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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. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. Rocket engine condition monitoring system

    SciTech Connect

    Hagar, S.K.; Alcock, J.F.

    1989-01-01

    It is expected that the Rocket Engine Condition Monitoring System (RECMS) program will define engine monitoring technologies and an integration approach which can be applied to engine development in support of advanced launch system objectives. The RECMS program approaches engine monitoring as a system which is fully integrated with the engine controller, vehicle monitoring system, and ground processing systems to ensure mission success in addition to engine reliability. The system components are monitored through health and performance sensors; they are analyzed with the diagnostic and prognostic algorithms and demonstrated by system testing with hardware from other advanced development programs.

  15. Solar rocket system concept analysis

    NASA Technical Reports Server (NTRS)

    Boddy, J. A.

    1980-01-01

    The use of solar energy to heat propellant for application to Earth orbital/planetary propulsion systems is of interest because of its performance capabilities. The achievable specific impulse values are approximately double those delivered by a chemical rocket system, and the thrust is at least an order of magnitude greater than that produced by a mercury bombardment ion propulsion thruster. The primary advantage the solar heater thruster has over a mercury ion bombardment system is that its significantly higher thrust permits a marked reduction in mission trip time. The development of the space transportation system, offers the opportunity to utilize the full performance potential of the solar rocket. The requirements for transfer from low Earth orbit (LEO) to geosynchronous equatorial orbit (GEO) was examined as the return trip, GEO to LEO, both with and without payload. Payload weights considered ranged from 2000 to 100,000 pounds. The performance of the solar rocket was compared with that provided by LO2-LH2, N2O4-MMH, and mercury ion bombardment systems.

  16. Pressurization systems for liquid rockets

    NASA Technical Reports Server (NTRS)

    1975-01-01

    Guidelines for the successful design of pressurization systems for main propulsion, auxiliary propulsion, and attitude control systems for boosters, upper stages, and spacecraft were presented, drawing on the wealth of design experience that has accumulated in the development of pressurization systems for liquid rockets operational in the last 15 years. The design begins with a preliminary phase in which the system requirements are received and evaluated. Next comes a detail-design and integration phase in which the controls and the hardware components that make up the system are determined. The final phase, design evaluation, provides analysis of problems that may arise at any point in the design when components are combined and considered for operation as a system. Throughout the monograph, the design tasks are considered in the order and manner in which the designer must handle them.

  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. 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.

  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. 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.

  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. 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.

  3. 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.

  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. 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.

  6. Turbopump systems for liquid rocket engines

    NASA Technical Reports Server (NTRS)

    1974-01-01

    The turbopump system, from preliminary design through rocket engine testing is examined. Selection of proper system type for each application and integration of the components into a working system are dealt with. Details are also given on the design of various components including inducers, pumps, turbines, gears, and bearings.

  7. 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.

  8. 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.

  9. 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

  10. 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.

  11. 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.

  12. 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.

  13. Rocket engine control and monitoring expert system

    NASA Technical Reports Server (NTRS)

    Ali, Moonis; Crawford, Roger

    1988-01-01

    This paper focuses on the application of expert systems technology to the automatic detection, verification and correction of anomalous rocket engine operations through interfacing with an intelligent adaptive control system. The design of a reliable and intelligent propulsion control and monitoring system is outlined which includes the architecture of an Integrated Expert System (IES) serving as the core component. The IES functions include automatic knowledge acquisition, integrated knowledge base, and fault diagnosis and prediction methodology. The results of fault analysis and diagnostic techniques are presented for an example fault in the SSME main combustion chamber injectors.

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

    NASA Technical Reports Server (NTRS)

    1989-01-01

    The purpose of this study was to determine the feasibility of Liquid Rocket Boosters (LRBs) replacing Solid Rocket Boosters on the Space Shuttle program. The major findings are given. The most significant conclusion is that LRBs offer significantly safety and performance advantages over the SRBs currently used by the STS without major impact to the ongoing program.

  15. 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.

  16. 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.

  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. 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.

  19. Reusable rocket engine turbopump health management system

    NASA Astrophysics Data System (ADS)

    Surko, Pamela

    1994-10-01

    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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  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) 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.

  7. 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.

  8. 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.

  9. 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.

  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.

    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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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

  17. 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

  18. 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

  19. 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

  20. 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.

  1. Design study of RL10 derivatives. Volume 2: Engine design characteristics, appendices. [development of rocket engine for application to space tug propulsion system

    NASA Technical Reports Server (NTRS)

    1973-01-01

    Calculations, curves, and substantiating data which support the engine design characteristics of the RL-10 engines are presented. A description of the RL-10 ignition system is provided. The performance calculations of the RL-10 derivative engines and the performance results obtained are reported. The computer simulations used to establish the control system requirements and to define the engine transient characteristics are included.

  2. 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.

  3. Integrated System Test of an Airbreathing Rocket (ISTAR)

    NASA Technical Reports Server (NTRS)

    Faulkner, Robert F.; Lyles, Garry (Technical Monitor)

    2001-01-01

    Rocket Based Combined Cycle (RBCC) propulsion system development and ground test is being conducted as part of the NASA Marshall Space Flight Center Integrated System Test of an Airbreathing Rocket (ISTAR) program. Rocketdyne, Aerojet and Pratt & Whitney have teamed as the Rocket Based Combined Cycle Consortium (RBC3) to work the propulsion system development. Each company offered unique RBCC propulsion concepts as candidates for the ISTAR propulsion system. A team of engine contractor, vehicle contractor and NASA representatives reviewed the concepts proposed by each company, reviewed the available data and selected the Aerojet RBCC propulsion system concept as the team propulsion system baseline for the ISTAR program. The ISTAR program is currently in a "Jumpstart" phase for development of the engine system leading to ground test of a thermally and power balanced RBCC propulsion system at Stennis Space Center in 2005. A parallel flight test demonstration of this propulsion system is anticipated to lead to first flight in the 2007 timeframe.

  4. 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

  5. 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.

  6. Fluid thrust control system. [for liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Howell, W. L.; Jansen, H. B.; Lehmann, E. N. (Inventor)

    1968-01-01

    A pure fluid thrust control system is described for a pump-fed, regeneratively cooled liquid propellant rocket engine. A proportional fluid amplifier and a bistable fluid amplifier control overshoot in the starting of the engine and take it to a predetermined thrust. An ejector type pump is provided in the line between the liquid hydrogen rocket nozzle heat exchanger and the turbine driving the fuel pump to aid in bringing the fluid at this point back into the regular system when it is not bypassed. The thrust control system is intended to function in environments too severe for mechanical controls.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. Mean Line Pump Flow Model in Rocket Engine System Simulation

    NASA Technical Reports Server (NTRS)

    Veres, Joseph P.; Lavelle, Thomas M.

    2000-01-01

    A mean line pump flow modeling method has been developed to provide a fast capability for modeling turbopumps of rocket engines. Based on this method, a mean line pump flow code PUMPA has been written that can predict the performance of pumps at off-design operating conditions, given the loss of the diffusion system at the design point. The pump code can model axial flow inducers, mixed-flow and centrifugal pumps. The code can model multistage pumps in series. The code features rapid input setup and computer run time, and is an effective analysis and conceptual design tool. The map generation capability of the code provides the map information needed for interfacing with a rocket engine system modeling code. The off-design and multistage modeling capabilities of the code permit parametric design space exploration of candidate pump configurations and provide pump performance data for engine system evaluation. The PUMPA code has been integrated with the Numerical Propulsion System Simulation (NPSS) code and an expander rocket engine system has been simulated. The mean line pump flow code runs as an integral part of the NPSS rocket engine system simulation and provides key pump performance information directly to the system model at all operating conditions.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. Single Stage Rocket Technology's real time data system

    NASA Technical Reports Server (NTRS)

    Voglewede, Steven D.

    1994-01-01

    The Single Stage Rocket Technology (SSRT) Delta Clipper Experimental (DC-X) Program is a United States Air Force Ballistic Missile Defense Organization (BMDO) rapid prototyping initiative that is currently demonstrating technology readiness for reusable suborbital rockets. The McDonnell Douglas DC-X rocket performed technology demonstrations at the U.S. Army White Sands Missile Range in New Mexico from April-October in 1993. The DC-X Flight Operations Control Center (FOCC) contains the ground control system that is used to monitor and control the DC-X vehicle and its Ground Support Systems (GSS). The FOCC is operated by a flight crew of three operators. Two operators manage the DC-X Flight Systems and one operator is the Ground Systems Manager. A group from McDonnell Douglas Aerospace at KSC developed the DC-X ground control system for the FOCC. This system is known as the Real Time Data System (RTDS). The RTDS is a distributed real time control and monitoring system that utilizes the latest available commercial off-the-shelf computer technology. The RTDS contains front end interfaces for the DC-X RF uplink/downlink and fiber optic interfaces to the GSS equipment. This paper describes the RTDS architecture and FOCC layout. The DC-X applications and ground operations are covered.

  18. Advanced active health monitoring system of liquid rocket engines

    NASA Astrophysics Data System (ADS)

    Qing, Xinlin P.; Wu, Zhanjun; Beard, Shawn; Chang, Fu-Kuo

    2008-11-01

    An advanced SMART TAPE system has been developed for real-time in-situ monitoring and long term tracking of structural integrity of pressure vessels in liquid rocket engines. The practical implementation of the structural health monitoring (SHM) system including distributed sensor network, portable diagnostic hardware and dedicated data analysis software is addressed based on the harsh operating environment. Extensive tests were conducted on a simulated large booster LOX-H2 engine propellant duct to evaluate the survivability and functionality of the system under the operating conditions of typical liquid rocket engines such as cryogenic temperature, vibration loads. The test results demonstrated that the developed SHM system could survive the combined cryogenic temperature and vibration environments and effectively detect cracks as small as 2 mm.

  19. Analytical concepts for health management systems of liquid rocket engines

    NASA Technical Reports Server (NTRS)

    Williams, Richard; Tulpule, Sharayu; Hawman, Michael

    1990-01-01

    Substantial improvement in health management systems performance can be realized by implementing advanced analytical methods of processing existing liquid rocket engine sensor data. In this paper, such techniques ranging from time series analysis to multisensor pattern recognition to expert systems to fault isolation models are examined and contrasted. The performance of several of these methods is evaluated using data from test firings of the Space Shuttle main engines.

  20. 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.

  1. 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.

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

    NASA Technical Reports Server (NTRS)

    Vonderesch, A. H.

    1972-01-01

    The factors affecting the choice of the 156 inch diameter, parallel burn, solid propellant rocket engine for use with the space shuttle booster are presented. Primary considerations leading to the selection are: (1) low booster vehicle cost, (2) the largest proven transportable system, (3) a demonstrated design, (4) recovery/reuse is feasible, (5) abort can be easily accomplished, and (6) ecological effects are minor.

  3. 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.

  4. 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.

  5. 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.

  6. A multidisciplinary optimization methodology for rocket vehicle systems

    NASA Astrophysics Data System (ADS)

    Colonno, Michael Richard

    Rocket vehicles have traditionally been designed in an iterative fashion, beginning with system requirements before proceeding sequentially through requisite analytical disciplines until resources are exhausted. A sequentially designed system, while adequate, is not an optimum due to the approximations and loss of fidelity inherent in separating analytical disciplines which are, in fact, coupled. Recently, increased computational power and advances in algorithms have allowed multidisciplinary optimization (MDO) to emerge as a system-level design tool accessible to industry. To date, MDO has primarily been applied to some facets of aircraft systems and, to a lesser extent, rocket vehicles in literature but has not yet met with widespread industry use. To this end, four obstacles have been identified: (1) MDO efforts to date have focused on system-level parameters rather than physical dimensions and hence have not yielded a preliminary design which includes manufacturing, cost, and other constraints, (2) Prohibitive computational performance requirements associated with high-fidelity analyses such as computational fluid mechanics (CFD) and finite element analysis (FEA), (3) Lack of an integrated design environment which incorporates computational tools already widely used in industry while remaining accessible to individual users without high-level expertise in the individual tools, and (4) The widely-varying and tightly-coupled environments to which rocket vehicles are typically exposed, including analyses not required for aircraft applications. Here, an MDO method for rocket systems has been formulated which simultaneously overcomes the challenges listed above. First, a response surface-based approach to modeling computationally expensive analyses with arbitrary dimensionality and general constraints was developed. This method focused on an evenly-distributed representation of the entire feasible region at any fidelity level, including combinations of discrete and

  7. Modernization of the multiple launch rocket system embedded system software

    NASA Astrophysics Data System (ADS)

    Mockensturm, Jeffrey J.

    1995-03-01

    Weapon systems in the Department of Defense (DOD) are becoming increasingly reliant on embedded software. As the size and level of complexity of these software development efforts have increased, the management of these programs has become more challenging. Additionally, as the Army strives to digitize the future battlefield, the demand for software will only increase. This thesis reviews the software development efforts associated with modernizing the Army's Multiple Launch Rocket System (MLRS). The thesis begins by presenting a background discussion of the Army's Fire Direction Data Manager (FDDM) development. After the FDDM background discussion, a case study of the troubled FDDM software development effort is presented. The FDDM case study follows the general format presented in the May 1992 General Accounting Office report on the FDDM software development difficulties. Following the FDDM review, the current MLRS software development effort, the Improved Fire Control System (IFCS), is presented. Next, the FDDM case study is reviewed to determine the software development lessons learned. Using the FDDM software lessons learned, the IFCS program is analyzed to determine the software risks, and to review the risk mitigation strategies of that program. The objective of the thesis is to provide insight into the use of modern software management methods in reducing software development program risk.

  8. 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.

  9. Synthesis of calculational methods for design and analysis of radiation shields for nuclear rocket systems

    NASA Technical Reports Server (NTRS)

    Capo, M. A.; Disney, R. K.; Jordan, T. A.; Soltesz, R. G.; Woodsum, H. C.

    1969-01-01

    Eight computer programs make up a nine volume synthesis containing two design methods for nuclear rocket radiation shields. The first design method is appropriate for parametric and preliminary studies, while the second accomplishes the verification of a final nuclear rocket reactor design.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  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. Determination of burning area and port volume in complex burning regions of a solid rocket motor

    NASA Technical Reports Server (NTRS)

    Kingsbury, J. A.

    1977-01-01

    An analysis of the geometry of the burning in both star-cylindrical port interface regions and regions of partially inhibited slots is presented. Some characteristics parameters are defined and illustrated. Methods are proposed for calculating burning areas which functionally depend only on the total distance burned. According to this method, several points are defined where abrupt changes in geometry occur, and these are tracked throughout the burn. Equations are developed for computing port perimeter and port area at pre-established longitudinal positions. Some common formulas and some newly developed formulas are then used to compute burning surface area and port volume. Some specific results are presented for the solid rocket motor committed to the space shuttle project.

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

    NASA Technical Reports Server (NTRS)

    1972-01-01

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

  17. 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.

  18. 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.

  19. 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.

  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. An expert system for spectroscopic analysis of rocket engine plumes

    NASA Technical Reports Server (NTRS)

    Reese, Greg; Valenti, Elizabeth; Alphonso, Keith; Holladay, Wendy

    1991-01-01

    The expert system described in this paper analyzes spectral emissions of rocket engine exhaust plumes and shows major promise for use in engine health diagnostics. Plume emission spectroscopy is an important tool for diagnosing engine anomalies, but it is time-consuming and requires highly skilled personnel. The expert system was created to alleviate such problems. The system accepts a spectral plot in the form of wavelength vs intensity pairs and finds the emission peaks in the spectrum, lists the elemental emitters present in the data and deduces the emitter that produced each peak. The system consists of a conventional language component and a commercially available inference engine that runs on an Apple Macintosh computer. The expert system has undergone limited preliminary testing. It detects elements well and significantly decreases analysis time.

  2. RS-88 Rocket Engine Tested for Pad Abort Escape System

    NASA Technical Reports Server (NTRS)

    2003-01-01

    In this photo, an RS-88 development rocket engine is being test fired at NASA's Marshall Space Flight Center in Huntsville, Alabama, in support of the Pad Abort Demonstration (PAD) test flights for NASA's Orbital Space Plane (OSP). The tests could be instrumental in developing the first crew launch escape system in almost 30 years. Paving the way for a series of integrated PAD test flights, the engine tests support development of a system that could pull a crew safely away from danger during liftoff. A series of 16 hot fire tests of a 50,000-pound thrust RS-88 rocket engine were conducted, resulting in a total of 55 seconds of successful engine operation. The engine is being developed by the Rocketdyne Propulsion and Power unit of the Boeing Company. Integrated launch abort demonstration tests in 2005 will use four RS-88 engines to separate a test vehicle from a test platform, simulating pulling a crewed vehicle away from an aborted launch. Four 156-foot parachutes will deploy and carry the vehicle to landing. Lockheed Martin is building the vehicles for the PAD tests. Seven integrated tests are plarned for 2005 and 2006.

  3. RS-88 Rocket Engine Tested for Pad Abort Escape System

    NASA Technical Reports Server (NTRS)

    2003-01-01

    This photo gives an overhead look at an RS-88 development rocket engine being test fired at NASA's Marshall Space Flight Center in Huntsville, Alabama, in support of the Pad Abort Demonstration (PAD) test flights for NASA's Orbital Space Plane (OSP). The tests could be instrumental in developing the first crew launch escape system in almost 30 years. Paving the way for a series of integrated PAD test flights, the engine tests support development of a system that could pull a crew safely away from danger during liftoff. A series of 16 hot fire tests of a 50,000-pound thrust RS-88 rocket engine were conducted, resulting in a total of 55 seconds of successful engine operation. The engine is being developed by the Rocketdyne Propulsion and Power unit of the Boeing Company. Integrated launch abort demonstration tests in 2005 will use four RS-88 engines to separate a test vehicle from a test platform, simulating pulling a crewed vehicle away from an aborted launch. Four 156-foot parachutes will deploy and carry the vehicle to landing. Lockheed Martin is building the vehicles for the PAD tests. Seven integrated tests are plarned for 2005 and 2006.

  4. 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.

  5. Simulations of Transient Phenomena in Liquid Rocket Feed Systems

    NASA Technical Reports Server (NTRS)

    Ahuja, V.; Hosangadi, A.; Cavallo, P. A.; Daines, R.

    2006-01-01

    Valve systems in rocket propulsion systems and testing facilities are constantly subject to dynamic events resulting from the timing of valve motion leading to unsteady fluctuations in pressure and mass flow. Such events can also be accompanied by cavitation, resonance, system vibration leading to catastrophic failure. High-fidelity dynamic computational simulations of valve operation can yield important information of valve response to varying flow conditions. Prediction of transient behavior related to valve motion can serve as guidelines for valve scheduling, which is of crucial importance in engine operation and testing. Feed components operating in cryogenic regimes can also experience cavitation based instabilities leading to large scale shedding of vapor clouds and pressure oscillations. In this paper, we present simulations of the diverse unsteady phenomena related to valve and feed systems that include valve stall, valve timing studies as well as two different forms of cavitation instabilities in components utilized in the test loop.

  6. 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

  7. Multiple dopant injection system for small rocket engines

    NASA Astrophysics Data System (ADS)

    Sakala, G. G.; Raines, N. G.

    1992-07-01

    The Diagnostics Test Facility (DTF) at NASA's Stennis Space Center (SSC) was designed and built to provide a standard rocket engine exhaust plume for use in the research and development of engine health monitoring instrumentation. A 1000 lb thrust class liquid oxygen (LOX)-gaseous hydrogen (GH2) fueled rocket engine is used as the subscale plume source to simulate the SSME during experimentation and instrument development. The ability of the DTF to provide efficient, and low cost test operations makes it uniquely suited for plume diagnostic experimentation. The most unique feature of the DTF is the Multiple Dopant Injection System (MDIS) that is used to seed the exhaust plume with the desired element or metal alloy. The dopant injection takes place at the fuel injector, yielding a very uniform and homogeneous distribution of the seeding material in the exhaust plume. The MDIS allows during a single test firing of the DTF, the seeding of the exhaust plume with up to three different dopants and also provides distilled water base lines between the dopants. A number of plume diagnostic-related experiments have already utilized the unique capabilities of the DTF.

  8. 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.

  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. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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

  16. 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.

  17. 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.

  18. Rocket based combined cycle (RBCC) propulsion systems offer additional options

    NASA Astrophysics Data System (ADS)

    Czysz, Paul A.

    The propulsion cycles presented at the 1991 IAF Congress in Montreal, and at The World Hydrogen Conference 1992 in Paris were the subject of an IAF paper for the 1992 World Space Conference in Washington DC. RBCC propulsion systems from several nations were analyzed in terms of a SSTO space launcher with a 7-Mg payload. The RBCC concept emerged from the advanced injector ramjet research of the early 1960s. The performance of the current RBCC propulsion systems such that the specific thrust of a rocket is combined with the specific impulse of an airbreather. This performance offers a new perspective to the options available. In a brief review of the present RBCC the reasons for these options are developed. The spectrum of the system options is presented in three examples, a LACE VTOL SSTO, an HTOL SSTO and a HTOL TSTO. Results using the present RBCC are dramatically different from the past concept of the Conventional Combined Cycle propulsion system, i.e., combinations of separate engines. The integration of the engine cycles into a single thermodynamically integrated system significantly changes the propulsion performance.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  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. 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.

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

    NASA Technical Reports Server (NTRS)

    Yoshida, Kinjiro; Hayashi, Kengo; Takami, Hiroshi

    1996-01-01

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

  7. Heat exchanger. [rocket combustion chambers and cooling systems

    NASA Technical Reports Server (NTRS)

    Sokolowski, D. E. (Inventor)

    1978-01-01

    A heat exchanger, as exemplified by a rocket combustion chamber, is constructed by stacking thin metal rings having microsized openings therein at selective locations to form cooling passages defined by an inner wall, an outer wall and fins. Suitable manifolds are provided at each end of the rocket chamber. In addition to the cooling channel openings, coolant feed openings may be formed in each of rings. The coolant feed openings may be nested or positioned within generally U-shaped cooling channel openings. Compression on the stacked rings may be maintained by welds or the like or by bolts extending through the stacked rings.

  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. 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.

  10. 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.

  11. 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.

  12. 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.

  13. A system level model for preliminary design of a space propulsion solid rocket motor

    NASA Astrophysics Data System (ADS)

    Schumacher, Daniel M.

    Preliminary design of space propulsion solid rocket motors entails a combination of components and subsystems. Expert design tools exist to find near optimal performance of subsystems and components. Conversely, there is no system level preliminary design process for space propulsion solid rocket motors that is capable of synthesizing customer requirements into a high utility design for the customer. The preliminary design process for space propulsion solid rocket motors typically builds on existing designs and pursues feasible rather than the most favorable design. Classical optimization is an extremely challenging method when dealing with the complex behavior of an integrated system. The complexity and combinations of system configurations make the number of the design parameters that are traded off unreasonable when manual techniques are used. Existing multi-disciplinary optimization approaches generally address estimating ratios and correlations rather than utilizing mathematical models. The developed system level model utilizes the Genetic Algorithm to perform the necessary population searches to efficiently replace the human iterations required during a typical solid rocket motor preliminary design. This research augments, automates, and increases the fidelity of the existing preliminary design process for space propulsion solid rocket motors. The system level aspect of this preliminary design process, and the ability to synthesize space propulsion solid rocket motor requirements into a near optimal design, is achievable. The process of developing the motor performance estimate and the system level model of a space propulsion solid rocket motor is described in detail. The results of this research indicate that the model is valid for use and able to manage a very large number of variable inputs and constraints towards the pursuit of the best possible design.

  14. 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.

  15. 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.

  16. 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.

  17. Liquid Rocket Propulsion Technology: An evaluation of NASA's program. [for space transportation systems

    NASA Technical Reports Server (NTRS)

    1981-01-01

    The liquid rocket propulsion technology needs to support anticipated future space vehicles were examined including any special action needs to be taken to assure that an industrial base in substained. Propulsion system requirements of Earth-to-orbit vehicles, orbital transfer vehicles, and planetary missions were evaluated. Areas of the fundamental technology program undertaking these needs discussed include: pumps and pump drives; combustion heat transfer; nozzle aerodynamics; low gravity cryogenic fluid management; and component and system life reliability, and maintenance. The primary conclusion is that continued development of the shuttle main engine system to achieve design performance and life should be the highest priority in the rocket engine program.

  18. Detailed modal testing of a solid rocket motor using a portable test system

    NASA Technical Reports Server (NTRS)

    Glozman, Vladimir; Brillhart, Ralph D.

    1990-01-01

    Modern analytical techniques have expended the ability to evaluate solid rocket motors used in launch vehicles. As more detailed models of solid rocket motors were developed, testing methods were required to verify the models. Experimental modal analysis (modal testing) of space structures and launch vehicles has been a requirement for model validation for many years. However, previous testing of solid rocket motors has not typically involved dynamic modal testing of full scale motors for verification of solid propellant or system assembly properties. Innovative approaches to the testing of solid rocket motors were developed and modal testing of a full scale, two segment Titan 34D Solid Rocket Motor (SRM) was performed to validate detailed computer modeling. Special modifications were made to convert an existing facility into a temporary modal test facility which would accommodate the test article. The assembly of conventional data acquisition equipment into a multiple channel count portable system has made modal testing in the field feasible. Special purpose hydraulic exciters were configured to apply the dynamic driving forces required. All instrumentation and data collection equipment were installed at the test site for the duration of the test program and removed upon completion. Conversion of an existing test facility into a temporary modal test facility, and use of a multiple channel count portable test data acquisition system allowed all test objectives to be met and resulted in validation of the computer model in a minimum time.

  19. 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.

  20. The development of a fiber optic Raman temperature measurement system for rocket flows

    NASA Technical Reports Server (NTRS)

    De Groot, Wim A.

    1991-01-01

    A fiber-optic Raman diagnostic system for H2/O2 rocket flows is currently under development. This system is designed for measurements of temperature and major species concentration in the combustion chamber and part of the nozzle of a 100 Newton thrust rocket currently undergoing tests. 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 concentration by comparing theoretical Raman scattering spectra with experimentally obtained spectra. Equipment selection and experimental approach are summarized. This experimental effort is part of a program, which is in progress, to evaluate Navier-Stokes based analyses for this class of rockets.

  1. 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.

  2. 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.

  3. 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).

  4. 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.

  5. 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.

  6. System for imposing directional stability on a rocket-propelled vehicle

    NASA Technical Reports Server (NTRS)

    Perkins, H. (Inventor)

    1976-01-01

    An improved system for use in imposing directional stability on a rocket-propelled vehicle is described. The system includes a pivotally supported engine-mounting platform, a gimbal ring mounted on the platform and adapted to pivotally support a rocket engine and an hydraulic actuator connected to the platform for imparting selected pivotal motion. An accelerometer and a signal comparator circuit for providing error intelligence indicative of aberration in vehicle acceleration is included along with an actuator control circuit connected with the actuator and responsive to error intelligence for imparting pivotal motion to the platform. Relocation of the engine's thrust vector is thus achieved for imparting directional stability to the vehicle.

  7. 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.

  8. 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

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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).

  14. 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.

  15. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1971-01-01

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

  17. Solid rocket booster thermal protection system materials development. [space shuttle boosters

    NASA Technical Reports Server (NTRS)

    Dean, W. G.

    1978-01-01

    A complete run log of all tests conducted in the NASA-MSFC hot gas test facility during the development of materials for the space shuttle solid rocket booster thermal protection system are presented. Lists of technical reports and drawings generated under the contract are included.

  18. 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…

  19. 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.

  20. 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.

  1. Mining volume measurement system

    NASA Technical Reports Server (NTRS)

    Heyman, Joseph Saul (Inventor)

    1988-01-01

    In a shaft with a curved or straight primary segment and smaller off-shooting segments, at least one standing wave is generated in the primary segment. The shaft has either an open end or a closed end and approximates a cylindrical waveguide. A frequency of a standing wave that represents the fundamental mode characteristic of the primary segment can be measured. Alternatively, a frequency differential between two successive harmonic modes that are characteristic of the primary segment can be measured. In either event, the measured frequency or frequency differential is characteristic of the length and thus the volume of the shaft based on length times the bore area.

  2. 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.

  3. 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.

  4. 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).

  5. 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.

  6. 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.

  7. 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.

  8. 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.

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

    NASA Astrophysics Data System (ADS)

    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.

  10. 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. PMID:23387673

  11. 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.

  12. 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.

  13. 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. PMID:11898833

  14. 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.

  15. 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.

  16. 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

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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)

  6. 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.

  7. 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

  8. 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.

  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. 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.

  12. 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.

  13. 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.

  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 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.

  17. 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.

  18. 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.

  19. 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.

  20. Reverse engineering of the multiple launch rocket system. Human factors, manpower, personnel, and training in the weapons system acquisition process

    NASA Astrophysics Data System (ADS)

    Arabian, J. M.; Hartel, C. R.; Kaplan, J. D.; Marcus, A.; Promisel, D. M.

    1984-06-01

    In a briefing format, this report on the Multiple Launch Rocket System summarizes an examination of human factors, manpower, personnel and training (HMPT) issues during the systems acquisition process. The report is one of four reverse engineering studies prepared at the request of Gen. M. R. Thurman, Army Vice Chief of Staff. The four systems were studied as a representative sample of Army weapons systems. They serve as the basis for drawing conclusions about aspects of the weapons system acquisition process which most affect HMPT considerations. A synthesis of the four system studies appears in the final report of the Reverse Engineering Task Force U.S. Army Research Institute.

  1. Computational Analysis of an LOx Supply Line System of an Liquid Rocket Engine

    NASA Astrophysics Data System (ADS)

    Moon, Insang; Moon, Il Yoon; Lee, Soo Yong

    2009-12-01

    A computational fluid analysis was performed on an LOx line system of a liquid rocket engine. The model was created with 3D CAD and imbedded to the 3D CFD program. Before the full scale analysis on the system was carried out, each components with simplified models was analyzed to save time and cost. As a result, the inlet pressure of the gas generator should be compensated with a certain device unless the inlet pressure of the line system is sufficiently high. The flow pattern of the exit of the system was dependant upon the location of the orifice as well as the size. As a whole the line system analyzed met the requirements, and will be tested and confirmed after being manufactured.

  2. Integrated system of test data management and monitoring for the ground test of liquid rocket engine

    NASA Astrophysics Data System (ADS)

    Yang, Xue; Zhang, Zhenpeng; Zhang, Jun

    2008-10-01

    An integrated system of test data management and monitoring (ISTDMM) for liquid rocket engine (LRE) ground test is designed to meet the demand of the LRE test station and development unit according to the LRE test information and test process. It is an opening, distributing and highly integrating application platform, mainly includes the test data management systems, the real-time fault detection systems and data display and playback system. It can manage and analyze the test data and simulation data of the LRE, can monitor the LRE test condition in real-time and the test process in long-distance by network, and can playback the engine test process and simulate the engine work process, and can test and evaluate the fault detection algorithms and systems of LRE. It is well advanced, reliable, and practical.

  3. Analytical investigation of two hydrogen oxygen rocket engine systems for low-thrust application

    NASA Technical Reports Server (NTRS)

    Scheer, D. D.

    1980-01-01

    Two hydrogen oxygen rocket engine system concepts were analyzed parametrically over a thrust range from 100 to 1000 pounds and a chamber pressure range from 175 to 1000 psia. Both concepts were regeneratively cooled with hydrogen and were pump fed by electric motor driven positive displacement pumps. Electric power was provided by either a turboalternator (turboalternator concept) or some means external to the engine system (auxiliary power concept). The turboalternator concept is discussed. The computer program used to conduct the analyses along with the design characteristics of the major engine system components is described. The feasible design range of the systems over the parametric range of thrust is discussed in terms of allowable chamber pressure. Engine system estimated performance, mass, and dimensional envelope parametric data within the feasible design range are presented.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. Rocket measurements of electrons in a system of multiple auroral arcs

    NASA Technical Reports Server (NTRS)

    Boyd, J. S.; Davis, T. N.

    1977-01-01

    A Nike-Tomahawk rocket was launched into a system of auroral arcs northward of Poker Flat Research Range, Fairbanks, Alaska. The pitch-angle distribution of electrons was measured at 2.5, 5, and 10 keV and also at 10 keV on a separating forward section of the payload. The auroral activity appeared to be the extension of substorm activity centered to the east. The rocket crossed a westward-propagating fold in the brightest band. The electron spectrum was relatively hard through most of the flight, showing a peak in the range from 2.5 to 10 keV in the weaker aurora and below 5 keV in the brightest arc. The detailed structure of the pitch-angle distribution suggested that, at times, a very selective process was accelerating some electrons in the magnetic field direction, so that a narrow field-aligned component appeared superimposed on a more isotropic distribution. It is concluded that this process could not be a near-ionosphere field-aligned potential drop, although the more isotropic component may have been produced by a parallel electric field extending several thousand kilometers along the field line above the ionosphere.

  9. 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.

  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. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. Highlights of NASA's Special ETO Program Planning Workshop on rocket-based combined-cycle propulsion system technologies

    NASA Technical Reports Server (NTRS)

    Escher, W. J. D.

    1992-01-01

    A NASA workshop on rocket-based combined-cycle propulsion technologies is described emphasizing the development of a starting point for earth-to-orbit (ETO) rocket technologies. The tutorial is designed with attention given to the combined development of aeronautical airbreathing propulsion and space rocket propulsion. The format, agenda, and group deliberations for the tutorial are described, and group deliberations include: (1) mission and space transportation infrastructure; (2) vehicle-integrated propulsion systems; (3) development operations, facilities, and human resource needs; and (4) spaceflight fleet applications and operations. Although incomplete the workshop elevates the subject of combined-cycle hypersonic propulsion and develops a common set of priniciples regarding the development of these technologies.

  17. ASTRID rocket flight test

    SciTech Connect

    Whitehead, J.C.; Pittenger, L.C.; Colella, N.J.

    1994-07-01

    On February 4, 1994, we successfully flight tested the ASTRID rocket from Vandenberg Air Force Base. The technology for this rocket originated in the Brilliant Pebbles program and represents a five-year development effort. This rocket demonstrated how our new pumped-propulsion technology-which reduced the total effective engine mass by more than one half and cut the tank mass to one fifth previous requirements-would perform in atmospheric flight. This demonstration paves the way for potential cost-effective uses of the new propulsion system in commercial aerospace vehicles, exploration of the planets, and defense applications.

  18. 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.

  19. Russian Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1998-01-01

    NASA engineers successfully tested a Russian-built rocket engine on November 4, 1998 at the Marshall Space Flight Center (MSFC) Advanced Engine Test Facility, which had been used for testing the Saturn V F-1 engines and Space Shuttle Main engines. The MSFC was under a Space Act Agreement with Lockheed Martin Astronautics of Denver to provide a series of test firings of the Atlas III propulsion system configured with the Russian-designed RD-180 engine. The tests were designed to measure the performance of the Atlas III propulsion system, which included avionics and propellant tanks and lines, and how these components interacted with the RD-180 engine. The RD-180 is powered by kerosene and liquid oxygen, the same fuel mix used in Saturn rockets. The RD-180, the most powerful rocket engine tested at the MSFC since Saturn rocket tests in the 1960s, generated 860,000 pounds of thrust.

  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. NDE of thermal protection system for space shuttle solid rocket booster

    NASA Technical Reports Server (NTRS)

    Myers, R. S.

    1990-01-01

    Potential nondestructive test (NDE) methods were evaluated for detecting debonds and weak bonds in the thermal protection system (TPS) for the space shuttle solid rocket boosters. The primary thermal protection material is a sprayable, thick epoxy coating that is filled with lightweight and thermal insulating materials. Test panels were fabricated with a wide variety of hidden realistic defects, including contact debonds and weak bonds. Nondestructive test results were obtained. Candidate NDE methods evaluated for booster production applications include laser interferometry (e.g., electronic shearography), infrared thermography, radiography (e.g., computed tomography), acousto-ultrasonics, mechanical/acoustic impedance, ultrasonics, acoustic emission, and the tap test. Capabilities, advantages, disadvantages, and relative performances in defect detection of each test method for TPS bonding applications are reported. Electronic shearography NDE was technically the superior method for detecting debonds.

  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. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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

  9. 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.

  10. 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.

  11. 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

  12. Guidance, navigation & control systems for sounding rockets - flight results, current status and the future

    NASA Astrophysics Data System (ADS)

    Ljunge, Lars

    2005-08-01

    At the 16th ESA Symposium on European Rockets and Balloons, two newly developed guidance and control systems by Saab Ericsson Space were presented: The S19D guidance and control system, which uses DS19 hardware to execute S19 type guidance and control. The GCS/DMARS guidance, navigation and control system, which is a modernisation of the GCS/RIINS. These two and the third recent system, the DS19, were developed as replacements for the analog S19 and the GCS/RIINS, both of which use very old technology. The design drivers or the DS19, the S19D and the GCS/DMARS are: User requirements. New technology with improved performance capability becoming available. Current technology becoming old and replacement parts hard to come by. This paper first lists some guidance related user requirements, and then discusses the performance that has been achieved in the various guidance systems, including the S19, for comparison. This is first done from a theoretical point of view and then by analyzing actual flight data. The ability of the systems to fulfil the user requirements is also discussed and finally, a look is taken into the future.

  13. Rocket Engine Oscillation Diagnostics

    NASA Technical Reports Server (NTRS)

    Nesman, Tom; Turner, James E. (Technical Monitor)

    2002-01-01

    Rocket engine oscillating data can reveal many physical phenomena ranging from unsteady flow and acoustics to rotordynamics and structural dynamics. Because of this, engine diagnostics based on oscillation data should employ both signal analysis and physical modeling. This paper describes an approach to rocket engine oscillation diagnostics, types of problems encountered, and example problems solved. Determination of design guidelines and environments (or loads) from oscillating phenomena is required during initial stages of rocket engine design, while the additional tasks of health monitoring, incipient failure detection, and anomaly diagnostics occur during engine development and operation. Oscillations in rocket engines are typically related to flow driven acoustics, flow excited structures, or rotational forces. Additional sources of oscillatory energy are combustion and cavitation. Included in the example problems is a sampling of signal analysis tools employed in diagnostics. The rocket engine hardware includes combustion devices, valves, turbopumps, and ducts. Simple models of an oscillating fluid system or structure can be constructed to estimate pertinent dynamic parameters governing the unsteady behavior of engine systems or components. In the example problems it is shown that simple physical modeling when combined with signal analysis can be successfully employed to diagnose complex rocket engine oscillatory phenomena.

  14. 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.

  15. 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).

  16. 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)

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. NASA Safety Manual. Volume 3: System Safety

    NASA Technical Reports Server (NTRS)

    1970-01-01

    This Volume 3 of the NASA Safety Manual sets forth the basic elements and techniques for managing a system safety program and the technical methods recommended for use in developing a risk evaluation program that is oriented to the identification of hazards in aerospace hardware systems and the development of residual risk management information for the program manager that is based on the hazards identified. The methods and techniques described in this volume are in consonance with the requirements set forth in NHB 1700.1 (VI), Chapter 3. This volume and future volumes of the NASA Safety Manual shall not be rewritten, reprinted, or reproduced in any manner. Installation implementing procedures, if necessary, shall be inserted as page supplements in accordance with the provisions of Appendix A. No portion of this volume or future volumes of the NASA Safety Manual shall be invoked in contracts.

  6. 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.

  7. 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)

  8. Foil chaff ejection systems for rocket-borne measurement of neutral winds in the mesosphere and lower thermosphere

    NASA Astrophysics Data System (ADS)

    Koizumi, Yoshiko; Shimoyama, Manabu; Oyama, Koh-Ichiro; Murayama, Yasuhiro; Tsuda, Toshitaka; Nakamura, Takuji

    2004-07-01

    The foil chaff technique has been used on microrockets such as "Viper" for a long time to measure neutral winds with high altitude resolution in the mesosphere and lower thermosphere. We have developed two new foil chaff storage and ejection systems for muti-instrumented sounding rockets. The first system uses a spring loaded split cylinder which holds the foil chaff, housed in an outer cylinder. The shaft of the split cylinder is kept in place by a lock plate and a stainless steel wire. The split cylinder is ejected by cutting the wire. The second system is of differential pressure type. The cap of an airtight cylinder has a shaft and a sponge piece for sweeping out the foil chaff. The cylinder is sealed at ground level and at the desired height of release, the cap comes out due to differential pressure and brings out the foil chaff. Both these systems were successfully tested on a Japanese sounding rocket in January 2000, releasing about 20 000 pieces of foil chaff during the rocket's descent. Neutral winds were measured in the height range of 85.5-95.0 km with a height resolution of 300 m.

  9. 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.

  10. 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.