to use based on their specific needs. The NDAS code is primarily written in LabVIEW, a graphical, data-flow driven language. Although LabVIEW is a general-purpose programming language; large-scale software development in the language is relatively rare compared to more commonly used languages. The NDAS software suite also makes extensive use of a new, advanced development framework called the Actor Framework. The Actor Framework provides a level of code reuse and extensibility that has previously been difficult

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

    NASA Technical Reports Server (NTRS)

    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.

  10. Wash Solution Bath Life Extension for the Space Shuttle Rocket Motor Aqueous Cleaning System

    NASA Technical Reports Server (NTRS)

    Saunders, Chad; Evans, Kurt; Sagers, Neil

    1999-01-01

    A spray-in-air aqueous cleaning system, which replaced 1,1,1 trichloroethane (TCA) vapor degreasing, is used for critical cleaning of Space Shuttle Redesigned Solid Rocket Motor (RSRM) metal parts. Small-scale testing demonstrated that the alkaline-based wash solution possesses adequate soil loading and cleaning properties. However, full-scale testing exhibited unexpected depletion of some primary components of the wash solution. Specifically, there was a significant decrease in the concentration of sodium metasilicate which forced change-out of the wash solution after eight days. Extension of wash solution bath life was necessary to ease the burden of frequent change-out on manufacturing. A laboratory study supports a depletion mechanism that is initiated by the hydrolysis of sodium tripolyphosphate (STPP) lowering the pH of the solution. The decrease in pH causes polymerization and subsequent precipitation of sodium metasilicate (SM). Further investigation showed that maintaining the pH was the key to preventing the precipitation of the sodium metasilicate. Implementation to the full scale operation demonstrated that periodic additions of potassium hydroxide (KOH) extended the useful bath life to more than four months.

  11. Development of an On-board Failure Diagnostics and Prognostics System for Solid Rocket Booster

    NASA Technical Reports Server (NTRS)

    Smelyanskiy, Vadim N.; Luchinsky, Dmitry G.; Osipov, Vyatcheslav V.; Timucin, Dogan A.; Uckun, Serdar

    2009-01-01

    We develop a case breach model for the on-board fault diagnostics and prognostics system for subscale solid-rocket boosters (SRBs). The model development was motivated by recent ground firing tests, in which a deviation of measured time-traces from the predicted time-series was observed. A modified model takes into account the nozzle ablation, including the effect of roughness of the nozzle surface, the geometry of the fault, and erosion and burning of the walls of the hole in the metal case. The derived low-dimensional performance model (LDPM) of the fault can reproduce the observed time-series data very well. To verify the performance of the LDPM we build a FLUENT model of the case breach fault and demonstrate a good agreement between theoretical predictions based on the analytical solution of the model equations and the results of the FLUENT simulations. We then incorporate the derived LDPM into an inferential Bayesian framework and verify performance of the Bayesian algorithm for the diagnostics and prognostics of the case breach fault. It is shown that the obtained LDPM allows one to track parameters of the SRB during the flight in real time, to diagnose case breach fault, and to predict its values in the future. The application of the method to fault diagnostics and prognostics (FD&P) of other SRB faults modes is discussed.

  12. Nuclear Thermal Rocket (NTR) Propulsion and Power Systems for Outer Planetary Exploration Missions

    NASA Technical Reports Server (NTRS)

    Borowski, S. K.; Cataldo, R. L.

    2001-01-01

    The high specific impulse (I (sub sp)) and engine thrust generated using liquid hydrogen (LH2)-cooled Nuclear Thermal Rocket (NTR) propulsion makes them attractive for upper stage applications for difficult robotic science missions to the outer planets. Besides high (I (sub sp)) and thrust, NTR engines can also be designed for "bimodal" operation allowing substantial amounts of electrical power (10's of kWe ) to be generated for onboard spacecraft systems and high data rate communications with Earth during the course of the mission. Two possible options for using the NTR are examined here. A high performance injection stage utilizing a single 15 klbf thrust engine can inject large payloads to the outer planets using a 20 t-class launch vehicle when operated in an "expendable mode". A smaller bimodal NTR stage generating approx. 1 klbf of thrust and 20 to 40 kWe for electric propulsion can deliver approx. 100 kg using lower cost launch vehicles. Additional information is contained in the original extended abstract.

  13. Seasat. Volume 3: Ground systems

    NASA Technical Reports Server (NTRS)

    Pounder, E. (Editor)

    1980-01-01

    The Seasat Project was a feasibility demonstration of the use of orbital remote sensing for global ocean observation. The satellite was launched in June of 1978 and was operated successfully until October 1978. A massive electrical failure occurred in the power system, terminating the mission prematurely. The ground systems using during the mission life are discussed. Descriptions of the operating organization, the system elements, and the testing program are included. The various phases of the mission: launch and orbit insertion; cruise; and calibration are discussed. A special section is included on the orbit maneuver activites. Operations during the satellite failure are reviewed and summarized.

  14. Multifunction display system, volume 1

    NASA Technical Reports Server (NTRS)

    1973-01-01

    The design and construction of a multifunction display man/machine interface for use with a 4 pi IBM-360 System are described. The system is capable of displaying superimposed volatile alphanumeric and graphical data on a 512 x 512 element plasma panel, and holographically stored multicolor archival information. The volatile data may be entered from a keyboard or by means of an I/O interface to the 360 system. A 2-page memory local to the display is provided for storing the entered data. The archival data is stored as a phase hologram on a vinyl tape strip. This data is accessible by means of a rapid transport system which responds to inputs provided by the I/O channel on the keyboard. As many as 500 frames may be stored on a tape strip for access in under 6 seconds.

  15. Imaging System for a Sub-Orbital Sounding Rocket Mission Based Upon Next Generation Detector Technology

    NASA Astrophysics Data System (ADS)

    Veach, Todd; Scowen, P.; Beasley, M.; Nikzad, S.

    2011-05-01

    We present the design and preliminary results from the fabrication of a charge-coupled device (CCD) based imaging system designed using a modified modular imager cell (MIC) for use in a sounding rocket mission. The heart of the imaging system is the modified MIC, which provides the video pre-amplifier circuitry and CCD clock level filtering. The MIC is designed with a four-layer FR4 printed circuit board (PCB) with surface mount and through-hole components for ease of testing and lower fabrication cost. The imager is a delta doped 3.5k by 3.5k LBNL CCD. Delta doping the detector provides for enhanced QE response in the UV. Detector readout is performed by the recently released PCIe/104 Small-Cam imager controller from Astronomical Research Cameras, Inc (ARC). The PCIe/104 Small-Cam system has the same capabilities as its larger PCIe brethren, but in a smaller form factor, which makes it ideally suited for sub-orbital ballistic missions. The overall control is then accomplished using a PCIe/104 computer from RTD Embedded Technologies, Inc. For laboratory testing and calibration, the modified MIC is placed inside an IR Labs ND5 liquid nitrogen cooled dewar. Upon flight, the modified MIC is placed within a 6.75” diameter 10” long ultra-high vacuum (UHV) vessel. The design, fabrication, and testing is being done at the Laboratory for Astronomical and Space Instrumentation (LASI) at Arizona State University. The LASI Lab is a state of the art detector calibration facility providing calibration from the 300 nm to 2.3 microns with further capability for designing hardware for use in suborbital ballistic missions.

  16. The United States sounding rocket program

    NASA Technical Reports Server (NTRS)

    1971-01-01

    The United States sounding rocket program is discussed. The program is concerned with the fields of solar physics, galactic astronomy, fields and particles, ionospheric physics, aeronomy, and meteorology. Sounding rockets are described with respect to propulsion systems, gross weight, and capabilities. Instruments used to conduct ionospheric probing missions are examined. Results of previously conducted sounding rocket missions are included.

  17. Array Of Rockets For Multicrewmember Evacuation

    NASA Technical Reports Server (NTRS)

    Allen, Margaret A.

    1990-01-01

    Emergency egress system undergoing development for aircraft and aerospace vehicles uses fixed array of tractor rockets to eject crewmembers. Crewmembers hook up to tractor rockets and fire them unaided. Positioned as unit in ready-to-use orientation during flight operations. On ground, swung out of way. Rocket array also mounts under exit hatch, serving as egress ramp.

  18. Seasat. Volume 2: Flight systems

    NASA Technical Reports Server (NTRS)

    Pounder, E. (Editor)

    1980-01-01

    Flight systems used in the Seasat Project are described. Included are (1) launch operation; (2) satellite performance after launch; (3) sensors that collected data; and (4) the launch vehicle that placed the satellite into Earth orbit. Techniques for sensor management are explained.

  19. Conceptual Design for a Dual-Bell Rocket Nozzle System Using a NASA F-15 Airplane as the Flight Testbed

    NASA Technical Reports Server (NTRS)

    Jones, Daniel S.; Ruf, Joseph H.; Bui, Trong T.; Martinez, Martel; St. John, Clinton W.

    2014-01-01

    The dual-bell rocket nozzle was first proposed in 1949, offering a potential improvement in rocket nozzle performance over the conventional-bell nozzle. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. In 2013 a proposal was constructed that offered a NASA F-15 airplane as the flight testbed, with the plan to operate a dual-bell rocket nozzle during captive-carried flight. If implemented, this capability will permit nozzle operation into an external flow field similar to that of a launch vehicle, and facilitate an improved understanding of dual-bell nozzle plume sensitivity to external flow-field effects. More importantly, this flight testbed can be utilized to help quantify the performance benefit with the dual-bell nozzle, as well as to advance its technology readiness level. Toward this ultimate goal, this paper provides plans for future flights to quantify the external flow field of the airplane near the nozzle experiment, as well as details on the conceptual design for the dual-bell nozzle cold-flow propellant feed system integration within the NASA F-15 Propulsion Flight Test Fixture. The current study shows that this concept of flight research is feasible, and could result in valuable flight data for the dual-bell nozzle.

  20. Conceptual Design for a Dual-Bell Rocket Nozzle System Using a NASA F-15 Airplane as the Flight Testbed

    NASA Technical Reports Server (NTRS)

    Jones, Daniel S.; Ruf, Joseph H.; Bui, Trong T.; Martinez, Martel; St. John, Clinton W.

    2014-01-01

    The dual-bell rocket nozzle was first proposed in 1949, offering a potential improvement in rocket nozzle performance over the conventional-bell nozzle. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. In 2013 a proposal was constructed that offered a National Aeronautics and Space Administration (NASA) F-15 airplane as the flight testbed, with the plan to operate a dual-bell rocket nozzle during captive-carried flight. If implemented, this capability will permit nozzle operation into an external flow field similar to that of a launch vehicle, and facilitate an improved understanding of dual-bell nozzle plume sensitivity to external flow-field effects. More importantly, this flight testbed can be utilized to help quantify the performance benefit with the dual-bell nozzle, as well as to advance its technology readiness level. Toward this ultimate goal, this report provides plans for future flights to quantify the external flow field of the airplane near the nozzle experiment, as well as details on the conceptual design for the dual-bell nozzle cold-flow propellant feed system integration within the NASA F-15 Propulsion Flight Test Fixture. The current study shows that this concept of flight research is feasible, and could result in valuable flight data for the dual-bell nozzle.

  1. Conceptual Design for a Dual-Bell Rocket Nozzle System Using a NASA F-15 Airplane as the Flight Testbed

    NASA Technical Reports Server (NTRS)

    Jones, Daniel S.; Ruf, Joseph H.; Bui, Trong T.; Martinez, Martel; St. John, Clinton W.

    2014-01-01

    The dual-bell rocket nozzle was first proposed in 1949, offering a potential improvement in rocket nozzle performance over the conventional-bell nozzle. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. In 2013 a proposal was constructed that offered a NASA F-15 airplane as the flight testbed, with the plan to operate a dual-bell rocket nozzle during captive-carried flight. If implemented, this capability will permit nozzle operation into an external flow field similar to that of a launch vehicle, and facilitate an improved understanding of dual-bell nozzle plume sensitivity to external flow-field effects. More importantly, this flight testbed can be utilized to help quantify the performance benefit with the dual-bell nozzle, as well as to advance its technology readiness level. This presentation provides highlights of a technical paper that outlines this ultimate goal, including plans for future flights to quantify the external flow field of the airplane near the nozzle experiment, as well as details on the conceptual design for the dual-bell nozzle cold-flow propellant feed system integration within the NASA F-15 Propulsion Flight Test Fixture. The current study shows that this concept of flight research is feasible, and could result in valuable flight data for the dual-bell nozzle.

  2. Rocket plume burn hazard.

    PubMed

    Stoll, A M; Piergallini, J R; Chianta, M A

    1980-05-01

    By use of miniature rocket engines, the burn hazard posed by exposure to ejection seat rocket plume flames was determined in the anaesthetized rat. A reference chart is provided for predicting equivalent effects in human skin based on extrapolation of earlier direct measurements of heat input for rat and human burns. The chart is intended to be used in conjunction with thermocouple temperature measurements of the plume environment for design and modification of escape seat system to avoid thermal injury on ejection from multiplace aircraft. PMID:7387571

  3. Ignition and Performance Tests of Rocket-Based Combined Cycle Propulsion System

    NASA Technical Reports Server (NTRS)

    Anderson, William E.

    2005-01-01

    The ground testing of a Rocket Based Combined Cycle engine implementing the Simultaneous Mixing and Combustion scheme was performed at the direct-connect facility of Purdue University's High Pressure Laboratory. The fuel-rich exhaust of a JP-8/H2O2 thruster was mixed with compressed, metered air in a constant area, axisymmetric duct. The thruster was similar in design and function to that which will be used in the flight test series of Dryden's Ducted-Rocket Experiment. The determination of duct ignition limits was made based on the variation of secondary air flow rates and primary thruster equivalence ratios. Thrust augmentation and improvements in specific impulse were studied along with the pressure and temperature profiles of the duct to study mixing lengths and thermal choking. The occurrence of ignition was favored by lower rocket equivalence ratios. However, among ignition cases, better thrust and specific impulse performance were seen with higher equivalence ratios owing to the increased fuel available for combustion. Thrust and specific impulse improvements by factors of 1.2 to 1.7 were seen. The static pressure and temperature profiles allowed regions of mixing and heat addition to be identified. The mixing lengths were found to be shorter at lower rocket equivalence ratios. Total pressure measurements allowed plume-based calculation of thrust, which agreed with load-cell measured values to within 6.5-8.0%. The corresponding Mach Number profile indicated the flow was not thermally choked for the highest duct static pressure case.

  4. Reusable sounding-rocket design

    NASA Astrophysics Data System (ADS)

    Woo, Dick L. Y.; Martin, James A.

    1995-03-01

    As a result of the reduction of budgets for flights, the ideas of reusability and cost-effectiveness in launch vehicles are becoming more and more important. One class of rockets, in particular the sounding rockets operating in a less demanding environment, has many potentials for many more flights. By augmenting the basic rocket configuration with wings, landing gear, flight controls and guidance systems, the vehicle can be made to glide and land back at the launch site or at a specific recovery site. In this paper, the design of such a reusable rocket is presented. This design can be extended and adapted to larger vehicles, thus attaining higher altitudes required in some of the applications of sounding rockets.

  5. Satellite voice broadcast. Volume 2: System study

    NASA Technical Reports Server (NTRS)

    Bachtell, E. E.; Bettadapur, S. S.; Coyner, J. V.; Farrell, C. E.

    1985-01-01

    The Technical Volume of the Satellite Broadcast System Study is presented. Designs are synthesized for direct sound broadcast satellite systems for HF-, VHF-, L-, and Ku-bands. Methods are developed and used to predict satellite weight, volume, and RF performance for the various concepts considered. Cost and schedule risk assessments are performed to predict time and cost required to implement selected concepts. Technology assessments and tradeoffs are made to identify critical enabling technologies that require development to bring technical risk to acceptable levels for full scale development.

  6. A Review of Propulsion Industrial Base Studies and an Introduction to the National Institute of Rocket Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Doreswamy, Rajiv; Fry, Emma K.

    2012-01-01

    Over the past decade there have been over 40 studies that have examined the state of the industrial base and infrastructure that supports propulsion systems development in the United States. This paper offers a comprehensive, systematic review of these studies and develops conclusions and recommendations in the areas of budget, policy, sustainment, infrastructure, workforce retention and development and mission/vision and policy. The National Institute for Rocket Propulsion System (NIRPS) is a coordinated, national organization that is responding to the key issues highlighted in these studies. The paper outlines the case for NIRPS and the specific actions that the Institute is taking to address these issues.

  7. Turbo Pump Fed Micro-Rocket Engine

    NASA Astrophysics Data System (ADS)

    Miotti, P.; Tajmar, M.; Seco, F.; Guraya, C.; Perennes, F.; Soldati, A.; Lang, M.

    2004-10-01

    Micro-satellites (from 10kg up to 100kg) have mass, volume, and electrical power constraints due to their low dimensions. These limitations lead to the lack in currently available active orbit control systems in micro-satellites. Therefore, a micro-propulsion system with a high thrust to mass ratio is required to increase the potential functionality of small satellites. Mechatronic is presently working on a liquid bipropellant micro-rocket engine under contract with ESA (Contract No.16914/NL/Sfe - Micro-turbo-machinery Based Bipropellant System Using MNT). The advances in Mechatronic's project are to realise a micro-rocket engine with propellants pressurised by micro-pumps. The energy for driving the pumps would be extracted from a micro-turbine. Cooling channels around the nozzle would be also used in order to maintain the wall material below its maximum operating temperature. A mass budget comparison with more traditional pressure-fed micro-rockets shows a real benefit from this system in terms of mass reduction. In the paper, an overview of the project status in Mechatronic is presented.

  8. Hybrid rocket performance

    NASA Technical Reports Server (NTRS)

    Frederick, Robert A., Jr.

    1992-01-01

    A hybrid rocket is a system consisting of a solid fuel grain and a gaseous or liquid oxidizer. Figure 1 shows three popular hybrid propulsion cycles that are under current consideration. NASA MSFC has teamed with industry to test two hybrid propulsion systems that will allow scaling to motors of potential interest for Titan and Atlas systems, as well as encompassing the range of interest for SEI lunar ascent stages and National Launch System Cargo Transfer Vehicle (NLS CTV) and NLS deorbit systems. Hybrid systems also offer advantages as moderate-cost, environmentally acceptable propulsion system. The objective of this work was to recommend a performance prediction methodology for hybrid rocket motors. The scope included completion of: a literature review, a general methodology, and a simplified performance model.

  9. Transient Mathematical Modeling for Liquid Rocket Engine Systems: Methods, Capabilities, and Experience

    NASA Technical Reports Server (NTRS)

    Seymour, David C.; Martin, Michael A.; Nguyen, Huy H.; Greene, William D.

    2005-01-01

    The subject of mathematical modeling of the transient operation of liquid rocket engines is presented in overview form from the perspective of engineers working at the NASA Marshall Space Flight Center. The necessity of creating and utilizing accurate mathematical models as part of liquid rocket engine development process has become well established and is likely to increase in importance in the future. The issues of design considerations for transient operation, development testing, and failure scenario simulation are discussed. An overview of the derivation of the basic governing equations is presented along with a discussion of computational and numerical issues associated with the implementation of these equations in computer codes. Also, work in the field of generating usable fluid property tables is presented along with an overview of efforts to be undertaken in the future to improve the tools use for the mathematical modeling process.

  10. Transient Mathematical Modeling for Liquid Rocket Engine Systems: Methods, Capabilities, and Experience

    NASA Technical Reports Server (NTRS)

    Martin, Michael A.; Nguyen, Huy H.; Greene, William D.; Seymout, David C.

    2003-01-01

    The subject of mathematical modeling of the transient operation of liquid rocket engines is presented in overview form from the perspective of engineers working at the NASA Marshall Space Flight Center. The necessity of creating and utilizing accurate mathematical models as part of liquid rocket engine development process has become well established and is likely to increase in importance in the future. The issues of design considerations for transient operation, development testing, and failure scenario simulation are discussed. An overview of the derivation of the basic governing equations is presented along with a discussion of computational and numerical issues associated with the implementation of these equations in computer codes. Also, work in the field of generating usable fluid property tables is presented along with an overview of efforts to be undertaken in the future to improve the tools use for the mathematical modeling process.

  11. Performance and heat transfer characteristics of the laser-heated rocket - A future space transportation system

    NASA Technical Reports Server (NTRS)

    Shoji, J. M.; Larson, V. R.

    1976-01-01

    The application of advanced liquid-bipropellant rocket engine analysis techniques has been utilized for prediction of the potential delivered performance and the design of thruster wall cooling schemes for laser-heated rocket thrusters. Delivered specific impulse values greater than 1000 lbf-sec/lbm are potentially achievable based on calculations for thrusters designed for 10-kW and 5000-kW laser beam power levels. A thruster wall-cooling technique utilizing a combination of regenerative cooling and a carbon-seeded hydrogen boundary layer is presented. The flowing carbon-seeded hydrogen boundary layer provides radiation absorption of the heat radiated from the high-temperature plasma. Also described is a forced convection thruster wall cooling design for an experimental test thruster.

  12. Small Solid Rocket Motor Test

    NASA Video Gallery

    It was three-two-one to brilliant fire as NASA's Marshall Space Flight Center tested a small solid rocket motor designed to mimic NASA's Space Launch System booster. The Mar. 14 test provides a qui...

  13. Study of solid rocket motors for a space shuttle booster. Volume 2 book 2: Supporting research and technology

    NASA Technical Reports Server (NTRS)

    Vonderesch, A. H.

    1972-01-01

    The baseline SRM design for the space shuttle employs proven technology based on actual motor firings. Supporting research and technology are therefore required only to address system technology that is specific to the shuttle requirements, and that is needed for optimization of design features. Eight programs are recommended to meet these requirements.

  14. Operationally Efficient Propulsion System Study (OEPSS) data book. Volume 4: OEPSS design concepts

    NASA Technical Reports Server (NTRS)

    Wong, George S.; Ziese, James M.; Farhangi, Shahram

    1990-01-01

    This study was initiated to identify operations problems and cost drivers for current propulsion systems and to identify technology and design approaches to increase the operational efficiency and reduce operations costs for future propulsion systems. To provide readily usable data for the Advanced Launch System (ALS) program, the results of the OEPSS study have been organized into a series of OEPSS Data Books. This volume describes three propulsion concepts that will simplify the propulsion system design and significantly reduce operational requirements. The concepts include: (1) a fully integrated, booster propulsion module concept for the ALS that avoids the complex system created by using autonomous engines with numerous artificial interfaces; (2) an LOX tank aft concept which avoids potentially dangerous geysering in long LOX propellant lines; and (3) an air augmented, rocket engine nozzle afterburning propulsion concept that will significantly reduce LOX propellant requirements, reduce vehicle size and simplify ground operations and ground support equipment and facilities.

  15. Advanced vehicle systems assessment. Volume V. Appendices

    SciTech Connect

    Not Available

    1985-03-01

    This report, which is divided into five volumes, documents the evaluation of advanced electric and hybrid vehicles for potential development by the early 1990s. The primary objective of the assessment is to recommend subsystem research priorities based on a comparison of alternatives as part of complete vehicle systems with equivalent performance. The assessment includes evaluations of candidate technologies as well as technical and economic comparisons of vehicle systems for specified missions. The availability of nonpetroleum fuel is also addressed, and preference analyses are used to assist in the evaluation of the relative merits of competing systems. Volume V, the Appendices, includes reports on battery design, battery cost, aluminum vehicle construction, IBM PC computer programs, and battery discharge models.

  16. Nuclear Engine System Simulation (NESS). Volume 1: Program user's guide

    NASA Technical Reports Server (NTRS)

    Pelaccio, Dennis G.; Scheil, Christine M.; Petrosky, Lyman J.

    1993-01-01

    A Nuclear Thermal Propulsion (NTP) engine system design analysis tool is required to support current and future Space Exploration Initiative (SEI) propulsion and vehicle design studies. Currently available NTP engine design models are those developed during the NERVA program in the 1960's and early 1970's and are highly unique to that design or are modifications of current liquid propulsion system design models. To date, NTP engine-based liquid design models lack integrated design of key NTP engine design features in the areas of reactor, shielding, multi-propellant capability, and multi-redundant pump feed fuel systems. Additionally, since the SEI effort is in the initial development stage, a robust, verified NTP analysis design tool could be of great use to the community. This effort developed an NTP engine system design analysis program (tool), known as the Nuclear Engine System Simulation (NESS) program, to support ongoing and future engine system and stage design study efforts. In this effort, Science Applications International Corporation's (SAIC) NTP version of the Expanded Liquid Engine Simulation (ELES) program was modified extensively to include Westinghouse Electric Corporation's near-term solid-core reactor design model. The ELES program has extensive capability to conduct preliminary system design analysis of liquid rocket systems and vehicles. The program is modular in nature and is versatile in terms of modeling state-of-the-art component and system options as discussed. The Westinghouse reactor design model, which was integrated in the NESS program, is based on the near-term solid-core ENABLER NTP reactor design concept. This program is now capable of accurately modeling (characterizing) a complete near-term solid-core NTP engine system in great detail, for a number of design options, in an efficient manner. The following discussion summarizes the overall analysis methodology, key assumptions, and capabilities associated with the NESS presents an

  17. Liquid Rocket Boosters for Shuttle

    NASA Astrophysics Data System (ADS)

    Hughes, James E.

    The Liquid Rocket Booster study was initiated by NASA to define an alternative to the Solid Rocket Boosters used on the STS. These studies have involved MSFC, JSC and KSC and their contractors. The prime study contractors, Martin Marietta Corporation and General Dynamics Space Systems, have identified Liquid Booster configurations which would replace the SRB's in the Shuttle stack. The Liquid Rocket Booster increases Shuttle performance to 70 K LBS, provides improved reliability, hold down and verification prior to vehicle release, engine out and improved abort capability, and is phased into the STS launch operations without adversely affecting flight rate.

  18. Real-Time Rocket/Vehicle System Integrated Health Management Laboratory For Development and Testing of Health Monitoring/Management Systems

    NASA Technical Reports Server (NTRS)

    Aguilar, R.

    2006-01-01

    Pratt & Whitney Rocketdyne has developed a real-time engine/vehicle system integrated health management laboratory, or testbed, for developing and testing health management system concepts. This laboratory simulates components of an integrated system such as the rocket engine, rocket engine controller, vehicle or test controller, as well as a health management computer on separate general purpose computers. These general purpose computers can be replaced with more realistic components such as actual electronic controllers and valve actuators for hardware-in-the-loop simulation. Various engine configurations and propellant combinations are available. Fault or failure insertion capability on-the-fly using direct memory insertion from a user console is used to test system detection and response. The laboratory is currently capable of simulating the flow-path of a single rocket engine but work is underway to include structural and multiengine simulation capability as well as a dedicated data acquisition system. The ultimate goal is to simulate as accurately and realistically as possible the environment in which the health management system will operate including noise, dynamic response of the engine/engine controller, sensor time delays, and asynchronous operation of the various components. The rationale for the laboratory is also discussed including limited alternatives for demonstrating the effectiveness and safety of a flight system.

  19. Flight Testing a Real-Time Hazard Detection System for Safe Lunar Landing on the Rocket-Powered Morpheus Vehicle

    NASA Technical Reports Server (NTRS)

    Trawny, Nikolas; Huertas, Andres; Luna, Michael E.; Villalpando, Carlos Y.; Martin, Keith E.; Carson, John M.; Johnson, Andrew E.; Restrepo, Carolina; Roback, Vincent E.

    2015-01-01

    The Hazard Detection System (HDS) is a component of the ALHAT (Autonomous Landing and Hazard Avoidance Technology) sensor suite, which together provide a lander Guidance, Navigation and Control (GN&C) system with the relevant measurements necessary to enable safe precision landing under any lighting conditions. The HDS consists of a stand-alone compute element (CE), an Inertial Measurement Unit (IMU), and a gimbaled flash LIDAR sensor that are used, in real-time, to generate a Digital Elevation Map (DEM) of the landing terrain, detect candidate safe landing sites for the vehicle through Hazard Detection (HD), and generate hazard-relative navigation (HRN) measurements used for safe precision landing. Following an extensive ground and helicopter test campaign, ALHAT was integrated onto the Morpheus rocket-powered terrestrial test vehicle in March 2014. Morpheus and ALHAT then performed five successful free flights at the simulated lunar hazard field constructed at the Shuttle Landing Facility (SLF) at Kennedy Space Center, for the first time testing the full system on a lunar-like approach geometry in a relevant dynamic environment. During these flights, the HDS successfully generated DEMs, correctly identified safe landing sites and provided HRN measurements to the vehicle, marking the first autonomous landing of a NASA rocket-powered vehicle in hazardous terrain. This paper provides a brief overview of the HDS architecture and describes its in-flight performance.

  20. Numerical Modeling of Fluid Transient in Cryogenic Fluid Network of Rocket Propulsion System

    NASA Technical Reports Server (NTRS)

    Majumdar, Alok; Flachbart, Robin

    2003-01-01

    Fluid transients, also known as water hammer, can have a significant impact on the design and operation of both spacecraft and launch vehicles propulsion systems. These transients often occur at system activation and shut down. For ground safety reasons, many spacecrafts are launched with the propellant lines dry. These lines are often evacuated by the time the spacecraft reaches orbit. When the propellant isolation valve opens during propulsion system activation, propellant rushes into lines creating a pressure surge. During propellant system shutdown, a pressure surge is created due to sudden closure of a valve. During both activation and shutdown, pressure surges must be predicted accurately to ensure structural integrity of the propulsion system fluid network. The method of characteristics is the most widely used method of calculating fluid transients in pipeline [ 1,2]. The method of characteristics, however, has limited applications in calculating flow distribution in complex flow circuits with phase change, heat transfer and rotational effects. A robust cryogenic propulsion system analyzer must have the capability to handle phase change, heat transfer, chemical reaction, rotational effects and fluid transients in conjunction with subsystem flow model for pumps, valves and various pipe fittings. In recent years, such a task has been undertaken at Marshall Space Flight Center with the development of the Generalized Fluid System Simulation Program (GFSSP), which is based on finite volume method in fluid network [3]. GFSSP has been extensively verified and validated by comparing its predictions with test data and other numerical methods for various applications such as internal flow of turbo-pump [4], propellant tank pressurization [5,6], chilldown of cryogenic transfer line [7] and squeeze film damper rotordynamics [8]. The purpose of the present paper is to investigate the applicability of the finite volume method to predict fluid transient in cryogenic flow

  1. Testing of Environmentally Preferable Aluminum Pretreatments and Coating Systems for Use on Space Shuttle Solid Rocket Boosters (SRB)

    NASA Technical Reports Server (NTRS)

    Clayton, C.; Raley, R.; Zook, L.

    2001-01-01

    The solid rocket booster (SRB) has historically used a chromate conversion coating prior to protective finish application. After conversion coating, an organic paint system consisting of a chromated epoxy primer and polyurethane topcoat is applied. An overall systems approach was selected to reduce waste generation from the coatings application and removal processes. While the most obvious waste reduction opportunity involved elimination of the chromate conversion coating, several other coating system configurations were explored in an attempt to reduce the total waste. This paper will briefly discuss the use of a systems view to reduce waste generation from the coating process and present the results of the qualification testing of nonchromated aluminum pretreatments and alternate coating systems configurations.

  2. Analysis and Results from a Flush Airdata Sensing (FADS) System in Close Proximity to Firing Rocket Nozzles

    NASA Technical Reports Server (NTRS)

    Ali, Aliyah N.; Borrer, Jerry L.

    2013-01-01

    This presentation presents information regarding the nose-cap flush airdata sensing (FADS) system on Orion's Pad Abort 1 (PA-1) vehicle. The purpose of the nose-cap FADS system was to test whether or not useful data could be obtained from a FADS system if it was placed in close proximity to firing rockets nozzles like the attitude control motor (ACM) nozzles on the PA-1 launch abort system (LAS). The nose-cap FADS systems use pressure measurements from a series of pressure ports which are arranged in a cruciform pattern and flush with the surface of the vehicle to estimate values of angle of attack, angle of side-slip, Mach number, impact pressure and free-stream static pressure.

  3. Hybrid Rocket Propulsion for Sounding Rocket Applications

    NASA Technical Reports Server (NTRS)

    1991-01-01

    A discussion of the H-225K hybrid rocket motor, produced by the American Rocket Company, is given. The H-225K motor is presented in terms of the following topics: (1) hybrid rocket fundamentals; (2) hybrid characteristics; and (3) hybrid advantages.

  4. Redesign of solid rocket booster/external tank attachment ring for the space transportation system

    NASA Technical Reports Server (NTRS)

    Mccomb, Harvey G., Jr. (Compiler)

    1987-01-01

    An improved design concept is presented for the Space Shuttle solid rocket booster (SRB)/external tank (ET) attachment ring structural component. This component picks up three struts which attach the aft end of each SRB to the ET. The concept is a partial ring with carefully tapered ends to distribute fastener loads safely into the SRB. Extensive design studies and analyses were performed to arrive at the concept. Experiments on structural elements were performed to determine material strength and stiffness characteristics. Materials and fabrication studies were conducted to determine acceptable tolerances for the design concept. An overview is provided of the work along with conclusions and major recommendations.

  5. Advanced Computing Technologies for Rocket Engine Propulsion Systems: Object-Oriented Design with C++

    NASA Technical Reports Server (NTRS)

    Bekele, Gete

    2002-01-01

    This document explores the use of advanced computer technologies with an emphasis on object-oriented design to be applied in the development of software for a rocket engine to improve vehicle safety and reliability. The primary focus is on phase one of this project, the smart start sequence module. The objectives are: 1) To use current sound software engineering practices, object-orientation; 2) To improve on software development time, maintenance, execution and management; 3) To provide an alternate design choice for control, implementation, and performance.

  6. Solid Rocket Motor Acoustic Testing

    SciTech Connect

    Rogers, J.D.

    1999-03-31

    Acoustic data are often required for the determination of launch and powered flight loads for rocket systems and payloads. Such data are usually acquired during test firings of the solid rocket motors. In the current work, these data were obtained for two tests at a remote test facility where we were visitors. This paper describes the data acquisition and the requirements for working at a remote site, interfacing with the test hosts.

  7. Computational simulation of liquid rocket injector anomalies

    NASA Technical Reports Server (NTRS)

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

    1986-01-01

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

  8. Computerized 50 liter volume calibration system

    SciTech Connect

    Proffitt, T.H.

    1990-01-01

    A system has been designed for the Savannah River Site that will be used to calibrate product shipping containers. For accountability purposes, it is necessary that these containers be calibrated to a very high precision. The Computerized 50 Liter Volume Calibration System (CVCS), which is based on the Ideal Gas Law (IGL), will use reference volumes with precision of no less {plus minus}0.03%, and helium to calibrate the containers to have a total error of no greater than {plus minus}0.10%. A statistical interpretation of the system has given a theoretical total calculated error of {plus minus}0.08%. Tests with the system will be performed once fabrication is complete to experimentally verify the calculated error. Since the total error was calculated using the worst case scenario, the actual error should be significantly less than the calculated value. The computer controlled, totally automated system is traceable to the National Institute of Standards and Technology. The design, calibration procedure, and statistical interpretation of the system will be discussed. 1 ref.

  9. Design and Analysis of a Getter-Based Vacuum Pumping System for a Rocket-Borne Mass Spectrometer

    NASA Astrophysics Data System (ADS)

    Everett, E. A.; Syrstad, E. A.; Dyer, J. S.

    2010-12-01

    The mesosphere / lower thermosphere (MLT) is a transition region where the turbulent mixing of earth’s lower atmosphere gives way to the molecular diffusion of space. This region hosts a rich array of chemical processes and atmospheric phenomena, and serves to collect and distribute particles of all sizes in thin layers. Spatially resolved in situ characterization of these layers is very difficult, due to the elevated pressure of the MLT, limited access via high-speed sounding rockets, and the enormous variety of charged and neutral species that range in size from atoms to smoke and dust particles. In terrestrial applications, time-of-flight mass spectrometry (TOF-MS) is the technique of choice for performing fast, sensitive composition measurements with extremely large mass range. However, because of its reliance on high voltages and microchannel plate (MCP) detectors prone to discharge at elevated pressures, TOF-MS has rarely been employed for measurements of the MLT, where ambient pressures approach 10 mTorr. We present a novel, compact mass spectrometer design appropriate for deployment aboard sounding rockets. This Hadamard transform time-of-flight mass spectrometer (HT-TOF-MS) applies a multiplexing technique through pseudorandom beam modulation and spectral deconvolution to achieve very high measurement duty cycles (50%), with a theoretically unlimited mass range. The HT-TOF-MS employs a simple, getter-based vacuum pumping system and pressure-tolerant MCP to allow operation in the MLT. The HT-TOF-MS must provide sufficient vacuum pumping to 1) maintain a minimum mean free path inside the instrument, to avoid spectral resolution loss, and 2) to avoid MCP failure through electrostatic discharge. The design incorporates inexpensive, room temperature tube getters loaded with nano-structured barium to meet these pumping speed requirements, without the use of cryogenics or mechanical pumping systems. We present experimental results for gettering rates and

  10. Investigation of Post-Flight Solid Rocket Booster Thermal Protection System

    NASA Technical Reports Server (NTRS)

    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.

  11. Development of a prototype fluid volume measurement system. [for urine volume measurement on space missions

    NASA Technical Reports Server (NTRS)

    Poppendiek, H. F.; Sabin, C. M.; Meckel, P. T.

    1974-01-01

    The research is reported in applying the axial fluid temperature differential flowmeter to a urine volume measurement system for space missions. The fluid volume measurement system is described along with the prototype equipment package. Flowmeter calibration, electronic signal processing, and typical void volume measurements are also described.

  12. Safe testing nuclear rockets economically

    SciTech Connect

    Howe, S. D.; Travis, B. J.; Zerkle, D. K.

    2002-01-01

    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 RoverMERVA 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.

  13. Focused Rocket-Ejector RBCC Experiments

    NASA Technical Reports Server (NTRS)

    Santoro, Robert J.; Pal, Sibtosh

    2003-01-01

    This document reports the results of additional efforts for the Rocket Based Combined Cycle (RBCC) rocket-ejector mode research work carried out at the Perm State Propulsion Engineering Research Center in support of NASA s technology development efforts for enabling 3rd generation Reusable Launch Vehicles (RLV). The two tasks conducted under this program build on earlier NASA MSFC funded research program on rocket ejector investigations. The first task continued a systematic investigation of the improvements provided by a gaseous hydrogen (GHz)/oxygen (GO2) twin thruster RBCC rocket ejector system over a single rocket system. In a similar vein, the second task continued investigations into 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. For the GH2/GO2 propellant rocket ejector experiments, high frequency measurements of the pressure field within the system were also made to understand the unsteady behavior of the flowfield.

  14. High altitude chemically reacting gas particle mixtures. Volume 2: Program manual for RAMP2. [rocket nozzle and orbital plume flow fields

    NASA Technical Reports Server (NTRS)

    Smith, S. D.

    1984-01-01

    All of the elements used in the Reacting and Multi-Phase (RAMP2) computer code are described in detail. The code can be used to model the dominant phenomena which affect the prediction of liquid and solid rocket nozzle and orbital plume flow fields.

  15. Liquid rocket engine injectors

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  16. Sirius-5 experimental rocket

    NASA Astrophysics Data System (ADS)

    Kerstein, A.; Omersel, P.; Goljuf, L.; Zidaric, M.

    1981-09-01

    After giving a historical account of multistage rocket development in Yugoslavia, a status report is presented for the three-stage Sirius-5 program. The rocket is composed of: (1) a solid-propellant first stage, consisting of a cluster of eight standard motors yielding 220 kN thrust for 1.3 sec; (2) a mixed amines/inhibited red fuming nitric acid, bipropellant second stage generating 50 kN thrust; and (3) a third stage of the same design as the second but with only 62 kg of fuel, by contrast to 168 kg. Among the design principles adhered to are: minimization of the number of components, conservative design margins, and specifications for key subsystems based on demonstration programs. The primary use of this system is in amateur rocketry, being able to carry a 20 kg payload to 150 km.

  17. A miniature solid propellant rocket motor

    SciTech Connect

    Grubelich, M.C.; Hagan, M.; Mulligan, E.

    1997-08-01

    A miniature solid-propellant rocket motor has been developed to impart a specific motion to an object deployed in space. This rocket motor effectively eliminated the need for a cold-gas thruster system or mechanical spin-up system. A low-energy igniter, an XMC4397, employing a semiconductor bridge was used to ignite the rocket motor. The rocket motor was ground-tested in a vacuum tank to verify predicted space performance and successfully flown in a Sandia National Laboratories flight vehicle program.

  18. Bleed cycle propellant pumping in a gas-core nuclear rocket engine system

    NASA Technical Reports Server (NTRS)

    Kascak, A. F.; Easley, A. J.

    1972-01-01

    The performance of ideal and real staged primary propellant pumps and bleed-powered turbines was calculated for gas-core nuclear rocket engines over a range of operating pressures from 500 to 5000 atm. This study showed that for a required engine operating pressure of 1000 atm the pump work was about 0.8 hp/(lb/sec), the specific impulse penalty resulting from the turbine propellant bleed flow as about 10 percent; and the heat required to preheat the propellant was about 7.8 MN/(lb/sec). For a specific impulse above 2400 sec, there is an excess of energy available in the moderator due to the gamma and neutron heating that occurs there. Possible alternative pumping cycles are the Rankine or Brayton cycles.

  19. Calculation of Dynamic Loads Due to Random Vibration Environments in Rocket Engine Systems

    NASA Technical Reports Server (NTRS)

    Christensen, Eric R.; Brown, Andrew M.; Frady, Greg P.

    2007-01-01

    An important part of rocket engine design is the calculation of random dynamic loads resulting from internal engine "self-induced" sources. These loads are random in nature and can greatly influence the weight of many engine components. Several methodologies for calculating random loads are discussed and then compared to test results using a dynamic testbed consisting of a 60K thrust engine. The engine was tested in a free-free condition with known random force inputs from shakers attached to three locations near the main noise sources on the engine. Accelerations and strains were measured at several critical locations on the engines and then compared to the analytical results using two different random response methodologies.

  20. Microfabricated Liquid Rocket Motors

    NASA Technical Reports Server (NTRS)

    Epstein, Alan H.; Joppin, C.; Kerrebrock, J. L.; Schneider, Steven J. (Technical Monitor)

    2003-01-01

    Under NASA Glenn Research Center sponsorship, MIT has developed the concept of micromachined, bipropellant, liquid rocket engines. This is potentially a breakthrough technology changing the cost-performance tradeoffs for small propulsion systems, enabling new applications, and redefining the meaning of the term low-cost-access-to-space. With this NASA support, a liquid-cooled, gaseous propellant version of the thrust chamber and nozzle was designed, built, and tested as a first step. DARPA is currently funding MIT to demonstrate turbopumps and controls. The work performed herein was the second year of a proposed three-year effort to develop the technology and demonstrate very high power density, regeneratively cooled, liquid bipropellant rocket engine thrust chamber and nozzles. When combined with the DARPA turbopumps and controls, this work would enable the design and demonstration of a complete rocket propulsion system. The original MIT-NASA concept used liquid oxygen-ethanol propellants. The military applications important to DARPA imply that storable liquid propellants are needed. Thus, MIT examined various storable propellant combinations including N2O4 and hydrazine, and H2O2 and various hydrocarbons. The latter are preferred since they do not have the toxicity of N2O4 and hydrazine. In reflection of the newfound interest in H2O2, it is once again in production and available commercially. A critical issue for the microrocket engine concept is cooling of the walls in a regenerative design. This is even more important at microscale than for large engines due to cube-square scaling considerations. Furthermore, the coolant behavior of rocket propellants has not been characterized at microscale. Therefore, MIT designed and constructed an apparatus expressly for this purpose. The report details measurements of two candidate microrocket fuels, JP-7 and JP-10.

  1. The Detector and Readout Systems of the Micro-X High Resolution Microcalorimeter X-Ray Imaging Rocket

    NASA Astrophysics Data System (ADS)

    Wikus, P.; Doriese, W. B.; Eckart, M. E.; Adams, J. S.; Bandler, S. R.; Brekosky, R. P.; Chervenak, J. A.; Ewin, A. J.; Figueroa-Feliciano, E.; Finkbeiner, F. M.; Galeazzi, M.; Hilton, G.; Irwin, K. D.; Kelley, R. L.; Kilbourne, C. A.; Leman, S. W.; McCammon, D.; Porter, F. S.; Reintsema, C. D.; Rutherford, J. M.; Trowbridge, S. N.

    2009-12-01

    The Micro-X sounding rocket experiment will deploy an imaging transition-edge-sensor (TES) microcalorimeter spectrometer to observe astrophysical sources in the 0.2-3.0 keV band. The instrument has been designed at a systems level, and the first items of flight hardware are presently being built. In the first flight, planned for January 2011, the spectrometer will observe a recently discovered Silicon knot in the Puppis-A supernova remnant. Here we describe the design of the Micro-X science instrument, focusing on the instrument's detector and detector assembly. The current design of the 2-dimensional spectrometer array contains 128 close-packed pixels with a pitch of 600 μm. The conically approximated Wolter-1 mirror will map each of these pixels to a 0.95 arcmin region on the sky; the field of view will be 11.4 arcmin. Targeted energy resolution of the TESs is about 2 eV over the full observing band. A SQUID time-division multiplexer (TDM) will read out the array. The detector time constants will be engineered to approximately 2 ms to match the TDM, which samples each pixel at 32.6 kHz, limited only by the telemetry system of the rocket. The detector array and two SQUID stages of the TDM readout system are accommodated in a lightweight Mg enclosure, which is mounted to the 50 mK stage of an adiabatic demagnetization refrigerator. A third SQUID amplification stage is located on the 1.6 K liquid He stage of the cryostat. An on-board 55-Fe source will fluoresce a Ca target, providing 3.69 and 4.01 keV calibration lines that will not interfere with the scientifically interesting energy band.

  2. ISRO's solid rocket motors

    NASA Astrophysics Data System (ADS)

    Nagappa, R.; Kurup, M. R.; Muthunayagam, A. E.

    1989-08-01

    Solid rocket motors have been the mainstay of ISRO's sounding rockets and the first generation satellite launch vehicles. For the new launch vehicle under development also, the solid rocket motors contribute significantly to the vehicle's total propulsive power. The rocket motors in use and under development have been developed for a variety of applications and range in size from 30 mm dia employing 450 g of solid propellant—employed for providing a spin to the apogee motors—to the giant 2.8 m dia motor employing nearly 130 tonnes of solid propellant. The initial development, undertaken in 1967 was of small calibre motor of 75 mm dia using a double base charge. The development was essentially to understand the technological elements. Extruded aluminium tubes were used as a rocket motor casing. The fore and aft closures were machined from aluminium rods. The grain was a seven-pointed star with an enlargement of the port at the aft end and was charged into the chamber using a polyester resin system. The nozzle was a metallic heat sink type with graphite throat insert. The motor was ignited with a black powder charge and fired for 2.0 s. Subsequent to this, further developmental activities were undertaken using PVC plastisol based propellants. A class of sounding rockets ranging from 125 to 560 mm calibre were realized. These rocket motors employed improved designs and had delivered lsp ranging from 2060 to 2256 Ns/kg. Case bonding could not be adopted due to the higher cure temperatures of the plastisol propellants but improvements were made in the grain charging techniques and in the design of the igniters and the nozzle. Ablative nozzles based on asbestos phenolic and silica phenolic with graphite inserts were used. For the larger calibre rocket motors, the lsp could be improved by metallic additives. In the early 1970s designs were evolved for larger and more efficient motors. A series of 4 motors for the country's first satellite launch vehicle SLV-3 were

  3. General purpose rocket furnace

    NASA Technical Reports Server (NTRS)

    Aldrich, B. R.; Whitt, W. D. (Inventor)

    1979-01-01

    A multipurpose furnace for space vehicles used for material processing experiments in an outer space environment is described. The furnace contains three separate cavities designed to process samples of the widest possible range of materials and thermal requirements. Each cavity contains three heating elements capable of independent function under the direction of an automatic and programmable control system. A heat removable mechanism is also provided for each cavity which operates in conjunction with the control system for establishing an isothermally heated cavity or a wide range of thermal gradients and cool down rates. A monitoring system compatible with the rocket telemetry provides furnace performance and sample growth rate data throughout the processing cycle.

  4. Rocket + Science = Dialogue

    NASA Technical Reports Server (NTRS)

    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.

  5. Rocket borne instrument to measure electric fields inside electrified clouds

    NASA Technical Reports Server (NTRS)

    Ruhnke, L. H.

    1973-01-01

    Simple electric field measuring system is mounted on small rocket and consists of two voltage probes, one extending from nose and other on tail fin. Electric field through which rocket passes is determined by potential difference between probes.

  6. CIRMIS Data system. Volume 2. Program listings

    SciTech Connect

    Friedrichs, D.R.

    1980-01-01

    The Assessment of Effectiveness of Geologic Isolation Systems (AEGIS) Program is developing and applying the methodology for assessing the far-field, long-term post-closure safety of deep geologic nuclear waste repositories. AEGIS is being performed by Pacific Northwest Laboratory (PNL) under contract with the Office of Nuclear Waste Isolation (OWNI) for the Department of Energy (DOE). One task within AEGIS is the development of methodology for analysis of the consequences (water pathway) from loss of repository containment as defined by various release scenarios. Analysis of the long-term, far-field consequences of release scenarios requires the application of numerical codes which simulate the hydrologic systems, model the transport of released radionuclides through the hydrologic systems, model the transport of released radionuclides through the hydrologic systems to the biosphere, and, where applicable, assess the radiological dose to humans. The various input parameters required in the analysis are compiled in data systems. The data are organized and prepared by various input subroutines for utilization by the hydraulic and transport codes. The hydrologic models simulate the groundwater flow systems and provide water flow directions, rates, and velocities as inputs to the transport models. Outputs from the transport models are basically graphs of radionuclide concentration in the groundwater plotted against time. After dilution in the receiving surface-water body (e.g., lake, river, bay), these data are the input source terms for the dose models, if dose assessments are required.The dose models calculate radiation dose to individuals and populations. CIRMIS (Comprehensive Information Retrieval and Model Input Sequence) Data System is a storage and retrieval system for model input and output data, including graphical interpretation and display. This is the second of four volumes of the description of the CIRMIS Data System.

  7. Automated Rocket Propulsion Test Management

    NASA Technical Reports Server (NTRS)

    Walters, Ian; Nelson, Cheryl; Jones, Helene

    2007-01-01

    The Rocket Propulsion Test-Automated Management System provides a central location for managing activities associated with Rocket Propulsion Test Management Board, National Rocket Propulsion Test Alliance, and the Senior Steering Group business management activities. A set of authorized users, both on-site and off-site with regard to Stennis Space Center (SSC), can access the system through a Web interface. Web-based forms are used for user input with generation and electronic distribution of reports easily accessible. Major functions managed by this software include meeting agenda management, meeting minutes, action requests, action items, directives, and recommendations. Additional functions include electronic review, approval, and signatures. A repository/library of documents is available for users, and all items are tracked in the system by unique identification numbers and status (open, closed, percent complete, etc.). The system also provides queries and version control for input of all items.

  8. A new large-volume multianvil system

    NASA Astrophysics Data System (ADS)

    Frost, D. J.; Poe, B. T.; Trønnes, R. G.; Liebske, C.; Duba, A.; Rubie, D. C.

    2004-06-01

    A scaled-up version of the 6-8 Kwai-type multianvil apparatus has been developed at the Bayerisches Geoinstitut for operation over ranges of pressure and temperature attainable in conventional systems but with much larger sample volumes. This split-cylinder multianvil system is used with a hydraulic press that can generate loads of up to 5000 t (50 MN). The six tool-steel outer-anvils define a cubic cavity of 100 mm edge-length in which eight 54 mm tungsten carbide cubic inner-anvils are compressed. Experiments are performed using Cr 2O 3-doped MgO octahedra and pyrophyllite gaskets. Pressure calibrations at room temperature and high temperature have been performed with 14/8, 18/8, 18/11, 25/17 and 25/15 OEL/TEL (octahedral edge-length/anvil truncation edge-length, in millimetre) configurations. All configurations tested reach a limiting plateau where the sample-pressure no longer increases with applied load. Calibrations with different configurations show that greater sample-pressure efficiency can be achieved by increasing the OEL/TEL ratio. With the 18/8 configuration the GaP transition is reached at a load of 2500 t whereas using the 14/8 assembly this pressure cannot be reached even at substantially higher loads. With an applied load of 2000 t the 18/8 can produce MgSiO 3 perovskite at 1900 °C with a sample volume of ˜20 mm 3, compared with <3 mm 3 in conventional multianvil systems at the same conditions. The large octahedron size and use of a stepped LaCrO 3 heater also results in significantly lower thermal gradients over the sample.

  9. High altitude chemically reacting gas particle mixtures. Volume 1: A theoretical analysis and development of the numerical solution. [rocket nozzle and orbital plume flow fields

    NASA Technical Reports Server (NTRS)

    Smith, S. D.

    1984-01-01

    The overall contractual effort and the theory and numerical solution for the Reacting and Multi-Phase (RAMP2) computer code are described. The code can be used to model the dominant phenomena which affect the prediction of liquid and solid rocket nozzle and orbital plume flow fields. Fundamental equations for steady flow of reacting gas-particle mixtures, method of characteristics, mesh point construction, and numerical integration of the conservation equations are considered herein.

  10. Transient characteristics of rocket turbopumps

    NASA Astrophysics Data System (ADS)

    Shimura, Takashi; Watanabe, Mitsuo

    Transient characteristics of high-speed, high-pressure cryogenic turbopumps for rocket engines were examined experimentally by starting and stopping several rocket turbopumps rapidly by gas turbine drive. Aspects of transient characteristics studied were instantaneous pump head coefficient and pump efficiency, cavitation growth in the inducer during rapid acceleration, effects of the starting mode related to the propulsion system mission, and transient behavior of a self-balancing type axial thrust balancing system. Based on the test results, the feasibility of definite starting modes for the LE-5 turbopump was confirmed and the self-balancing type axial thrust balancing system for the LE-7 LOX turbopump was completed.

  11. Plume flowfield analysis of the shuttle primary Reaction Control System (RCS) rocket engine

    NASA Technical Reports Server (NTRS)

    Hueser, J. E.; Brock, F. J.

    1990-01-01

    A solution was generated for the physical properties of the Shuttle RCS 4000 N (900 lb) rocket engine exhaust plume flowfield. The modeled exhaust gas consists of the five most abundant molecular species, H2, N2, H2O, CO, and CO2. The solution is for a bare RCS engine firing into a vacuum; the only additional hardware surface in the flowfield is a cylinder (=engine mount) which coincides with the nozzle lip outer corner at X = 0, extends to the flowfield outer boundary at X = -137 m and is coaxial with the negative symmetry axis. Continuum gas dynamic methods and the Direct Simulation Monte Carlo (DSMC) method were combined in an iterative procedure to produce a selfconsistent solution. Continuum methods were used in the RCS nozzle and in the plume as far as the P = 0.03 breakdown contour; the DSMC method was used downstream of this continuum flow boundary. The DSMC flowfield extends beyond 100 m from the nozzle exit and thus the solution includes the farfield flow properties, but substantial information is developed on lip flow dynamics and thus results are also presented for the flow properties in the vicinity of the nozzle lip.

  12. Simulation of supercritical flows in rocket-motor engines: application to cooling channel and injection system

    NASA Astrophysics Data System (ADS)

    Ribert, G.; Taieb, D.; Petit, X.; Lartigue, G.; Domingo, P.

    2013-03-01

    To address physical modeling of supercritical multicomponent fluid flows, ideal-gas law must be changed to real-gas equation of state (EoS), thermodynamic and transport properties have to incorporate dense fluid corrections, and turbulence modeling has to be reconsidered compared to classical approaches. Real-gas thermodynamic is presently investigated with validation by NIST (National Institute of Standards and Technology) data. Two major issues of Liquid Rocket Engines (LRE) are also presented. The first one is the supercritical fluid flow inside small cooling channels. In a context of LRE, a strong heat flux coming from the combustion chamber (locally Φ ≈ 80 MW/m2) may lead to very steep density gradients close to the wall. These gradients have to be thermodynamically and numerically captured to properly reproduce in the simulation the mechanism of heat transfer from the wall to the fluid. This is done with a shock-capturing weighted essentially nonoscillatory (WENO) numerical discretization scheme. The second issue is a supercritical fluid injection following experimental conditions [1] in which a trans- or supercritical nitrogen is injected into warm nitrogen. The two-dimensional results show vortex structures with high fluid density detaching from the main jet and persisting in the low-speed region with low fluid density.

  13. Images in the rocket ultraviolet - Star formation and extinction in the M51 system

    NASA Technical Reports Server (NTRS)

    Bohlin, Ralph C.; Cornett, Robert H.; Hill, Jesse K.; O'Connell, Robert W.; Stecher, Theodore P.

    1990-01-01

    UV images of M51 with up to 15 arcsec resolution were obtained by a rocket-borne telescope. The bandwidth was 970 A with maximum response at 2650 A. The two most prominent features of M51 in the UV are the bright central region and a region of intense star formation about 2.5 NE of the nucleus toward the companion NGC 5195. This complex is the source of 20 percent of the total UV flux in M51. The companion is much less prominent in the UV than in optical bands. Spiral arms show much higher contrast in UV and U bands than in the R band. The nuclear region in the UV band shows clumped emission from probable star-forming regions, possibly associated with the inner Lindblad resonance. These regions are probably the UV counterparts of FIR sources discovered in an EW scan across the nucleus. UV/U colors of most bright H II regions with known extinction are consistent with O star spectra reddened by amounts estimated from radio and Balmer line measurements. However, some of the brightest H II regions have redder UV/U colors, probably indicating the presence of cooler stars which contribute in the U band. CO and IR observations indicate that the companion NGC 5195 may be a reddened starburst in spite of its faintness in the UV.

  14. Technology status of a liquid fluorine-hydrazine rocket engine for a planetary spacecraft propulsion system

    NASA Technical Reports Server (NTRS)

    Appel, M. A.; Kruger, G. W.

    1980-01-01

    This paper discusses the current status of a fluorine-hydrazine rocket engine development program. Incorporation of a thin rhenium inner liner successfully eliminated corrosion of the carbon/carbon composite thrust chamber wall experienced during a previous test program. The results of hot-fire tests utilizing reworked and new injectors which provide increased fuel film cooling showed that thrust chamber head-end temperatures could be maintained at an acceptable level. As expected, the accompanying specific impulse performance loss requires optimizing the amount of film cooling to minimize the loss. The efforts to refine the rhenium liner vapor deposition process culminated in a carbon/carbon composite thrust chamber total test duration of 1008 seconds. Tasks presently in process include: (1) fabrication of two carbon/carbon composite thrust chambers incorporating 60:1 expansion ratio nozzles; (2) injector tests to optimize performance and cooling; (3) additional refinements to the rhenium lining process; and (4) fabrication and test of a freestanding rhenium thrust chamber.

  15. Rapid Construction of Hardware-in-the-Loop Simulation and Control System Validation for the THX Rocket

    NASA Astrophysics Data System (ADS)

    Zhang, Y.; Yang, H.; Jiang, Z.; Hu, F.; Zhang, W.

    2015-09-01

    The rapid construction of hardware-in-the-loop simulation(HILS) system and validation of the flight control approach for the TianHang eXperimental(THX) rocket are investigated. Firstly, the six degree of freedom simulation system is accomplished using MATLAB/Simulink, and the simulation models are classified and masked into various blocks by their functions. And then, the integrated design of testing and debugging of the flight control devices and the HILS system construction is proposed based on the dSPACE real-time platform. The test and calibration of the gyroscope and rudder system and the flight control code automatically generation are carried out, and various leve's of HILS are achieved. Finally, the optimal linear quadratic regulator with the integrate part is proposed for the three channels of the flight vehicle. The numerical and HILS results demonstrated the effective performance of the proposed control approach. As the control system design implements only on the MATLAB/Simulink platform, a large of repeated work is overleaped. Therefore, the rapid construction of the HILS is accomplished, and the flight control design and validation period is decreased, which implies the remarkable value in engineering application.

  16. Catalytic Microtube Rocket Igniter

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.; Deans, Matthew C.

    2011-01-01

    Devices that generate both high energy and high temperature are required to ignite reliably the propellant mixtures in combustion chambers like those present in rockets and other combustion systems. This catalytic microtube rocket igniter generates these conditions with a small, catalysis-based torch. While traditional spark plug systems can require anywhere from 50 W to multiple kW of power in different applications, this system has demonstrated ignition at less than 25 W. Reactants are fed to the igniter from the same tanks that feed the reactants to the rest of the rocket or combustion system. While this specific igniter was originally designed for liquid methane and liquid oxygen rockets, it can be easily operated with gaseous propellants or modified for hydrogen use in commercial combustion devices. For the present cryogenic propellant rocket case, the main propellant tanks liquid oxygen and liquid methane, respectively are regulated and split into different systems for the individual stages of the rocket and igniter. As the catalyst requires a gas phase for reaction, either the stored boil-off of the tanks can be used directly or one stream each of fuel and oxidizer can go through a heat exchanger/vaporizer that turns the liquid propellants into a gaseous form. For commercial applications, where the reactants are stored as gases, the system is simplified. The resulting gas-phase streams of fuel and oxidizer are then further divided for the individual components of the igniter. One stream each of the fuel and oxidizer is introduced to a mixing bottle/apparatus where they are mixed to a fuel-rich composition with an O/F mass-based mixture ratio of under 1.0. This premixed flow then feeds into the catalytic microtube device. The total flow is on the order of 0.01 g/s. The microtube device is composed of a pair of sub-millimeter diameter platinum tubes connected only at the outlet so that the two outlet flows are parallel to each other. The tubes are each

  17. Closeup view of the Solid Rocket Booster (SRB) Forward Skirt ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Close-up view of the Solid Rocket Booster (SRB) Forward Skirt sitting on ground support equipment in the Solid Rocket Booster Assembly and Refurbishment Facility at Kennedy Space Center while being prepared for mating with the Frustum-Nose Cap Assembly and the Forward Rocket Motor Segment. The prominent feature in this view is the electrical, data, telemetry and safety systems terminal which connects to the Aft Skirt Assembly systems via the Systems Tunnel that runs the length of the Rocket Motor. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  18. Low-thrust rocket trajectories

    SciTech Connect

    Keaton, P.W.

    1986-01-01

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

  19. Emergency egress fixed rocket package

    NASA Technical Reports Server (NTRS)

    Allen, Margaret A. (Inventor)

    1989-01-01

    A method of effecting the in-flight departure of an astronaut from a shuttle craft, and apparatus is presented. A plurality of removeable compartment covers are provided, behind which rocket assemblies are stowed. To actuate the system, the astronaut pulls off a tab from one of the compartments which exposes a cannister having a lanyard with a hook. The lanyard extends around a spring biased sleeve with a safety lever preventing rocket ignition until the hook is moved by the astronaut. Upward movement of the hook allows the trigger mechanism to actuate the system resulting in the rods projecting out of the hatch. When the lanyard becomes taut, a lanyard elongation detector transmits a signal to the firing mechanisms to fire the rocket.

  20. Low-thrust rocket trajectories

    SciTech Connect

    Keaton, P.W.

    1987-03-01

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

  1. Small rocket research and technology

    NASA Technical Reports Server (NTRS)

    Schneider, Steven; Biaglow, James

    1993-01-01

    Small chemical rockets are used on nearly all space missions. The small rocket program provides propulsion technology for civil and government space systems. Small rocket concepts are developed for systems which encompass reaction control for launch and orbit transfer systems, as well as on-board propulsion for large space systems and earth orbit and planetary spacecraft. Major roles for on-board propulsion include apogee kick, delta-V, de-orbit, drag makeup, final insertions, north-south stationkeeping, orbit change/trim, perigee kick, and reboost. The program encompasses efforts on earth-storable, space storable, and cryogenic propellants. The earth-storable propellants include nitrogen tetroxide (NTO) as an oxidizer with monomethylhydrazine (MMH) or anhydrous hydrazine (AH) as fuels. The space storable propellants include liquid oxygen (LOX) as an oxidizer with hydrazine or hydrocarbons such as liquid methane, ethane, and ethanol as fuels. Cryogenic propellants are LOX or gaseous oxygen (GOX) as oxidizers and liquid or gaseous hydrogen as fuels. Improved performance and lifetime for small chemical rockets are sought through the development of new predictive tools to understand the combustion and flow physics, the introduction of high temperature materials to eliminate fuel film cooling and its associated combustion inefficiency, and improved component designs to optimize performance. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Results indicate that modeling of the injector and combustion process in small rockets needs improvement. High temperature materials require the development of fabrication processes, a durability data base in both laboratory and rocket environments, and basic engineering property data such as strength, creep, fatigue, and work hardening properties at both room and elevated temperature. Promising materials under development include iridium-coated rhenium and a

  2. Thermal Fault Tolerance Analysis of Carbon Fiber Rope Barrier Systems for Use in the Reusable Solid Rocket Motor ( RSRM) Nozzle Joints

    NASA Technical Reports Server (NTRS)

    Clayton, J. Louie; Phelps, Lisa (Technical Monitor)

    2001-01-01

    Carbon Fiber Rope (CFR) thermal barrier systems are being considered for use in several RSRM (Reusable Solid Rocket Motor) nozzle joints as a replacement for the current assembly gap close-out process/design. This study provides for development and test verification of analysis methods used for flow-thermal modeling of a CFR thermal barrier subject to fault conditions such as rope combustion gas blow-by and CFR splice failure. Global model development is based on a 1-D (one dimensional) transient volume filling approach where the flow conditions are calculated as a function of internal 'pipe' and porous media 'Darcy' flow correlations. Combustion gas flow rates are calculated for the CFR on a per-linear inch basis and solved simultaneously with a detailed thermal-gas dynamic model of a local region of gas blow by (or splice fault). Effects of gas compressibility, friction and heat transfer are accounted for the model. Computational Fluid Dynamic (CFD) solutions of the fault regions are used to characterize the local flow field, quantify the amount of free jet spreading and assist in the determination of impingement film coefficients on the nozzle housings. Gas to wall heat transfer is simulated by a large thermal finite element grid of the local structure. The employed numerical technique loosely couples the FE (Finite Element) solution with the gas dynamics solution of the faulted region. All free constants that appear in the governing equations are calibrated by hot fire sub-scale test. The calibrated model is used to make flight predictions using motor aft end environments and timelines. Model results indicate that CFR barrier systems provide a near 'vented joint' style of pressurization. Hypothetical fault conditions considered in this study (blow by, splice defect) are relatively benign in terms of overall heating to nozzle metal housing structures.

  3. Lateral pulse jet control of a direct fire atmospheric rocket using an inertial measurement unit sensor system

    NASA Astrophysics Data System (ADS)

    Jitpraphai, Thanat

    Impact point dispersion of a direct fire rocket can be drastically reduced with a ring of appropriately sized lateral pulse jets coupled to a trajectory tracking flight control system. The system is shown to work well against uncertainty in the form of initial off-axis angular velocity perturbations as well as atmospheric winds. For an example case examined, dispersion was reduced by a factor of one hundred. Dispersion reduction and mean miss distance are strong functions of the number of individual pulse jets, the pulse jet impulse, and the trajectory tracking window size. Proper selection of these parameters for a particular rocket and launcher combination is required to achieve optimum dispersion reduction to the pulse jet control mechanism. For the lateral pulse jet control mechanism that falls into the category of an impulse control mechanism, the trajectory tracking flight control law provides better reduction in dispersion and mean miss distance than the proportional navigation guidance law especially when small number of individual pulse jets is used. Estimation of body frame components of angular velocity and angular acceleration of a rigid body projectile undergoing general three-dimensional motion using linear acceleration measurements is considered. The results are comparable to those obtained from a conventional Inertial Measurement Unit (IMU) that composes of accelerometers and gyroscopes. From the study of the effect of sensor errors to the measurement and the control performance, the sensitivity of the angular rate estimation to the sensor noise is a strong function of the constellation of these three accelerometers. When more than three point measurements are used, the most effective method to fuse data is with one cluster that contains all sensors. In the conventional IMU, the dispersion and miss distance are less sensitive to the errors from accelerometers than to the gyroscopes. The estimation of angular rates plays essential roles in the

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

    NASA Technical Reports Server (NTRS)

    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.

  5. Common Data Acquisition Systems (DAS) Software Development for Rocket Propulsion Test (RPT) Test Facilities - A General Overview

    NASA Technical Reports Server (NTRS)

    Hebert, Phillip W., Sr.; Hughes, Mark S.; Davis, Dawn M.; Turowski, Mark P.; Holladay, Wendy T.; Marshall, PeggL.; Duncan, Michael E.; Morris, Jon A.; Franzl, Richard W.

    2012-01-01

    The advent of the commercial space launch industry and NASA's more recent resumption of operation of Stennis Space Center's large test facilities after thirty years of contractor control resulted in a need for a non-proprietary data acquisition system (DAS) software to support government and commercial testing. The software is designed for modularity and adaptability to minimize the software development effort for current and future data systems. An additional benefit of the software's architecture is its ability to easily migrate to other testing facilities thus providing future commonality across Stennis. Adapting the software to other Rocket Propulsion Test (RPT) Centers such as MSFC, White Sands, and Plumbrook Station would provide additional commonality and help reduce testing costs for NASA. Ultimately, the software provides the government with unlimited rights and guarantees privacy of data to commercial entities. The project engaged all RPT Centers and NASA's Independent Verification & Validation facility to enhance product quality. The design consists of a translation layer which provides the transparency of the software application layers to underlying hardware regardless of test facility location and a flexible and easily accessible database. This presentation addresses system technical design, issues encountered, and the status of Stennis' development and deployment.

  6. Liquid atomization by coaxial rocket injectors

    NASA Technical Reports Server (NTRS)

    Sankar, S. V.; Brena De La Rosa, A.; Isakovic, A.; Bachalo, W. D.

    1991-01-01

    The atomization characteristics of a scaled-down version of a coaxial rocket injector was investigated using a phase Doppler particle analyzer (PDPA). The injector was operated in the conventional mode with liquid being injected through its inner orifice and gas being injected through its outer annulus. The shearing action occurring at the liquid-gas interface causes the liquid jet to atomize. In this study, two different liquid-air systems, namely a water-air system and a liquid nitrogen-gaseous nitrogen system, were chosen for detailed investigation. This paper discusses the performance characteristics of the coaxial injector under different flow and geometric conditions. Specifically, the effects of injection gas pressure and the injector cavity size on variables such as the mean particle diameter, Sauter mean diameter, number density, volume flux, and velocity have been presented.

  7. Rotating and positive-displacement pumps for low-thrust rocket engines. Volume 1: Pump Evaluation and design. [of centrifugal pumps

    NASA Technical Reports Server (NTRS)

    Macgregor, C.; Csomor, A.

    1974-01-01

    Rotating and positive displacement pumps of various types were studied for pumping liquid fluorine for low-thrust, high-performance rocket engines. Included in the analysis were: centrifugal, pitot, Barske, Tesla, drag, gear, vane, axial piston, radial piston, diaphragm, and helirotor pump concepts. The centrifugal pump and the gear pump were selected and these were carried through detailed design and fabrication. Mechanical difficulties were encountered with the gear pump during the preliminary tests in Freon-12. Further testing and development was therefore limited to the centrifugal pump. Tests on the centrifugal pump were conducted in Freon-12 to determine the hydrodynamic performance and in liquid fluorine to demonstrate chemical compatibility.

  8. High altitude chemically reacting gas particle mixtures. Volume 3: Computer code user's and applications manual. [rocket nozzle and orbital plume flow fields

    NASA Technical Reports Server (NTRS)

    Smith, S. D.

    1984-01-01

    A users manual for the RAMP2 computer code is provided. The RAMP2 code can be used to model the dominant phenomena which affect the prediction of liquid and solid rocket nozzle and orbital plume flow fields. The general structure and operation of RAMP2 are discussed. A user input/output guide for the modified TRAN72 computer code and the RAMP2F code is given. The application and use of the BLIMPJ module are considered. Sample problems involving the space shuttle main engine and motor are included.

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

    NASA Technical Reports Server (NTRS)

    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.

  10. Integrated approach for hybrid rocket technology development

    NASA Astrophysics Data System (ADS)

    Barato, Francesco; Bellomo, Nicolas; Pavarin, Daniele

    2016-11-01

    Hybrid rocket motors tend generally to be simple from a mechanical point of view but difficult to optimize because of their complex and still not well understood cross-coupled physics. This paper addresses the previous issue presenting the integrated approach established at University of Padua to develop hybrid rocket based systems. The methodology tightly combines together system analysis and design, numerical modeling from elementary to sophisticated CFD, and experimental testing done with incremental philosophy. As an example of the approach, the paper presents the experience done in the successful development of a hybrid rocket booster designed for rocket assisted take off operations. It is thought that following the proposed approach and selecting carefully the most promising applications it is possible to finally exploit the major advantages of hybrid rocket motors as safety, simplicity, low cost and reliability.

  11. The propulsion system for the “Ludion” one-man hopper: An Anglo-French rocket engine cooperation 40 years ago

    NASA Astrophysics Data System (ADS)

    Rothmund, Christophe

    2010-07-01

    The Ludion was a French concept for a military one-man rocket-powered "hopper" aimed at giving troopers the ability to cross rivers or trenches. Sud Aviation was awarded in 1965 the development contract together with Bertin (for the jet pumps) and SEPR was to develop the rocket propulsion system. The propulsion system selected for the "Ludion" was based on the British extensive experience in handling and using isopropyl nitrate as a way of obtaining large amounts of gas able to drive jet engines at start. The idea was to use such a fluid as monopropellant for the propulsion system's gas generator. SEPR integrated this system with pipes and Bertin ejector nozzles and developed a thermally insulated generator able to fulfil all requirements. The main challenges laid in the successful integration of all components and the management of a multi-partners programme and in the design of a highly efficient thermal insulation protecting both generator and pilot from the intense heat generated by the thermal decomposition of isopropyl nitrate (used as a propellant). The first ground test of a Ludion with its rocket propulsion system took place on January 24, 1968, on the test airfield of Villaroche. After 59 tethered and five free flights, the programme was abandoned. It was deemed to be too impractical due to its size, the noise generated and changes in tactical considerations in favour of helicopters and amphibious transports. The first prototype is now displayed at the Paris Le Bourget Air & Space Museum.

  12. Advanced vehicle systems assessment. Volume 2: Subsystems assessment

    NASA Technical Reports Server (NTRS)

    Hardy, K.

    1985-01-01

    Volume 2 (Subsystems Assessment) is part of a five-volume report entitled Advanced Vehicle Systems Assessment. Volume 2 presents the projected performance capabilities and cost characteristics of applicable subsystems, considering an additional decade of development. Subsystems of interest include energy storage and conversion devices as well as the necessary powertrain components and vehicle subsystems. Volume 2 also includes updated battery information based on the assessment of an independent battery review board (with the aid of subcontractor reports on advanced battery characteristics).

  13. Consort 1 sounding rocket flight

    NASA Technical Reports Server (NTRS)

    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.

  14. The Advanced Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

    Mitchell, Royce E.

    1992-01-01

    The paper describes the Advanced Solid Rocket Motor (ASRM) that is being developed to replace, in 1997, the Redesigned Solid Rocket Motor which currently boosts the Space Shuttle. The ASRM will contain features to improve motor safety (fewer potential leak paths, improved seal materials, stronger case material, and fewer nozzle and case joints), an improved ignition system using through-bulkhead initiators, and highly reproducible manufacturing and inspection techniques with a large number of automated procedures. The ASRM will be able to deliver 12,000 lbs greater payloads to any given orbit of the Shuttle. There are also environmental improvements, realized by waste propellant recovery.

  15. Design of a dual port volume measuring system

    SciTech Connect

    Klevgard, P.A.

    1990-09-01

    A volume measuring system is described which uses the ideal gas law and pressure measurements to determine an unknown vessel's volume when a gas expands into that vessel from a known volume. The design, the engineering principles, the calibration, and the accuracy of this computer-controlled system are all discussed. A set of electrical and mechanical drawings of the system is included. 3 refs., 6 figs.

  16. Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

    NASA Technical Reports Server (NTRS)

    Schoenfeld, Michael

    2009-01-01

    A detailed description of the Nuclear Thermal Rocket Element Environmental Simulator (NTREES) is presented. The contents include: 1) Design Requirements; 2) NTREES Layout; 3) Data Acquisition and Control System Schematics; 4) NTREES System Schematic; and 5) NTREES Setup.

  17. Replacement of chemical rocket launchers by beamed energy propulsion.

    PubMed

    Fukunari, Masafumi; Arnault, Anthony; Yamaguchi, Toshikazu; Komurasaki, Kimiya

    2014-11-01

    Microwave Rocket is a beamed energy propulsion system that is expected to reach space at drastically lower cost. This cost reduction is estimated by replacing the first-stage engine and solid rocket boosters of the Japanese H-IIB rocket with Microwave Rocket, using a recently developed thrust model in which thrust is generated through repetitively pulsed microwave detonation with a reed-valve air-breathing system. Results show that Microwave Rocket trajectory, in terms of velocity versus altitude, can be designed similarly to the current H-IIB first stage trajectory. Moreover, the payload ratio can be increased by 450%, resulting in launch-cost reduction of 74%. PMID:25402933

  18. Replacement of chemical rocket launchers by beamed energy propulsion.

    PubMed

    Fukunari, Masafumi; Arnault, Anthony; Yamaguchi, Toshikazu; Komurasaki, Kimiya

    2014-11-01

    Microwave Rocket is a beamed energy propulsion system that is expected to reach space at drastically lower cost. This cost reduction is estimated by replacing the first-stage engine and solid rocket boosters of the Japanese H-IIB rocket with Microwave Rocket, using a recently developed thrust model in which thrust is generated through repetitively pulsed microwave detonation with a reed-valve air-breathing system. Results show that Microwave Rocket trajectory, in terms of velocity versus altitude, can be designed similarly to the current H-IIB first stage trajectory. Moreover, the payload ratio can be increased by 450%, resulting in launch-cost reduction of 74%.

  19. Low gravity investigations in suborbital rockets

    NASA Technical Reports Server (NTRS)

    Wessling, Francis C.; Lundquist, Charles A.

    1990-01-01

    Two series of suborbital rocket missions are outlined which are intended to support materials and biotechnology investigations under microgravity conditions and enhance commercial rocket activity. The Consort series of missions employs the two-stage Starfire I rocket and recovery systems as well as a payload of three sealed or vented cylindrical sections. The Consort 1 and 2 missions are described which successfully supported six classes of experiments each. The Joust program is the second series of rocket missions, and the Prospector rocket is employed to provide comparable payload masses with twice as much microgravity time as the Consort series. The Joust and Consort missions provide 6-8 and 13-15 mins, respectively, of microgravity flight to support such experiments as polymer processing, scientific apparatus testing, and electrodeposition.

  20. Computational modeling of nuclear thermal rockets

    NASA Technical Reports Server (NTRS)

    Peery, Steven D.

    1993-01-01

    The topics are presented in viewgraph form and include the following: rocket engine transient simulation (ROCETS) system; ROCETS performance simulations composed of integrated component models; ROCETS system architecture significant features; ROCETS engineering nuclear thermal rocket (NTR) modules; ROCETS system easily adapts Fortran engineering modules; ROCETS NTR reactor module; ROCETS NTR turbomachinery module; detailed reactor analysis; predicted reactor power profiles; turbine bypass impact on system; and ROCETS NTR engine simulation summary.

  1. Advanced vehicle systems assessment. Volume 3: Systems assessment

    NASA Technical Reports Server (NTRS)

    Hardy, K.

    1985-01-01

    The systems analyses integrate the advanced component and vehicle characteristics into conceptual vehicles with identical performance (for a given application) and evaluates the vehicles in typical use patterns. Initial and life-cycle costs are estimated and compared to conventional reference vehicles with comparable technological advances, assuming the vehicles will be in competition in the early 1990s. Electric vans, commuter vehicles, and full-size vehicles, in addition to electric/heat-engine hybrid and fuel-cell powered vehicles, are addressed in terms of performance and economics. System and subsystem recommendations for vans and two-passenger commuter vehicles are based on the economic analyses in this volume.

  2. Mars power system concept definition study. Volume 2: Appendices

    NASA Technical Reports Server (NTRS)

    Littman, Franklin D.

    1994-01-01

    This report documents the work performed by Rockwell International's Rocketdyne Division on NASA Contract No. NAS3-25808 (Task Order No. 16) entitled 'Mars Power System Definition Study'. This work was performed for NASA's Lewis Research Center (LeRC). The report is divided into two volumes as follows: Volume 1 - Study Results; and Volume 2 - Appendices. The results of the power system characterization studies, operations studies, and technology evaluations are summarized in Volume 1. The appendices include complete, standalone technology development plans for each candidate power system that was investigated.

  3. High Speed Photographic Studies Of Rocket Engine Combustion

    NASA Astrophysics Data System (ADS)

    Uyemura, Tsuneyoshi; Ozono, Shigeo; Mizunuma, Toshio; Yamamoto, Yoshitaka; Kikusato, Yutaka; Eiraku, Masamitsu; Uchida, Yubu

    1983-03-01

    The high speed cameras were used to develop the new sounding rocket motor and to check the safety operation system. The new rocket motor was designed as a single stage rocket and its power was greater than the multi-stage K-9M rocket motor. The test combustion of this new type rocket engine was photographed by the high speed cameras to analyze the burning process. At the outside of rocket chamber, the cable which connect the detector of an engine nozzle with the telemeter system was fixed. To check the thero.,a1 influences of combustion flame to the cable, the thermo-tapes and high speed cameras were used Safety operation system was tested and photographed with high speed cameras using a S0-1510 model rocket.

  4. Theoretical and Experimental Analysis of the Physics of Water Rockets

    ERIC Educational Resources Information Center

    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…

  5. Commercial Development Suborbital Rocket Program

    NASA Technical Reports Server (NTRS)

    1993-01-01

    The enclosed report provides information on the sixth flight of the Consort suborbital rocket series. Consort 6 is currently scheduled for launch on February 19, 1993, with lift off at 11:00 a.m., Mountain Time. It will carry seven materials and biotechnology experiments, two accelerometer systems, a controller and battery packs in a module nearly 12 feet tall and weighing approximately 1,004 pounds. Consort 6 will reach an apogee of approximately 200 miles providing about 7 minutes of microgravity time. The entire mission, from launch to touchdown, is expected to last approximately 15 minutes. The Consort series is part of a unique suborbital rocket launch services program conducted by the Office of Advanced Concepts and Technology (OACT) in conjunction with its Centers for the Commercial Development of Space (CCDS). This service is managed through the Consortium for Materials Development in Space (CMDS), a CCDS based University of Alabama in Huntsville (UAH). at the This suborbital rocket program provides CCDS investigators with a microgravity environment to achieve commercial development objectives, or to test developmental hardware or techniques in preparation for orbital flights or additional follow-on work. Rocket and launch services for Consort 6, including use of the Starfire 1 launch vehicle, are provided by EER Systems Corporation. Integration of the payload into Starfire 1 will be handled by McDonnell Douglas Space Systems Company.

  6. Rhenium Rocket Manufacturing Technology

    NASA Technical Reports Server (NTRS)

    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. Collaborative Sounding Rocket launch in Alaska and Development of Hybrid Rockets

    NASA Astrophysics Data System (ADS)

    Ono, Tomohisa; Tsutsumi, Akimasa; Ito, Toshiyuki; Kan, Yuji; Tohyama, Fumio; Nakashino, Kyouichi; Hawkins, Joseph

    Tokai University student rocket project (TSRP) was established in 1995 for a purpose of the space science and engineering hands-on education, consisting of two space programs; the one is sounding rocket experiment collaboration with University of Alaska Fairbanks and the other is development and launch of small hybrid rockets. In January of 2000 and March 2002, two collaborative sounding rockets were successfully launched at Poker Flat Research Range in Alaska. In 2001, the first Tokai hybrid rocket was successfully launched at Alaska. After that, 11 hybrid rockets were launched to the level of 180-1,000 m high at Hokkaido and Akita in Japan. Currently, Tokai students design and build all parts of the rockets. In addition, they are running the organization and development of the project under the tight budget control. This program has proven to be very effective in providing students with practical, real-engineering design experience and this program also allows students to participate in all phases of a sounding rocket mission. Also students learn scientific, engineering subjects, public affairs and system management through experiences of cooperative teamwork. In this report, we summarize the TSRP's hybrid rocket program and discuss the effectiveness of the program in terms of educational aspects.

  8. Simultaneous measurements of auroral particles and electric currents by a rocket-borne instrument system - Introductory remarks

    NASA Technical Reports Server (NTRS)

    Anderson, H. R.; Cloutier, P. A.

    1975-01-01

    A rocket-borne experiment package has been designed to obtain simultaneous in situ measurements of the pitch angle distributions and energy spectra of primary auroral particles, the flux of neutral hydrogen at auroral energies, the electric currents flowing in the vicinity of the auroral arc as determined from vector magnetic data, and the modulation of precipitating electrons in the frequency range 0.5-10 MHz. The experiment package was launched by a Nike-Tomahawk rocket from Poker Flat, Alaska, at 0722 UT on Feb. 25, 1972, over a bright auroral band. This paper is intended to serve as an introduction to the detailed discussion of results given in the companion papers. As such it includes a brief review of the general problem, a discussion of the rocket instrumentation, a delineation of the auroral and geomagnetic conditions at the time of launch, and comments on the overall payload performance.

  9. Determination of gas volume trapped in a closed fluid system

    NASA Technical Reports Server (NTRS)

    Hunter, W. F.; Jolley, J. E.

    1971-01-01

    Technique involves extracting known volume of fluid and measuring system before and after extraction, volume of entrapped gas is then computed. Formula derived from ideal gas laws is basis of this method. Technique is applicable to thermodynamic cycles and hydraulic systems.

  10. Advanced Transportation System Studies. Technical Area 3: Alternate Propulsion Subsystem Concepts. Volume 1; Executive Summary

    NASA Technical Reports Server (NTRS)

    Levack, Daniel J. H.

    2000-01-01

    The Alternate Propulsion Subsystem Concepts contract had seven tasks defined that are reported under this contract deliverable. The tasks were: FAA Restart Study, J-2S Restart Study, Propulsion Database Development. SSME Upper Stage Use. CERs for Liquid Propellant Rocket Engines. Advanced Low Cost Engines, and Tripropellant Comparison Study. The two restart studies, F-1A and J-2S, generated program plans for restarting production of each engine. Special emphasis was placed on determining changes to individual parts due to obsolete materials, changes in OSHA and environmental concerns, new processes available, and any configuration changes to the engines. The Propulsion Database Development task developed a database structure and format which is easy to use and modify while also being comprehensive in the level of detail available. The database structure included extensive engine information and allows for parametric data generation for conceptual engine concepts. The SSME Upper Stage Use task examined the changes needed or desirable to use the SSME as an upper stage engine both in a second stage and in a translunar injection stage. The CERs for Liquid Engines task developed qualitative parametric cost estimating relationships at the engine and major subassembly level for estimating development and production costs of chemical propulsion liquid rocket engines. The Advanced Low Cost Engines task examined propulsion systems for SSTO applications including engine concept definition, mission analysis. trade studies. operating point selection, turbomachinery alternatives, life cycle cost, weight definition. and point design conceptual drawings and component design. The task concentrated on bipropellant engines, but also examined tripropellant engines. The Tripropellant Comparison Study task provided an unambiguous comparison among various tripropellant implementation approaches and cycle choices, and then compared them to similarly designed bipropellant engines in the

  11. NASA Rocket Propulsion Test Replacement Effort for Oxygen System Cleaner - Hydrochlorofluorocarbon (HCFC) 225

    NASA Technical Reports Server (NTRS)

    DeWitt Burns, H.; Mitchell, Mark A.; Lowrey, Nikki M.; Farner, Bruce R.; Ross, H. Richard

    2014-01-01

    Gaseous and liquid oxygen are extremely reactive materials used in bipropellant propulsion systems. Both flight and ground oxygen systems require a high level of cleanliness to support engine performance, testing, and prevent mishaps. Solvents used to clean and verify the cleanliness of oxygen systems and supporting test hardware must be compatible with the system's materials of construction and effective at removing or reducing expected contaminants to an acceptable level. This paper will define the philosophy and test approach used for evaluating replacement solvents for the current Marshall Space Flight Center/Stennis Space Center baseline HCFC-225 material that will no longer be available for purchase after 2014. MSFC/SSC applications in cleaning / sampling oxygen propulsion components, support equipment, and test system were reviewed then candidate replacement cleaners and test methods selected. All of these factors as well as testing results will be discussed.

  12. NASA, ATK Successfully Test Solid Rocket Motor

    NASA Video Gallery

    With a loud roar and mighty column of flame, NASA and ATK Aerospace Systems successfully completed a two-minute, full-scale test of the largest and most powerful solid rocket motor designed for fli...

  13. Pattern classification approach to rocket engine diagnostics

    SciTech Connect

    Tulpule, S.

    1989-01-01

    This paper presents a systems level approach to integrate state-of-the-art rocket engine technology with advanced computational techniques to develop an integrated diagnostic system (IDS) for future rocket propulsion systems. The key feature of this IDS is the use of advanced diagnostic algorithms for failure detection as opposed to the current practice of redline-based failure detection methods. The paper presents a top-down analysis of rocket engine diagnostic requirements, rocket engine operation, applicable diagnostic algorithms, and algorithm design techniques, which serve as a basis for the IDS. The concepts of hierarchical, model-based information processing are described, together with the use uf signal processing, pattern recognition, and artificial intelligence techniques which are an integral part of this diagnostic system. 27 refs.

  14. Nitrous Oxide/Paraffin Hybrid Rocket Engines

    NASA Technical Reports Server (NTRS)

    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.

  15. Rocket having barium release system to create ion clouds in the upper atmosphere

    NASA Technical Reports Server (NTRS)

    Lewis, B. W.; Stokes, C. S.; Smith, E. W.; Murphy, W. J. (Inventor)

    1974-01-01

    A chemical system for releasing a good yield of free barium atoms and barium ions to create ion clouds in the upper atmosphere and interplanetary space for the study of the geophysical properties of the medium is presented.

  16. Satellite communications systems and technology. Volume 2: Site reports

    NASA Technical Reports Server (NTRS)

    Edelson, Burton I. (Editor); Pelton, Joseph N. (Editor); Bostian, Charles W.; Brandon, William T.; Chan, Vincent W. S.; Hager, E. Paul; Helm, Neil R.; Jennings, Raymond D.; Kwan, Robert K.; Mahle, Christoph E.

    1993-01-01

    This is volume 2 of the final report of the NASA/NSF Panel on Satellite Communications Systems and Technology. It consists of the site reports from the panel's visits to satellite communications facilities and laboratories in Europe, Japan, and Russia. The Executive Summary of the panel's final report is published separately. Volume 1, also published separately, consists of the panel's analytical chapters. Information on ordering the Executive Summary and Volume 1 from the National Technical Information Service is included.

  17. Kennedy Space Center's Command and Control System - "Toasters to Rocket Ships"

    NASA Technical Reports Server (NTRS)

    Lougheed, Kirk; Mako, Cheryle

    2011-01-01

    This slide presentation reviews the history of the development of the command and control system at Kennedy Space Center. From a system that could be brought to Florida in the trunk of a car in the 1950's. Including the development of larger and more complex launch vehicles with the Apollo program where human launch controllers managed the launch process with a hardware only system that required a dedicated human interface to perform every function until the Apollo vehicle lifted off from the pad. Through the development of the digital computer that interfaced with ground launch processing systems with the Space Shuttle program. Finally, showing the future control room being developed to control the missions to return to the moon and Mars, which will maximize the use of Commercial-Off-The Shelf (COTS) hardware and software which was standards based and not tied to a single vendor. The system is designed to be flexible and adaptable to support the requirements of future spacecraft and launch vehicles.

  18. Identification of a physically idealized human rated rocket based interplanetary transportation system

    NASA Astrophysics Data System (ADS)

    Ewig, Ralph

    Every system engineering trade study has to address the challenge of eliminating unintentional bias towards one of the available system options. This challenge becomes especially difficult when trading conceptual options, where the amount and fidelity of data available to characterize the options is highly variable. This dissertation introduces the methodology of Physical Idealization as a tool to remove unintentional bias from conceptual trade studies. The premise is that (1) given the options available based on our understanding of physics, and (2) within the set of constraints necessary to define the problem, it is possible to identify the optimal physically idealized solution. This solution can then be used as a benchmark for technology development and real world system implementation. The methodology of Physical Idealization is developed to support a study of Interplanetary Transportation Systems (ITS). The ITS is modeled as consisting of payload, power, and propulsion subsystems, and optimized using a simplified two-dimensional equation of motion set. Both a genetic algorithm and gradient based optimization methods are used in a nested loop process. The presented results illustrate both the strengths and weaknesses associated with using physical idealization in a trade study, showing the methodology to be a useful addition to the system engineer's selection of tools.

  19. Performance of a shaft seal system for the LE-7 rocket engine oxidizer turbopump

    NASA Astrophysics Data System (ADS)

    Oike, Mamoru; Nosaka, Masataka; Kikuchi, Masataka; Watanabe, Yoshiaki

    An experimental study on a rotating-shaft seal system for a high-pressure liquid oxygen (LOX) turbopump has been conducted to develop the LE-7 engine for the Japanese H-II launch vehicle. The LOX turbopump rotating-shaft seal system, which prevents LOX (4.9 MPa) and the high-pressure turbine-drive gas (16.6 MPa, 970 K) from mixing during high-speed operations (18,000 to 20,000 rpm), consists of the following seals: an LOX seal comprising a floating-ring and a wear-ring, a turbine gas seal comprising two floating-rings, and a helium purge seal comprising two segmented circumferential seal-rings. This report describes experimental and observational results concerning the rotating-shaft seal system obtained in the LOX turbopump operations and the seal tests. Based on comparisons between the measurements and the analytical results, sealing characteristics of the seal system are discussed. Inspections of the sealing surfaces after the engine firing tests demonstrated that the LOX turbopump rotating-shaft seal system has sufficient durability for use in the LE-7 engine for the H-II launch vehicle.

  20. Distributed Parameter Analysis of Pressure and Flow Disturbances in Rocket Propellant Feed Systems

    NASA Technical Reports Server (NTRS)

    Dorsch, Robert G.; Wood, Don J.; Lightner, Charlene

    1966-01-01

    A digital distributed parameter model for computing the dynamic response of propellant feed systems is formulated. The analytical approach used is an application of the wave-plan method of analyzing unsteady flow. Nonlinear effects are included. The model takes into account locally high compliances at the pump inlet and at the injector dome region. Examples of the calculated transient and steady-state periodic responses of a simple hypothetical propellant feed system to several types of disturbances are presented. Included are flow disturbances originating from longitudinal structural motion, gimbaling, throttling, and combustion-chamber coupling. The analytical method can be employed for analyzing developmental hardware and offers a flexible tool for the calculation of unsteady flow in these systems.

  1. Orbit transfer rocket engine integrated control and health monitoring system technology readiness assessment

    NASA Technical Reports Server (NTRS)

    Bickford, R. L.; Collamore, F. N.; Gage, M. L.; Morgan, D. B.; Thomas, E. R.

    1992-01-01

    The objectives of this task were to: (1) estimate the technology readiness of an integrated control and health monitoring (ICHM) system for the Aerojet 7500 lbF Orbit Transfer Vehicle engine preliminary design assuming space based operations; and (2) estimate the remaining cost to advance this technology to a NASA defined 'readiness level 6' by 1996 wherein the technology has been demonstrated with a system validation model in a simulated environment. The work was accomplished through the conduct of four subtasks. In subtask 1 the minimally required functions for the control and monitoring system was specified. The elements required to perform these functions were specified in Subtask 2. In Subtask 3, the technology readiness level of each element was assessed. Finally, in Subtask 4, the development cost and schedule requirements were estimated for bringing each element to 'readiness level 6'.

  2. Development of the Functional Flow Block Diagram for the J-2X Rocket Engine System

    NASA Technical Reports Server (NTRS)

    White, Thomas; Stoller, Sandra L.; Greene, WIlliam D.; Christenson, Rick L.; Bowen, Barry C.

    2007-01-01

    The J-2X program calls for the upgrade of the Apollo-era Rocketdyne J-2 engine to higher power levels, using new materials and manufacturing techniques, and with more restrictive safety and reliability requirements than prior human-rated engines in NASA history. Such requirements demand a comprehensive systems engineering effort to ensure success. Pratt & Whitney Rocketdyne system engineers performed a functional analysis of the engine to establish the functional architecture. J-2X functions were captured in six major operational blocks. Each block was divided into sub-blocks or states. In each sub-block, functions necessary to perform each state were determined. A functional engine schematic consistent with the fidelity of the system model was defined for this analysis. The blocks, sub-blocks, and functions were sequentially numbered to differentiate the states in which the function were performed and to indicate the sequence of events. The Engine System was functionally partitioned, to provide separate and unique functional operators. Establishing unique functional operators as work output of the System Architecture process is novel in Liquid Propulsion Engine design. Each functional operator was described such that its unique functionality was identified. The decomposed functions were then allocated to the functional operators both of which were the inputs to the subsystem or component performance specifications. PWR also used a novel approach to identify and map the engine functional requirements to customer-specified functions. The final result was a comprehensive Functional Flow Block Diagram (FFBD) for the J-2X Engine System, decomposed to the component level and mapped to all functional requirements. This FFBD greatly facilitates component specification development, providing a well-defined trade space for functional trades at the subsystem and component level. It also provides a framework for function-based failure modes and effects analysis (FMEA), and a

  3. Radiation hardening of components and systems for nuclear rocket vehicle applications

    NASA Technical Reports Server (NTRS)

    Greenhow, W. A.; Cheever, P. R.

    1972-01-01

    The results of the analysis of the S-2 and S-4B components, although incomplete, indicate that many Saturn 5 components and subsystems, e.g., pumps, valves, etc., can be radiation hardened to meet NRV requirements by material substitution and minor design modifications. Results of these analyses include (1) recommended radiation tolerance limits for over 100 material applications; (2) design data which describes the components of each system; (3) presentation of radiation hardening examples of systems; and (4) designing radiation effects tests to supply data for selecting materials.

  4. Hydrogen plasma tests of some insulating coating systems for the nuclear rocket thrust chamber

    NASA Technical Reports Server (NTRS)

    Current, A. N.; Grisaffe, S. J.; Wycoff, K. C.

    1972-01-01

    Several plasma-sprayed and slurry-coated insulating coating systems were evaluated for structural stability in a low-pressure hot hydrogen environment at a maximum heat flux of 19.6 million watts/sq meter. The heat was provided by an electric-arc plasma generator. The coating systems consisted of a number of thin layers of metal oxides and/or metals. The materials included molybdenum, nichrome, tungsten, alumina, zirconia, and chromia. The study indicates potential usefulness in this environment for some coatings, and points up the need for improved coating application techniques.

  5. Parametric Study of Water Rocket for Optimum Flight

    NASA Astrophysics Data System (ADS)

    Ota, Takayuki; Umemura, Akira

    Parametric study is conducted to find the optimum condition of water rocket for long flight, provided that the tank volume is prescribed. The parameters considered in the present study are the initial air pressure, water volume fraction, empty rocket mass, launching angle and bottle diameter which significantly affect the flight performance of water rocket. First, we calculate the temporal changes in tank pressure, water and air issue speeds and thrust, on the basis of a simple physical model which has been experimentally validated. Then, this model is incorporated into the equation of motion to calculate the ballistic flight of water rocket with various parameter values. As a result, it is found that PET bottles in the market are one of the most suitable for use as the pressure tank of water rocket.

  6. Pressure-Equalizing Cradle for Booster Rocket Mounting

    NASA Technical Reports Server (NTRS)

    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.

  7. Submillimetre-sized dust aggregate collision and growth properties. Experimental study of a multi-particle system on a suborbital rocket

    NASA Astrophysics Data System (ADS)

    Brisset, J.; Heißelmann, D.; Kothe, S.; Weidling, R.; Blum, J.

    2016-08-01

    Context. In the very first steps of the formation of a new planetary system, dust agglomerates grow inside the protoplanetary disk that rotates around the newly formed star. In this disk, collisions between the dust particles, induced by interactions with the surrounding gas, lead to sticking. Aggregates start growing until their sizes and relative velocities are high enough for collisions to result in bouncing or fragmentation. With the aim of investigating the transitions between sticking and bouncing regimes for colliding dust aggregates and the formation of clusters from multiple aggregates, the Suborbital Particle and Aggregation Experiment (SPACE) was flown on the REXUS 12 suborbital rocket. Aims: The collisional and sticking properties of sub-mm-sized aggregates composed of protoplanetary dust analogue material are measured, including the statistical threshold velocity between sticking and bouncing, their surface energy and tensile strength within aggregate clusters. Methods: We performed an experiment on the REXUS 12 suborbital rocket. The protoplanetary dust analogue materials were micrometre-sized monodisperse and polydisperse SiO2 particles prepared into aggregates with sizes around 120 μm and 330 μm, respectively and volume filling factors around 0.37. During the experimental run of 150 s under reduced gravity conditions, the sticking of aggregates and the formation and fragmentation of clusters of up to a few millimetres in size was observed. Results: The sticking probability of the sub-mm-sized dust aggregates could be derived for velocities decreasing from ~22 to 3 cm s-1. The transition from bouncing to sticking collisions happened at 12.7+2.1-1.4 cm s-1 for the smaller aggregates composed of monodisperse particles and at 11.5+1.9-1.3 and 11.7+1.9-1.3 cm s-1 for the larger aggregates composed of mono- and polydisperse dust particles, respectively. Using the pull-off force of sub-mm-sized dust aggregates from the clusters, the surface energy of the

  8. DataRocket: Interactive Visualisation of Data Structures

    NASA Astrophysics Data System (ADS)

    Parkes, Steve; Ramsay, Craig

    2010-08-01

    CodeRocket is a software engineering tool that provides cognitive support to the software engineer for reasoning about a method or procedure and for documenting the resulting code [1]. DataRocket is a software engineering tool designed to support visualisation and reasoning about program data structures. DataRocket is part of the CodeRocket family of software tools developed by Rapid Quality Systems [2] a spin-out company from the Space Technology Centre at the University of Dundee. CodeRocket and DataRocket integrate seamlessly with existing architectural design and coding tools and provide extensive documentation with little or no effort on behalf of the software engineer. Comprehensive, abstract, detailed design documentation is available early on in a project so that it can be used for design reviews with project managers and non expert stakeholders. Code and documentation remain fully synchronised even when changes are implemented in the code without reference to the existing documentation. At the end of a project the press of a button suffices to produce the detailed design document. Existing legacy code can be easily imported into CodeRocket and DataRocket to reverse engineer detailed design documentation making legacy code more manageable and adding substantially to its value. This paper introduces CodeRocket. It then explains the rationale for DataRocket and describes the key features of this new tool. Finally the major benefits of DataRocket for different stakeholders are considered.

  9. Large space systems technology, 1980, volume 1

    NASA Technical Reports Server (NTRS)

    Kopriver, F., III (Compiler)

    1981-01-01

    The technological and developmental efforts in support of the large space systems technology are described. Three major areas of interests are emphasized: (1) technology pertient to large antenna systems; (2) technology related to large space systems; and (3) activities that support both antenna and platform systems.

  10. Sounding rocket lessons learned

    NASA Technical Reports Server (NTRS)

    Wessling, Francis C.; Maybee, George W.

    1991-01-01

    Programmatic, applicatory, developmental, and operational aspects of sounding rocket utilization for materials processing studies are discussed. Lessons learned through the experience of 10 sounding rocket missions are described. Particular attention is given to missions from the SPAR, Consort, and Joust programs. Successful experiments on Consort include the study of polymer membranes and resins, biological processes, demixing of immiscible liquids, and electrodeposition.

  11. The Rocket Project.

    ERIC Educational Resources Information Center

    Winemiller, Jake; And Others

    1991-01-01

    Describes an extra credit science project in which students compete to see who can build the most efficient water rocket out of a two-liter pop bottle. Provides instructions on how to build a demonstration rocket and launching pad. (MDH)

  12. Life Saving Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

    By 1870, American and British inventors had found other ways to use rockets. For example, the Congreve rocket was capable of carrying a line over 1,000 feet to a stranded ship. In 1914, an estimated 1,000 lives were saved by this technique.

  13. Model Rockets and Microchips.

    ERIC Educational Resources Information Center

    Fitzsimmons, Charles P.

    1986-01-01

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

  14. Rockets -- Part II.

    ERIC Educational Resources Information Center

    Leitner, Alfred

    1982-01-01

    If two rockets are identical except that one engine burns in one-tenth the time of the other (total impulse and initial fuel mass of the two engines being the same), which rocket will rise higher? Why? The answer to this question (part 1 response in v20 n6, p410, Sep 1982) is provided. (Author/JN)

  15. Laser power conversion system analysis, volume 1

    NASA Technical Reports Server (NTRS)

    Jones, W. S.; Morgan, L. L.; Forsyth, J. B.; Skratt, J. P.

    1979-01-01

    The orbit-to-orbit laser energy conversion system analysis established a mission model of satellites with various orbital parameters and average electrical power requirements ranging from 1 to 300 kW. The system analysis evaluated various conversion techniques, power system deployment parameters, power system electrical supplies and other critical supplies and other critical subsystems relative to various combinations of the mission model. The analysis show that the laser power system would not be competitive with current satellite power systems from weight, cost and development risk standpoints.

  16. Infrared Imagery of Solid Rocket Exhaust Plumes

    NASA Technical Reports Server (NTRS)

    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.

  17. Computer Sciences and Data Systems, volume 1

    NASA Technical Reports Server (NTRS)

    1987-01-01

    Topics addressed include: software engineering; university grants; institutes; concurrent processing; sparse distributed memory; distributed operating systems; intelligent data management processes; expert system for image analysis; fault tolerant software; and architecture research.

  18. Modeling the Earth System, volume 3

    NASA Technical Reports Server (NTRS)

    Ojima, Dennis (Editor)

    1992-01-01

    The topics covered fall under the following headings: critical gaps in the Earth system conceptual framework; development needs for simplified models; and validating Earth system models and their subcomponents.

  19. Two-Nucleon Systems in a Finite Volume

    SciTech Connect

    Briceno, Raul

    2014-11-01

    I present the formalism and methodology for determining the nucleon-nucleon scattering parameters from the finite volume spectra obtained from lattice quantum chromodynamics calculations. Using the recently derived energy quantization conditions and the experimentally determined scattering parameters, the bound state spectra for finite volume systems with overlap with the 3S1-3D3 channel are predicted for a range of volumes. It is shown that the extractions of the infinite-volume deuteron binding energy and the low-energy scattering parameters, including the S-D mixing angle, are possible from Lattice QCD calculations of two-nucleon systems with boosts of |P| <= 2pi sqrt{3}/L in volumes with spatial extents L satisfying fm <~ L <~ 14 fm.

  20. Multiple IMU system development, volume 1

    NASA Technical Reports Server (NTRS)

    Landey, M.; Mckern, R.

    1974-01-01

    A redundant gimballed inertial system is described. System requirements and mechanization methods are defined and hardware and software development is described. Failure detection and isolation algorithms are presented and technology achievements described. Application of the system as a test tool for shuttle avionics concepts is outlined.

  1. SIRU development. Volume 1: System development

    NASA Technical Reports Server (NTRS)

    Gilmore, J. P.; Cooper, R. J.

    1973-01-01

    A complete description of the development and initial evaluation of the Strapdown Inertial Reference Unit (SIRU) system is reported. System development documents the system mechanization with the analytic formulation for fault detection and isolation processing structure; the hardware redundancy design and the individual modularity features; the computational structure and facilities; and the initial subsystem evaluation results.

  2. Space shuttle system program definition. Volume 4: Cost and schedule report

    NASA Technical Reports Server (NTRS)

    1972-01-01

    The supporting cost and schedule data for the second half of the Space Shuttle System Phase B Extension Study is summarized. The major objective for this period was to address the cost/schedule differences affecting final selection of the HO orbiter space shuttle system. The contending options under study included the following booster launch configurations: (1) series burn ballistic recoverable booster (BRB), (2) parallel burn ballistic recoverable booster (BRB), (3) series burn solid rocket motors (SRM's), and (4) parallel burn solid rocket motors (SRM's). The implications of varying payload bay sizes for the orbiter, engine type for the ballistics recoverable booster, and SRM motors for the solid booster were examined.

  3. Study of aircraft in intraurban transportation systems. Volume 4: Appendix

    NASA Technical Reports Server (NTRS)

    Stout, E. G.; Kesling, P. H.; Matteson, H. C.; Sherwood, D. E.; Tuck, W. R., Jr.; Vaughn, L. A.

    1971-01-01

    An appendix of the supporting data leading to conclusions and recommendations for an effective intraurban transportation system from volumes 1, 2, and 3 is presented. The data are given in tables and graphs.

  4. Rocket/Nimbus Sounder Comparison (RNSC)

    NASA Technical Reports Server (NTRS)

    1972-01-01

    The experimental results for radiance and temperature differences in the Wallops Island comparisons indicate that the differences between satellite and rocket systems are of the same order of magnitude as the differences among the various satellite and rocket sounders. The Arcasondes produced usable data to about 50 km, while the Datasondes require design modification. The SIRS and IRIS soundings provided usable data to 30 mb; extension of these soundings was also investigated.

  5. Challenger Rocket Booster

    NASA Technical Reports Server (NTRS)

    1986-01-01

    At about 76 seconds, fragments of the Orbiter can be seen tumbling against a background of fire, smoke and vaporized propellants from the External Tank. The left Solid Rocket Booster (SRB) flys rampant, still thrusting. The reddish-brown cloud envelops the disintergrating Orbiter. The color is indicative of the nitrogen tetroxide oxidizer propellant in the Orbiter Reaction Control System. On January 28, 1986 frigid overnight temperatures caused normally pliable rubber O-ring seals and putty that are designed to seal and establish joint integrity between the Solid Rocket Booster (SRB) joint segments, to become hard and non- flexible. At the instant of SRB ignition, tremendous stresses and pressures occur within the SRB casing and especially at the joint attachment points. The failure of the O-rings and putty to 'seat' properly at motor ignition, caused hot exhaust gases to blow by the seals and putty. During Challenger's ascent, this hot gas 'blow by' ultimately cut a swath completely through the steel booster casing; and like a welder's torch, began cutting into the External Tank (ET). It is believed that the ET was compromised in several locations starting in the aft at the initial point where SRB joint failure occured. The ET hydrogen tank is believed to have been breached first, with continuous rapid incremental failure of both the ET and SRB. The chain reaction of events occurring in milliseconds culminated in a massive explosion. The orbiter Challenger was instantly ejected by the blast and went askew into the supersonic air flow. These aerodynamic forces caused structural shattering and complete destruction of the orbiter. Though it was concluded that the G-forces experienced during orbiter ejection and break-up were survivable, impact with the ocean surface was not. Tragically, all seven crewmembers perished.

  6. Individual Global Navigation Satellite Systems in the Space Service Volume

    NASA Technical Reports Server (NTRS)

    Force, Dale A.

    2015-01-01

    Besides providing position, navigation, and timing (PNT) to terrestrial users, GPS is currently used to provide for precision orbit determination, precise time synchronization, real-time spacecraft navigation, and three-axis control of Earth orbiting satellites. With additional Global Navigation Satellite Systems (GNSS) coming into service (GLONASS, Beidou, and Galileo), it will be possible to provide these services by using other GNSS constellations. The paper, "GPS in the Space Service Volume," presented at the ION GNSS 19th International Technical Meeting in 2006 (Ref. 1), defined the Space Service Volume, and analyzed the performance of GPS out to 70,000 km. This paper will report a similar analysis of the performance of each of the additional GNSS and compare them with GPS alone. The Space Service Volume, defined as the volume between 3,000 km altitude and geosynchronous altitude, as compared with the Terrestrial Service Volume between the surface and 3,000 km. In the Terrestrial Service Volume, GNSS performance will be similar to performance on the Earth's surface. The GPS system has established signal requirements for the Space Service Volume. A separate paper presented at the conference covers the use of multiple GNSS in the Space Service Volume.

  7. The Learning System. Volume 5, Number 2

    ERIC Educational Resources Information Center

    Crow, Tracy, Ed.

    2009-01-01

    Ensuring quality teaching in every classroom across an entire system of schools--that's what a district leader's job is all about. A district leader's challenges are unique so "The Learning System" was created with that in mind. This issue contains: (1) Imagine the Possibilities: 2020 Forecast Explores 6 Change Forces that Will Shape the Future of…

  8. Television broadcast relay system, volume 2

    NASA Technical Reports Server (NTRS)

    Graf, E. R.

    1972-01-01

    A statistical analysis of propagation characteristics at 12 GHz for the design of a worldwide communication satellite system is reported. High power spaceborne TV transmitter design and the development of a low cost, low noise ground receiver for domestic reception are considered most important for implementation of a TVBS system.

  9. Preliminary guided rocket feasibility study.

    NASA Technical Reports Server (NTRS)

    Nolan, M. B.; Celmer, J. J.

    1973-01-01

    The feasibility of actively guiding sounding rockets to reduce impact dispersion has been investigated. The theoretical probability of Range Safety thrust termination for several high performance rockets was combined with the cost of acquiring the extended range at White Sands Missile Range (WSMR) to establish a guidance system price ceiling of $20K per flight. Guiding the Black Brant VC for the first five seconds of flight results in sufficient dispersion reduction to impact within the standard range boundaries at WSMR. The guidance system thrust level required to statically control the vehicle to a nominal-wind weighted trajectory for five seconds is between 150-200 pounds. A six-degree-of-freedom trajectory program with guidance simulation capability has been developed and the equations are delineated in this paper.

  10. Preliminary guided rocket feasibility study

    NASA Technical Reports Server (NTRS)

    Nolan, M. B.; Celmer, J. J.

    1973-01-01

    The feasibility of actively guiding sounding rockets to reduce impact dispersion has been investigated. The theoretical probability of range safety thrust termination for several high performance rockets was combined with the cost of acquiring the extended range at White Sands Missile Range (WSMR) to establish a guidance system price ceiling of $20K per flight. Guiding the Black Brant VC (BBVC) for the first five seconds of flight results in sufficient dispersion reduction to impact within the standard range boundaries at WSMR. The guidance system thrust level required to statically control the vehicle to a nominal-wind weighted trajectory for five seconds is between 150-200 pounds. A six-degree-of-freedom trajectory program with guidance simulation capability has been developed and the equations are presented.

  11. Low thrust rocket test facility

    NASA Technical Reports Server (NTRS)

    Arrington, Lynn A.; Schneider, Steven J.

    1990-01-01

    A low thrust chemical rocket test facility has recently become operational at the NASA-Lewis. The new facility is used to conduct both long duration and performance tests at altitude over a thruster's operating envelope using hydrogen and oxygen gas for propellants. The facility provides experimental support for a broad range of objectives, including fundamental modeling of fluids and combustion phenomena, the evaluation of thruster components, and life testing of full rocket designs. The major mechanical and electrical systems are described along with aspects of the various optical diagnostics available in the test cell. The electrical and mechanical systems are designed for low down time between tests and low staffing requirements for test operations. Initial results are also presented which illustrate the various capabilities of the cell.

  12. Advanced Solid Rocket Launcher and Its Evolution

    NASA Astrophysics Data System (ADS)

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

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

  13. Indians Repulse British With Rocket

    NASA Technical Reports Server (NTRS)

    2004-01-01

    During the early introduction of rockets to Europe, they were used only as weapons. Enemy troops in India repulsed the British with rockets. Later, in Britain, Sir William Congreve developed a rocket that could fire to about 9,000 feet. The British fired Congreve rockets against the United States in the War of 1812.

  14. Laser power conversion system analysis, volume 2

    NASA Technical Reports Server (NTRS)

    Jones, W. S.; Morgan, L. L.; Forsyth, J. B.; Skratt, J. P.

    1979-01-01

    The orbit-to-ground laser power conversion system analysis investigated the feasibility and cost effectiveness of converting solar energy into laser energy in space, and transmitting the laser energy to earth for conversion to electrical energy. The analysis included space laser systems with electrical outputs on the ground ranging from 100 to 10,000 MW. The space laser power system was shown to be feasible and a viable alternate to the microwave solar power satellite. The narrow laser beam provides many options and alternatives not attainable with a microwave beam.

  15. Ionospheric modification by rocket effluents. Final report

    SciTech Connect

    Bernhardt, P.A.; Price, K.M.; da Rosa, A.V.

    1980-06-01

    This report describes experimental and theoretical studies related to ionospheric disturbances produced by rocket exhaust vapors. The purpose of our research was to estimate the ionospheric effects of the rocket launches which will be required to place the Satellite Power System (SPS) in operation. During the past year, we have developed computational tools for numerical simulation of ionospheric changes produced by the injection of rocket exhaust vapors. The theoretical work has dealt with (1) the limitations imposed by condensation phenomena in rocket exhaust; (2) complete modeling of the ionospheric depletion process including neutral gas dynamics, plasma physics, chemistry and thermal processes; and (3) the influence of the modified ionosphere on radio wave propagation. We are also reporting on electron content measurements made during the launch of HEAO-C on Sept. 20, 1979. We conclude by suggesting future experiments and areas for future research.

  16. Liquid rocket engine turbopump gears

    NASA Technical Reports Server (NTRS)

    1974-01-01

    Design and fabrication of gear drives for rocket engine turbopumps are described in the sequence encountered during the design process as follows: (1) selection of overall arrangement; (2) selection of gear type; (3) preliminary sizing; (4) lubrication system design; (5) detail tooth design; (6) selection of gear materials; and (7) gear fabrication and testing as it affects the design. The description is oriented towards the use of involute spur gears, although reference material for helical gears is also cited.

  17. Human Transportation System (HTS) study, volume 1

    NASA Technical Reports Server (NTRS)

    Lance, N.; Geyer, M. S.; Gaunce, M. T.

    1993-01-01

    Work completed under the Human Transportation System Study is summarized. This study was conducted by the New Initiatives Office at JSC with the technical support of Boeing, General Dynamics, Lockheed, McDonnell-Douglas, Martin Marietta, and Rockwell. The study was designed to generate information on determining the appropriate path to follow for new system development to meet the Nation's space transportation needs. The study evaluates 18 transportation architecture options using a parametric set of mission requirements. These options include use of current systems as well as proposed systems to assess the impact of various considerations, such as the cost of alternate access, or the benefit of separating people and cargo. The architecture options are compared to each other with six measurable evaluation criteria or attributes. They are the following: funding profile, human safety, probability of mission success, architecture cost risk, launch schedule confidence, and environmental impact. Values for these attributes are presented for the architecture options, with pertinent conclusions and recommendations.

  18. Advanced vehicle systems assessment. Volume 5: Appendices

    NASA Technical Reports Server (NTRS)

    Hardy, K.

    1985-01-01

    An appendix to the systems assessment for the electric hybrid vehicle project is presented. Included are battery design, battery cost, aluminum vehicle construction, IBM PC computer programs and battery discharge models.

  19. Astrionics system designers handbook, volume 1

    NASA Technical Reports Server (NTRS)

    1973-01-01

    Hardware elements in new and advanced astrionics system designs are discussed. This cost effective approach has as its goal the reduction of R&D and testing costs through the application of proven and tested astrionics components. The ready availability to the designer of data facts for applicable system components is highly desirable. The astrionics System Designers Handbook has as its objective this documenting of data facts to serve the anticipated requirements of the astrionics system designer. Eleven NASA programs were selected as the reference base for the document. These programs are: ATS-F, ERTS-B, HEAO-A, OSO-I, Viking Orbiter, OAO-C, Skylab AM/MDA, Skylab ATM, Apollo 17 CSM, Apollo 17 LM and Mariner Mars 71. Four subsystems were chosen for documentation: communications, data management, electrical power and guidance, navigation and control.

  20. Combined Global Navigation Satellite Systems in the Space Service Volume

    NASA Technical Reports Server (NTRS)

    Force, Dale A.; Miller, James J.

    2015-01-01

    Besides providing position, navigation, and timing (PNT) services to traditional terrestrial and airborne users, GPS is also being increasingly used as a tool to enable precision orbit determination, precise time synchronization, real-time spacecraft navigation, and three-axis attitude control of Earth orbiting satellites. With additional Global Navigation Satellite System (GNSS) constellations being replenished and coming into service (GLONASS, Beidou, and Galileo), it will become possible to benefit from greater signal availability and robustness by using evolving multi-constellation receivers. The paper, "GPS in the Space Service Volume," presented at the ION GNSS 19th International Technical Meeting in 2006 (Ref. 1), defined the Space Service Volume, and analyzed the performance of GPS out to seventy thousand kilometers. This paper will report a similar analysis of the signal coverage of GPS in the space domain; however, the analyses will also consider signal coverage from each of the additional GNSS constellations noted earlier to specifically demonstrate the expected benefits to be derived from using GPS in conjunction with other foreign systems. The Space Service Volume is formally defined as the volume of space between three thousand kilometers altitude and geosynchronous altitude circa 36,000 km, as compared with the Terrestrial Service Volume between 3,000 km and the surface of the Earth. In the Terrestrial Service Volume, GNSS performance is the same as on or near the Earth's surface due to satellite vehicle availability and geometry similarities. The core GPS system has thereby established signal requirements for the Space Service Volume as part of technical Capability Development Documentation (CDD) that specifies system performance. Besides the technical discussion, we also present diplomatic efforts to extend the GPS Space Service Volume concept to other PNT service providers in an effort to assure that all space users will benefit from the enhanced

  1. Theoretical Astrophysics, Volume 2: Stars and Stellar Systems

    NASA Astrophysics Data System (ADS)

    Padmanabhan, T.

    2001-04-01

    This second volume of a comprehensive three-volume work on theoretical astrophysics deals with stellar physics. After reviewing the key observational results and nomenclature used in stellar astronomy, the book develops a solid understanding of central concepts including stellar structure and evolution, the physics of stellar remnants, pulsars, binary stars, the sun and planetary systems, interstellar medium and globular clusters. Throughout, the reader's comprehension is developed and tested with more than seventy-five exercises. This indispensable volume will allow graduate students to master the material sufficiently to read and engage in research with heightened understanding. It can be used alone or in conjunction with Volume 1, which covers a wide range of astrophysical processes, and the forthcoming Volume 3, on galaxies and cosmology.

  2. Baking Soda and Vinegar Rockets

    NASA Astrophysics Data System (ADS)

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

    2009-02-01

    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 vendors1,2 that combine baking soda and vinegar for propulsion. The experiment compared two analytical approximations for the maximum rocket height to the experimentally measured rocket height. Baking soda and vinegar rockets present fewer safety concerns and require a smaller launch area than rapid combustion chemical rockets. Both kits were of nearly identical design, costing ˜20. The rockets required roughly 30 minutes of assembly time consisting of mostly taping the soft plastic fuselage to the Styrofoam nose cone.

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

    NASA Astrophysics Data System (ADS)

    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.

    2000-04-01

    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.

  4. Transportation systems analyses. Volume 1: Executive summary

    NASA Astrophysics Data System (ADS)

    1992-11-01

    The principal objective is to accomplish a systems engineering assessment of the nation's space transportation infrastructure. This analysis addresses the necessary elements to perform crew delivery and return, cargo transfer, cargo delivery and return, payload servicing, and the exploration of the Moon and Mars. Specific elements analyzed, but not limited to, include: the Space Exploration Initiative (SEI), the National Launch System (NLS), the current expendable launch vehicle (ELV) fleet, ground facilities, the Space Station Freedom (SSF), and other civil, military and commercial payloads. The performance of this study entails maintaining a broad perspective on the large number of transportation elements that could potentially comprise the U.S. space infrastructure over the next several decades. To perform this systems evaluation, top-level trade studies are conducted to enhance our understanding of the relationship between elements of the infrastructure. This broad 'infrastructure-level perspective' permits the identification of preferred infrastructures. Sensitivity analyses are performed to assure the credibility and usefulness of study results. Conceptual studies of transportation elements contribute to the systems approach by identifying elements (such as ETO node and transfer/excursion vehicles) needed in current and planned transportation systems. These studies are also a mechanism to integrate the results of relevant parallel studies.

  5. Integrated safety management system verification: Volume 2

    SciTech Connect

    Christensen, R.F.

    1998-08-10

    Department of Energy (DOE) Policy (P) 450.4, Safety Management System Policy, commits to institutionalization of an Integrated Safety Management System (ISMS) throughout the DOE complex. The DOE Acquisition Regulations (DEAR, 48 CFR 970) requires contractors to manage and perform work in accordance with a documented Integrated Safety Management System (ISMS). Guidance and expectations have been provided to PNNL by incorporation into the operating contract (Contract DE-ACM-76FL0 1830) and by letter. The contract requires that the contractor submit a description of their ISMS for approval by DOE. PNNL submitted their proposed Safety Management System Description for approval on November 25,1997. RL tentatively approved acceptance of the description pursuant to a favorable recommendation from this review. The Integrated Safety Management System Verification is a review of the adequacy of the ISMS description in fulfilling the requirements of the DEAR and the DOE Policy. The purpose of this review is to provide the Richland Operations Office Manager with a recommendation for approval of the ISMS description of the Pacific Northwest Laboratory based upon compliance with the requirements of 49 CFR 970.5204(-2 and -78); and to verify the extent and maturity of ISMS implementation within the Laboratory. Further the review will provide a model for other DOE laboratories managed by the Office of Assistant Secretary for Energy Research.

  6. Radial Internal Material Handling System (RIMS) for Circular Habitat Volumes

    NASA Technical Reports Server (NTRS)

    Howe, Alan S.; Haselschwardt, Sally; Bogatko, Alex; Humphrey, Brian; Patel, Amit

    2013-01-01

    On planetary surfaces, pressurized human habitable volumes will require a means to carry equipment around within the volume of the habitat, regardless of the partial gravity (Earth, Moon, Mars, etc.). On the NASA Habitat Demonstration Unit (HDU), a vertical cylindrical volume, it was determined that a variety of heavy items would need to be carried back and forth from deployed locations to the General Maintenance Work Station (GMWS) when in need of repair, and other equipment may need to be carried inside for repairs, such as rover parts and other external equipment. The vertical cylindrical volume of the HDU lent itself to a circular overhead track and hoist system that allows lifting of heavy objects from anywhere in the habitat to any other point in the habitat interior. In addition, the system is able to hand-off lifted items to other material handling systems through the side hatches, such as through an airlock. The overhead system consists of two concentric circle tracks that have a movable beam between them. The beam has a hoist carriage that can move back and forth on the beam. Therefore, the entire system acts like a bridge crane curved around to meet itself in a circle. The novelty of the system is in its configuration, and how it interfaces with the volume of the HDU habitat. Similar to how a bridge crane allows coverage for an entire rectangular volume, the RIMS system covers a circular volume. The RIMS system is the first generation of what may be applied to future planetary surface vertical cylinder habitats on the Moon or on Mars.

  7. Advanced extravehicular protective systems study, volume 1

    NASA Technical Reports Server (NTRS)

    Sutton, J. G.; Heimlich, P. F.; Tepper, E. H.

    1972-01-01

    An appraisal was made of advanced portable and emergency life support systems concepts for space station, space shuttle, lunar base, and Mars EVA missions. Specifications are given, and the methodology is described. Subsystem studies and systems integration efforts are summarized. Among the conclusions are the following: (1) For long duration missions, a configuration incorporating a regenerable CO2 control subsystem and a thermal control subsystem utilizing a minimum of expendables decreases the vehicle penalty of present configurations. (2) For shorter duration missions, a configuration incorporating an expendable water thermal control subsystem is the most competitive subsystem; regenerable CO2 control subsystems if properly developed are competitive with nonregenerable counterparts. (3) The CO2 reduction and oxygen reclamation withing the parent vehicle is only competitive when there are three or more parent vehicle resupply periods. (4) For long duration emergency systems of one hour or more, inherent redundancy within the primary configuration to provide emergency thermal control is the most competitive approach.

  8. Space shuttle system program definition. Volume 2: Technical report

    NASA Technical Reports Server (NTRS)

    1972-01-01

    The Phase B Extension of the Space Shuttle System Program Definition study was redirected to apply primary effort to consideration of space shuttle systems utilizing either recoverable pressure fed liquids or expendable solid rocket motor boosters. Two orbiter configurations were to be considered, one with a 15x60 foot payload bay with a 65,000 lb, due East, up-payload capability and the other with a 14x45 payload bay with 45,000 lb, of due East, up-payload. Both were to use three SSME engines with 472,000 lb of vacuum thrust each. Parallel and series burn ascent modes were to be considered for the launch configurations of primary interest. A recoverable pump-fed booster is included in the study in a series burn configuration with the 15x60 orbiter. To explore the potential of the swing engine orbiter configuration in the pad abort case, it is included in the study matrix in two launch configurations, a series burn pressure fed BRB and a parallel burn SRM. The resulting matrix of configuration options is shown. The principle objectives of this study are to evaluate the cost and technical differences between the liquid and solid propellant booster systems and to assess the development and operational cost savings available with a smaller orbiter.

  9. 2003 NASA Seal/Secondary Air System Workshop. Volume 1

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M. (Editor); Hendricks, Robert C. (Editor)

    2004-01-01

    The following reports were included in the 2003 NASA Seal/Secondary Air System Workshop:Low Emissions Alternative Power (LEAP); Overview of NASA Glenn Seal Developments; NASA Ultra Efficient Engine Technology Project Overview; Development of Higher Temperature Abradable Seals for Industrial Gas Turbines; High Misalignment Carbon Seals for the Fan Drive Gear System Technologies; Compliant Foil Seal Investigations; Test Rig for Evaluating Active Turbine Blade Tip Clearance Control Concepts; Controls Considerations for Turbine Active Clearance Control; Non-Contacting Finger Seal Developments and Design Considerations; Effect of Flow-Induced Radial Load on Brush Seal/Rotor Contact Mechanics; Seal Developments at Flowserve Corporation; Investigations of High Pressure Acoustic Waves in Resonators With Seal-Like Features; Numerical Investigations of High Pressure Acoustic Waves in Resonators; Feltmetal Seal Material Through-Flow; "Bimodal" Nuclear Thermal Rocket (BNTR) Propulsion for Future Human Mars Exploration Missions; High Temperature Propulsion System Structural Seals for Future Space Launch Vehicles; Advanced Control Surface Seal Development for Future Space Vehicles; High Temperature Metallic Seal Development for Aero Propulsion and Gas Turbine Applications; and BrazeFoil Honeycomb.

  10. Control system retrofit guidelines. Volume 2: Technical assessment, Final report

    SciTech Connect

    Stoddard, L.E.; Gil, L.F.

    1992-07-01

    This report has been prepared by Black & Veatch for the Electric Power Research Institute under Contract RP2710-16. This project has been a part of an EPRI program to enhance computer and control technology applications for the utility industry. It is estimated that the utility industry will spend 500 million dollars in the next two years for control system enhancement projects. The need for such enhancement results from a combination of the age and obsolescence of the control systems in existing fossil-fired plants, the decision by utilities to extend the lifetime of those plants, the changes in operating strategies, and the continued development of control systems with expanded capabilities. Opportunities for control system retrofits are myriad; therefore, utilities are faced with complex decisions as to how best to use the limited financial resources available for such projects. To assist utilities in making these decisions, the Electric Power Research Institute (EPRI) has retained Black & Veatch and Sargent & Lundy to develop a set of control system retrofit guideline documents. The guidelines are in a three-volume format. Volume I--Control System Retrofit Project Methodology; Volume II--Control System Technical Assessment; and Volume III--Utility Case Studies. Volume II (this document) gives guidance in the technical assessment for a retrofit, and has been prepared by Black & Veatch under EPRI RP2710-16. The combination of these guideline documents gives utilities the necessary guidance to perform a retrofit project in a straightforward, cost-effective manner.

  11. The Learning System. Volume 5, Number 1

    ERIC Educational Resources Information Center

    von Frank, Valerie, Ed.

    2009-01-01

    "The Learning System" is an eight-page newsletter published eight times a year. Designed for superintendents and central office staff with professional learning responsibilities. This issue includes: (1) Superintendent Stays on Course with Personal Learning Plan (Valerie von Frank); (2) District Leadership: Permit, Don't Proscribe, to Build…

  12. Human Transportation System (HTS) study, volume 2

    NASA Technical Reports Server (NTRS)

    Lance, N.; Geyer, M. S.; Gaunce, M. T.

    1993-01-01

    This report summarizes work completed under the Human Transportation System Study. This study was conducted by the New Initiatives Office at JSC with the technical support of Boeing, General Dynamics, Lockheed, McDonnell-Douglas, Martin Marietta, and Rockwell. The study was designed to generate information on determining the appropriate path to follow for new system development to meet the Nation's space transportation needs. The study evaluates 18 transportation architecture options using a parametric set of mission requirements. These options include use of current systems (e.g., Shuttle, Titan, etc. ) as well as proposed systems (e.g., PLS, Single-Stage-to-Orbit, etc.) to assess the impact of various considerations, such as the cost of alternate access, or the benefit of separating people and cargo. The architecture options are compared to each other with six measurable evaluation criteria or attributes. They are: funding profile, human safety, probability of mission success, architecture cost risk, launch schedule confidence, and environmental impact. Values for these attributes are presented for the architecture options, with pertinent conclusions and recommendations.

  13. Human Transportation System (HTS) study, volume 2

    NASA Astrophysics Data System (ADS)

    Lance, N.; Geyer, M. S.; Gaunce, M. T.

    1993-10-01

    This report summarizes work completed under the Human Transportation System Study. This study was conducted by the New Initiatives Office at JSC with the technical support of Boeing, General Dynamics, Lockheed, McDonnell-Douglas, Martin Marietta, and Rockwell. The study was designed to generate information on determining the appropriate path to follow for new system development to meet the Nation's space transportation needs. The study evaluates 18 transportation architecture options using a parametric set of mission requirements. These options include use of current systems (e.g., Shuttle, Titan, etc. ) as well as proposed systems (e.g., PLS, Single-Stage-to-Orbit, etc.) to assess the impact of various considerations, such as the cost of alternate access, or the benefit of separating people and cargo. The architecture options are compared to each other with six measurable evaluation criteria or attributes. They are: funding profile, human safety, probability of mission success, architecture cost risk, launch schedule confidence, and environmental impact. Values for these attributes are presented for the architecture options, with pertinent conclusions and recommendations.

  14. The Learning System. Volume 4, Number 7

    ERIC Educational Resources Information Center

    von Frank, Valerie, Ed.

    2009-01-01

    "The Learning System" is a newsletter designed for superintendents and central office staff with professional learning responsibilities. This issue includes: (1) District Pulls Together in Pursuit of Excellence: Creating Collaboration Systemwide Requires Commitment (Valerie von Frank); (2) Scheduling Time for Teacher Learning Is Key for Both…

  15. Transportation systems analyses: Volume 1: Executive Summary

    NASA Astrophysics Data System (ADS)

    1993-05-01

    The principal objective of this study is to accomplish a systems engineering assessment of the nation's space transportation infrastructure. This analysis addresses the necessary elements to perform man delivery and return, cargo transfer, cargo delivery, payload servicing, and the exploration of the Moon and Mars. Specific elements analyzed, but not limited to, include the Space Exploration Initiative (SEI), the National Launch System (NLS), the current expendable launch vehicle (ELV) fleet, ground facilities, the Space Station Freedom (SSF), and other civil, military and commercial payloads. The performance of this study entails maintaining a broad perspective on the large number of transportation elements that could potentially comprise the U.S. space infrastructure over the next several decades. To perform this systems evaluation, top-level trade studies are conducted to enhance our understanding of the relationships between elements of the infrastructure. This broad 'infrastructure-level perspective' permits the identification of preferred infrastructures. Sensitivity analyses are performed to assure the credibility and usefulness of study results. This executive summary of the transportation systems analyses (TSM) semi-annual report addresses the SSF logistics resupply. Our analysis parallels the ongoing NASA SSF redesign effort. Therefore, there could be no SSF design to drive our logistics analysis. Consequently, the analysis attempted to bound the reasonable SSF design possibilities (and the subsequent transportation implications). No other strategy really exists until after a final decision is rendered on the SSF configuration.

  16. Satellite voice broadcast system study, volume 2

    NASA Technical Reports Server (NTRS)

    Horstein, M.

    1985-01-01

    This study investigates the feasibility of providing Voice of America (VOA) broadcasts by satellite relay, rather than via terrestrial relay stations. Satellite voice broadcast systems are described for three different frequency bands: HF (26 MHz), VHF (68 MHz), and L-band (1.5 GHz). The geographical areas of interest at HF and L-band include all major land masses worldwide with the exception of the U.S., Canada, and Australia. Geostationary satellite configurations are considered for both frequency bands. In addition, a system of subsynchronous, circular satellites with an orbit period of 8 hours is developed for the HF band. VHF broadcasts, which are confined to the Soviet Union, are provied by a system of Molniya satellites. Satellites intended for HF or VHF broadcastinbg are extremely large and heavy. Satellite designs presented here are limited in size and weight to the capability of the STS/Centaur launch vehicle combination. Even so, at HF it would take 47 geostationary satellites or 20 satellites in 8-hour orbits to fully satisfy the voice-channel requirements of the broadcast schedule provided by VOA. On the other hand, three Molniya satellites suffice for the geographically restricted schedule at VHF. At L-band, only four geostationary satellites are needed to meet the requirements of the complete broadcast schedule. Moreover, these satellites are comparable in size and weight to current satellites designed for direct broadcast of video program material.

  17. System and Method for Wirelessly Determining Fluid Volume

    NASA Technical Reports Server (NTRS)

    Woodard, Stanley E. (Inventor); Taylor, Bryant D. (Inventor)

    2009-01-01

    A system and method are provided for determining the volume of a fluid in container. Sensors are positioned at distinct locations in a container of a fluid. Each sensor is sensitive to an interface defined by the top surface of the fluid. Interfaces associated with at least three of the sensors are determined and used to find the volume of the fluid in the container in a geometric process.

  18. Antares Rocket Lifts Off!

    NASA Video Gallery

    NASA commercial space partner Orbital Sciences Corp. of Dulles, Va., launched its Cygnus cargo spacecraft aboard its Antares rocket at 10:58 a.m. EDT Wednesday from the Mid-Atlantic Regional Spacep...

  19. Rocketing into Adaptive Inquiry.

    ERIC Educational Resources Information Center

    Farenga, Stephen J.; Joyce, Beverly A.; Dowling, Thomas W.

    2002-01-01

    Defines adaptive inquiry and argues for employing this method which allows lessons to be shaped in response to student needs. Illustrates this idea by detailing an activity in which teams of students build rockets. (DDR)

  20. The History of Rockets.

    ERIC Educational Resources Information Center

    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)

  1. Rocket engine numerical simulator

    NASA Technical Reports Server (NTRS)

    Davidian, Ken

    1993-01-01

    The topics are presented in viewgraph form and include the following: a rocket engine numerical simulator (RENS) definition; objectives; justification; approach; potential applications; potential users; RENS work flowchart; RENS prototype; and conclusion.

  2. Rocket engine numerical simulation

    NASA Technical Reports Server (NTRS)

    Davidian, Ken

    1993-01-01

    The topics are presented in view graph form and include the following: a definition of the rocket engine numerical simulator (RENS); objectives; justification; approach; potential applications; potential users; RENS work flowchart; RENS prototype; and conclusions.

  3. Design requirements for SRB production control system. Volume 1: Study background and overview

    NASA Technical Reports Server (NTRS)

    1981-01-01

    The solid rocket boosters assembly environment is described in terms of the contraints it places upon an automated production control system. The business system generated for the SRB assembly and the computer system which meets the business system requirements are described. The selection software process and modifications required to the recommended software are addressed as well as the hardware and configuration requirements necessary to support the system.

  4. STARPAHC systems report. Volume 1: Executive summary

    NASA Technical Reports Server (NTRS)

    1977-01-01

    A joint NASA and Department of Health, Education, and Welfare/Indian Health Services demonstration project entitled Space Technology Applied to Rural Papago Advanced Health Care (STARPAHC) was conducted to develop a solution for delivering quality health care to people in remote geographical areas. The STARPAHC concept verified the feasibility of telemedicine plus physician assistant - under the direction of a physician as a means of delivering quality health care. The two years of operational evaluation have provided considerable medical and engineering data which will be valuable to the designers and planners of future health care systems on earth and in space.

  5. STARPAHC systems report. Volume 2: Operational performance

    NASA Technical Reports Server (NTRS)

    1977-01-01

    The Space Technology Applied to Rural Papago Advanced Health Care (STARPAHC) demonstrated the value and potential of telemedicine using physician's assistants for providing quality health care delivery to people in a remote area. Generally, the program's achievements were to: (1) establish the feasibility of the STARPAHC concept in the delivery of health care; (2) gain information for developing health care systems for future manned spacecraft; (3) determine the constraints and capabilities involved in the interaction between physicians and non-physician health care personnel; (4) determine effectiveness of the STARPAHC technique; and (5) define the additional developments that are needed and/or most valuable to improving telemedicine and its exportable potential.

  6. Closeup view of the Solid Rocket Booster (SRB) Forward Skirt, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Close-up view of the Solid Rocket Booster (SRB) Forward Skirt, Frustum and Nose Cap mated assembly undergoing final preparations in the Solid Rocket Booster Assembly and Refurbishment Facility at Kennedy Space Center. In this view the access panel on the Forward Skirt is removed and you can see a small portion of the interior of the Forward Skirt. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  7. Closeup view of the Solid Rocket Booster Frustum and Nose ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Close-up view of the Solid Rocket Booster Frustum and Nose Cap assembly undergoing preparations and close-out procedures in the Solid Rocket Booster Assembly and Refurbishment Facility at Kennedy Space Center. The Nose Cap contains the Pilot and Drogue Chutes and the Frustum contains the three Main Parachutes, Altitude Switches and forward booster Separation Motors. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  8. Improved hybrid rocket fuel

    NASA Technical Reports Server (NTRS)

    Dean, David L.

    1995-01-01

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

  9. The Advanced Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

    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.

  10. The Advanced Solid Rocket Motor

    NASA Astrophysics Data System (ADS)

    Mitchell, Royce E.

    1992-08-01

    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.

  11. The KamLAND Full-Volume Calibration System

    SciTech Connect

    KamLAND Collaboration; Berger, B. E.; Busenitz, J.; Classen, T.; Decowski, M. P.; Dwyer, D. A.; Elor, G.; Frank, A.; Freedman, S. J.; Fujikawa, B. K.; Galloway, M.; Gray, F.; Heeger, K. M.; Hsu, L.; Ichimura, K.; Kadel, R.; Keefer, G.; Lendvai, C.; McKee, D.; O'Donnell, T.; Piepke, A.; Steiner, H. M.; Syversrud, D.; Wallig, J.; Winslow, L. A.; Ebihara, T.; Enomoto, S.; Furuno, K.; Gando, Y.; Ikeda, H.; Inoue, K.; Kibe, Y.; Kishimoto, Y.; Koga, M.; Minekawa, Y.; Mitsui, T.; Nakajima, K.; Nakajima, K.; Nakamura, K.; Owada, K.; Shimizu, I.; Shimizu, Y.; Shirai, J.; Suekane, F.; Suzuki, A.; Tamae, K.; Yoshida, S.; Kozlov, A.; Murayama, H.; Grant, C.; Leonard, D. S.; Luk, K.-B.; Jillings, C.; Mauger, C.; McKeown, R. D.; Zhang, C.; Lane, C. E.; Maricic, J.; Miletic, T.; Batygov, M.; Learned, J. G.; Matsuno, S.; Pakvasa, S.; Foster, J.; Horton-Smith, G. A.; Tang, A.; Dazeley, S.; Downum, K. E.; Gratta, G.; Tolich, K.; Bugg, W.; Efremenko, Y.; Kamyshkov, Y.; Perevozchikov, O.; Karwowski, H. J.; Markoff, D. M.; Tornow, W.; Piquemal, F.; Ricol, J.-S.

    2009-03-05

    We have successfully built and operated a source deployment system for the KamLAND detector. This system was used to position radioactive sources throughout the delicate 1-kton liquid scintillator volume, while meeting stringent material cleanliness, material compatibility, and safety requirements. The calibration data obtained with this device were used to fully characterize detector position and energy reconstruction biases. As a result, the uncertainty in the size of the detector fiducial volume was reduced by a factor of two. Prior to calibration with this system, the fiducial volume was the largest source of systematic uncertainty in measuring the number of antineutrinos detected by KamLAND. This paper describes the design, operation and performance of this unique calibration system.

  12. Low volume packaging for a microinstrumentation system

    NASA Technical Reports Server (NTRS)

    Mason, A.; Wise, K.

    1995-01-01

    A folding, multi-platform assembly has been developed for the packaging of a multi-chip microinstrumentation system. The assembly includes three solid platforms connected by flexible micromachined ribbon cables and it can be populated by sensors and electronics from a variety of technologies including surface mount and integrated circuit (IC) processes. The entire structure is fabricated from a four-inch silicon wafer using a simple four mask process and a post-process EDP etch. The micromachined ribbon cables allow the platform assembly to be folded into a three level structure with control electronics on the bottom level, microsensors and interface electronics on the second level, and sensors that need environmental access on the top level. Using the silicon multi-platform assembly, a prototype microinstrumentation system has been developed that includes a microcontroller unit and sensors for measuring temperature, barometric pressure, humidity, altitude, and acceleration as well as a telemetry device for wireless communication. When folded in the three level structure, this microsystem occupies only 5 cc and can be placed in an outer case the size of a wristwatch.

  13. Rocket Motor Microphone Investigation

    NASA Technical Reports Server (NTRS)

    Pilkey, Debbie; Herrera, Eric; Gee, Kent L.; Giraud, Jerom H.; Young, Devin J.

    2010-01-01

    At ATK's facility in Utah, large full-scale solid rocket motors are tested. The largest is a five-segment version of the reusable solid rocket motor, which is for use on the Ares I launch vehicle. As a continuous improvement project, ATK and BYU investigated the use of microphones on these static tests, the vibration and temperature to which the instruments are subjected, and in particular the use of vent tubes and the effects these vents have at low frequencies.

  14. A reusable rocket engine intelligent control

    NASA Technical Reports Server (NTRS)

    Merrill, Walter C.; Lorenzo, Carl F.

    1988-01-01

    An intelligent control system for reusable space propulsion systems for future launch vehicles is described. The system description includes a framework for the design. The framework consists of an execution level with high-speed control and diagnostics, and a coordination level which marries expert system concepts with traditional control. A comparison is made between air breathing and rocket engine control concepts to assess the relative levels of development and to determine the applicability of air breathing control concepts ot future reusable rocket engine systems.

  15. Radial Internal Material Handling System (RIMS) for Circular Habitat Volumes

    NASA Technical Reports Server (NTRS)

    Howe, A. Scott; Haselschwardt, Sally

    2012-01-01

    A Radial Internal Material Handling System (RIMS) has been developed to service a circular floor area in variable gravity. On planetary surfaces, pressurized human habitable volumes will require a means to carry heavy equipment between various locations within the volume of the habitat, regardless of the partial gravity (Earth, moon, Mars, etc). On the NASA Habitat Demonstration Unit (HDU), a vertical cylindrical volume, it was determined that a variety of heavy items would need to be carried back and forth from deployed locations to the General Maintenance Work Station (GMWS) when in need of repair, and other equipment may need to be carried inside for repairs, such as rover parts and other external equipment. The vertical cylindrical volume of the HDU lent itself to a circular overhead track and hoist system that allows lifting of heavy objects from anywhere in the habitat to any other point in the habitat interior. In addition, the system is able to hand off lifted items to other material handling systems through the side hatches, such as through an airlock. This paper describes the RIMS system which is scalable for application in a variety of circular habitat volumes.

  16. Dynamic gas temperature measurement system, volume 1

    NASA Technical Reports Server (NTRS)

    Elmore, D. L.; Robinson, W. W.; Watkins, W. B.

    1983-01-01

    A gas temperature measurement system with compensated frequency response of 1 kHz and capability to operate in the exhaust of a gas turbine engine combustor was developed. A review of available technologies which could attain this objective was done. The most promising method was identified as a two wire thermocouple, with a compensation method based on the responses of the two different diameter thermocouples to the fluctuating gas temperature field. In a detailed design of the probe, transient conduction effects were identified as significant. A compensation scheme was derived to include the effects of gas convection and wire conduction. The two wire thermocouple concept was tested in a laboratory burner exhaust to temperatures of about 3000 F and in a gas turbine engine to combustor exhaust temperatures of about 2400 F. Uncompensated and compensated waveforms and compensation spectra are presented.

  17. Space Tug systems study. Volume 2: Compendium

    NASA Technical Reports Server (NTRS)

    1974-01-01

    Possible storable propellant configurations and program plans are evaluated for the space tug. Alternatives examined include: use of existing expendable stages modified for use with shuttle, followed by a space tug at a later date; use of a modified growth version of existing expendable stages for greater performance and potential reuse, followed by a space tug at a later date; use of a low development cost, reusable, interim space tug available at shuttle initial operational capability (IOC) that could be evolved to greater system capabilities at a later date; and use a direct developed tug with maximum potential to be available at some specified time after space shuttle IOC. The capability options were narrowed down to three final options for detailed program definition.

  18. Infrasound Rocket Signatures

    NASA Astrophysics Data System (ADS)

    Olson, J.

    2012-09-01

    This presentation reviews the work performed by our research group at the Geophysical Institute as we have applied the tools of infrasound research to rocket studies. This report represents one aspect of the effort associated with work done for the National Consortium for MASINT Research (NCMR) program operated by the National MASINT Office (NMO) of the Defense Intelligence Agency (DIA). Infrasound, the study of acoustic signals and their propagation in a frequency band below 15 Hz, enables an investigator to collect and diagnose acoustic signals from distant sources. Absorption of acoustic energy in the atmosphere decreases as the frequency is reduced. In the infrasound band signals can propagate hundreds and thousands of kilometers with little degradation. We will present an overview of signatures from rockets ranging from small sounding rockets such as the Black Brandt and Orion series to larger rockets such as Delta 2,4 and Atlas V. Analysis of the ignition transients provides information that can uniquely identify the motor type. After the rocket ascends infrasound signals can be used to characterize the rocket and identify the various events that take place along a trajectory such as staging and maneuvering. We have also collected information on atmospheric shocks and sonic booms from the passage of supersonic vehicles such as the shuttle. This review is intended to show the richness of the unique signal set that occurs in the low-frequency infrasound band.

  19. 21 CFR 862.1130 - Blood volume test system.

    Code of Federal Regulations, 2013 CFR

    2013-04-01

    ... 21 Food and Drugs 8 2013-04-01 2013-04-01 false Blood volume test system. 862.1130 Section 862.1130 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) MEDICAL DEVICES CLINICAL CHEMISTRY AND CLINICAL TOXICOLOGY DEVICES Clinical Chemistry Test Systems §...

  20. 21 CFR 862.1130 - Blood volume test system.

    Code of Federal Regulations, 2014 CFR

    2014-04-01

    ... 21 Food and Drugs 8 2014-04-01 2014-04-01 false Blood volume test system. 862.1130 Section 862.1130 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) MEDICAL DEVICES CLINICAL CHEMISTRY AND CLINICAL TOXICOLOGY DEVICES Clinical Chemistry Test Systems §...

  1. 21 CFR 862.1130 - Blood volume test system.

    Code of Federal Regulations, 2010 CFR

    2010-04-01

    ... 21 Food and Drugs 8 2010-04-01 2010-04-01 false Blood volume test system. 862.1130 Section 862.1130 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) MEDICAL DEVICES CLINICAL CHEMISTRY AND CLINICAL TOXICOLOGY DEVICES Clinical Chemistry Test Systems §...

  2. 21 CFR 862.1130 - Blood volume test system.

    Code of Federal Regulations, 2011 CFR

    2011-04-01

    ... 21 Food and Drugs 8 2011-04-01 2011-04-01 false Blood volume test system. 862.1130 Section 862.1130 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) MEDICAL DEVICES CLINICAL CHEMISTRY AND CLINICAL TOXICOLOGY DEVICES Clinical Chemistry Test Systems §...

  3. 21 CFR 862.1130 - Blood volume test system.

    Code of Federal Regulations, 2012 CFR

    2012-04-01

    ... 21 Food and Drugs 8 2012-04-01 2012-04-01 false Blood volume test system. 862.1130 Section 862.1130 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) MEDICAL DEVICES CLINICAL CHEMISTRY AND CLINICAL TOXICOLOGY DEVICES Clinical Chemistry Test Systems §...

  4. Spray Diagnostics in Rocket Engines Using Phase Doppler Analyzer

    NASA Technical Reports Server (NTRS)

    Isaac, Kakkattukuzhy M.

    1996-01-01

    Characteristics such as drop velocity, drop size, number density, volume flow rate, volume flux, and evaporation rate of the fuel spray in a rocket engine are directly related to engine performance. Several studies of shear coaxial atomization have been done in the past. However, additional work related to sprays at supercritical and transcritical conditions would be useful. The author undertook a study of the feasibility of using a phase Doppler particle analyzer (PDPA) for spray measurements in rocket engines as a part of the Summer Faculty Fellowship Program at the NASA Marshall Space Flight Center. The PDPA is a single particle counter (SPC) system based on light scattering from spherical particles. The PDPA instrument is based on refractive and reflective scattering as opposed to other instruments based on diffractive scattering. The main advantage of the PDPA instrument is its ability to provide point-wise information. The following sections describe the principle of operation of the PDPA system and the data obtained using a PDPA instrument in a spray formed by a commercially available fuel injection nozzle.

  5. Rocket Engine Altitude Simulation Technologies

    NASA Technical Reports Server (NTRS)

    Woods, Jody L.; Lansaw, John

    2010-01-01

    John C. Stennis Space Center is embarking on a very ambitious era in its rocket engine propulsion test history. The first new large rocket engine test stand to be built at Stennis Space Center in over 40 years is under construction. The new A3 Test Stand is designed to test very large (294,000 Ibf thrust) cryogenic propellant rocket engines at a simulated altitude of 100,000 feet. A3 Test Stand will have an engine testing chamber where the engine will be fired after the air in the chamber has been evacuated to a pressure at the simulated altitude of less than 0.16 PSIA. This will result in a very unique environment with extremely low pressures inside a very large chamber and ambient pressures outside this chamber. The test chamber is evacuated of air using a 2-stage diffuser / ejector system powered by 5000 lb/sec of steam produced by 27 chemical steam generators. This large amount of power and flow during an engine test will result in a significant acoustic and vibrational environment in and around A3 Test Stand.

  6. Closeup view of the Solid Rocket Booster (SRB) Forward Skirt ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Close-up view of the Solid Rocket Booster (SRB) Forward Skirt sitting on ground support equipment in the Solid Rocket Booster Assembly and Refurbishment Facility at Kennedy Space Center while being prepared for mating with the Frustum-Nose Cap Assembly and the Forward Rocket Motor Segment. The prominent feature in this view is the Forward Thrust Attach Fitting which mates up with the Forward Thrust Attach Fitting of the External Tank (ET) at the ends of the SRB Beam that runs through the ET's Inter Tank Assembly. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  7. Preliminary study of a hydrogen peroxide rocket for use in moving source jet noise tests

    NASA Technical Reports Server (NTRS)

    Plencner, R. M.

    1977-01-01

    A preliminary investigation was made of using a hydrogen peroxide rocket to obtain pure moving source jet noise data. The thermodynamic cycle of the rocket was analyzed. It was found that the thermodynamic exhaust properties of the rocket could be made to match those of typical advanced commercial supersonic transport engines. The rocket thruster was then considered in combination with a streamlined ground car for moving source jet noise experiments. When a nonthrottlable hydrogen peroxide rocket was used to accelerate the vehicle, propellant masses and/or acceleration distances became too large. However, when a throttlable rocket or an auxiliary system was used to accelerate the vehicle, reasonable propellant masses could be obtained.

  8. Rocket Engine Numerical Simulator (RENS)

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth O.

    1997-01-01

    Work is being done at three universities to help today's NASA engineers use the knowledge and experience of their Apolloera predecessors in designing liquid rocket engines. Ground-breaking work is being done in important subject areas to create a prototype of the most important functions for the Rocket Engine Numerical Simulator (RENS). The goal of RENS is to develop an interactive, realtime application that engineers can utilize for comprehensive preliminary propulsion system design functions. RENS will employ computer science and artificial intelligence research in knowledge acquisition, computer code parallelization and objectification, expert system architecture design, and object-oriented programming. In 1995, a 3year grant from the NASA Lewis Research Center was awarded to Dr. Douglas Moreman and Dr. John Dyer of Southern University at Baton Rouge, Louisiana, to begin acquiring knowledge in liquid rocket propulsion systems. Resources of the University of West Florida in Pensacola were enlisted to begin the process of enlisting knowledge from senior NASA engineers who are recognized experts in liquid rocket engine propulsion systems. Dr. John Coffey of the University of West Florida is utilizing his expertise in interviewing and concept mapping techniques to encode, classify, and integrate information obtained through personal interviews. The expertise extracted from the NASA engineers has been put into concept maps with supporting textual, audio, graphic, and video material. A fundamental concept map was delivered by the end of the first year of work and the development of maps containing increasing amounts of information is continuing. Find out more information about this work at the Southern University/University of West Florida. In 1996, the Southern University/University of West Florida team conducted a 4day group interview with a panel of five experts to discuss failures of the RL10 rocket engine in conjunction with the Centaur launch vehicle. The

  9. Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Appendix F: Performance and trajectory for ALS/LRB launch vehicles

    NASA Technical Reports Server (NTRS)

    1989-01-01

    By simply combining two baseline pump-fed LOX/RP-1 Liquid Rocket Boosters (LRBs) with the Denver core, a launch vehicle (Option 1 Advanced Launch System (ALS)) is obtained that can perform both the 28.5 deg (ALS) mission and the polar orbit ALS mission. The Option 2 LRB was obtained by finding the optimum LOX/LH2 engine for the STS/LRB reference mission (70.5 K lb payload). Then this engine and booster were used to estimate ALS payload for the 28.5 deg inclination ALS mission. Previous studies indicated that the optimum number of STS/LRB engines is four. When the engine/booster sizing was performed, each engine had 478 K lb sea level thrust and the booster carried 625,000 lb of useable propellant. Two of these LRBs combined with the Denver core provided a launch vehicle that meets the payload requirements for both the ALS and STS reference missions. The Option 3 LRB uses common engines for the cores and boosters. The booster engines do not have the nozzle extension. These engines were sized as common ALS engines. An ALS launch vehicle that has six core engines and five engines per booster provides 109,100 lb payload for the 28.5 deg mission. Each of these LOX/LH2 LRBs carries 714,100 lb of useable propellant. It is estimated that the STS/LRB reference mission payload would be 75,900 lb.

  10. Liquid rocket engine nozzles

    NASA Technical Reports Server (NTRS)

    1976-01-01

    The nozzle is a major component of a rocket engine, having a significant influence on the overall engine performance and representing a large fraction of the engine structure. The design of the nozzle consists of solving simultaneously two different problems: the definition of the shape of the wall that forms the expansion surface, and the delineation of the nozzle structure and hydraulic system. This monography addresses both of these problems. The shape of the wall is considered from immediately upstream of the throat to the nozzle exit for both bell and annular (or plug) nozzles. Important aspects of the methods used to generate nozzle wall shapes are covered for maximum-performance shapes and for nozzle contours based on criteria other than performance. The discussion of structure and hydraulics covers problem areas of regeneratively cooled tube-wall nozzles and extensions; it treats also nozzle extensions cooled by turbine exhaust gas, ablation-cooled extensions, and radiation-cooled extensions. The techniques that best enable the designer to develop the nozzle structure with as little difficulty as possible and at the lowest cost consistent with minimum weight and specified performance are described.

  11. Outbrief - Long Life Rocket Engine Panel

    NASA Technical Reports Server (NTRS)

    Quinn, Jason Eugene

    2004-01-01

    This white paper is an overview of the JANNAF Long Life Rocket Engine (LLRE) Panel results from the last several years of activity. The LLRE Panel has met over the last several years in order to develop an approach for the development of long life rocket engines. Membership for this panel was drawn from a diverse set of the groups currently working on rocket engines (Le. government labs, both large and small companies and university members). The LLRE Panel was formed in order to determine the best way to enable the design of rocket engine systems that have life capability greater than 500 cycles while meeting or exceeding current performance levels (Specific Impulse and Thrust/Weight) with a 1/1,OOO,OOO likelihood of vehicle loss due to rocket system failure. After several meetings and much independent work the panel reached a consensus opinion that the primary issues preventing LLRE are a lack of: physics based life prediction, combined loads prediction, understanding of material microphysics, cost effective system level testing. and the inclusion of fabrication process effects into physics based models. With the expected level of funding devoted to LLRE development, the panel recommended that fundamental research efforts focused on these five areas be emphasized.

  12. Rocket noise - A review

    NASA Astrophysics Data System (ADS)

    McInerny, S. A.

    1990-10-01

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

  13. ROCKET PORT CLOSURE

    DOEpatents

    Mattingly, J.T.

    1963-02-12

    This invention provides a simple pressure-actuated closure whereby windowless observation ports are opened to the atmosphere at preselected altitudes. The closure comprises a disk which seals a windowless observation port in rocket hull. An evacuated instrument compartment is affixed to the rocket hull adjacent the inner surface of the disk, while the outer disk surface is exposed to the atmosphere through which the rocket is traveling. The pressure differential between the evacuated instrument compartment and the relatively high pressure external atmosphere forces the disk against the edge of the observation port, thereby effecting a tight seai. The instrument compartment is evacuated to a pressure equal to the atmospheric pressure existing at the altitude at which it is desiretl that the closure should open. When the rocket reaches this preselected altitude, the inwardly directed atmospheric force on the disk is just equaled by the residual air pressure force within the instrument compartment. Consequently, the closure disk falls away and uncovers the open observation port. The separation of the disk from the rocket hull actuates a switch which energizes the mechanism of a detecting instrument disposed within the instrument compartment. (AE C)

  14. Verification and validation guidelines for high integrity systems. Volume 1

    SciTech Connect

    Hecht, H.; Hecht, M.; Dinsmore, G.; Hecht, S.; Tang, D.

    1995-03-01

    High integrity systems include all protective (safety and mitigation) systems for nuclear power plants, and also systems for which comparable reliability requirements exist in other fields, such as in the process industries, in air traffic control, and in patient monitoring and other medical systems. Verification aims at determining that each stage in the software development completely and correctly implements requirements that were established in a preceding phase, while validation determines that the overall performance of a computer system completely and correctly meets system requirements. Volume I of the report reviews existing classifications for high integrity systems and for the types of errors that may be encountered, and makes recommendations for verification and validation procedures, based on assumptions about the environment in which these procedures will be conducted. The final chapter of Volume I deals with a framework for standards in this field. Volume II contains appendices dealing with specific methodologies for system classification, for dependability evaluation, and for two software tools that can automate otherwise very labor intensive verification and validation activities.

  15. Satellite communications systems and technology. Volume 2; Site Reports

    NASA Technical Reports Server (NTRS)

    Edelson, Burton I.; Pelton, Joseph N.; Bostian, Carles W.; Brandon, William T.; Chan, Vincent W. S.; Hager, E. Paul; Helm, Neil R.; Jennings, Raymond D.; Kwan, Robert K.; Mahle, Christoph E.; Miller, Edward F.; Riley, Lance

    1993-01-01

    Volume 2 of the final report of the NASA/NSF Panel on Satellite Communications Systems and Technology is presented. It consists of the site reports from the panel's visits to satellite communications facilities and laboratories in Europe, Japan, and Russia.

  16. Space shuttle/food system study. Volume 1: Technical volume, oven study

    NASA Technical Reports Server (NTRS)

    1975-01-01

    The baseline space shuttle galley was designed to utilize lightweight rehydratable foods, to be prepared for consumption by rehydration with chilled or hot water. The impact is examined of an extension of food types to include thermostabilized food, at ambient temperature, and frozen foods on the baseline design of the shuttle galley. Weight, volume, and power penalities associated with heating thermostabilized and frozen foods by means of a hot air convection heating system and a conduction heating system are determined along with the impact on crew/galley interface and meal preparation.

  17. Automated and Manual Rocket Crater Measurement Software

    NASA Technical Reports Server (NTRS)

    Metzger, Philip; Immer, Christopher

    2012-01-01

    An update has been performed to software designed to do very rapid automated measurements of craters created in sandy substrates by rocket exhaust on liftoff. The previous software was optimized for pristine lab geometry and lighting conditions. This software has been enhanced to include a section for manual measurements of crater parameters; namely, crater depth, crater full width at half max, and estimated crater volume. The tools provide a very rapid method to measure these manual parameters to ease the burden of analyzing large data sets. This software allows for rapid quantization of the rocket crater parameters where automated methods may not work. The progress of spreadsheet data is continuously saved so that data is never lost, and data can be copied to clipboards and pasted to other software for analysis. The volume estimation of a crater is based on the central max depth axis line, and the polygonal shape of the crater is integrated around that axis.

  18. 2000 NASA Seal/Secondary Air System Workshop. Volume 1

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M. (Editor); Hendricks, Robert C. (Editor)

    2001-01-01

    The 2000 NASA Seal/Secondary Air System Workshop covered four main areas: (1) overviews of NASA-sponsored Ultra-Efficient Engine Technology (UEET) and Access to Space Programs, with emphasis on program goals and seal needs; (2) review of turbine engine seal issues from the perspective of end users such as United Airlines; (3) reviews of sealing concepts, test results, experimental facilities, and numerical predictions; and (4) reviews of material development programs relevant to advanced seals development. The NASA UEET overview illustrates for the reader the importance of advanced technologies, including seals, in meeting future engine system efficiency and emission goals. GE, Pratt & Whitney, and Honeywell presented advanced seal development work being performed within their organizations. The NASA-funded GE/Stein Seal team has successfully demonstrated a large (3-ft. diam) aspirating seal that can withstand all anticipated pressures, speeds, and rotor runouts anticipated for a GE90 L.P. turbine balance piston location. GE/Stein Seal are fabricating a full-scale seal to be tested in a GE-90 ground test engine in early 2002. Pratt & Whitney and Stein Seal are investigating carbon seals to accommodate large radial movements anticipated in future geared-fan gearbox locations. Honeywell presented a finger seal design being considered for a high-temperature static combustor location incorporating ceramic finger elements. Successful demonstration of the braided carbon rope thermal barriers to extreme temperatures (5500 F) for short durations provide a new form of very high temperature thermal barrier for future Shuttle solid rocket motor nozzle joints. The X-37, X-38, and future highly reusable launch vehicles pose challenging control surface seal demands that require new seal concepts made from emerging high temperature ceramics and other materials.

  19. Rockets in World War I

    NASA Technical Reports Server (NTRS)

    2004-01-01

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

  20. Advanced vehicle systems assessment. Volume 4: Supporting analyses

    NASA Technical Reports Server (NTRS)

    Hardy, K.

    1985-01-01

    Volume 4 (Supporting Analyses) is part of a five-volume report, Advanced Vehicle Systems Assessment. Thirty-nine individuals, knowledgeable in advanced technology, were interviewed to obtain their preferences. Rankings were calculated for the eight groups they represented, using multiplicative and additive utility models. The four topics for consideration were: (1) preferred range for various battery technologies; (2) preferred battery technology for each of a variety of travel ranges; (3) most promising battery technology, vehicle range combination; and (4) comparison of the most preferred electric vehicle with the methanol-fuled, spark-ignition engine vehicle and with the most preferred of the hybrid vehicles.

  1. General view of a fully assembled Solid Rocket Booster sitting ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    General view of a fully assembled Solid Rocket Booster sitting atop the Mobile Launch Platform in the Vehicle Assembly Building at Kennedy Space Center - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  2. General view of the Aft Rocket Motor mated with the ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    General view of the Aft Rocket Motor mated with the External Tank Attach Ring and Aft Skirt Assembly being transported from the Rotation Processing and Surge Facility to the Vehicle Assembly Building at Kennedy Space Center. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  3. General view of a Solid Rocket Motor Forward Segment in ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    General view of a Solid Rocket Motor Forward Segment in the process of being offloaded from it's railcar inside the Rotation Processing and Surge Facility at Kennedy Space Center. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  4. General view of the Aft Rocket Motor mated with the ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    General view of the Aft Rocket Motor mated with the External Tank Attach Ring and Aft Skirt Assembly in the process of being mounted onto the Mobile Launch Platform in the Vehicle Assembly Building at Kennedy Space Center. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

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

    NASA Technical Reports Server (NTRS)

    Ganapathi, Gani; Narayan, Sri

    2010-01-01

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

  6. Baking Soda and Vinegar Rockets

    ERIC Educational Resources Information Center

    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…

  7. Rocket center Peenemuende - Personal memories

    NASA Technical Reports Server (NTRS)

    Dannenberg, Konrad; Stuhlinger, Ernst

    1993-01-01

    A brief history of Peenemuende, the rocket center where Von Braun and his team developed the A-4 (V-2) rocket under German Army auspices, and the Air Force developed the V-1 (buzz bomb), wire-guided bombs, and rocket planes, is presented. Emphasis is placed on the expansion of operations beginning in 1942.

  8. Mars power system concept definition study. Volume 1: Study results

    NASA Technical Reports Server (NTRS)

    Littman, Franklin D.

    1994-01-01

    A preliminary top level study was completed to define power system concepts applicable to Mars surface applications. This effort included definition of power system requirements and selection of power systems with the potential for high commonality. These power systems included dynamic isotope, Proton Exchange Membrane (PEM) regenerative fuel cell, sodium sulfur battery, photovoltaic, and reactor concepts. Design influencing factors were identified. Characterization studies were then done for each concept to determine system performance, size/volume, and mass. Operations studies were done to determine emplacement/deployment maintenance/servicing, and startup/shutdown requirements. Technology development roadmaps were written for each candidate power system (included in Volume 2). Example power system architectures were defined and compared on a mass basis. The dynamic isotope power system and nuclear reactor power system architectures had significantly lower total masses than the photovoltaic system architectures. Integrated development and deployment time phasing plans were completed for an example DIPS and reactor architecture option to determine the development strategies required to meet the mission scenario requirements.

  9. Preliminary systems design study assessment report. Volume 7, Subsystem concepts

    SciTech Connect

    Mayberry, J.L.; Feizollahi, F.; Del Signore, J.C.

    1992-01-01

    The System Design Study (SDS), part of the Waste Technology Development Department at the Idaho National Engineering Laboratory (INEL), examined techniques available for the remediation of hazardous and transuranic waste stored at the Radioactive Waste Management Complex`s Subsurface Disposal Area at the INEL. Using specific technologies, system concepts for treating the buried waste and the surrounding contaminated soil were evaluated. Evaluation included implementability, effectiveness, and cost. The SDS resulted in the development of technology requirements including demonstration, testing, and evaluation activities needed for implementing each. This volume contains the descriptions and other relevant information of the four subsystems required for most of the ex situ processing systems. This volume covers the metal decontamination and sizing subsystem, soils processing subsystem, low-level waste subsystem, and retrieval subsystem.

  10. JANNAF 17th Propulsion Systems Hazards Subcommittee Meeting. Volume 1

    NASA Technical Reports Server (NTRS)

    Cocchiaro, James E. (Editor); Gannaway, Mary T. (Editor); Rognan, Melanie (Editor)

    1998-01-01

    Volume 1, the first of two volumes is a compilation of 16 unclassified/unlimited technical papers presented at the 17th meeting of the Joint Army-Navy-NASA-Air Force (JANNAF) Propulsion Systems Hazards Subcommittee (PSHS) held jointly with the 35th Combustion Subcommittee (CS) and Airbreathing Propulsion Subcommittee (APS). The meeting was held on 7 - 11 December 1998 at Raytheon Systems Company and the Marriott Hotel, Tucson, AZ. Topics covered include projectile and shaped charge jet impact vulnerability of munitions; thermal decomposition and cookoff behavior of energetic materials; damage and hot spot initiation mechanisms with energetic materials; detonation phenomena of solid energetic materials; and hazard classification, insensitive munitions, and propulsion systems safety.

  11. Rivaroxaban in the Prevention of Stroke and Systemic Embolism in Patients with Non-Valvular Atrial Fibrillation: Clinical Implications of the ROCKET AF Trial and Its Subanalyses.

    PubMed

    Spencer, Ryan J; Amerena, John V

    2015-12-01

    Atrial fibrillation (AF) is an increasingly common cause of stroke and systemic embolism. While warfarin has been the mainstay of stroke prevention in patients with AF, newer novel oral anticoagulant medications are now available. Rivaroxaban, a direct factor Xa inhibitor with a rapid onset and offset after oral administration, offers potential advantages over warfarin, predominantly due to its predictable pharmacokinetics across wide patient populations. It requires no coagulation monitoring, and only two different doses are needed (20 mg daily for patients with normal renal function and 15 mg daily in those with reduced renal function). A large randomized trial (ROCKET AF) has shown non-inferiority to warfarin for preventing stroke or systemic embolism in the per-protocol population and superiority to warfarin in the on-treatment safety population. Several subanalyses confirm that the treatment effect of rivaroxaban is consistent across different patient subgroups, including those with reduced renal function. The tolerability of rivaroxaban appears similar to that of warfarin, with comparable overall bleeding rates in clinical trials. In ROCKET AF, significantly lower rates of fatal and intracranial bleeding were seen with rivaroxaban, while lower rates of gastrointestinal bleeding were seen with warfarin. Important contraindications to rivaroxaban include valvular AF, the presence of a prosthetic valve (mechanical or bioprosthetic) or valve repair, the need for concurrent dual antiplatelet therapy, and creatinine clearance <30 ml/min. Once-daily dosing and the lack of coagulation monitoring may increase utilization and adherence compared with warfarin, potentially decreasing the large burden of care associated with stroke secondary to AF. Overall, rivaroxaban offers a useful alternative to warfarin for stroke prevention in patients with AF.

  12. Large Liquid Rocket Testing: Strategies and Challenges

    NASA Technical Reports Server (NTRS)

    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

  13. Autonomous Attitude Determination System (AADS). Volume 1: System description

    NASA Technical Reports Server (NTRS)

    Saralkar, K.; Frenkel, Y.; Klitsch, G.; Liu, K. S.; Lefferts, E.; Tasaki, K.; Snow, F.; Garrahan, J.

    1982-01-01

    Information necessary to understand the Autonomous Attitude Determination System (AADS) is presented. Topics include AADS requirements, program structure, algorithms, and system generation and execution.

  14. The National Space Science Data Center guide to international rocket data

    NASA Technical Reports Server (NTRS)

    Dubach, L. L.

    1972-01-01

    Background information is given which briefly describes the mission of the National Space Science Data Center (NSSDC), including its functions and systems, along with its policies and purposes for collecting rocket data. The operation of a machine-sensible rocket information system, which allows the Data Center to have convenient access to information and data concerning all rocket flights carrying scientific experiments, is also described. The central feature of this system, an index of rocket flights maintained on magnetic tape, is described. Standard outputs for NSSDC and for the World Data Center A (WDC-A) for Rockets and Satellites are described.

  15. Water Rocket Workout.

    ERIC Educational Resources Information Center

    Esler, William K.; Sanford, Daniel

    1989-01-01

    Water rockets are used to present Newton's three laws of motion to high school physics students. Described is an outdoor activity which uses four students per group. Provides a launch data sheet to record height, angle of elevation, amount of water used, and launch number. (MVL)

  16. This "Is" Rocket Science!

    ERIC Educational Resources Information Center

    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…

  17. The Relativistic Rocket

    ERIC Educational Resources Information Center

    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…

  18. Liquid rocket engine turbines

    NASA Technical Reports Server (NTRS)

    1974-01-01

    Criteria for the design and development of turbines for rocket engines to meet specific performance, and installation requirements are summarized. The total design problem, and design elements are identified, and the current technology pertaining to these elements is described. Recommended practices for achieving a successful design are included.

  19. Solid rocket motors

    NASA Technical Reports Server (NTRS)

    Carpenter, Ronn L.

    1993-01-01

    Structural requirements, materials and, especially, processing are critical issues that will pace the introduction of new types of solid rocket motors. Designers must recognize and understand the drivers associated with each of the following considerations: (1) cost; (2) energy density; (3) long term storage with use on demand; (4) reliability; (5) safety of processing and handling; (6) operability; and (7) environmental acceptance.

  20. Thiokol Solid Rocket Motors

    NASA Technical Reports Server (NTRS)

    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.

  1. Solid rocket motors

    NASA Astrophysics Data System (ADS)

    Carpenter, Ronn L.

    1993-02-01

    Structural requirements, materials and, especially, processing are critical issues that will pace the introduction of new types of solid rocket motors. Designers must recognize and understand the drivers associated with each of the following considerations: (1) cost; (2) energy density; (3) long term storage with use on demand; (4) reliability; (5) safety of processing and handling; (6) operability; and (7) environmental acceptance.

  2. Dr. Goddard Transports Rocket

    NASA Technical Reports Server (NTRS)

    1974-01-01

    Dr. Robert H. Goddard tows his rocket to the launching tower behind a Model A Ford truck, 15 miles northwest of Roswell, New Mexico. 1930- 1932. Dr. Goddard has been recognized as the 'Father of American Rocketry' and as one of three pioneers in the theoretical exploration of space. Robert Hutchings Goddard was born in Worcester, Massachusetts, on October 15, 1882. He was a theoretical scientist as well as a practical engineer. His dream was the conquest of the upper atmosphere and ultimately space through the use of rocket propulsion. Dr. Goddard, who died in 1945, was probably as responsible for the dawning of the Space Age as the Wright Brothers were for the begining of the Air Age. Yet his work attracted little serious attention during his lifetime. When the United States began to prepare for the conquest of space in the 1950's, American rocket scientists began to recognize the debt owed to the New England professor. They discovered that it was virtually impossible to construct a rocket or launch a satellite without acknowledging the work of Dr. Goddard. This great legacy was covered by more than 200 patents, many of which were issued after his death.

  3. Liquid rocket valve assemblies

    NASA Technical Reports Server (NTRS)

    1973-01-01

    The design and operating characteristics of valve assemblies used in liquid propellant rocket engines are discussed. The subjects considered are as follows: (1) valve selection parameters, (2) major design aspects, (3) design integration of valve subassemblies, and (4) assembly of components and functional tests. Information is provided on engine, stage, and spacecraft checkout procedures.

  4. Liquid rocket valve components

    NASA Technical Reports Server (NTRS)

    1973-01-01

    A monograph on valves for use with liquid rocket propellant engines is presented. The configurations of the various types of valves are described and illustrated. Design criteria and recommended practices for the various valves are explained. Tables of data are included to show the chief features of valve components in use on operational vehicles.

  5. This Is Rocket Science!

    NASA Astrophysics Data System (ADS)

    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

  6. Liquid Rocket Engine Testing

    NASA Technical Reports Server (NTRS)

    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.

  7. New York State Educational Information System (NYSEIS) Systems Design. Volume I, Phase II. Final Report.

    ERIC Educational Resources Information Center

    Price Waterhouse and Co., New York, NY.

    This volume on Phase II of the New York State Educational Information System (NYSEIS) describes the Gross Systems Analysis and Design, which includes the general flow diagram and processing chart for each of the student, personnel, and financial subsystems. Volume II, Functional Specifications, includes input/output requirements and file…

  8. Transpiration cooled throat for hydrocarbon rocket engines

    NASA Technical Reports Server (NTRS)

    May, Lee R.; Burkhardt, Wendel M.

    1991-01-01

    The objective for the Transpiration Cooled Throat for Hydrocarbon Rocket Engines Program was to characterize the use of hydrocarbon fuels as transpiration coolants for rocket nozzle throats. The hydrocarbon fuels investigated in this program were RP-1 and methane. To adequately characterize the above transpiration coolants, a program was planned which would (1) predict engine system performance and life enhancements due to transpiration cooling of the throat region using analytical models, anchored with available data; (2) a versatile transpiration cooled subscale rocket thrust chamber was designed and fabricated; (3) the subscale thrust chamber was tested over a limited range of conditions, e.g., coolant type, chamber pressure, transpiration cooled length, and coolant flow rate; and (4) detailed data analyses were conducted to determine the relationship between the key performance and life enhancement variables.

  9. Place of solar thermal rockets in space

    SciTech Connect

    Selph, C.C.

    1981-05-01

    The harnessing of sunlight for propulsive energy is a recurring theme in space propulsion, particularly for applications requiring large velocity increments, such as planetary exploration or comet rendezvous. Characteristically, it is viewed in terms of the solar sail and the solar cell, but for operations in Earth orbit these approaches are less desirable because the very low thrust leads to undesirably long maneuver times. Thrust levels several orders of magnitude higher are available with solar thermal rockets, while preserving a specific impulse advantage over chemical systems. The performance advantages, penalties, technological problems, and approaches were examined for solar thermal rockets. Its suitability in several Earth orbit missions is assessed. The peculiarities of vehicle design, the nature of the thruster and the solar concentrator are presented, and AF plans to implement the development of solar rockets are outlined.

  10. Metallic Hydrogen: A Game Changing Rocket Propellant

    NASA Technical Reports Server (NTRS)

    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.

  11. JANNAF 18th Propulsion Systems Hazards Subcommittee Meeting. Volume 1

    NASA Technical Reports Server (NTRS)

    Cocchiaro, James E. (Editor); Gannaway, Mary T. (Editor)

    1999-01-01

    This volume, the first of two volumes is a compilation of 18 unclassified/unlimited-distribution technical papers presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 18th Propulsion Systems Hazards Subcommittee (PSHS) meeting held jointly with the 36th Combustion Subcommittee (CS) and 24th Airbreathing Propulsion Subcommittee (APS) meetings. The meeting was held 18-21 October 1999 at NASA Kennedy Space Center and The DoubleTree Oceanfront Hotel, Cocoa Beach, Florida. Topics covered at the PSHS meeting include: shaped charge jet and kinetic energy penetrator impact vulnerability of gun propellants; thermal decomposition and cookoff behavior of energetic materials; violent reaction; detonation phenomena of solid energetic materials subjected to shock and impact stimuli; and hazard classification, insensitive munitions, and propulsion systems safety.

  12. Shared Communications: Volume 2. In-Depth Systems Research

    SciTech Connect

    Truett, LF

    2004-09-22

    This report is the second of two documents that examine the literature for actual examples of organizations and agencies that share communications resources. While the primary emphasis is on rural, intelligent transportation system (ITS) communications involving transit, examples will not be limited to rural activities, nor to ITS implementation, nor even to transit. In addition, the term ''communication'' will be broadly applied to include all information resources. The first document of this series, ''Shared Communications: Volume I. A Summary and Literature Review'', defines the meaning of the term ''shared communication resources'' and provides many examples of agencies that share resources. This document, ''Shared Communications: Volume II. In-Depth Systems Research'', reviews attributes that contributed to successful applications of the sharing communication resources concept. A few examples of each type of communication sharing are provided. Based on the issues and best practice realworld examples, recommendations for potential usage and recommended approaches for field operational tests are provided.

  13. JANNAF 19th Propulsion Systems Hazards Subcommittee Meeting. Volume 1

    NASA Technical Reports Server (NTRS)

    Cocchiaro, James E. (Editor); Kuckels, Melanie C. (Editor)

    2000-01-01

    This volume, the first of two volumes is a compilation of 25 unclassified/unlimited-distribution technical papers presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 19th Propulsion Systems Hazards Subcommittee (PSHS) meeting held jointly with the 37th Combustion Subcommittee (CS) and 25th Airbreathing Propulsion Subcommittee (APS), and 1st Modeling and Simulation Subcommittee (MSS) meetings. The meeting was held 13-17 November 2000 at the Naval Postgraduate School and Hyatt Regency Hotel, Monterey, California. Topics covered at the PSHS meeting include: impact and thermal vulnerability of gun propellants; thermal decomposition and cookoff behavior of energetic materials; violent reaction and detonation phenomena of solid energetic materials subjected to shock and impact loading; and hazard classification, and insensitive munitions testing of propellants and propulsion systems.

  14. Rockets using Liquid Oxygen

    NASA Technical Reports Server (NTRS)

    Busemann, Adolf

    1947-01-01

    It is my task to discuss rocket propulsion using liquid oxygen and my treatment must be highly condensed for the ideas and experiments pertaining to this classic type of rocket are so numerous that one could occupy a whole morning with a detailed presentation. First, with regard to oxygen itself as compared with competing oxygen carriers, it is known that the liquid state of oxygen, in spite of the low boiling point, is more advantageous than the gaseous form of oxygen in pressure tanks, therefore only liquid oxygen need be compared with the oxygen carriers. The advantages of liquid oxygen are absolute purity and unlimited availability at relatively small cost in energy. The disadvantages are those arising from the impossibility of absolute isolation from heat; consequently, allowance must always be made for a certain degree of vaporization and only vented vessels can be used for storage and transportation. This necessity alone eliminates many fields of application, for example, at the front lines. In addition, liquid oxygen has a lower specific weight than other oxygen carriers, therefore many accessories become relatively larger and heavier in the case of an oxygen rocket, for example, the supply tanks and the pumps. The advantages thus become effective only in those cases where definitely scheduled operation and a large ground organization are possible and when the flight requires a great concentration of energy relative to weight. With the aim of brevity, a diagram of an oxygen rocket will be presented and the problem of various component parts that receive particularly thorough investigation in this classic case but which are also often applicable to other rocket types will be referred to.

  15. Advanced Solid Rocket Motor case design status

    NASA Technical Reports Server (NTRS)

    Palmer, G. L.; Cash, S. F.; Beck, J. P.

    1993-01-01

    The Advanced Solid Rocket Motor (ASRM) case design aimed at achieving a safer and more reliable solid rocket motor for the Space Shuttle system is considered. The ASRM case has a 150.0 inch diameter, three equal length segment, and 9Ni-4CO-0.3C steel alloy. The major design features include bolted casebolted case joints which close during pressurization, plasma arc welded factory joints, integral stiffener for splash down and recovery, and integral External Tank attachment rings. Each mechanical joint has redundant and verifiable o-ring seals.

  16. Analytical investigation of solid rocket nozzle failure

    NASA Technical Reports Server (NTRS)

    Mccoy, K. E.; Hester, J.

    1985-01-01

    On April 5, 1983, an Inertial Upper Stage (IUS) spacecraft experienced loss of control during the burn of the second of two solid rocket motors. The anomaly investigation showed the cause to be a malfunction of the solid rocket motor. This paper presents a description of the IUS system, a failure analysis summary, an account of the thermal testing and computer modeling done at Marshall Space Flight Center, a comparison of analysis results with thermal data obtained from motor static tests, and describes some of the design enhancement incorporated to prevent recurrence of the anomaly.

  17. Composite Material Application to Liquid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Judd, D. C.

    1982-01-01

    The substitution of reinforced plastic composite (RPC) materials for metal was studied. The major objectives were to: (1) determine the extent to which composite materials can be beneficially used in liquid rocket engines; (2) identify additional technology requirements; and (3) determine those areas which have the greatest potential for return. Weight savings, fabrication costs, performance, life, and maintainability factors were considered. Two baseline designs, representative of Earth to orbit and orbit to orbit engine systems, were selected. Weight savings are found to be possible for selected components with the substitution of materials for metal. Various technology needs are identified before RPC material can be used in rocket engine applications.

  18. Space aging of solid rocket materials

    NASA Technical Reports Server (NTRS)

    Lester, Dean M.; Jones, Leon L.; Smalley, R. B., Jr.; Ord, R. Neil

    1991-01-01

    Solid rocket propellant and rocket motor components were aged in a vented container on the interior of the LDEF. The results of aging IPSM-II/PAM-DII space motor components are presented. Ballistic and mechanical properties of the space aged main propellant, igniter propellant, and ignition system were compared with similar data from preflight and ground aged samples. Mechanical properties of the composite materials and bonded joints used in the motor case, insulation, liner, nozzle, exit cone, and skirt were similarly evaluated. The space aging results are compared to data collected in a ground based vacuum aging program on similar components.

  19. Calibrated Ancillary System (CAS) user's guide, volume 6

    NASA Technical Reports Server (NTRS)

    1986-01-01

    The Calibrated Ancillary System (CAS) provides real-time calibrated parameters from the orbiter downlink (ancillary data) to the Goddard Space Flight Center (GSFC). This user's guide contains the introduction to the equipment, operation, general procedures, and specific procedures of CAS. Volume 6 describes ancillary products procedures, enhancement menu and processing task procedures for SDT/TDT (shuttle data tape/telemetry descriptor tape), database errors and network data driver (NDD) product menu procedures, and utility menu procedures.

  20. Combined Global Navigation Satellite Systems in the Space Service Volume

    NASA Technical Reports Server (NTRS)

    Force, Dale A.; Miller, James J.

    2013-01-01

    Besides providing position, velocity, and timing (PVT) for terrestrial users, the Global Positioning System (GPS) is also being used to provide PVT information for earth orbiting satellites. In 2006, F. H. Bauer, et. al., defined the Space Service Volume in the paper GPS in the Space Service Volume , presented at ION s 19th international Technical Meeting of the Satellite Division, and looked at GPS coverage for orbiting satellites. With GLONASS already operational, and the first satellites of the Galileo and Beidou/COMPASS constellations already in orbit, it is time to look at the use of the new Global Navigation Satellite Systems (GNSS) coming into service to provide PVT information for earth orbiting satellites. This presentation extends GPS in the Space Service Volume by examining the coverage capability of combinations of the new constellations with GPS GPS was first explored as a system for refining the position, velocity, and timing of other spacecraft equipped with GPS receivers in the early eighties. Because of this, a new GPS utility developed beyond the original purpose of providing position, velocity, and timing services for land, maritime, and aerial applications. GPS signals are now received and processed by spacecraft both above and below the GPS constellation, including signals that spill over the limb of the earth. Support of GPS space applications is now part of the system plan for GPS, and support of the Space Service Volume by other GNSS providers has been proposed to the UN International Committee on GNSS (ICG). GPS has been demonstrated to provide decimeter level position accuracy in real-time for satellites in low Earth orbit (centimeter level in non-real-time applications). GPS has been proven useful for satellites in geosynchronous orbit, and also for satellites in highly elliptical orbits. Depending on how many satellites are in view, one can keep time locked to the GNSS standard, and through that to Universal Time as long as at least one