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Sample records for rocket based combined

  1. The Strutjet Rocket Based Combined Cycle Engine

    NASA Technical Reports Server (NTRS)

    Siebenhaar, A.; Bulman, M. J.; Bonnar, D. K.

    1998-01-01

    The multi stage chemical rocket has been established over many years as the propulsion System for space transportation vehicles, while, at the same time, there is increasing concern about its continued affordability and rather involved reusability. Two broad approaches to addressing this overall launch cost problem consist in one, the further development of the rocket motor, and two, the use of airbreathing propulsion to the maximum extent possible as a complement to the limited use of a conventional rocket. In both cases, a single-stage-to-orbit (SSTO) vehicle is considered a desirable goal. However, neither the "all-rocket" nor the "all-airbreathing" approach seems realizable and workable in practice without appreciable advances in materials and manufacturing. An affordable system must be reusable with minimal refurbishing on-ground, and large mean time between overhauls, and thus with high margins in design. It has been suggested that one may use different engine cycles, some rocket and others airbreathing, in a combination over a flight trajectory, but this approach does not lead to a converged solution with thrust-to-mass, specific impulse, and other performance and operational characteristics that can be obtained in the different engines. The reason is this type of engine is simply a combination of different engines with no commonality of gas flowpath or components, and therefore tends to have the deficiencies of each of the combined engines. A further development in this approach is a truly combined cycle that incorporates a series of cycles for different modes of propulsion along a flight path with multiple use of a set of components and an essentially single gas flowpath through the engine. This integrated approach is based on realizing the benefits of both a rocket engine and airbreathing engine in various combinations by a systematic functional integration of components in an engine class usually referred to as a rocket-based combined cycle (RBCC) engine

  2. Computational investigation of rocket based combined cycle

    NASA Astrophysics Data System (ADS)

    Zhang, Xiao-bo; Wang, Zhan-xue; Liu, Zeng-wen

    2013-03-01

    Based on Computational Fluid Dynamic technology, the mixing process of Rocket Based Combined Cycle (RBCC) propulsion system is researched. The idea of RBCC propulsion system means combining rocket engine with ramjet engine effectively, which can flight from sea level to high altitude in wide Mach ranges. In order to analyze how the length of the mixing part affects mixing process, different length of mixing part are researched. As it is indicated, with a constant Mach number, increasing the length of mixing part makes main flow and second flow mix more evenly. Moreover, the length of mixing part has a slight impact on the thrust. Obviously the main consequence of increasing the length of mixing part is promoting the mix of main flow and second flow. Therefore, in order to decrease the weight of aircraft, it is of importance to reduce the length. Through comparing distribution of different cases, when working in the situation of maximum power, the flow in the nozzle of rocket engine is under expansion, while that in the nozzle is fully expanded. Nevertheless, in the case of high altitude and high Mach number, there exists a vortex in the nozzle of rocket engine because of over expansion; meanwhile, the flow in the nozzle is under expansion. Therefore, it is necessary to adjust nozzle throat area in order to increase the thrust of RBCC at high altitude.

  3. 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.

  4. Rocket-Based Combined Cycle Engine Concept Development

    NASA Technical Reports Server (NTRS)

    Ratekin, G.; Goldman, Allen; Ortwerth, P.; Weisberg, S.; McArthur, J. Craig (Technical Monitor)

    2001-01-01

    The development of rocket-based combined cycle (RBCC) propulsion systems is part of a 12 year effort under both company funding and contract work. The concept is a fixed geometry integrated rocket, ramjet, scramjet, which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals, seal purge gas, and closeout side attachments. Engine A5 is the current configuration for NASA Marshall Space Flight Center (MSFC) for the ART program. Engine A5 models the complete flight engine flowpath of inlet, isolator, airbreathing combustor, and nozzle. High-performance rocket thrusters are integrated into the engine enabling both low speed air-augmented rocket (AAR) and high speed pure rocket operation. Engine A5 was tested in GASL's new Flight Acceleration Simulation Test (FAST) facility in all four operating modes, AAR, RAM, SCRAM, and Rocket. Additionally, transition from AAR to RAM and RAM to SCRAM was also demonstrated. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. SCRAM and rocket mode performance was above predictions. For the first time, testing also demonstrated transition between operating modes.

  5. Analysis of a Rocket Based Combined Cycle Engine during Rocket Only Operation

    NASA Technical Reports Server (NTRS)

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

    1998-01-01

    The all rocket mode of operation is a critical factor in the overall performance of a rocket based combined cycle (RBCC) vehicle. However, outside of performing experiments or a full three dimensional analysis, there are no first order parametric models to estimate performance. As a result, an axisymmetric RBCC engine was used to analytically determine specific impulse efficiency values based upon both full flow and gas generator configurations. Design of experiments methodology was used to construct a test matrix and statistical regression analysis was used to build parametric models. The main parameters investigated in this study were: rocket chamber pressure, rocket exit area ratio, percent of injected secondary flow, mixer-ejector inlet area, mixer-ejector area ratio, and mixer-ejector length-to-inject diameter ratio. A perfect gas computational fluid dynamics analysis was performed to obtain values of vacuum specific impulse. Statistical regression analysis was performed based on both full flow and gas generator engine cycles. Results were also found to be dependent upon the entire cycle assumptions. The statistical regression analysis determined that there were five significant linear effects, six interactions, and one second-order effect. Two parametric models were created to provide performance assessments of an RBCC engine in the all rocket mode of operation.

  6. 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.

  7. Design Rules and Issues with Respect to Rocket Based Combined Cycles

    DTIC Science & Technology

    2010-09-01

    Section Analysis As we seek for the accelerator, the inlet design is quite art of compromise. To make benefits due to air- breathing propulsion, the...Design Rules and Issues with Respect to Rocket Based Combined Cycles 3 - 4 RTO-EN-AVT-185 2.1.2 Combustor Section Analysis Embedded rocket chamber...cause thrust augmentation due to the ejector effects, which in turn, can reduce the requirement for the rocket engine output. In the speed regime with

  8. Performance of an Axisymmetric Rocket Based Combined Cycle Engine During Rocket Only Operation Using Linear Regression Analysis

    NASA Technical Reports Server (NTRS)

    Smith, Timothy D.; Steffen, Christopher J., Jr.; Yungster, Shaye; Keller, Dennis J.

    1998-01-01

    The all rocket mode of operation is shown to be a critical factor in the overall performance of a rocket based combined cycle (RBCC) vehicle. An axisymmetric RBCC engine was used to determine specific impulse efficiency values based upon both full flow and gas generator configurations. Design of experiments methodology was used to construct a test matrix and multiple linear regression analysis was used to build parametric models. The main parameters investigated in this study were: rocket chamber pressure, rocket exit area ratio, injected secondary flow, mixer-ejector inlet area, mixer-ejector area ratio, and mixer-ejector length-to-inlet diameter ratio. A perfect gas computational fluid dynamics analysis, using both the Spalart-Allmaras and k-omega turbulence models, was performed with the NPARC code to obtain values of vacuum specific impulse. Results from the multiple linear regression analysis showed that for both the full flow and gas generator configurations increasing mixer-ejector area ratio and rocket area ratio increase performance, while increasing mixer-ejector inlet area ratio and mixer-ejector length-to-diameter ratio decrease performance. Increasing injected secondary flow increased performance for the gas generator analysis, but was not statistically significant for the full flow analysis. Chamber pressure was found to be not statistically significant.

  9. Combustion oscillation study in a kerosene fueled rocket-based combined-cycle engine combustor

    NASA Astrophysics Data System (ADS)

    Huang, Zhi-Wei; He, Guo-Qiang; Qin, Fei; Xue, Rui; Wei, Xiang-Geng; Shi, Lei

    2016-12-01

    This study reports the combustion oscillation features in a three-dimensional (3D) rocket-based combined-cycle (RBCC) engine combustor under flight Mach number (Mflight) 3.0 conditions both experimentally and numerically. Experiment is performed on a direct-connect ground test facility, which measures the wall pressure along the flow-path. High-speed imaging of the flame luminosity and schlieren is carried out at exit of the primary rocket. Compressible reactive large eddy simulation (LES) with reduced chemical kinetics of a surrogate model for kerosene is performed to further understand the combustion oscillation mechanisms in the combustor. LES results are validated with experimental data by the time-averaged and root mean square (RMS) pressure values, and show acceptable agreement. Effects of the primary rocket jet on pressure oscillation in the combustor are analyzed. Relation of the high speed rocket jet oscillation, which is thought to among the most probable sources of combustion oscillation, with the RBCC combustor is recognized. Results reveal that the unsteady over-expanded rocket jet has significant impacts on the combustion oscillation feature of the RBCC combustor, which is different from a thermo-acoustics type oscillation. The rocket jet/air inflow physical interactions under different rocket jet expansion degrees are experimentally studied.

  10. A retrospective on early cryogenic primary rocket subsystem designs as integrated into rocket-based combined-cycle (RBCC) engines

    NASA Technical Reports Server (NTRS)

    Escher, William J. D.; Schnurstein, Robert E.

    1993-01-01

    A study (Escher and Flornes, 1966) of aerospace propulsion systems for a fully reusable earth-to-orbit space transport application that was performed in 1965-67 is reviewed. The present review provides a detailed, subject-focused technical retrospective on a key subsystem element of the rocket-based combined-cycle (RBCC) class of aerospace propulsion systems. The RBCC concept is considered to be a leading candidate propulsion approach for either SSTO or two-stage-to-orbit space transportaion applications.

  11. A retrospective on early cryogenic primary rocket subsystem designs as integrated into rocket-based combined-cycle (RBCC) engines

    NASA Astrophysics Data System (ADS)

    Escher, William J. D.; Schnurstein, Robert E.

    1993-06-01

    A study (Escher and Flornes, 1966) of aerospace propulsion systems for a fully reusable earth-to-orbit space transport application that was performed in 1965-67 is reviewed. The present review provides a detailed, subject-focused technical retrospective on a key subsystem element of the rocket-based combined-cycle (RBCC) class of aerospace propulsion systems. The RBCC concept is considered to be a leading candidate propulsion approach for either SSTO or two-stage-to-orbit space transportaion applications.

  12. Multidisciplinary design of a rocket-based combined cycle SSTO launch vehicle using Taguchi methods

    NASA Technical Reports Server (NTRS)

    Olds, John R.; Walberg, Gerald D.

    1993-01-01

    Results are presented from the optimization process of a winged-cone configuration SSTO launch vehicle that employs a rocket-based ejector/ramjet/scramjet/rocket operational mode variable-cycle engine. The Taguchi multidisciplinary parametric-design method was used to evaluate the effects of simultaneously changing a total of eight design variables, rather than changing them one at a time as in conventional tradeoff studies. A combination of design variables was in this way identified which yields very attractive vehicle dry and gross weights.

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

    NASA Astrophysics Data System (ADS)

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

    2016-02-01

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

  14. Computational Fluid Dynamics Analysis Method Developed for Rocket-Based Combined Cycle Engine Inlet

    NASA Technical Reports Server (NTRS)

    1997-01-01

    Renewed interest in hypersonic propulsion systems has led to research programs investigating combined cycle engines that are designed to operate efficiently across the flight regime. The Rocket-Based Combined Cycle Engine is a propulsion system under development at the NASA Lewis Research Center. This engine integrates a high specific impulse, low thrust-to-weight, airbreathing engine with a low-impulse, high thrust-to-weight rocket. From takeoff to Mach 2.5, the engine operates as an air-augmented rocket. At Mach 2.5, the engine becomes a dual-mode ramjet; and beyond Mach 8, the rocket is turned back on. One Rocket-Based Combined Cycle Engine variation known as the "Strut-Jet" concept is being investigated jointly by NASA Lewis, the U.S. Air Force, Gencorp Aerojet, General Applied Science Labs (GASL), and Lockheed Martin Corporation. Work thus far has included wind tunnel experiments and computational fluid dynamics (CFD) investigations with the NPARC code. The CFD method was initiated by modeling the geometry of the Strut-Jet with the GRIDGEN structured grid generator. Grids representing a subscale inlet model and the full-scale demonstrator geometry were constructed. These grids modeled one-half of the symmetric inlet flow path, including the precompression plate, diverter, center duct, side duct, and combustor. After the grid generation, full Navier-Stokes flow simulations were conducted with the NPARC Navier-Stokes code. The Chien low-Reynolds-number k-e turbulence model was employed to simulate the high-speed turbulent flow. Finally, the CFD solutions were postprocessed with a Fortran code. This code provided wall static pressure distributions, pitot pressure distributions, mass flow rates, and internal drag. These results were compared with experimental data from a subscale inlet test for code validation; then they were used to help evaluate the demonstrator engine net thrust.

  15. Options for flight testing rocket-based combined-cycle (RBCC) engines

    NASA Technical Reports Server (NTRS)

    Olds, John

    1996-01-01

    While NASA's current next-generation launch vehicle research has largely focused on advanced all-rocket single-stage-to-orbit vehicles (i.e. the X-33 and it's RLV operational follow-on), some attention is being given to advanced propulsion concepts suitable for 'next-generation-and-a-half' vehicles. Rocket-based combined-cycle (RBCC) engines combining rocket and airbreathing elements are one candidate concept. Preliminary RBCC engine development was undertaken by the United States in the 1960's. However, additional ground and flight research is required to bring the engine to technological maturity. This paper presents two options for flight testing early versions of the RBCC ejector scramjet engine. The first option mounts a single RBCC engine module to the X-34 air-launched technology testbed for test flights up to about Mach 6.4. The second option links RBCC engine testing to the simultaneous development of a small-payload (220 lb.) two-stage-to-orbit operational vehicle in the Bantam payload class. This launcher/testbed concept has been dubbed the W vehicle. The W vehicle can also serve as an early ejector ramjet RBCC launcher (albeit at a lower payload). To complement current RBCC ground testing efforts, both flight test engines will use earth-storable propellants for their RBCC rocket primaries and hydrocarbon fuel for their airbreathing modes. Performance and vehicle sizing results are presented for both options.

  16. Rocket-Based Combined Cycle Flowpath Testing for Modes 1 and 4

    NASA Technical Reports Server (NTRS)

    Rice, Tharen

    2002-01-01

    Under sponsorship of the NASA Glenn Research Center (NASA GRC), the Johns Hopkins University Applied Physics Laboratory (JHU/APL) designed and built a five-inch diameter, Rocket-Based Combined Cycle (RBCC) engine to investigate mode 1 and mode 4 engine performance as well as Mach 4 inlet performance. This engine was designed so that engine area and length ratios were similar to the NASA GRC GTX engine is shown. Unlike the GTX semi-circular engine design, the APL engine is completely axisymmetric. For this design, a traditional rocket thruster was installed inside of the scramjet flowpath, along the engine centerline. A three part test series was conducted to determine Mode I and Mode 4 engine performance. In part one, testing of the rocket thruster alone was accomplished and its performance determined (average Isp efficiency = 90%). In part two, Mode 1 (air-augmented rocket) testing was conducted at a nominal chamber pressure-to-ambient pressure ratio of 100 with the engine inlet fully open. Results showed that there was neither a thrust increment nor decrement over rocket-only thrust during Mode 1 operation. In part three, Mode 4 testing was conducted with chamber pressure-to-ambient pressure ratios lower than desired (80 instead of 600) with the inlet fully closed. Results for this testing showed a performance decrease of 20% as compared to the rocket-only testing. It is felt that these results are directly related to the low pressure ratio tested and not the engine design. During this program, Mach 4 inlet testing was also conducted. For these tests, a moveable centerbody was tested to determine the maximum contraction ratio for the engine design. The experimental results agreed with CFD results conducted by NASA GRC, showing a maximum geometric contraction ratio of approximately 10.5. This report details the hardware design, test setup, experimental results and data analysis associated with the aforementioned tests.

  17. Analysis of a New Rocket-Based Combined-Cycle Engine Concept at Low Speed

    NASA Technical Reports Server (NTRS)

    Yungster, S.; Trefny, C. J.

    1999-01-01

    An analysis of the Independent Ramjet Stream (IRS) cycle is presented. The IRS cycle is a variation of the conventional ejector-Ramjet, and is used at low speed in a rocket-based combined-cycle (RBCC) propulsion system. In this new cycle, complete mixing between the rocket and ramjet streams is not required, and a single rocket chamber can be used without a long mixing duct. Furthermore, this concept allows flexibility in controlling the thermal choke process. The resulting propulsion system is intended to be simpler, more robust, and lighter than an ejector-ramjet. The performance characteristics of the IRS cycle are analyzed for a new single-stage-to-orbit (SSTO) launch vehicle concept, known as "Trailblazer." The study is based on a quasi-one-dimensional model of the rocket and air streams at speeds ranging from lift-off to Mach 3. The numerical formulation is described in detail. A performance comparison between the IRS and ejector-ramjet cycles is also presented.

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

    NASA Technical Reports Server (NTRS)

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

    2002-01-01

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

  19. Studies of an extensively axisymmetric rocket based combined cycle (RBCC) engine powered single-stage-to-orbit (SSTO) vehicle

    SciTech Connect

    Foster, R.W.; Escher, W.J.D.; Robinson, J.W.

    1989-01-01

    The present comparative performance study has established that rocket-based combined cycle (RBCC) propulsion systems, when incorporated by essentially axisymmetric SSTO launch vehicle configurations whose conical forebody maximizes both capture-area ratio and total capture area, are capable of furnishing payload-delivery capabilities superior to those of most multistage, all-rocket launchers. Airbreathing thrust augmentation in the rocket-ejector mode of an RBCC powerplant is noted to make a major contribution to final payload capability, by comparison to nonair-augmented rocket engine propulsion systems. 16 refs.

  20. Studies of an extensively axisymmetric rocket based combined cycle (RBCC) engine powered single-stage-to-orbit (SSTO) vehicle

    NASA Technical Reports Server (NTRS)

    Foster, Richard W.; Escher, William J. D.; Robinson, John W.

    1989-01-01

    The present comparative performance study has established that rocket-based combined cycle (RBCC) propulsion systems, when incorporated by essentially axisymmetric SSTO launch vehicle configurations whose conical forebody maximizes both capture-area ratio and total capture area, are capable of furnishing payload-delivery capabilities superior to those of most multistage, all-rocket launchers. Airbreathing thrust augmentation in the rocket-ejector mode of an RBCC powerplant is noted to make a major contribution to final payload capability, by comparison to nonair-augmented rocket engine propulsion systems.

  1. Turbulent Mixing of Primary and Secondary Flow Streams in a Rocket-Based Combined Cycle Engine

    NASA Technical Reports Server (NTRS)

    Cramer, J. M.; Greene, M. U.; Pal, S.; Santoro, R. J.; Turner, Jim (Technical Monitor)

    2002-01-01

    This viewgraph presentation gives an overview of the turbulent mixing of primary and secondary flow streams in a rocket-based combined cycle (RBCC) engine. A significant RBCC ejector mode database has been generated, detailing single and twin thruster configurations and global and local measurements. On-going analysis and correlation efforts include Marshall Space Flight Center computational fluid dynamics modeling and turbulent shear layer analysis. Potential follow-on activities include detailed measurements of air flow static pressure and velocity profiles, investigations into other thruster spacing configurations, performing a fundamental shear layer mixing study, and demonstrating single-shot Raman measurements.

  2. Preliminary Sizing of Vertical Take-off Rocket-based Combined-cycle Powered Launch Vehicles

    NASA Technical Reports Server (NTRS)

    Roche, Joseph M.; McCurdy, David R.

    2001-01-01

    The task of single-stage-to-orbit has been an elusive goal due to propulsion performance, materials limitations, and complex system integration. Glenn Research Center has begun to assemble a suite of relationships that tie Rocket-Based Combined-Cycle (RBCC) performance and advanced material data into a database for the purpose of preliminary sizing of RBCC-powered launch vehicles. To accomplish this, a near optimum aerodynamic and structural shape was established as a baseline. The program synthesizes a vehicle to meet the mission requirements, tabulates the results, and plots the derived shape. A discussion of the program architecture and an example application is discussed herein.

  3. Propulsion/ASME Rocket-Based Combined Cycle Activities in the Advanced Space Transportation Program Office

    NASA Technical Reports Server (NTRS)

    Hueter, Uwe; Turner, James

    1998-01-01

    NASA's Office Of Aeronautics and Space Transportation Technology (OASTT) has establish three major coals. "The Three Pillars for Success". The Advanced Space Transportation Program Office (ASTP) at the NASA's Marshall Space Flight Center in Huntsville,Ala. focuses on future space transportation technologies under the "Access to Space" pillar. The Advanced Reusable Technologies (ART) Project, part of ASTP, focuses on the reusable technologies beyond those being pursued by X-33. The main activity over the past two and a half years has been on advancing the rocket-based combined cycle (RBCC) technologies. In June of last year, activities for reusable launch vehicle (RLV) airframe and propulsion technologies were initiated. These activities focus primarily on those technologies that support the year 2000 decision to determine the path this country will take for Space Shuttle and RLV. In February of this year, additional technology efforts in the reusable technologies were awarded. The RBCC effort that was completed early this year was the initial step leading to flight demonstrations of the technology for space launch vehicle propulsion. Aerojet, Boeing-Rocketdyne and Pratt & Whitney were selected for a two-year period to design, build and ground test their RBCC engine concepts. In addition, ASTROX, Pennsylvania State University (PSU) and University of Alabama in Huntsville also conducted supporting activities. The activity included ground testing of components (e.g., injectors, thrusters, ejectors and inlets) and integrated flowpaths. An area that has caused a large amount of difficulty in the testing efforts is the means of initiating the rocket combustion process. All three of the prime contractors above were using silane (SiH4) for ignition of the thrusters. This follows from the successful use of silane in the NASP program for scramjet ignition. However, difficulties were immediately encountered when silane (an 80/20 mixture of hydrogen/silane) was used for rocket

  4. Aero-Thermo-Structural Analysis of Inlet for Rocket Based Combined Cycle Engines

    NASA Technical Reports Server (NTRS)

    Shivakumar, K. N.; Challa, Preeti; Sree, Dave; Reddy, Dhanireddy R. (Technical Monitor)

    2000-01-01

    NASA has been developing advanced space transportation concepts and technologies to make access to space less costly. One such concept is the reusable vehicles with short turn-around times. The NASA Glenn Research Center's concept vehicle is the Trailblazer powered by a rocket-based combined cycle (RBCC) engine. Inlet is one of the most important components of the RBCC engine. This paper presents fluid flow, thermal, and structural analysis of the inlet for Mach 6 free stream velocity for fully supersonic and supercritical with backpressure conditions. The results concluded that the fully supersonic condition was the most severe case and the largest stresses occur in the ceramic matrix composite layer of the inlet cowl. The maximum tensile and the compressive stresses were at least 3.8 and 3.4, respectively, times less than the associated material strength.

  5. 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.

  6. Rocket-Based Combined-Cycle Propulsion Technology for Access-to-Space Applications

    NASA Technical Reports Server (NTRS)

    Hueter, Uwe

    1999-01-01

    NASA's Office of Aero-Space Technology (OAST) established three major goals, referred to as, "The Three Pillars for Success". The Advanced Space Transportation Program Office (ASTP) at the NASA's Marshall Space Flight Center (MSFC) in Huntsville, Ala. focuses on future space transportation technologies under the "Access to Space" pillar. One of the main activities over the past three years has been on advancing the hydrogen fueled rocket-based combined cycle (RBCC) technologies. The RBCC effort that was completed early this year was the initial step leading to flight demonstrations of the technology for space launch vehicle propulsion. Aerojet and Boeing-Rocketdyne designed, built and ground tested their RBCC engine concepts. In addition, ASTROX, Georgia Institute of Technology, McKinney Associates, Pennsylvania State University (PSU), and University of Alabama in Huntsville conducted supporting activities. The RBCC activity included ground testing of components (e.g., injectors, thrusters, ejectors and inlets) and integrated flowpaths. Inlet testing was performed at the Lewis Research Center's 1 x 1 wind tunnel. All direct connect and free-jet engine testing were conducted at the GASL facilities on Long Island, New York. Testing spanned the Mach range from sea level static to Mach 8. Testing of the rocket-only mode, simulating the final phase of the ascent mission profile, was also performed. The originally planned work on these contracts was completed in 1999. Follow-on activities have been initiated for both hydrogen and hydrocarbon fueled RBCC concepts. Studies to better understand system level issues with the integration of RBCC propulsion with earth-to-orbit vehicles have also been conducted. This paper describes the status, progress and future plans of the RBCC activities funded by NASA/MSFC with a major focus on the benefits of utilizing air-breathing combined-cycle propulsion in access-to-space applications.

  7. Rocket-Based Combined-Cycle (RBCC) Propulsion Technology Workshop. Tutorial session

    NASA Technical Reports Server (NTRS)

    1992-01-01

    The goal of this workshop was to illuminate the nation's space transportation and propulsion engineering community on the potential of hypersonic combined cycle (airbreathing/rocket) propulsion systems for future space transportation applications. Four general topics were examined: (1) selections from the expansive advanced propulsion archival resource; (2) related propulsion systems technical backgrounds; (3) RBCC engine multimode operations related subsystem background; and (4) focused review of propulsion aspects of current related programs.

  8. Rocket-Based Combined Cycle Engine Technology Development: Inlet CFD Validation and Application

    NASA Technical Reports Server (NTRS)

    DeBonis, J. R.; Yungster, S.

    1996-01-01

    A CFD methodology has been developed for inlet analyses of Rocket-Based Combined Cycle (RBCC) Engines. A full Navier-Stokes analysis code, NPARC, was used in conjunction with pre- and post-processing tools to obtain a complete description of the flow field and integrated inlet performance. This methodology was developed and validated using results from a subscale test of the inlet to a RBCC 'Strut-Jet' engine performed in the NASA Lewis 1 x 1 ft. supersonic wind tunnel. Results obtained from this study include analyses at flight Mach numbers of 5 and 6 for super-critical operating conditions. These results showed excellent agreement with experimental data. The analysis tools were also used to obtain pre-test performance and operability predictions for the RBCC demonstrator engine planned for testing in the NASA Lewis Hypersonic Test Facility. This analysis calculated the baseline fuel-off internal force of the engine which is needed to determine the net thrust with fuel on.

  9. Three Dimensional Numerical Simulation of Rocket-based Combined-cycle Engine Response During Mode Transition Events

    NASA Technical Reports Server (NTRS)

    Edwards, Jack R.; McRae, D. Scott; Bond, Ryan B.; Steffan, Christopher (Technical Monitor)

    2003-01-01

    The GTX program at NASA Glenn Research Center is designed to develop a launch vehicle concept based on rocket-based combined-cycle (RBCC) propulsion. Experimental testing, cycle analysis, and computational fluid dynamics modeling have all demonstrated the viability of the GTX concept, yet significant technical issues and challenges still remain. Our research effort develops a unique capability for dynamic CFD simulation of complete high-speed propulsion devices and focuses this technology toward analysis of the GTX response during critical mode transition events. Our principal attention is focused on Mode 1/Mode 2 operation, in which initial rocket propulsion is transitioned into thermal-throat ramjet propulsion. A critical element of the GTX concept is the use of an Independent Ramjet Stream (IRS) cycle to provide propulsion at Mach numbers less than 3. In the IRS cycle, rocket thrust is initially used for primary power, and the hot rocket plume is used as a flame-holding mechanism for hydrogen fuel injected into the secondary air stream. A critical aspect is the establishment of a thermal throat in the secondary stream through the combination of area reduction effects and combustion-induced heat release. This is a necessity to enable the power-down of the rocket and the eventual shift to ramjet mode. Our focus in this first year of the grant has been in three areas, each progressing directly toward the key initial goal of simulating thermal throat formation during the IRS cycle: CFD algorithm development; simulation of Mode 1 experiments conducted at Glenn's Rig 1 facility; and IRS cycle simulations. The remainder of this report discusses each of these efforts in detail and presents a plan of work for the next year.

  10. A rocket-based combined-cycle engine prototype demonstrating comprehensive component compatibility and effective mode transition

    NASA Astrophysics Data System (ADS)

    Shi, Lei; He, Guoqiang; Liu, Peijin; Qin, Fei; Wei, Xianggeng; Liu, Jie; Wu, Lele

    2016-11-01

    A rocket-based combined cycle (RBCC) engine was designed to demonstrate its broad applicability in the ejector and ramjet modes within the flight range from Mach 0 to Mach 4.5. To validate the design, a prototype was fabricated and tested as a freejet engine operating at flight Mach 3 using hydrocarbon fuel. The proposed design was a single module, heat sink steel alloy model with an interior fuel supply and active control system and a fully integrated flowpath that was comprehensively instrumented with pressure sensors. The mass capture and back pressure resistance of the inlet were numerically investigated and experimentally calibrated. The combustion process and rocket operation during mode transition were investigated by direct-connect tests. Finally, the comprehensive component compatibility and multimodal operational capability of the RBCC engine prototype was validated through freejet tests. This paper describes the design of the RBCC engine prototype, reviews the testing procedures, and discusses the experimental results of these efforts in detail.

  11. Numerical analysis of flow features and operation characteristics of a rocket-based combined-cycle inlet in ejector mode

    NASA Astrophysics Data System (ADS)

    Shi, Lei; Liu, Xiaowei; He, Guoqiang; Qin, Fei; Wei, Xianggeng; Yang, Bin; Liu, Jie

    2016-10-01

    A ready-made central strut-based rocket-based combined-cycle (RBCC) engine was numerically investigated in the ejector mode. The flow features in the RBCC inlet and the matching characteristics between the inlet and the embedded rocket during different flight regimes were examined in detail. It was necessary to perform integrated numerical simulations in the ejector mode within considerable pressure far fields around the inlet/exhaust system. The observed flow features and operation characteristics in the RBCC inlet were strongly correlated with the flight conditions, inlet configuration, and operation of the embedded rocket. It was further found that the integrated function status of multiple factors significantly influenced the performance of the RBCC engine in the ejector mode. The two parameters that macroscopically affected the performance most were the air entrainment mass and the drag of the RBCC inlet. To improve these parameters, it is vital to employ an appropriate design of the RBCC inlet and establish the optimal flight trajectory of the flight vehicle.

  12. Reuse fo a Cold War Surveillance Drone to Flight Test a NASA Rocket Based Combined Cycle Engine

    NASA Technical Reports Server (NTRS)

    Brown, T. M.; Smith, Norm

    1999-01-01

    Plans for and early feasibility investigations into the modification of a Lockheed D21B drone to flight test the DRACO Rocket Based Combined Cycle (RBCC) engine are discussed. Modifications include the addition of oxidizer tanks, modern avionics systems, actuators, and a vehicle recovery system. Current study results indicate that the D21B is a suitable candidate for this application and will allow demonstrations of all DRACO engine operating modes at Mach numbers between 0.8 and 4.0. Higher Mach numbers may be achieved with more extensive modification. Possible project risks include low speed stability and control, and recovery techniques.

  13. Inlet Development for a Rocket Based Combined Cycle, Single Stage to Orbit Vehicle Using Computational Fluid Dynamics

    NASA Technical Reports Server (NTRS)

    DeBonis, J. R.; Trefny, C. J.; Steffen, C. J., Jr.

    1999-01-01

    Design and analysis of the inlet for a rocket based combined cycle engine is discussed. Computational fluid dynamics was used in both the design and subsequent analysis. Reynolds averaged Navier-Stokes simulations were performed using both perfect gas and real gas assumptions. An inlet design that operates over the required Mach number range from 0 to 12 was produced. Performance data for cycle analysis was post processed using a stream thrust averaging technique. A detailed performance database for cycle analysis is presented. The effect ot vehicle forebody compression on air capture is also examined.

  14. Rocket-based combined-cycle (RBCC) powered spaceliner class vehicle can advantageously employ vertical takeoff and landing (VTOL)

    NASA Technical Reports Server (NTRS)

    Escher, William J. D.

    1995-01-01

    The subject is next generation orbital space transporation, taken to be fully reusable non-staged 'aircraft like' systems targeted for routine, affordable access to space. Specifically, the takeoff and landing approach to be selected for such systems is considered, mainly from a propulsion viewpoint. Conventional wisdom has it that any transatmospheric-class vehicle which uses high-speed airbreathing propulsion modes (e.g., scramjet) intrinsically must utilize horizontal takeoff and landing, HTOHL. Although this may be true for all-airbreathing propulsion (i.e., no rocket content as in turboramjet propulsion), that emerging class of powerplant which integrally combines airbreathing and rocket propulsion, referred to as rocket-based combined-cycle (RBCC) propulsion, is considerably more flexible with respect to selecting takeoff/landing modes. In fact, it is proposed that any of the modes of interest may potentially be selected: HTOHL, VTOHL, VTOVL. To illustrate this surmise, the case of a previously documented RBCC-powered 'Spaceliner' class space transport concept, which is designed for vertical takeoff and landing, is examined. The 'RBCC' and 'Spaceliner' categories are first described for background. Departing form an often presumed HTOHL baseline, the leading design and operational advantages of moving to VTOVL are then elucidated. Technical substantiation that the RBCC approach, in fact, enables this capability (but also that of HTOHL and VTOVL) is provided, with extensive reference to case-in-point supporting studies. The paper closes with a set of conditional surmises bearing on its set of conclusions, which point up the operational cost advantages associated with selecting the vertical takeoff and landing mode combination (VTOL), uniquely offered by RBCC propulsion.

  15. Ongoing Analyses of Rocket Based Combined Cycle Engines by the Applied Fluid Dynamics Analysis Group at Marshall Space Flight Center

    NASA Technical Reports Server (NTRS)

    Ruf, Joseph H.; Holt, James B.; Canabal, Francisco

    2001-01-01

    This paper presents the status of analyses on three Rocket Based Combined Cycle (RBCC) configurations underway in the Applied Fluid Dynamics Analysis Group (TD64). TD64 is performing computational fluid dynamics (CFD) analysis on a Penn State RBCC test rig, the proposed Draco axisymmetric RBCC engine and the Trailblazer engine. The intent of the analysis on the Penn State test rig is to benchmark the Finite Difference Navier Stokes (FDNS) code for ejector mode fluid dynamics. The Draco analysis was a trade study to determine the ejector mode performance as a function of three engine design variables. The Trailblazer analysis is to evaluate the nozzle performance in scramjet mode. Results to date of each analysis are presented.

  16. Ongoing Analysis of Rocket Based Combined Cycle Engines by the Applied Fluid Dynamics Analysis Group at Marshall Space Flight Center

    NASA Technical Reports Server (NTRS)

    Ruf, Joseph; Holt, James B.; Canabal, Francisco

    1999-01-01

    This paper presents the status of analyses on three Rocket Based Combined Cycle configurations underway in the Applied Fluid Dynamics Analysis Group (TD64). TD64 is performing computational fluid dynamics analysis on a Penn State RBCC test rig, the proposed Draco axisymmetric RBCC engine and the Trailblazer engine. The intent of the analysis on the Penn State test rig is to benchmark the Finite Difference Navier Stokes code for ejector mode fluid dynamics. The Draco engine analysis is a trade study to determine the ejector mode performance as a function of three engine design variables. The Trailblazer analysis is to evaluate the nozzle performance in scramjet mode. Results to date of each analysis are presented.

  17. Experimental Analysis of a Rocket Based Combined Cycle (RBCC) Engine in a Direct-Connect Test Facility

    NASA Technical Reports Server (NTRS)

    Nelson, K.; Hawk, Clark W.

    1997-01-01

    The object of this study is to investigate the operation of a RBCC at ramjet and scramjet flight conditions using a direct-connect test facility. The apparatus being tested is a single strut-rocket within a dual-mode ram/scramjet combustor. The gaseous hydrogen/oxygen, linear strut-rocket was supplied by Aerojet Propulsion Company. The hardware is being tested in the Direct Connect Supersonic Combustion Test Facility at NASA Langley Research Center. The test facilities hydrogen/oxygen vitiated heater is capable of flight total enthalpies to Mach 8. A Mach 2.5 facility nozzle mates the heater to the combustor duct. The rocket ejector will ordinarily operate in a fuel-rich mode. Additional fuel injection is provided by a pair of parallel injectors located at the base of the strut body. Instrumentation on the test apparatus includes a unique, direct thrust measurement system. Performance predictions for the anticipated test conditions have been made using a one-dimensional, thermodynamic analysis code. Results from the code show the dependence of overall thrust and specific impulse on rocket chamber pressure, rocket fuel equivalence ratio, and overall fuel equivalence ratio. Once the experimental test series begins, the inferred combustion efficiency as a function of axial location and the thermal choke region (where applicable) can also be determined using this code. Upon completion of the experimental test series, measurements will be used to calculate thrust, specific impulse, etc. Measured and calculated values will be compared to those found analytically. If appropriate, the code will be tailored to better predict hardware operation. Conclusions will be drawn as to the fuel-rich rocket's overall effect on ramjet and scramjet performance. Also, comparisons will be made between the integrated thrust calculated from the static pressure taps located along the duct and the thrust measured by the direct thrust measurement system.

  18. Numerical Investigation of Cowl Lip Adjustments for a Rocket-Based Combined-Cycle Inlet in Takeoff Regime

    NASA Astrophysics Data System (ADS)

    Shi, Lei; Liu, Xiaowei; He, Guoqiang; Qin, Fei; Wei, Xianggeng; Yang, Bing; Wu, Lele

    2016-09-01

    Numerical integration simulations were performed on a ready-made central strut-based rocket-based combined-cycle (RBCC) engine operating in the ejector mode during the takeoff regime. The effective principles of various cowl lip positions and shapes on the inlet operation and the overall performance of the entire engine were investigated in detail. Under the static condition, reverse cowl lip rotation in a certain range was found to contribute comprehensive improvement to the RBCC inlet and the entire engine. However, the reverse rotation of the cowl lip contributed very little enhancement of the RBCC inlet under the low subsonic flight regime and induced extremely negative impacts in the high subsonic flight regime, especially in terms of a significant increase in the drag of the inlet. Changes to the cowl lip shape provided little improvement to the overall performance of the RBCC engine, merely shifting the location of the leeward area inside the RBCC inlet, as well as the flow separation and eddy, but not relieving or eliminating those phenomena. The results of this study indicate that proper cowl lip rotation offers an efficient variable geometry scheme for a RBCC inlet in the takeoff regime.

  19. Parametric Data from a Wind Tunnel Test on a Rocket-Based Combined-Cycle Engine Inlet

    NASA Technical Reports Server (NTRS)

    Fernandez, Rene; Trefny, Charles J.; Thomas, Scott R.; Bulman, Mel J.

    2001-01-01

    A 40-percent scale model of the inlet to a rocket-based combined-cycle (RBCC) engine was tested in the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT). The full-scale RBCC engine is scheduled for test in the Hypersonic Tunnel Facility (HTF) at NASA Glenn's Plum Brook Station at Mach 5 and 6. This engine will incorporate the configuration of this inlet model which achieved the best performance during the present experiment. The inlet test was conducted at Mach numbers of 4.0, 5.0, 5.5, and 6.0. The fixed-geometry inlet consists of an 8 deg.. forebody compression plate, boundary layer diverter, and two compressive struts located within 2 parallel sidewalls. These struts extend through the inlet, dividing the flowpath into three channels. Test parameters investigated included strut geometry, boundary layer ingestion, and Reynolds number (Re). Inlet axial pressure distributions and cross-sectional Pitot-pressure surveys at the base of the struts were measured at varying back-pressures. Inlet performance and starting data are presented. The inlet chosen for the RBCC engine self-started at all Mach numbers from 4 to 6. Pitot-pressure contours showed large flow nonuniformity on the body-side of the inlet. The inlet provided adequate pressure recovery and flow quality for the RBCC cycle even with the flow separation.

  20. Large eddy simulation of combustion characteristics in a kerosene fueled rocket-based combined-cycle engine combustor

    NASA Astrophysics Data System (ADS)

    Huang, Zhi-wei; He, Guo-qiang; Qin, Fei; Cao, Dong-gang; Wei, Xiang-geng; Shi, Lei

    2016-10-01

    This study reports combustion characteristics of a rocket-based combined-cycle engine combustor operating at ramjet mode numerically. Compressible large eddy simulation with liquid kerosene sprayed and vaporized is used to study the intrinsic unsteadiness of combustion in such a propulsion system. Results for the pressure oscillation amplitude and frequency in the combustor as well as the wall pressure distribution along the flow-path, are validated using experimental data, and they show acceptable agreement. Coupled with reduced chemical kinetics of kerosene, results are compared with the simultaneously obtained Reynolds-Averaged Navier-Stokes results, and show significant differences. A flow field analysis is also carried out for further study of the turbulent flame structures. Mixture fraction is used to determine the most probable flame location in the combustor at stoichiometric condition. Spatial distributions of the Takeno flame index, scalar dissipation rate, and heat release rate reveal that different combustion modes, such as premixed and non-premixed modes, coexisted at different sections of the combustor. The RBCC combustor is divided into different regions characterized by their non-uniform features. Flame stabilization mechanism, i.e., flame propagation or fuel auto-ignition, and their relative importance, is also determined at different regions in the combustor.

  1. Preliminary Sizing Completed for Single- Stage-To-Orbit Launch Vehicles Powered By Rocket-Based Combined Cycle Technology

    NASA Technical Reports Server (NTRS)

    Roche, Joseph M.

    2002-01-01

    Single-stage-to-orbit (SSTO) propulsion remains an elusive goal for launch vehicles. The physics of the problem is leading developers to a search for higher propulsion performance than is available with all-rocket power. Rocket-based combined cycle (RBCC) technology provides additional propulsion performance that may enable SSTO flight. Structural efficiency is also a major driving force in enabling SSTO flight. Increases in performance with RBCC propulsion are offset with the added size of the propulsion system. Geometrical considerations must be exploited to minimize the weight. Integration of the propulsion system with the vehicle must be carefully planned such that aeroperformance is not degraded and the air-breathing performance is enhanced. Consequently, the vehicle's structural architecture becomes one with the propulsion system architecture. Geometrical considerations applied to the integrated vehicle lead to low drag and high structural and volumetric efficiency. Sizing of the SSTO launch vehicle (GTX) is itself an elusive task. The weight of the vehicle depends strongly on the propellant required to meet the mission requirements. Changes in propellant requirements result in changes in the size of the vehicle, which in turn, affect the weight of the vehicle and change the propellant requirements. An iterative approach is necessary to size the vehicle to meet the flight requirements. GTX Sizer was developed to do exactly this. The governing geometry was built into a spreadsheet model along with scaling relationships. The scaling laws attempt to maintain structural integrity as the vehicle size is changed. Key aerodynamic relationships are maintained as the vehicle size is changed. The closed weight and center of gravity are displayed graphically on a plot of the synthesized vehicle. In addition, comprehensive tabular data of the subsystem weights and centers of gravity are generated. The model has been verified for accuracy with finite element analysis. The

  2. Reynolds-averaged Navier-Stokes analysis of the flow through a model rocket-based combined-cycle engine with an independently-fueled ramjet stream

    NASA Astrophysics Data System (ADS)

    Bond, Ryan Bomar

    A new concept for the low speed propulsion mode in rocket based combined cycle (RBCC) engines has been developed as part of the NASA GTX program. This concept, called the independent ramjet stream (IRS) cycle, is a variation of the traditional ejector ramjet (ER) design and involves the injection of hydrogen fuel directly into the air stream, where it is ignited by the rocket plume. Experiments and computational fluid dynamics (CFD) are currently being used to evaluate the feasibility of the new design. In this work, a Navier-Stokes code valid for general reactive flows is applied to the model engine under cold flow, ejector ramjet, and IRS cycle operation. Pressure distributions corresponding to cold-flow and ejector ramjet operation are compared with experimental data. The engine response under independent ramjet stream cycle operation is examined for different reaction models and grid sizes. The engine response to variations in fuel injection is also examined. Mode transition simulations are also analyzed both with and without a nitrogen purge of the rocket. The solutions exhibit a high sensitivity to both grid resolution and reaction mechanism, but they do indicate that thermal throat ramjet operation is possible through the injection and burning of additional fuel into the air stream. The solutions also indicate that variations in fuel injection location can affect the position of the thermal throat. The numerical simulations predicted successful mode transition both with and without a nitrogen purge of the rocket; however, the reliability of the mode transition results cannot be established without experimental data to validate the reaction mechanism.

  3. Supersonic Wind Tunnel Tests of a Half-axisymmetric 12 Deg-spike Inlet to a Rocket-based Combined-cycle Propulsion System

    NASA Technical Reports Server (NTRS)

    DeBonis, J. R.; Trefny, C. J.

    2001-01-01

    Results of an isolated inlet test for NASA's GTX air-breathing launch vehicle concept are presented. The GTX is a Vertical Take-off/ Horizontal Landing reusable single-stage-to-orbit system powered by a rocket-based combined-cycle propulsion system. Tests were conducted in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel during two entries in October 1998 and February 1999. Tests were run from Mach 2.8 to 6. Integrated performance parameters and static pressure distributions are reported. The maximum contraction ratios achieved in the tests were lower than predicted by axisymmetric Reynolds-averaged Navier-Stokes computational fluid dynamics (CFD). At Mach 6, the maximum contraction ratio was roughly one-half of the CFD value of 16. The addition of either boundary-layer trip strips or vortex generators had a negligible effect on the maximum contraction ratio. A shock boundary-layer interaction was also evident on the end-walls that terminate the annular flowpath cross section. Cut-back end-walls, designed to reduce the boundary-layer growth upstream of the shock and minimize the interaction, also had negligible effect on the maximum contraction ratio. Both the excessive turning of low-momentum comer flows and local over-contraction due to asymmetric end-walls were identified as possible reasons for the discrepancy between the CFD predictions and the experiment. It is recommended that the centerbody spike and throat angles be reduced in order to lessen the induced pressure rise. The addition of a step on the cowl surface, and planar end-walls more closely approximating a plane of symmetry are also recommended. Provisions for end-wall boundary-layer bleed should be incorporated.

  4. US Rocket Propulsion Industrial Base Health Metrics

    NASA Technical Reports Server (NTRS)

    Doreswamy, Rajiv

    2013-01-01

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

  5. Design Study: Rocket Based MHD Generator

    NASA Technical Reports Server (NTRS)

    1997-01-01

    This report addresses the technical feasibility and design of a rocket based MHD generator using a sub-scale LOx/RP rocket motor. The design study was constrained by assuming the generator must function within the performance and structural limits of an existing magnet and by assuming realistic limits on (1) the axial electric field, (2) the Hall parameter, (3) current density, and (4) heat flux (given the criteria of heat sink operation). The major results of the work are summarized as follows: (1) A Faraday type of generator with rectangular cross section is designed to operate with a combustor pressure of 300 psi. Based on a magnetic field strength of 1.5 Tesla, the electrical power output from this generator is estimated to be 54.2 KW with potassium seed (weight fraction 3.74%) and 92 KW with cesium seed (weight fraction 9.66%). The former corresponds to a enthalpy extraction ratio of 2.36% while that for the latter is 4.16%; (2) A conceptual design of the Faraday MHD channel is proposed, based on a maximum operating time of 10 to 15 seconds. This concept utilizes a phenolic back wall for inserting the electrodes and inter-electrode insulators. Copper electrode and aluminum oxide insulator are suggested for this channel; and (3) A testing configuration for the sub-scale rocket based MHD system is proposed. An estimate of performance of an ideal rocket based MHD accelerator is performed. With a current density constraint of 5 Amps/cm(exp 2) and a conductivity of 30 Siemens/m, the push power density can be 250, 431, and 750 MW/m(sup 3) when the induced voltage uB have values of 5, 10, and 15 KV/m, respectively.

  6. A review of findings of a study of rocket based combined cycle engines applied to extensively axisymmetric single stage to orbit vehicles

    NASA Technical Reports Server (NTRS)

    Foster, Richard W.

    1992-01-01

    Extensively axisymmetric and non-axisymmetric Single Stage To Orbit (SSTO) vehicles are considered. The information is presented in viewgraph form and the following topics are presented: payload comparisons; payload as a percent of dry weight - a system hardware cost indicator; life cycle cost estimations; operations and support costs estimation; selected engine type; and rocket engine specific impulse calculation.

  7. 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.

  8. Rehabilitation of the Rocket Vehicle Integration Test Stand at Edwards Air Force Base

    NASA Technical Reports Server (NTRS)

    Jones, Daniel S.; Ray, Ronald J.; Phillips, Paul

    2005-01-01

    Since initial use in 1958 for the X-15 rocket-powered research airplane, the Rocket Engine Test Facility has proven essential for testing and servicing rocket-powered vehicles at Edwards Air Force Base. For almost two decades, several successful flight-test programs utilized the capability of this facility. The Department of Defense has recently demonstrated a renewed interest in propulsion technology development with the establishment of the National Aerospace Initiative. More recently, the National Aeronautics and Space Administration is undergoing a transformation to realign the organization, focusing on the Vision for Space Exploration. These initiatives provide a clear indication that a very capable ground-test stand at Edwards Air Force Base will be beneficial to support the testing of future access-to-space vehicles. To meet the demand of full integration testing of rocket-powered vehicles, the NASA Dryden Flight Research Center, the Air Force Flight Test Center, and the Air Force Research Laboratory have combined their resources in an effort to restore and upgrade the original X-15 Rocket Engine Test Facility to become the new Rocket Vehicle Integration Test Stand. This report describes the history of the X-15 Rocket Engine Test Facility, discusses the current status of the facility, and summarizes recent efforts to rehabilitate the facility to support potential access-to-space flight-test programs. A summary of the capabilities of the facility is presented and other important issues are discussed.

  9. Hybrid rockets - Combining the best of liquids and solids

    NASA Technical Reports Server (NTRS)

    Cook, Jerry R.; Goldberg, Ben E.; Estey, Paul N.; Wiley, Dan R.

    1992-01-01

    Hybrid rockets employing liquid oxidizer and solid fuel grain answers to cost, safety, reliability, and environmental impact concerns that have become as prominent as performance in recent years. The oxidizer-free grain has limited sensitivity to grain anomalies, such as bond-line separations, which can cause catastrophic failures in solid rocket motors. An account is presently given of the development effort associated with the AMROC commercial hybrid booster and component testing efforts at NASA-Marshall. These hybrid rockets can be fired, terminated, inspected, evaluated, and restarted for additional testing.

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

    NASA Technical Reports Server (NTRS)

    Ganapathi, Gani; Narayan, Sri

    2010-01-01

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

  11. Study Of Heating Of The Base Region Of A Rocket

    NASA Technical Reports Server (NTRS)

    Ascoli, Edward P.; Heiba, Adel A.; Hsu, Yann-Fu; Lagnado, Ronald R.; Lynch, Edward D.; Ungewitter, Ronald J.

    1994-01-01

    Report describes theoretical study of heating in base region of proposed rocket called "NLS 1.5 stage reference vehicle." Study employed approach based on computational fluid dynamics (CFD). Involved numerical simulations of flow field in base region and in main exhaust plume of cluster of six engines with heat shields.

  12. Spaceliner Class Operability Gains Via Combined Airbreathing/ Rocket Propulsion: Summarizing an Operational Assessment of Highly Reusable Space Transports

    NASA Technical Reports Server (NTRS)

    Nix, Michael B.; Escher, William J. d.

    1999-01-01

    In discussing a new NASA initiative in advanced space transportation systems and technologies, the Director of the NASA Marshall Space Flight Center, Arthur G. Stephenson, noted that, "It would use new propulsion technology, air-breathing engine so you don't have to carry liquid oxygen, at least while your flying through the atmosphere. We are calling it Spaceliner 100 because it would be 100 times cheaper, costing $ 100 dollars a pound to orbit." While airbreathing propulsion is directly named, rocket propulsion is also implied by, "... while you are flying through the atmosphere." In-space final acceleration to orbital speed mandates rocket capabilities. Thus, in this informed view, Spaceliner 100 will be predicated on combined airbreathing/rocket propulsion, the technical subject of this paper. Interestingly, NASA's recently concluded Highly Reusable Space Transportation (HRST) study focused on the same affordability goal as that of the Spaceliner 100 initiative and reflected the decisive contribution of combined propulsion as a way of expanding operability and increasing the design robustness of future space transports, toward "aircraft like" capabilities. The HRST study built on the Access to Space Study and the Reusable Launch Vehicle (RLV) development activities to identify and characterize space transportation concepts, infrastructure and technologies that have the greatest potential for reducing delivery cost by another order of magnitude, from $1,000 to $100-$200 per pound for 20,000 lb. - 40.000 lb. payloads to low earth orbit (LEO). The HRST study investigated a number of near-term, far-term, and very far-term launch vehicle concepts including all-rocket single-stage-to-orbit (SSTO) concepts, two-stage-to-orbit (TSTO) concepts, concepts with launch assist, rocket-based combined cycle (RBCC) concepts, advanced expendable vehicles, and more far term ground-based laser powered launchers. The HRST study consisted of preliminary concept studies, assessments

  13. Fiber-optic applications for space-based rocket engines

    NASA Astrophysics Data System (ADS)

    Sovie, Amy L.; Bewley, Douglas P.; Millis, Marc G.

    1991-09-01

    The use of fiber-optic technology is discussed with respect to the instrumentation systems for space-based rocket engines. Optical fiber technologies are reviewed with specific attention given to the reliability, light weight, small fiber diameter, and operating life of the components in the space environment. An optical system can facilitate the incorporation of an optical health-monitoring system, increase the space available for necessary redundancy, and safe high-bandwidth communications that are immune to the effects of electromagnetic radiation.

  14. Biodegradable protein-based rockets for drug transportation and light-triggered release.

    PubMed

    Wu, Zhiguang; Lin, Xiankun; Zou, Xian; Sun, Jianmin; He, Qiang

    2015-01-14

    We describe a biodegradable, self-propelled bovine serum albumin/poly-l-lysine (PLL/BSA) multilayer rocket as a smart vehicle for efficient anticancer drug encapsulation/delivery to cancer cells and near-infrared light controlled release. The rockets were constructed by a template-assisted layer-by-layer assembly of the PLL/BSA layers, followed by incorporation of a heat-sensitive gelatin hydrogel containing gold nanoparticles, doxorubicin, and catalase. These rockets can rapidly deliver the doxorubicin to the targeted cancer cell with a speed of up to 68 μm/s, through a combination of biocatalytic bubble propulsion and magnetic guidance. The photothermal effect of the gold nanoparticles under NIR irradiation enable the phase transition of the gelatin hydrogel for rapid release of the loaded doxorubicin and efficient killing of the surrounding cancer cells. Such biodegradable and multifunctional protein-based microrockets provide a convenient and efficient platform for the rapid delivery and controlled release of therapeutic drugs.

  15. Conceptual study of manned space transportation vehicle using laser thruster in combination with the H-II rocket

    NASA Astrophysics Data System (ADS)

    Minami, Yoshinari; Uchida, Shigeaki

    2013-02-01

    This paper describes the conceptual study of a Manned Space Transportation Vehicle (MSTV) using a laser thruster in combination with the H-II Rocket. By combining the use of a laser thruster and H-II Rocket, space trip to the International Space Station (ISS) or a round trip mission around the moon can be performed. Once MSTV with one crew achieves a circular orbit at an altitude of 200 km around the earth (parking orbit) by use of H-II Rocket, MSTV will then put into a circular orbit into an altitude of 400 km (ISS orbit) from 200 km circular orbit by use of the laser thruster. H-II Rocket has the following launch capability with payloads for LEO (300 km): 10 t (H-II A Rocket), 16.5 t (H-II B Rocket). Laser thruster using water propellant, power source for the laser, orbital transfer calculations (to ISS or the Moon) and other practical aspects are examined.

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

    NASA Astrophysics Data System (ADS)

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

    2004-10-01

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

  17. Combining MHD Airbreathing and Fusion Rocket Propulsion for Earth-to-Orbit Flight

    SciTech Connect

    Froning, H. D. Jr; Yang, Yang; Momota, H.; Burton, E.; Miley, G. H.; Luo, Nie

    2005-02-06

    Previous studies have shown that Single-State-to-Orbit (SSTO) vehicle propellant can be reduced by Magnets-Hydro-Dynamic (MHD) processes that minimize airbreathing propulsion losses and propellant consumption during atmospheric flight. Similarly additional reduction in SSTO propellant is enabled by Inertial Electrostatic Confinement (IEC) fusion, whose more energetic reactions reduce rocket propellant needs. MHD airbreathing propulsion during an SSTO vehicle's initial atmospheric flight phase and IEC fusion propulsion during its final exo-atmospheric flight phase is therefore being explored. Accomplished work is not yet sufficient for claiming such a vehicle's feasibility. But takeoff and propellant mass for an MHD airbreathing and IEC fusion vehicle could be as much as 25 and 40 percent less than one with ordinary airbreathing and IEC fusion; and as much as 50 and 70 percent less than SSTO takeoff and propellant mass with MHD airbreathing and chemical rocket propulsion. Thus this unusual combined cycle engine shows great promise for performance gains beyond contemporary combined-cycle airbreathing engines.

  18. Dual-mode Operation of a Rocket-Ramjet Combined Cycle Engine

    NASA Astrophysics Data System (ADS)

    Tomioka, Sadatake; Tani, Koichiro; Masumoto, Ryo; Ueda, Shuuichi

    One-dimensional evaluation of Ramjet-mode operation was carried out on a rocket-ramjet combined cycle engine model. For simplicity, instantaneous mixing between the airflow and rocket exhaust, instantaneous heat release, and pressure recovery by a normal-shock wave were assumed. Shock wave location was so decided that the heat release at the injection (heat addition) location was to thermally-choke the combustion gas flow. By changing the injection location, it was shown that a further downstream injection resulted in a further thrust production and a further fuel flow rate requirement for choking, and a lesser specific impulse. Balancing the thrust production and the specific impulse in terms of the launch vehicle acceleration performance should be pursued. The total pressure loss within the engine model was dominated by the shock wave location, not depended on injection location and fuel flow rate, so that having shock wave penetration to further upstream location was beneficial both for thrust production in the engine and at the external nozzle.

  19. Flight Measurements of Base Pressure on Bodies of Revolution with and Without Simulated Rocket Chambers

    NASA Technical Reports Server (NTRS)

    Peck, Robert F

    1955-01-01

    Base pressures were measured on fin-stabilized bodies of revolution with and without rocket chambers and with and without a converging afterbody. At Mach numbers between 0.7 and 1.2, the results show that the presence of a "cold" rocket chamber increased the pressure (less suction) over the center portion of the bases. The effects of rocket chambers on pressures near the edge of the bases were not as consistent throughout the Mach number range nor as appreciable at most speeds as were the effects of pressures measured on the center line.

  20. Experimental Study of Ballistic-Missile Base Heating with Operating Rocket

    NASA Technical Reports Server (NTRS)

    Nettle, J. Cary

    1958-01-01

    A rocket of the 1000-pound-thrust class using liquid oxygen and JP-4 fuel as propellant was installed in the Lewis 8- by 6-foot tunnel to permit a controlled study of some of the factors affecting the heating of a rocket-missile base. Temperatures measured in the base region are presented from findings of three motor extension lengths relative to the base. Data are also presented for two combustion efficiency levels in the rocket motor. Temperature as high as 1200 F was measured in the base region because of the ignition of burnable rocket gases. combustibles that are dumped into the base by accessories seriously aggravate the base-burning temperature rise.

  1. 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.

  2. Ground-based firing of a linear aerospike rocket engine

    NASA Technical Reports Server (NTRS)

    1990-01-01

    In this 23-second clip the Rocketdyne linear aerospike rocket engine is fired in a test stand at the Santa Susana facility of Rocketdyne in Chatsworth, California. Unlike conventional rocket engines the linear aerospike has no bell nozzle to contain and direct the expanding gases. Instead, the gases are aimed down a V-shaped ramp, which makes up one side of the 'nozzle,' with the side open to the atmosphere allowing optimum expansion at all altitudes while providing thrust.

  3. Air-Breathing Rocket Engines

    NASA Technical Reports Server (NTRS)

    1998-01-01

    This photograph depicts an air-breathing rocket engine prototype in the test bay at the General Applied Science Lab facility in Ronkonkoma, New York. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced Space Transportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  4. CFD-Based Design Optimization for Single Element Rocket Injector

    NASA Technical Reports Server (NTRS)

    Vaidyanathan, Rajkumar; Tucker, Kevin; Papila, Nilay; Shyy, Wei

    2003-01-01

    To develop future Reusable Launch Vehicle concepts, we have conducted design optimization for a single element rocket injector, with overall goals of improving reliability and performance while reducing cost. Computational solutions based on the Navier-Stokes equations, finite rate chemistry, and the k-E turbulence closure are generated with design of experiment techniques, and the response surface method is employed as the optimization tool. The design considerations are guided by four design objectives motivated by the consideration in both performance and life, namely, the maximum temperature on the oxidizer post tip, the maximum temperature on the injector face, the adiabatic wall temperature, and the length of the combustion zone. Four design variables are selected, namely, H2 flow angle, H2 and O2 flow areas with fixed flow rates, and O2 post tip thickness. In addition to establishing optimum designs by varying emphasis on the individual objectives, better insight into the interplay between design variables and their impact on the design objectives is gained. The investigation indicates that improvement in performance or life comes at the cost of the other. Best compromise is obtained when improvements in both performance and life are given equal importance.

  5. Gravity wave reflection: Case study based on rocket data

    NASA Astrophysics Data System (ADS)

    Wüst, Sabine; Bittner, Michael

    2008-03-01

    Since gravity waves significantly influence the atmosphere by transporting energy and momentum, it is important to study their wave spectrum and their energy dissipation rates. Besides that, knowledge about gravity wave sources and the propagation of the generated waves is essential. Originating in the lower atmosphere, gravity waves can move upwards; when the background wind field is equal to their phase speed a so-called critical layer is reached. Their breakdown and deposition of energy and momentum is possible. Another mechanism which can take place at critical layers is gravity wave reflection. In this paper, gravity waves which were observed by foil chaff measurements during the DYANA (DYnamics Adapted Network for the Atmosphere) campaign in 1990 in Biscarrosse (44°N, 1°W)--as reported by Wüst and Bittner [2006. Non-linear wave-wave interaction: case studies based on rocket data and first application to satellite data. Journal of Atmospheric and Solar-Terrestrial Physics 68, 959-976]--are investigated to look for gravity wave reflection processes. Following nonlinear theory, energy dissipation rates according to Weinstock [1980. Energy dissipation rates of turbulence in the stable free atmosphere. Journal of the Atmospheric Sciences 38, 880-883] are calculated from foil chaff cloud and falling sphere data and compared with the critical layer heights. Enhanced energy dissipation rates are found at those altitudes where the waves' phase speed matches the zonal background wind speeds. Indication of gravity wave trapping is found between two altitudes of around 95 and 86 km.

  6. Ablation study of tungsten-based nuclear thermal rocket fuel

    NASA Astrophysics Data System (ADS)

    Smith, Tabitha Elizabeth Rose

    The research described in this thesis has been performed in order to support the materials research and development efforts of NASA Marshall Space Flight Center (MSFC), of Tungsten-based Nuclear Thermal Rocket (NTR) fuel. The NTR was developed to a point of flight readiness nearly six decades ago and has been undergoing gradual modification and upgrading since then. Due to the simplicity in design of the NTR, and also in the modernization of the materials fabrication processes of nuclear fuel since the 1960's, the fuel of the NTR has been upgraded continuously. Tungsten-based fuel is of great interest to the NTR community, seeking to determine its advantages over the Carbide-based fuel of the previous NTR programs. The materials development and fabrication process contains failure testing, which is currently being conducted at MSFC in the form of heating the material externally and internally to replicate operation within the nuclear reactor of the NTR, such as with hot gas and RF coils. In order to expand on these efforts, experiments and computational studies of Tungsten and a Tungsten Zirconium Oxide sample provided by NASA have been conducted for this dissertation within a plasma arc-jet, meant to induce ablation on the material. Mathematical analysis was also conducted, for purposes of verifying experiments and making predictions. The computational method utilizes Anisimov's kinetic method of plasma ablation, including a thermal conduction parameter from the Chapman Enskog expansion of the Maxwell Boltzmann equations, and has been modified to include a tangential velocity component. Experimental data matches that of the computational data, in which plasma ablation at an angle shows nearly half the ablation of plasma ablation at no angle. Fuel failure analysis of two NASA samples post-testing was conducted, and suggestions have been made for future materials fabrication processes. These studies, including the computational kinetic model at an angle and the

  7. Experimental investigation of paraffin-based fuels for hybrid rocket propulsion

    NASA Astrophysics Data System (ADS)

    Galfetti, L.; Merotto, L.; Boiocchi, M.; Maggi, F.; DeLuca, L. T.

    2013-03-01

    Solid fuels for hybrid rockets were characterized in the framework of a research project aimed to develop a new generation of solid fuels, combining at the same time good mechanical and ballistic properties. Original techniques were implemented in order to improve paraffin-based fuels. The first strengthening technique involves the use of a polyurethane foam (PUF); a second technique is based on thermoplastic polymers mixed at molecular level with the paraffin binder. A ballistic characterization of paraffin-based hybrid rocket solid fuels was performed, considering pure wax-based fuels and fuels doped with suitable metal additives. Nano-Al powders and metal hydrides (magnesium hydride (MgH2), lithium aluminum hydride (LiAlH4 )) were used as fillers in paraffin matrices. The results of this investigation show a strong correlation between the measured viscosity of the melted paraffin layer and the regression rate: a decrease of viscosity increases the regression rate. This trend is due to the increasing development of entrainment phenomena, which strongly increase the regression rate. Addition of LiAlH4 (mass fraction 10%) can further increase the regression rate up to 378% with respect to the pure HTPB regression rate, taken as baseline reference fuel. The highest regression rates were found for the Solid Wax (SW) composition, added with 5% MgH2 mass fraction; at 350 kg/(m2s) oxygen mass flux, the measured regression rate, averaged in space and time, was 2.5 mm/s, which is approximately five times higher than that of the pure HTPB composition. Compositions added with nanosized aluminum powders were compared with those added with MgH2, using gel or solid wax.

  8. An autonomous receiver/digital signal processor applied to ground-based and rocket-borne wave experiments

    NASA Astrophysics Data System (ADS)

    Dombrowski, M. P.; LaBelle, J.; McGaw, D. G.; Broughton, M. C.

    2016-07-01

    The programmable combined receiver/digital signal processor platform presented in this article is designed for digital downsampling and processing of general waveform inputs with a 66 MHz initial sampling rate and multi-input synchronized sampling. Systems based on this platform are capable of fully autonomous low-power operation, can be programmed to preprocess and filter the data for preselection and reduction, and may output to a diverse array of transmission or telemetry media. We describe three versions of this system, one for deployment on sounding rockets and two for ground-based applications. The rocket system was flown on the Correlation of High-Frequency and Auroral Roar Measurements (CHARM)-II mission launched from Poker Flat Research Range, Alaska, in 2010. It measured auroral "roar" signals at 2.60 MHz. The ground-based systems have been deployed at Sondrestrom, Greenland, and South Pole Station, Antarctica. The Greenland system synchronously samples signals from three spaced antennas providing direction finding of 0-5 MHz waves. It has successfully measured auroral signals and man-made broadcast signals. The South Pole system synchronously samples signals from two crossed antennas, providing polarization information. It has successfully measured the polarization of auroral kilometric radiation-like signals as well as auroral hiss. Further systems are in development for future rocket missions and for installation in Antarctic Automatic Geophysical Observatories.

  9. Towards Understanding the Fluid Dynamic Phenomenon of Interest to Rocket Base Heating: A Review

    NASA Technical Reports Server (NTRS)

    Venkatapathy, E.; Park, C.; Palmer, G.; Arnold, James O. (Technical Monitor)

    1994-01-01

    The significance of the base heating problem for rockets during ascent is due to the complex interaction between the rocket nozzle plumes and the external-flow which can change the flow characteristics in the base region dramatically. At lower altitudes the external-flow merges with the plume-flow, without the formation of a large separated flow region, and the cooler external-flow promotes convective cooling of the base wall. Under these conditions the majority of the base heating is due to radiative heating from the shock heated plume gases. At higher altitudes, however, the process of base heating is not so straightforward. The plume and the base flow expands dramatically and separated flow regions occur in the base area. Hot exhaust gases from the rocket nozzle will be entrained into the separated flow regions and produce a convective component to the base wall heating. Further, if the rocket exhaust-gas contains soot, the soot can increase the emission from the gas and dramatically increase the wall absorption coefficient for radiative heating if it is deposited on the walls . In addition, if the rocket exhaust gas is fuel rich, the fuel can bum in the separated flow regions and further increase the base heating. The base burning phenomenon, and the increased base heating caused by it at higher altitudes, have been observed for the Space Shuttle and Saturn Rocket. Under these conditions, the total heating is significantly higher than the heating without separated flow in the base region, and the increase in heating is directly attributable to the fluid dynamic complexity of the base region. Realistic simulation of the base heating requires that the calculated flow environment reproduce the fluid dynamic flow features accurately. Thus, it will be necessary to introduce into the CFD codes the capability for the flow to respond to the complex vehicle geometry, the effect of turbulence, the ability to accurately reproduce the plume shock/shear layer structures and

  10. Focused Rocket-Ejector RBCC Experiments

    NASA Technical Reports Server (NTRS)

    Santoro, Robert J.; Pal, Sibtosh

    2003-01-01

    This document reports the results of additional efforts for the Rocket Based Combined Cycle (RBCC) rocket-ejector mode research work carried out at the Perm State Propulsion Engineering Research Center in support of NASA s technology development efforts for enabling 3rd generation Reusable Launch Vehicles (RLV). The two tasks conducted under this program build on earlier NASA MSFC funded research program on rocket ejector investigations. The first task continued a systematic investigation of the improvements provided by a gaseous hydrogen (GHz)/oxygen (GO2) twin thruster RBCC rocket ejector system over a single rocket system. In a similar vein, the second task continued investigations into the performance of a hydrocarbon (liquid JP-7)/gaseous oxygen single thruster rocket-ejector system. To gain a systematic understanding of the rocket-ejector s internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static diffusion and afterburning (DAB) configurations for a range of rocket operating conditions. For all experimental conditions, overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust. For the GH2/GO2 propellant rocket ejector experiments, high frequency measurements of the pressure field within the system were also made to understand the unsteady behavior of the flowfield.

  11. Baking Soda and Vinegar Rockets

    NASA Astrophysics Data System (ADS)

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

    2009-02-01

    Rocket experiments demonstrating conservation of momentum will never fail to generate enthusiasm in undergraduate physics laboratories. In this paper, we describe tests on rockets from two vendors1,2 that combine baking soda and vinegar for propulsion. The experiment compared two analytical approximations for the maximum rocket height to the experimentally measured rocket height. Baking soda and vinegar rockets present fewer safety concerns and require a smaller launch area than rapid combustion chemical rockets. Both kits were of nearly identical design, costing ˜20. The rockets required roughly 30 minutes of assembly time consisting of mostly taping the soft plastic fuselage to the Styrofoam nose cone.

  12. Report on Rocket Power Plants Based on T-Substance

    NASA Technical Reports Server (NTRS)

    Walter, Hellmuth

    1947-01-01

    In the search for an energy source independent of air for the propulsion of underwater craft, attention was early concentrated on T-substance. It was possible to convince the OKM [NACA comment: Navy High Command] very quickly of the importance of this material. In 1934, the first experiments were undertaken. A difficulty was at once presented by the limited concentration that had been attained. At first only 60 percent T-substance could be supplied; this amount was later increased to as much as 85 percent. Decomposition and combustion experiments conducted on the grounds of the CPVA in Kiel-Dietrichsdorf led to the first practical information as to the technical feasibility of the use of T-substance. New perspectives soon developed because a method of concentrated energy production had been found here, which was capable of many applications. The idea of using this energy for the propulsion of missiles either in guns or as rockets suggested itself and appropriate proposals, which quickly led to the construction of the first experimental devices, were made to the official quarters concerned. In January 1937, the first flight of a DVL aircraft with T-substance auxiliary propulsion took place at Alimbsmuhle in the presence of Colonel Udet, who piloted the third flight. In June 1937, the first T-substance rockets were fired (Altenwalde). Then in rapid succession take-off auxiliary, main propulsion, and other rocket drives were brought out in experimental versions. Hydrogen peroxide is a well known chemical, which is widely used in the textile industry. Its chemical and physical properties as well as the processes of manufacture and use are familiar and have been described in voluminous books and papers, Nevertheless, much developmental work was required to open the way for T-substance as a usable oxygen carrier. In fact the utilization of hydrogen peroxide as the oxygen carrier for energy production had hitherto been the subject only of isolated suggestions, which have

  13. Integrated approach for hybrid rocket technology development

    NASA Astrophysics Data System (ADS)

    Barato, Francesco; Bellomo, Nicolas; Pavarin, Daniele

    2016-11-01

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

  14. Mesospheric minor-species determinations from rocket and ground-based i. r. measurements

    SciTech Connect

    Ulwick, J.C.; Baker, K.D.; Baker, D.J.; Steed, A.J.; Pendleton, W.R.

    1987-01-01

    A wealth of ground-based, balloon and rocket measurements of the airglow hydroxyl (OH) emission was accumulated by various experiments. Indeed, since the first-reported rocket observations (HEPPNER and MEREDITH, 1958), there have been over 50 rocket flights from which OH emission profiles have been reported (BAKER and STAIR, 1987). The sensors initially measured in the visible range (PACKER, 1961) and later in the infrared (LOWE, 1960; BAKER, 1978). Models were constructed of the photochemistry (MOREELS et al., 1977; BATTANER and LOPEZ-MORENO, 1979) and of the vertical-species transport in an attempt to fit the measured hydroxyl radiance variations. However, uncertainties exit in quenching-rate coefficients, radiation-transition probabilities and number densities of participating species, such that results are not conclusive (COXON and FOSTER, 1982; DUPUIS et al., 1986). As part of the overall investigation of neutral atmospheric dynamics of the middle atmosphere and to investigate the excitation processes of the night airglow, a coordinated rocket and ground-based measurement program was conducted as part of the MAP/WINE campaign to observe simultaneously several key parameters.

  15. Combining Undergraduate Student Curriculum, Research, and Outreach: High-altitude Balloon and Rockets

    NASA Astrophysics Data System (ADS)

    Davis, E. J.; Nielsen, K.

    2015-12-01

    The Society of Physics Students chapter at Utah Valley University (UVU) recently established a high altitude balloon project to provide students with research opportunities. This highly successful program involves students not only from physics but also from other STEM fields and non-STEM subjects, and as such acts as a unique outreach program for the department of physics. Examples of experiments performed with the balloon project are: 3D-acceleration measurements, altitude/pressure/temperature measurements, ozone monitoring, bio-aerosol collection, and solar panel performance output. All these experiment are designed and build by groups of students either as part of research projects or through class participation as the projects link with the curriculum in several courses. Most recently, a group of UVU students have initiated the implementation of small rockets capable of carrying payloads to this high-altitude program. Both balloon and rocket platforms are fundamental in-situ measuring techniques for numerous geoscience subjects, and are arguably best illustrated by the NASA balloon and sounding rocket programs. In this presentation, we give an overview of the program and how it is 1) being implemented into the curriculum, 2) provide unique research opportunities for students, and 3) specific outreach activities.

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

    NASA Technical Reports Server (NTRS)

    Santoro, Robert J.; Pal, Sibtosh

    2003-01-01

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

  17. Experimental analysis of SiC-based refractory concrete in hybrid rocket nozzles

    NASA Astrophysics Data System (ADS)

    D'Elia, Raffaele; Bernhart, Gérard; Hijlkema, Jouke; Cutard, Thierry

    2016-09-01

    Hybrid propulsion represents a good alternative to the more widely used liquid and solid systems. This technology combines some important specifications of the latters, as the possibility of re-ignition, thrust modulation, a higher specific impulse than solid systems, a greater simplicity and a lower cost than liquid systems. Nevertheless the highly oxidizing environment represents a major problem as regards the thermo-oxidation and ablative behavior of nozzle materials. The main goal of this research is to characterize a silicon carbide based micro-concrete with a maximum aggregates size of 800 μm, in a hybrid propulsion environment. The nozzle throat has to resist to a highly oxidizing polyethylene/nitrous oxide hybrid environment, under temperatures up to 2900 K. Three tests were performed on concrete-based nozzles in HERA Hybrid Rocket Motor (HRM) test bench at ONERA. Pressure chamber evolution and observations before and after tests are used to investigate the ablated surface at nozzle throat. Ablation behavior and crack generation are discussed and some improvements are proposed.

  18. Plasma filamentation and shock wave enhancement in microwave rockets by combining low-frequency microwaves with external magnetic field

    NASA Astrophysics Data System (ADS)

    Takahashi, Masayuki; Ohnishi, Naofumi

    2016-08-01

    A filamentary plasma is reproduced based on a fully kinetic model of electron and ion transports coupled with electromagnetic wave propagation. The discharge plasma transits from discrete to diffusive patterns at a 110-GHz breakdown, with decrease in the ambient pressure, because of the rapid electron diffusion that occurs during an increase in the propagation speed of the ionization front. A discrete plasma is obtained at low pressures when a low-frequency microwave is irradiated because the ionization process becomes more dominant than the electron diffusion, when the electrons are effectively heated by the low-frequency microwave. The propagation speed of the plasma increases with decrease in the incident microwave frequency because of the higher ionization frequency and faster plasma diffusion resulting from the increase in the energy-absorption rate. An external magnetic field is applied to the breakdown volume, which induces plasma filamentation at lower pressures because the electron diffusion is suppressed by the magnetic field. The thrust performance of a microwave rocket is improved by the magnetic fields corresponding to the electron cyclotron resonance (ECR) and its higher-harmonic heating, because slower propagation of the ionization front and larger energy-absorption rates are obtained at lower pressures. It would be advantageous if the fundamental mode of ECR heating is coupled with a lower frequency microwave instead of combining the higher-harmonic ECR heating with the higher frequency microwave. This can improve the thrust performance with smaller magnetic fields even if the propagation speed increases because of the decrease in the incident microwave frequency.

  19. Air-breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1999-01-01

    This Quick Time movie depicts the Rocketdyne static test of an air-breathing rocket. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's advanced Transportation Program at the Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  20. Formulation and Testing of Paraffin-Based Solid Fuels Containing Energetic Additives for Hybrid Rockets

    NASA Technical Reports Server (NTRS)

    Larson, Daniel B.; Boyer, Eric; Wachs,Trevor; Kuo, Kenneth K.; Story, George

    2012-01-01

    Many approaches have been considered in an effort to improve the regression rate of solid fuels for hybrid rocket applications. One promising method is to use a fuel with a fast burning rate such as paraffin wax; however, additional performance increases to the fuel regression rate are necessary to make the fuel a viable candidate to replace current launch propulsion systems. The addition of energetic and/or nano-sized particles is one way to increase mass-burning rates of the solid fuels and increase the overall performance of the hybrid rocket motor.1,2 Several paraffin-based fuel grains with various energetic additives (e.g., lithium aluminum hydride (LiAlH4) have been cast in an attempt to improve regression rates. There are two major advantages to introducing LiAlH4 additive into the solid fuel matrix: 1) the increased characteristic velocity, 2) decreased dependency of Isp on oxidizer-to-fuel ratio. The testing and characterization of these solid-fuel grains have shown that continued work is necessary to eliminate unburned/unreacted fuel in downstream sections of the test apparatus.3 Changes to the fuel matrix include higher melting point wax and smaller energetic additive particles. The reduction in particle size through various methods can result in more homogeneous grain structure. The higher melting point wax can serve to reduce the melt-layer thickness, allowing the LiAlH4 particles to react closer to the burning surface, thus increasing the heat feedback rate and fuel regression rate. In addition to the formulation of LiAlH4 and paraffin wax solid-fuel grains, liquid additives of triethylaluminum and diisobutylaluminum hydride will be included in this study. Another promising fuel formulation consideration is to incorporate a small percentage of RDX as an additive to paraffin. A novel casting technique will be used by dissolving RDX in a solvent to crystallize the energetic additive. After dissolving the RDX in a solvent chosen for its compatibility

  1. Development and application of a reverse Monte Carlo radiative transfer code for rocket plume base heating

    NASA Technical Reports Server (NTRS)

    Everson, John; Nelson, H. F.

    1993-01-01

    A reverse Monte Carlo radiative transfer code to predict rocket plume base heating is presented. In this technique rays representing the radiation propagation are traced backwards in time from the receiving surface to the point of emission in the plume. This increases the computational efficiency relative to the forward Monte Carlo technique when calculating the radiation reaching a specific point, as only the rays that strike the receiving point are considered.

  2. A Comparison of Auroral In-Situ Rocket Electron Measurements and Ground-Based Multi-spectral EMCCD Imaging

    NASA Astrophysics Data System (ADS)

    Grubbs, G. A., II; Samara, M.; Michell, R.; Hampton, D.; Hecht, J. H.

    2015-12-01

    The Ground-to-Rocket Electrodynamics-Electrons Correlative Experiment (GREECE) mission successfully launched from Poker Flat, Alaska on 03 March 2014 at 11:09:50 UT and reached an apogee of approximately 335 km during a luminous auroral event. Multiple ground-based electron-multiplying charge-coupled device (EMCCD) imagers were positioned at Venetie, Alaska and aimed along magnetic zenith in order to observe the brightness of different auroral emission lines (427.8, 557.7, and 844.6 nm with a 47 degree field of view) at the magnetic footpoint of the payload, near apogee. Emission line brightness data are presented at the footpoint of the rocket flight and compared with electron characteristics taken by the Acute Precipitating Electron Spectrometer (APES) on-board instrument. Ratios of different auroral emission lines are combined with previously published models in order to estimate the characteristic energy of the incident electron population, which is directly compared to the APES data for validation. Our goal is to describe the auroral emissions produced from a known precipitating electron distribution, such that we can more accurately use ground-based imaging and photometry to infer the characteristics of the precipitating electrons. These techniques can then be applied over larger scales and longer times, when only multi-spectral imaging data are available with no corresponding in situ data.

  3. Baking Soda and Vinegar Rockets

    ERIC Educational Resources Information Center

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

    2009-01-01

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

  4. Development of an IgY-based rocket-immunoelectrophoresis for identity monitoring of Pertussis vaccines.

    PubMed

    Matheis, Walter; Schade, Rüdiger

    2011-06-30

    An important step in vaccine production and quality control is the analysis of identity of different lots. For that purpose chicken was immunized with acellular Pertussis components (Pertussis toxoid, Filamenteous haemagglutinin, Pertactin, Fimbriae 2/3 antigen). The resulting antibodies (IgY) were non-invasive extracted from egg yolk and used for rocket immunoelectrophoresis (RIE). We demonstrated that the Ab reacted with characteristic peaks ("rockets") with the corresponding antigen. The shape of the peaks varied depending on the manufacturer and the nature of antigen (adsorbed or non-adsorbed). The coefficients of variation was about 20% during a year period. In summary, our data illustrate that an IgY-based RIE is not only a cost-effective method but also proficient for monitoring Pertussis vaccines.

  5. Satellite-rocket docking ring recognition method based on mathematical morphology

    NASA Astrophysics Data System (ADS)

    Xu, Zhiqiang; Shang, Yang; Ma, Xuan

    2015-10-01

    Satellite-rocket docking ring recognition method based on mathematical morphology is presented in this paper, according to the geometric and grayscale characteristics of the docking ring typical structure. The docking ring used in this paper is a circle with a cross in the middle. Most of spacecrafts are transported into orbit by rocket, and they retain the connection component with the rocket. The tracing spacecraft should capture the target spacecraft first before operating the target spacecraft. The docking ring is one of the typical parts of a spacecraft, and it can be recognized automatically. Thereby we can capture the spacecraft through the information of the docking ring. Firstly a multi-step mathematical morphology processing is applied to the image of the target spacecraft with different structure element, followed by edge detection and line detection, and finally docking ring typical structure is located in the image by relative geometry analysis. The images used in this paper are taken of real satellite in lab. The docking ring can be recognized when the distance between the two spacecraft is different. The results of physical simulation experiment show that the method in this paper can recognize docking ring typical structure accurately when the tracing spacecraft is approaching the target spacecraft.

  6. Paraffin-based hybrid rocket engines applications: A review and a market perspective

    NASA Astrophysics Data System (ADS)

    Mazzetti, Alessandro; Merotto, Laura; Pinarello, Giordano

    2016-09-01

    Hybrid propulsion technology for aerospace applications has received growing attention in recent years due to its important advantages over competitive solutions. Hybrid rocket engines have a great potential for several aeronautics and aerospace applications because of their safety, reliability, low cost and high performance. As a consequence, this propulsion technology is feasible for a number of innovative missions, including space tourism. On the other hand, hybrid rocket propulsion's main drawback, i.e. the difficulty in reaching high regression rate values using standard fuels, has so far limited the maturity level of this technology. The complex physico-chemical processes involved in hybrid rocket engines combustion are of major importance for engine performance prediction and control. Therefore, further investigation is ongoing in order to achieve a more complete understanding of such phenomena. It is well known that one of the most promising solutions for overcoming hybrid rocket engines performance limits is the use of liquefying fuels. Such fuels can lead to notably increased solid fuel regression rate due to the so-called "entrainment phenomenon". Among liquefying fuels, paraffin-based formulations have great potentials as solid fuels due to their low cost, availability (as they can be derived from industrial waste), low environmental impact and high performance. Despite the vast amount of literature available on this subject, a precise focus on market potential of paraffins for hybrid propulsion aerospace applications is lacking. In this work a review of hybrid rocket engines state of the art was performed, together with a detailed analysis of the possible applications of such a technology. A market study was carried out in order to define the near-future foreseeable development needs for hybrid technology application to the aforementioned missions. Paraffin-based fuels are taken into account as the most promising segment for market development

  7. 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.

  8. 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

  9. Torpedo Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

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

  10. 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)

  11. 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.

  12. Focused Experimental and Analytical Studies of the RBCC Rocket-Ejector

    NASA Technical Reports Server (NTRS)

    Lehman, M.; Pal, S.; Schwes, D.; Chen, J. D.; Santoro, R. J.

    1999-01-01

    The rocket-ejector mode of a Rocket Based Combined Cycle Engine (RBCC) was studied through a joint experimental/analytical approach. A two-dimensional variable geometry rocket-ejector system with enhanced optical access was designed and fabricated for experimentation. The rocket-ejector system utilizes a single two-dimensional gaseous oxygen/gaseous hydrogen rocket as the ejector. To gain a systematic understanding of the rocket ejector's internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static configurations for a range of rocket operating conditions Overall system performance was obtained through Global measurements of wall static pressure profiles, heat flux profiles and engine thrust, whereas detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (gaseous oxygen, hydrogen. nitrogen and water vapor). These experimental efforts were complemented by Computational Fluid Dynamic (CFD) flowfield analyses.

  13. Experiment/Analytical Characterization of the RBCC Rocket-Ejector Mode

    NASA Technical Reports Server (NTRS)

    Ruf, J. H.; Lehman, M.; Pal, S.; Santoro, R. J.; West, J.; Turner, James E. (Technical Monitor)

    2000-01-01

    Experimental and complementary CFD results from the study of the rocket-ejector mode of a Rocket Based Combined Cycle (RBCC) engine are presented and discussed. The experiments involved systematic flowfield measurements in a two-dimensional, variable geometry rocket-ejector system. The rocket-ejector system utilizes a single two-dimensional, gaseous oxygen/gaseous hydrogen rocket as the ejector. To gain a thorough understanding of the rocket-ejector's internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static configurations for a range of rocket operating conditions. Overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust, whereas detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (oxygen, hydrogen, nitrogen and water vapor). The experimental results for both the direct-connect and sea-level static configurations are compared with CFD predictions of the flowfield.

  14. From the Peenemuende 'Aggregates' to the American moon rocket - The development of the Saturn V Apollo rocket by Werner von Braun's team based on the Peenemuende concept

    NASA Astrophysics Data System (ADS)

    Reisig, G. H. R.

    The origin of active space flight and the important role played in it by the innovations made in the Peenemuende rocket program are discussed. The rocket development carried out by the Werner von Braun team is chronologically recalled. The A4 rocket and its aerodynamic configuration are discussed, including the ballistic demands on its configuration, the development of the Peenemuende supersonic wind tunnel, and the pretesting of the A4 configuration using the A5 rocket. Rocket development in the U.S. from the Redstone to the Saturn is reviewed.

  15. The Solid Rocket Motor Slag Population: Results of a Radar-Based Regressive Statistical Evaluation

    NASA Astrophysics Data System (ADS)

    Horstman, Matthew

    Solid rocket motor (SRM) slag has been identified as a potential source of man-made orbital debris. The possibility that SRMs (in addition to generating dust particles in the sub-millimeter range) may generate particles up to centimeters in size has caused concern regarding their contribution to the debris environment. Returned surfaces from space do not have sufficient area or exposure time to provide a clear picture of the SRM millimeter and centimeter debris population. Currently, radar observation is probably the only way to collect data showing the debris contribution from SRMs. Such observation is used to sample the debris environment, but it is difficult to obtain accurate orbital elements for the detected debris objects. NASA has developed several models to describe the different orbital debris populations, based on assumed debris production mechanisms to create clouds of debris objects that can be propagated in time. The NASA model, LEGEND (LEO-to-GEO Environment Debris), functions as a timetested debris model for most debris sources. However, the current LEGEND model does not include contributions from the SRM population. An SRM model has recently been developed by NASA, based on purely theoretical details of SRM production and known SRM launches, but verification with hard data is needed. Because the detections of individual SRM objects cannot be deterministically separated from the total debris observed by radar, the validation of the SRM model can only be done by combining it with the LEGEND breakup model and comparing it with data. By applying observational constraints, the degree of SRM slag contribution to the environment may be estimated. This serves as an observationally sound method from which to calibrate a purely theoretical model into something more realistic. For this study, we use the populations observed by the Haystack radar from 1996 to present. For the SRM debris, we use a historical database of SRM launches, propellant masses, and

  16. The Solid Rocket Motor Slag Population: Results of a Radar-Based Regressive Statistical Evaluation

    NASA Technical Reports Server (NTRS)

    Horstman, Matthew F.; Xu, Yu-Lin

    2008-01-01

    Solid rocket motor (SRM) slag has been identified as a potential source of man-made orbital debris. The possibility that SRMs (in addition to generating dust particles in the sub-millimeter range) may generate particles up to centimeters in size has caused concern regarding their contribution to the debris environment. Returned surfaces from space do not have sufficient area or exposure time to provide a clear picture of the SRM millimeter and centimeter debris population. Currently, radar observation is probably the only way to collect data showing the debris contribution from SRMs. Such observation is used to sample the debris environment, but it is difficult to obtain accurate orbital elements for the detected debris objects. NASA has developed several models to describe the different orbital debris populations, based on assumed debris production mechanisms to create clouds of debris objects that can be propagated in time. The NASA model, LEGEND (LEO-to-GEO Environment Debris), functions as a time-tested debris model for most debris sources. However, the current LEGEND model does not include contributions from the SRM population. An SRM model has recently been developed by NASA, based on purely theoretical details of SRM production and known SRM launches, but verification with hard data is needed. Because the detections of individual SRM objects cannot be deterministically separated from the total debris observed by radar, the validation of the SRM model can only be done by combining it with the LEGEND breakup model and comparing it with data. By applying observational constraints, the degree of SRM slag contribution to the environment may be estimated. This serves as an observationally sound method from which to calibrate a purely theoretical model into something more realistic. For this study, we use the populations observed by the Haystack radar from 1996 to present. For the SRM debris, we use a historical database of SRM launches, propellant masses, and

  17. Robust Rocket Engine Concept

    NASA Technical Reports Server (NTRS)

    Lorenzo, Carl F.

    1995-01-01

    The potential for a revolutionary step in the durability of reusable rocket engines is made possible by the combination of several emerging technologies. The recent creation and analytical demonstration of life extending (or damage mitigating) control technology enables rapid rocket engine transients with minimum fatigue and creep damage. This technology has been further enhanced by the formulation of very simple but conservative continuum damage models. These new ideas when combined with recent advances in multidisciplinary optimization provide the potential for a large (revolutionary) step in reusable rocket engine durability. This concept has been named the robust rocket engine concept (RREC) and is the basic contribution of this paper. The concept also includes consideration of design innovations to minimize critical point damage.

  18. Air-Breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    2000-01-01

    This photograph depicts an air-breathing rocket engine that completed an hour or 3,600 seconds of testing at the General Applied Sciences Laboratory in Ronkonkoma, New York. Referred to as ARGO by its design team, the engine is named after the mythological Greek ship that bore Jason and the Argonauts on their epic voyage of discovery. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced SpaceTransportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

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

    NASA Technical Reports Server (NTRS)

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

    2000-01-01

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

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

    NASA Astrophysics Data System (ADS)

    Karabeyoglu, Arif; Tuncer, Onur; Inalhan, Gokhan

    2016-07-01

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

  1. A US History of Airbreathing/Rocket Combined-Cycle (RBCC) Propulsion for Powering Future Aerospace Transports, with a Look Ahead to the Year 2020

    NASA Technical Reports Server (NTRS)

    Escher, William J. D.

    1999-01-01

    A technohistorical and forward-planning overview of U.S. developments in combined airbreathing/rocket propulsion for advanced aerospace vehicle applications is presented. Such system approaches fall into one of two categories: (1) Combination propulsion systems (separate, non-interacting engines installed), and (2) Combined-Cycle systems. The latter, and main subject, comprises a large family of closely integrated engine types, made up of both airbreathing and rocket derived subsystem hardware. A single vehicle-integrated, multimode engine results, one capable of operating efficiently over a very wide speed and altitude range, atmospherically and in space. While numerous combination propulsion systems have reached operational flight service, combined-cycle propulsion development, initiated ca. 1960, remains at the subscale ground-test engine level of development. However, going beyond combination systems, combined-cycle propulsion potentially offers a compelling set of new and unique capabilities. These capabilities are seen as enabling ones for the evolution of Spaceliner class aerospace transportation systems. The following combined-cycle hypersonic engine developments are reviewed: (1) RENE (rocket engine nozzle ejector), (2) Cryojet and LACE, (3) Ejector Ramjet and its derivatives, (4) the seminal NASA NAS7-377 study, (5) Air Force/Marquardt Hypersonic Ramjet, (6) Air Force/Lockheed-Marquardt Incremental Scramjet flight-test project, (7) NASA/Garrett Hypersonic Research Engine (HRE), (8) National Aero-Space Plane (NASP), (9) all past projects; and such current and planned efforts as (10) the NASA ASTP-ART RBCC project, (11) joint CIAM/NASA DNSCRAM flight test,(12) Hyper-X, (13) Trailblazer,( 14) W-Vehicle and (15) Spaceliner 100. Forward planning programmatic incentives, and the estimated timing for an operational Spaceliner powered by combined-cycle engines are discussed.

  2. Investigation of Low-Reynolds-Number Rocket Nozzle Design Using PNS-Based Optimization Procedure

    NASA Technical Reports Server (NTRS)

    Hussaini, M. Moin; Korte, John J.

    1996-01-01

    An optimization approach to rocket nozzle design, based on computational fluid dynamics (CFD) methodology, is investigated for low-Reynolds-number cases. This study is undertaken to determine the benefits of this approach over those of classical design processes such as Rao's method. A CFD-based optimization procedure, using the parabolized Navier-Stokes (PNS) equations, is used to design conical and contoured axisymmetric nozzles. The advantage of this procedure is that it accounts for viscosity during the design process; other processes make an approximated boundary-layer correction after an inviscid design is created. Results showed significant improvement in the nozzle thrust coefficient over that of the baseline case; however, the unusual nozzle design necessitates further investigation of the accuracy of the PNS equations for modeling expanding flows with thick laminar boundary layers.

  3. 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.

  4. Comparison of vibrations of a combination of solid-rocket launch vehicle and payload during a ground firing and launching

    NASA Technical Reports Server (NTRS)

    Schoenster, J. A.; Pierce, H. B.

    1975-01-01

    The results of a study into the environmental vibrations of a payload mounted on the Nike rocket launch vehicle were presented. Data were obtained during the flight acceptance test of the payload, the firing of the total vehicle in a special test stand, and the powered and unpowered flights of the vehicle. The vibrational response of the structure was measured. Data were also obtained on the fluctuating pressure on the outside surface of the vehicle and inside the forward and after ends of the rocket chamber. A comparison of the data from the three test conditions indicated that external pressure fluctuations were the major source of vibrations in the payload area, and pressure fluctuations within the rocket motor were the major source of vibrations contiguous to the payload area.

  5. Rocket Power Plants Based on Nitric Acid and their Specific Propulsive Weights

    NASA Technical Reports Server (NTRS)

    Zborowski, Helmut

    1947-01-01

    Two fields are reserved for the application of rocket power plants. The first field is determined by the fact that the rocket power plant is the only type of power plant that can produce thrust without dependence upon environment. For this field,the rocket is therefore the only possible power plant and the limit of what may be done is determined by the status of the technical development of these power plants at the given moment. The second field is that in which the rocket power plant proves itself the most suitable as a high-power drive in free competition with other types of power plants. The exposition will be devoted to the demarcation of this field and its division among the various types of rocket power plants.

  6. Performance of a RBCC Engine in Rocket-Operation

    NASA Astrophysics Data System (ADS)

    Tomioka, Sadatake; Kubo, Takahiro; Noboru Sakuranaka; Tani, Koichiro

    Combination of a scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines (the combined system termed as Rocket-based Combined Cycle engine) are expected to be the most effective propulsion system for space launch vehicles. Either SSTO (Single Stage To Orbit) system or TSTO (Two Stage To Orbit) system with separation at high altitude needs final stage acceleration in space, so that the RBCC (Rocket Based Combined Cycle) engine should be operated as rocket engines. Performance of the scramjet combustor as the extension to the rocket nozzle, was experimentally evaluated by injecting inert gas at various pressure through the embedded rocket chamber while the whole sub-scaled model was placed in a low pressure chamber connected to an air-driven ejector system. The results showed that the thrust coefficient was about 1.2, the low value being found to mainly due to the friction force on the scramjet combustor wall, while blocking the scramjet flow pass’s opening to increase nozzle extension thrust surface, was found to have little effects on the thrust performance. The combustor was shortened to reduce the friction loss, however, degree of reduction was limited as friction decreased rapidly with distance from the onset of the scramjet combustor.

  7. Atomic-Based-Combined-Cycle Analysis

    NASA Technical Reports Server (NTRS)

    Han, Samuel S.

    1999-01-01

    Atomic-based-combined-cycle (ABCC) engine combines an air-breathing ramjet engine with an atomic reactor to increase the mission-averaged specific impulse and thereby increasing the dry-mass ratio. ABCC engine is similar to RBCC engine except that energy needed for the propulsive power is derived from nuclear reaction rather than chemical combustion used in the RBCC engine. The potential performance improvement of an ABCC engine over a RBCC engine comes from two factors. Firstly, the energy density of nuclear reaction is several order of magnitudes higher than the chemical combustion. Secondly, hydrogen can produce much higher nozzle exit velocity because of its small molecular weight. A one-dimensional, transient numerical model was used to analyze a generic RBCC engine and it is used as a baseline to evaluate an imaginary ABCC engine performance. A nuclear reactor is treated as a black box energy source that replaces the role of the primary rocket and the chemical combustion chamber in a RBCC engine. The performance of a generic ABCC engine along a flight path (q0 =10 (exp 3) lbf per square ft) shows that the mission averaged-specific impulse is about twice larger than RBCC engine and the dry mass-ratio is about 50% larger. Results of the present ABCC engine performance are based on the assumptions that the flow passage of working fluids is identical to that of RBCC engine and that a nuclear reactor is treated as an energy black box. Preliminary heat transfer calculation shows that the rate of heat transfer to the working fluids is within the limit of turbulent convective heat transfer regimes. The flow passage of realistic ABCC engine must be known for a better prediction of ABCC engine performance. Also, critical heat transfer calculations must be performed for the ejector mode and ramjet mode operations. This is possible only when the details of a reactor configuration are available.

  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. Outbrief - Long Life Rocket Engine Panel

    NASA Technical Reports Server (NTRS)

    Quinn, Jason Eugene

    2004-01-01

    This white paper is an overview of the JANNAF Long Life Rocket Engine (LLRE) Panel results from the last several years of activity. The LLRE Panel has met over the last several years in order to develop an approach for the development of long life rocket engines. Membership for this panel was drawn from a diverse set of the groups currently working on rocket engines (Le. government labs, both large and small companies and university members). The LLRE Panel was formed in order to determine the best way to enable the design of rocket engine systems that have life capability greater than 500 cycles while meeting or exceeding current performance levels (Specific Impulse and Thrust/Weight) with a 1/1,OOO,OOO likelihood of vehicle loss due to rocket system failure. After several meetings and much independent work the panel reached a consensus opinion that the primary issues preventing LLRE are a lack of: physics based life prediction, combined loads prediction, understanding of material microphysics, cost effective system level testing. and the inclusion of fabrication process effects into physics based models. With the expected level of funding devoted to LLRE development, the panel recommended that fundamental research efforts focused on these five areas be emphasized.

  10. 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.

  11. Atomic-Based-Combined-Cycle Analysis

    NASA Technical Reports Server (NTRS)

    Han, Sam; Bai, Don; Schmidt, George

    2000-01-01

    Atomic-based-combined-cycle (ABCC) engine combines an air-breathing ramjet engine with an atomic reactor to increase the mission-averaged specific impulse and thereby increasing the dry-mass ratio. ABCC engine is similar to RBCC engine except that energy needed for the propulsive power is derived from nuclear reaction rather than chemical combustion used in the RBCC engine. The potential performance improvement of an ABCC engine over a RBCC engine comes from two factors. Firstly, the energy density of nuclear reaction is several order of magnitudes higher than the chemical combustion. Secondly, hydrogen can produce much higher nozzle exit velocity because of its small molecular weight. A one-dimensional, transient numerical model was used to analyze a generic scramjet engine and it is used as a baseline to evaluate an imaginary ABCC engine performance. A nuclear reactor is treated as a black box energy source that replaces the role of the primary rocket and the chemical combustion chamber in a RBCC engine. Hydrogen is heated by the reactor and accelerated to produce high-speed ejection velocity. The ejection velocity up 10,000 m/sec is theoretically possible because of high energy density from the reactor and large gas constant of the hydrogen. Oxygen contained in the entrained air reacts with hydrogen and produces propulsive power for ejector mode operation. To provide enough thrust for initial acceleration, relatively large amount of hydrogen must be pumped through the reactor. Amount of oxygen contained in the entrained air may not be sufficient to burn all hydrogen and consequently combustion could occur at the end of exit nozzle. It is assumed that this combustion process is constant-pressure combustion at 1.0 atmospheric pressure and thus not affects the nozzle exit condition.

  12. Rockets Away!

    ERIC Educational Resources Information Center

    Kaahaaina, Nancy

    1997-01-01

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

  13. Base Heating Sensitivity Study for a 4-Cluster Rocket Motor Configuration in Supersonic Freestream

    NASA Technical Reports Server (NTRS)

    Mehta, Manish; Canabal, Francisco; Tashakkor, Scott B.; Smith, Sheldon D.

    2011-01-01

    In support of launch vehicle base heating and pressure prediction efforts using the Loci-CHEM Navier-Stokes computational fluid dynamics solver, 35 numerical simulations of the NASA TND-1093 wind tunnel test have been modeled and analyzed. This test article is composed of four JP-4/LOX 500 lbf rocket motors exhausting into a Mach 2 - 3.5 wind tunnel at various ambient pressure conditions. These water-cooled motors are attached to a base plate of a standard missile forebody. We explore the base heating profiles for fully coupled finite-rate chemistry simulations, one-way coupled RAMP (Reacting And Multiphase Program using Method of Characteristics)-BLIMPJ (Boundary Layer Integral Matrix Program - Jet Version) derived solutions and variable and constant specific heat ratio frozen flow simulations. Variations in turbulence models, temperature boundary conditions and thermodynamic properties of the plume have been investigated at two ambient pressure conditions: 255 lb/sq ft (simulated low altitude) and 35 lb/sq ft (simulated high altitude). It is observed that the convective base heat flux and base temperature are most sensitive to the nozzle inner wall thermal boundary layer profile which is dependent on the wall temperature, boundary layer s specific energy and chemical reactions. Recovery shock dynamics and afterburning significantly influences convective base heating. Turbulence models and external nozzle wall thermal boundary layer profiles show less sensitivity to base heating characteristics. Base heating rates are validated for the highest fidelity solutions which show an agreement within +/-10% with respect to test data.

  14. Ground based instruments and basic structures supporting rocket & balloon campaigns at Esrange

    NASA Astrophysics Data System (ADS)

    Widell, Ola

    2005-08-01

    Many campaigns at Esrange are involving validation of scientific instruments onboard satellites. The validation is often done by balloon borne flights within different stratospheric conditions. Several campaigns are also coordinated programs including rocket, balloon and ground-based instruments. For testing of unmanned vehicles and parachute systems we are taking advantage of the huge land recovery area near Esrange and the Vidsel test field 300km south of Esrange. Several flights within the NEAT concept have been performed. An optical observatory called KEOPS, located at Esrange, is the main site for ground based instruments. The observatory is mainly dedicated for optical instruments like photometers, cameras, FPIs and an IR interferometer. The major expansion of the launch pad for stratospheric balloons and the cooperation with NASA will result in long duration balloon flights from Esrange to Alaska carrying heavy astronomical payloads. First flight will start summer 2005 and with annual flights. The accommodation complex is also extended to a total of more than 100 rooms.

  15. Observations of the neutral atmosphere between 100 and 200 km using ARIA rocket-borne and ground-based instruments

    SciTech Connect

    Hecht, J.H.; Christensen, A.B.; Gutierrez, D.J.; Kayser, D.C.; Sharp, W.E.; Sharber, J.R.; Winningham, J.D.; Frahm, R.A.; Strickland, D.J.; Mcewen, D.J.

    1995-11-01

    The atmospheric response in the aurora (ARIA) rocket was launched at 1406 UT on March 3, 1992, from Poker Flat, Alaska, into a pulsating diffuse aurora; rocket-borne instruments included an eight-channel photometer, a far ultraviolet spectrometer, a 130.4-nm atomic oxygen resonance lamp, and two particle spectrometers covering the energy range of 1-400 eV and 10 eV to 20 keV. The photometer channels were isolated using narrow-band interference filters and included measurements of the strong permitted auroral emissions N2 (337.1-nm), N2(+) (391.4-nm), and O I (844.6-nm). A ground-based photometer measured the permitted N2(+) (427.8-nm), the forbidden O I (630.0-nm), and the permitted O I (844.6-nm) emissions. The ground-based instrument was pointed in the magnetic zenith. Also, the rocket payload was pointed in the magnetic zenith from 100 to 200 km on the upleg. The data were analyzed using the Strickland electron transport code, and the rocket and ground-based results were found to be in good agreement regarding the inferred characteristic energy (E(sub 0) is approximately equal to 3 keV) of the precipitating auroral flux and the composition of the neutral atmosphere during the rocket flight. In particular, it was found that the O/N2 density ratio in the neutral atmosphere diminished during the auroral substorm, which started about 2 hours before the ARIA rocket flight. The data showed that there was about a 10-minute delay between the onset of the substorm and the decrease of the O/N2 density ratio. At the time of the ARIA flight this ratio had nearly returned to its presubstorm value. However, the data also showed that the O/N2 density ration did not recover to its presubstorm value until nearly 30 minutes after the particle and joule heating had subsided. Both the photometer and oxygen resonance lamp data showed the presence of structure in the atomic oxygen densities in the region above 130 km.

  16. Observations of the neutral atmosphere between 100 and 200 km using ARIA rocket-borne and ground-based instruments

    SciTech Connect

    Hecht, J.H.; Christensen, A.B.; Gutierrez, D.J.

    1995-09-01

    The atmospheric response in the aurora (ARIA) rocket was launched at 1406 UT on March 3, 1992, from Poker Flat, Alaska, into a pulsating diffuse aurora; rocket-borne instruments included an eight-channel photometer, a far ultraviolet spectrometer, a 130.4-nm atomic oxygen resonance lamp, and two particle spectrometers covering the energy range of 1-400 eV and 10 eV to 20 keV. The photometer channels were isolated using narrow-band interference filters and included measurements of the strong permitted auroral emissions N{sub 2} (337.1 nm), N{sub 2}{sup +} (391.4 nm), and O I (844.6 nm). A ground-based photometer measured the premitted N{sub 2}{sup +} (427.8 nm), the forbidden O I (630.0 nm), and the premitted O I (844.6 nm) emissions. The ground-based instrument was pointed in the magnetic zenith. Also, the rocket payload was pointed in the magnetic zenith from 100 to 200 km on the upleg. The data were analyzed using the Strickland electron transport code, and the rocket and ground-based results were found to be in good agreement regarding the inferred characteristic energy of the precipitating auroral flux and the composition of the neutral atmosphere during the rocket flight. In particular, it was found that the O/N{sub 2} density ratio in the neutral atmosphere diminished during the auroral substorm, which started about 2 hours before the ARIA rocket flight. The data showed that there was about a 10-min delay between the onset of the substorm and the decrease of the O/N{sub 2} density ratio. At the time of the ARIA flight this ratio had nearly returned to its presubstorm value. However, the data also showed that the O/N{sub 2} density ratio did not recover to its presubstorm value until nearly 30 min after the particle and joule heating had subsided. Both the photometer and oxygen densities in the region above 130 km. The observed auroral brightness ratio B{sub 337.1}/B{sub 391.4} equaled 0.29 and was in agreement with other recent measurements.

  17. 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)

  18. Listeria's right-handed helical rocket-tail trajectories: mechanistic implications for force generation in actin-based motility.

    PubMed

    Zeile, William L; Zhang, Fangliang; Dickinson, Richard B; Purich, Daniel L

    2005-02-01

    Listeria monocytogenes forms right-handed helical rocket tail trajectories during actin-based motility in cell-free extracts, and this stereochemical feature is consistent with actoclampin's affinity-modulated, clamped-filament elongation model [Dickinson and Purich, 2002: Biophys J 82:605-617]. In that mechanism, right-handed torque is generated by an end-tracking molecular motor, each comprised of a filament barbed end and clamping protein that processively traces the right-handed helix of its filament partner. By contrast, torque is not a predicted property of those models (e.g., elastic propulsion, elastic Brownian ratchet, tethered ratchet, and insertional polymerization models) requiring filament barbed ends to depart/detach from the motile object's surface during/after each monomer-addition step. Helical trajectories also explain why Listeria undergoes longitudinal-axis rotation on a length-scale matching the helical periodicity of Listeria's rocket tails.

  19. Project-based introduction to aerospace engineering course: A model rocket

    NASA Astrophysics Data System (ADS)

    Jayaram, Sanjay; Boyer, Lawrence; George, John; Ravindra, K.; Mitchell, Kyle

    2010-05-01

    In this paper, a model rocket project suitable for sophomore aerospace engineering students is described. This project encompasses elements of drag estimation, thrust determination and analysis using digital data acquisition, statistical analysis of data, computer aided drafting, programming, team work and written communication skills. The student built rockets are launched in the university baseball field with the objective of carrying a specific amount of payload so that the rocket achieves a specific altitude before the parachute is deployed. During the course of the project, the students are introduced to real-world engineering practice through written report submission of their designs. Over the years, the project has proven to enhance the learning objectives, yet cost effective and has provided good outcome measures.

  20. Combined effect of Nitrogen, Phosphorus and Potassium fertilizers on the contents of glucosinolates in rocket salad (Eruca sativa Mill.).

    PubMed

    Chun, Jin-Hyuk; Kim, Silbia; Arasu, Mariadhas Valan; Al-Dhabi, Naif Abdullah; Chung, Doug Young; Kim, Sun-Ju

    2017-02-01

    Nitrogen (N), phosphorous (P) and potassium (K) are the most limiting factors in crop production. N often affects the amino acid composition of protein and in turn its nutritional quality. In Brassica plants, abundant supply of N fertilizer decreases the relative proportion of glucosinolates (GSLs), thus reducing the biological and medical values of the vegetables. Hence effort was made to evaluate the influence of different proportions of nutrient solutions containing N-P-K on the GSL profiles of rocket salad (Eruca sativa Mill.). Fifteen desulpho-(DS) GSLs were isolated and identified using liquid chromatography-mass spectrometry (LC/MS) analysis. Rocket salad plants supplied with lesser amount of N, P or higher concentrations of K showed a typical improvement in total GSL contents. In contrast, total GSL levels were less at higher N supply. Furthermore, with N concentrations above 5 mM and K concentrations less than 2.5 mM, the GSL amounts were on average 13.51 and 13.75 μmol/g dry weight (DW), respectively. Aliphatic GSLs predominated in all concentrations of NPK while indolyl GSLs made up marginally less amount of the total compositions. Five and 2 mM N and P possessed much higher levels of several types of aliphatic GSLs than other concentrations, including glucoerucin, glucoraphanin and dimeric 4-mercaptobutyl GSL. From this perspective, it is contended that supply of less N results in enhancing the metabolic pathway for the synthesis of GSLs in rocket salad.

  1. Upper atmospheric research at Woomera: The Australian-built sounding rockets

    NASA Astrophysics Data System (ADS)

    Dougherty, Kerrie

    2006-07-01

    Sounding rocket programs for upper atmospheric research (UAR) commenced at Woomera in 1957, with the British Skylark project and the Anglo-Australian HARP rockoon program. Shortly afterwards, the Weapons Research Establishment (WRE) inaugurated its own sounding rocket program with the first successful Australian-built sounding rocket, Long Tom, developed from surplus British motors. This research would eventually lead to the development of Australia's first satellite, WRESAT, launched in 1967. Long Tom became the first more than 10 Australian designed and built sounding rockets that would be employed at Woomera until the demise of the WRE upper atmosphere program in 1976, as the Range was winding down to the cessation of the Anglo-Australian Joint Project. Like Long Tom, most of the early Australian sounding rockets were constructed by the WRE from surplus British motors, but after 1967 the later vehicles had a much higher local content, based on combinations of British motors and/or rocket motors developed in Australia. These new motors were named after constellations of the southern sky, while the rockets themselves bore the names of Australian native birds. This paper will outline the technical and scientific history of the Australian sounding rocket program, examining its origins and the reasons for its demise. It will look at the sequential development of the various Australian rockets and consider the particular research projects with which they were associated, the relationship with the WRESAT project and the move to Australian production after 1967.

  2. Rheological, optical, and ballistic investigations of paraffin-based fuels for hybrid rocket propulsion using a two-dimensional slab-burner

    NASA Astrophysics Data System (ADS)

    Kobald, M.; Toson, E.; Ciezki, H.; Schlechtriem, S.; di Betta, S.; Coppola, M.; DeLuca, L.

    2016-07-01

    This paper describes combined rheological, ballistic, and optical analyses performed on paraffin-based mixtures that can be used as high regression rate hybrid rocket fuels. Experimental activities have been done at the DLR Institute of Space Propulsion in Lampoldshausen and at SPLab of Politecnico di Milano [1]. Herein, the experiments that were performed at the DLR are described in detail. Viscosity, surface tension, and regression rate of the fuels have been determined. Furthermore, the combustion was evaluated by optical measurements. Data collected so far indicate an increasing regression rate for decreasing viscosity of the liquid paraffin which is in accordance with the current theories. Droplet entrainment, which is related to high regression rates, is only visible for the low-viscosity paraffin-based fuels.

  3. Predicting performance of axial pump inducer of LOX booster turbo-pump of staged combustion cycle based rocket engine using CFD

    NASA Astrophysics Data System (ADS)

    Mishra, Arpit; Ghosh, Parthasarathi

    2015-12-01

    For low cost, high thrust, space missions with high specific impulse and high reliability, inert weight needs to be minimized and thereby increasing the delivered payload. Turbopump feed system for a liquid propellant rocket engine (LPRE) has the highest power to weight ratio. Turbopumps are primarily equipped with an axial flow inducer to achieve the high angular velocity and low suction pressure in combination with increased system reliability. Performance of the turbopump strongly depends on the performance of the inducer. Thus, for designing a LPRE turbopump, demands optimization of the inducer geometry based on the performance of different off-design operating regimes. In this paper, steady-state CFD analysis of the inducer of a liquid oxygen (LOX) axial pump used as a booster pump for an oxygen rich staged combustion cycle rocket engine has been presented using ANSYS® CFX. Attempts have been made to obtain the performance characteristic curves for the LOX pump inducer. The formalism has been used to predict the performance of the inducer for the throttling range varying from 80% to 113% of nominal thrust and for the different rotational velocities from 4500 to 7500 rpm. The results have been analysed to determine the region of cavitation inception for different inlet pressure.

  4. Study of midlatitude ionospheric irregularities and E- and F-region coupling based on rocket and radar observations from Japan

    NASA Astrophysics Data System (ADS)

    Yamamoto, M.

    2015-12-01

    We have been studying ionspheric irregularities in mid-latitude region by using radars, sounding rockets, etc. The mid-latitude ionosphere was considered much stable than those in the equatorial or polar region in the past, but our studies for years have revealed that there are much active variabilities. We found variety of wave-like structures that are specific in the mid-latitudes. One of the phenomena is quasi-periodic echoes (QP echoes) first observed by the MU radar that reflects horizontal plasma-density structures associated to sporadic-E layers. Another phenomenon is medium-scale traveling ionospheric disturbance (MSTID) in the F-region. In the generation mechanism we think that Ionospheric E- and F-region coupling process is important. In this presentation, we will discuss nature of mid-latitude ionosphere based on our observations; the MU radar, sounding rocket campaigns of SEEK-1/2, and recent MSTID rocket experiment from JAXA Uchinoura Space Center in July 2013.

  5. Rocket plume burn hazard.

    PubMed

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

    1980-05-01

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

  6. Note on some observed effects of rocket motor operation on the base pressures of bodies in free light

    NASA Technical Reports Server (NTRS)

    Purser, Paul E; Thibodaux, Joseph G; Jackson, H Herbert

    1950-01-01

    Some measurements of the effects of rocket-motor operation on base pressure were obtained incidental to other research on some bodies in free flight. These data are presented and qualitatively analyzed. The analysis indicates that jet effects on drag are of sufficient importance to deserve consideration in the design of jet motor nozzles, especially for aircraft and missiles where the thrust and drag are of the same order of magnitude. The base-pressure changes induced by the jet should be considered in the structural design of the outer body skin on the aft portion of fuselages containing jets. (author)

  7. Sounding rocket based investigations of HF waves in the auroral ionosphere

    NASA Astrophysics Data System (ADS)

    McAdams, Kristin Lynn

    1999-10-01

    The PHAZE II and Auroral Turbulence II sounding rockets were launched into active, pre-midnight aurora during the February 1997 sounding rocket campaign from Poker Flat, Alaska. Both rockets carried a full complement of plasma intruments including particle detectors and electric field instruments. The high frequency electric field instrument (HFE), flown on both rockets, was designed and built at Dartmouth College. This unusual instrument transmitted the full electric field waveform using a dedicated telemetry link. The unprecedented resolution in both frequency and time yielded the first identifiable observations of several HF wave phenomena. We investigated two of these phenomena, HF chirps in the region when fpe > fce and HF bands at higher altitudes where fpe < fce. HF chirps are extremely narrowband, short-lived emissions which occur when fpe > fce. We propose that these waves are created as Z-modes waves which are quasi- trapped in density cavities. HF bands have long durations and narrowband, constant frequency structure and are observed in regions where the local plasma density is varying. These emissions occur when fpe < fce and the whistler mode connects to the Langmuir mode. They are generated by an electron beam interaction which produces Langmuir waves which then move onto the whistler mode when the local plasma density increases. The HFE also provided a method for determining the local plasma density without relying on Langmuir probes or active plasma experiments. When the frequency cutoff of the background wideband emissions is evident, this cutoff is used as a track of the local plasma frequency, which is dependent on the plasma density. We used this technique to definitively correlate lower hybrid solitary structures with density gradients. The use of the HFE on both flights has allowed us to observe HF wave phenomena which have been inaccessible previously.

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

    NASA Technical Reports Server (NTRS)

    Doreswamy, Rajiv; Fry, Emma K.

    2012-01-01

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

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

    ERIC Educational Resources Information Center

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

    2012-01-01

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

  10. Base Flow and Heat Transfer Characteristics of a Four-Nozzle Clustered Rocket Engine: Effect of Nozzle Pressure Ratio

    NASA Technical Reports Server (NTRS)

    Nallasamy, R.; Kandula, M.; Duncil, L.; Schallhorn, P.

    2010-01-01

    The base pressure and heating characteristics of a four-nozzle clustered rocket configuration is studied numerically with the aid of OVERFLOW Navier-Stokes code. A pressure ratio (chamber pressure to freestream static pressure) range of 990 to 5,920 and a freestream Mach number range of 2.5 to 3.5 are studied. The qualitative trends of decreasing base pressure with increasing pressure ratio and increasing base heat flux with increasing pressure ratio are correctly predicted. However, the predictions for base pressure and base heat flux show deviations from the wind tunnel data. The differences in absolute values between the computation and the data are attributed to factors such as perfect gas (thermally and calorically perfect) assumption, turbulence model inaccuracies in the simulation, and lack of grid adaptation.

  11. Formulation, Casting, and Evaluation of Paraffin-Based Solid Fuels Containing Energetic and Novel Additives for Hybrid Rockets

    NASA Technical Reports Server (NTRS)

    Larson, Daniel B.; Desain, John D.; Boyer, Eric; Wachs, Trevor; Kuo, Kenneth K.; Borduin, Russell; Koo, Joseph H.; Brady, Brian B.; Curtiss, Thomas J.; Story, George

    2012-01-01

    This investigation studied the inclusion of various additives to paraffin wax for use in a hybrid rocket motor. Some of the paraffin-based fuels were doped with various percentages of LiAlH4 (up to 10%). Addition of LiAlH4 at 10% was found to increase regression rates between 7 - 10% over baseline paraffin through tests in a gaseous oxygen hybrid rocket motor. Mass burn rates for paraffin grains with 10% LiAlH4 were also higher than those of the baseline paraffin. RDX was also cast into a paraffin sample via a novel casting process which involved dissolving RDX into dimethylformamide (DMF) solvent and then drawing a vacuum on the mixture of paraffin and RDX/DMF in order to evaporate out the DMF. It was found that although all DMF was removed, the process was not conducive to generating small RDX particles. The slow boiling generated an inhomogeneous mixture of paraffin and RDX. It is likely that superheating the DMF to cause rapid boiling would likely reduce RDX particle sizes. In addition to paraffin/LiAlH4 grains, multi-walled carbon nanotubes (MWNT) were cast in paraffin for testing in a hybrid rocket motor, and assorted samples containing a range of MWNT percentages in paraffin were imaged using SEM. The fuel samples showed good distribution of MWNT in the paraffin matrix, but the MWNT were often agglomerated, indicating that a change to the sonication and mixing processes were required to achieve better uniformity and debundled MWNT. Fuel grains with MWNT fuel grains had slightly lower regression rate, likely due to the increased thermal conductivity to the fuel subsurface, reducing the burning surface temperature.

  12. 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.

  13. Ground and Space-Based Measurement of Rocket Engine Burns in the Ionosphere

    NASA Technical Reports Server (NTRS)

    Bernhardt, P. A.; Ballenthin, J. O.; Baumgardner, J. L.; Bhatt, A.; Boyd, I. D.; Burt, J. M.; Caton, R. G.; Coster, A.; Erickson, P. J.; Huba, J. D.; Earle, G. D.; Kaplan, C. R.; Foster, J. C.; Groves, K. M.; Haaser, R. A.; Heelis, R. A.; Hunton, D. E.; Hysell, D. L.; Klenzing, J. H.; Larsen, M. F.; Lind, F. D.; Pedersen, T. R.; Pfaff, R. F.; Stoneback, R. A.; Roddy, P. A.; Rodriguez, S. P.; San Antonio, G. S.; Schuck, P. W.; Siefring, C. L.; Selcher, C. A.; Smith, S. M.; Talaat, E. R.; Thomason, J. F.; Tsunoda, R. T.; Varney, R. H.

    2013-01-01

    On-orbit firings of both liquid and solid rocket motors provide localized disturbances to the plasma in the upper atmosphere. Large amounts of energy are deposited to ionosphere in the form of expanding exhaust vapors which change the composition and flow velocity. Charge exchange between the neutral exhaust molecules and the background ions (mainly O+) yields energetic ion beams. The rapidly moving pickup ions excite plasma instabilities and yield optical emissions after dissociative recombination with ambient electrons. Line-of-sight techniques for remote measurements rocket burn effects include direct observation of plume optical emissions with ground and satellite cameras, and plume scatter with UHF and higher frequency radars. Long range detection with HF radars is possible if the burns occur in the dense part of the ionosphere. The exhaust vapors initiate plasma turbulence in the ionosphere that can scatter HF radar waves launched from ground transmitters. Solid rocket motors provide particulates that become charged in the ionosphere and may excite dusty plasma instabilities. Hypersonic exhaust flow impacting the ionospheric plasma launches a low-frequency, electromagnetic pulse that is detectable using satellites with electric field booms. If the exhaust cloud itself passes over a satellite, in situ detectors measure increased ion-acoustic wave turbulence, enhanced neutral and plasma densities, elevated ion temperatures, and magnetic field perturbations. All of these techniques can be used for long range observations of plumes in the ionosphere. To demonstrate such long range measurements, several experiments were conducted by the Naval Research Laboratory including the Charged Aerosol Release Experiment, the Shuttle Ionospheric Modification with Pulsed Localized Exhaust experiments, and the Shuttle Exhaust Ionospheric Turbulence Experiments.

  14. High Strength Carbide-Based Fibrous Monolith Materials for Solid Rocket Nozzles

    DTIC Science & Technology

    2008-02-19

    by winding FM filaments around a steel mandrel. 5 Figure 3. Fabrication of radially aligned, “pie” nozzle. Figure 4. FM solid rocket...brittle transition temperature ( DBTT ) between 3000°F and 3500°F so a change in behavior of the Gen 3 material between these temperatures would be possible...Figure 18 shows the setup used during these evaluations. The 1.5” x 1.5” monolithic and fibrous monolithic specimens were mounted in a steel

  15. Current and Future Rocket Propulsion Testing at NASA Stennis Space Center

    NASA Technical Reports Server (NTRS)

    Ryan, H. M.; Rahman, S.; Gilbrech, R.

    2000-01-01

    Year 2000 has been an active one for large-scale propulsion testing at the NASA John C. Stennis Space Center. This paper highlights several of the current-year test programs conducted at the Stennis Space Center (SSC) including the X-33 Aerospike Engine, Ultra Low Cost Engine (ULCE) program, and the Hybrid Sounding Rocket (HYSR) program. Future directions in propulsion test are also introduced including the development of a large-scale Rocket Based Combined Cycle (RBCC) test facility.

  16. 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.

  17. Microfabricated Liquid Rocket Motors

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

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

  18. A fuzzy case based reasoning tool for model based approach to rocket engine health monitoring

    NASA Technical Reports Server (NTRS)

    Krovvidy, Srinivas; Nolan, Adam; Hu, Yong-Lin; Wee, William G.

    1992-01-01

    In this system we develop a fuzzy case based reasoner that can build a case representation for several past anomalies detected, and we develop case retrieval methods that can be used to index a relevant case when a new problem (case) is presented using fuzzy sets. The choice of fuzzy sets is justified by the uncertain data. The new problem can be solved using knowledge of the model along with the old cases. This system can then be used to generalize the knowledge from previous cases and use this generalization to refine the existing model definition. This in turn can help to detect failures using the model based algorithms.

  19. Colorado Hydrogen Imaging Rocket Payload

    NASA Astrophysics Data System (ADS)

    Burgh, Eric B.; France, K.

    2009-01-01

    We present the design for a rocket-borne narrow-band far-ultraviolet imaging telescope. It will measure the spatial distribution of photo-excited molecular hydrogen emission nearby hot stars by utilizing multi-layer reflection coatings, similar to those used in previous NASA experiments, to obtain two images during a flight: one with a narrow-band filter that captures the 1575/1608A emission features (the "on-band" filter), and a second one that measures the dust-scattered stellar continuum at 1800A (the "off-band" filter). The difference image will then isolate the molecular hydrogen emission by subtracting the underlying scattered-light background. This would be a large improvement over existing studies at ultraviolet wavelengths for which many individual pointings with spectroscopic apertures are required to map the region of interest. These data will complete the picture, combined with far-ultraviolet spectra and near-infrared observations of vibrational emission that we will obtain from ground-based instrumentation, of the physical conditions in sites of recent and on-going star formation. A sounding rocket payload such as this provides the opportunity to perform niche science that other facilities cannot as well as advances the readiness of junior researchers to assume leadership roles on future NASA space flight missions.

  20. Base Flow Characteristics for Several Four-Clustered Rocket Configurations at Mach Numbers from 2.0 to 3.5

    NASA Technical Reports Server (NTRS)

    Musial, Norman T.; Ward, James J.

    1961-01-01

    A generalized study of base flow phenomena has been conducted with four 500-pound-thrust JP-4 fuel-liquid-oxygen rocket motors installed in the base of a 12-inch-diameter cylindrical model. Data were obtained over a Mach number and nozzle pressure ratio range of 2.0 to 3.5 and 340 to 600, respectively. Base heat flux, gas temperature, and pressure were highest in the center of the cluster core and decreased in a radial direction. Although a maximum heat flux of 93 Btu per square foot per second was measured within the cluster core, peripheral heat fluxes were low, averaging about 5 Btu per square foot per second for all configurations. Generally base heat flux was found to be independent of Mach number over the range investigated. Base heat flux within the cluster core was decreased by increasing motor spacing, motor extension, a combination of increasing nozzle area ratio and decreasing exit angle and gimbaling the two side engines. Small amounts of nitrogen injected within the cluster core sharply reduced core heat flux.

  1. The rocket-and-wire triggering process: Channel-base currents and ground-level electric fields

    NASA Astrophysics Data System (ADS)

    Ngin, Terry Keo

    Rocket-and-wire triggered lightning flashes were studied from 2011 to 2013 at the International Center for Lightning Research and Testing. Ground-level electric fields and channel-base currents were recorded for 79 rocket launches. An eight-station network of electric field meters along with a milliampere-scale wire-base current measurement and a high-speed video record of the wire ascent allowed the calculation and analysis of the trigger-wire line charge density, generally found to be in the muC m-1 to hundreds of muC m-1 range and to increase quadratically with height. The wire-base currents collected during the wire ascent here are the most comprehensive in the literature to date. The trigger-wire line charge density, electric field at ground level, and characteristics of precursor pulses at the wire tip were examined to determine their usefulness in predicting the success or failure of a triggered-lightning attempt. The usefulness of the PICASSO model of space charge evolution from ground, originally developed by researchers at Paul Sabatier University in the 1980s, as a triggering criterion was also evaluated. An electrostatic model of the corona sheath around the trigger-wire was developed in order to estimate the radial extent of the corona sheath and the charge distribution within the corona sheath as a function of measured electric fields aloft taken from the published literature. The most sensitive measurements to date of channel-base current flowing prior to subsequent return strokes, where the current had generally been considered to be zero, were collected and are analyzed here. The channel-base current before return strokes was found to average 5.3 mA with a 2.8 mA standard deviation prior to 120 measured return strokes.

  2. Sounding rockets explore the ionosphere

    SciTech Connect

    Mendillo, M. )

    1990-08-01

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

  3. Aerodynamics of Sounding-Rocket Geometries

    NASA Technical Reports Server (NTRS)

    Barrowman, J.

    1982-01-01

    Theoretical aerodynamics program TAD predicts aerodynamic characteristics of vehicles with sounding-rocket configurations. These slender, Axisymmetric finned vehicles have a wide range of aeronautical applications from rockets to high-speed armament. TAD calculates characteristics of separate portions of vehicle, calculates interference between portions, and combines results to form total vehicle solution.

  4. CFD-based surrogate modeling of liquid rocket engine components via design space refinement and sensitivity assessment

    NASA Astrophysics Data System (ADS)

    Mack, Yolanda

    Computational fluid dynamics (CFD) can be used to improve the design and optimization of rocket engine components that traditionally rely on empirical calculations and limited experimentation. CFD based-design optimization can be made computationally affordable through the use of surrogate modeling which can then facilitate additional parameter sensitivity assessments. The present study investigates surrogate-based adaptive design space refinement (DSR) using estimates of surrogate uncertainty to probe the CFD analyses and to perform sensitivity assessments for complex fluid physics associated with liquid rocket engine components. Three studies were conducted. First, a surrogate-based preliminary design optimization was conducted to improve the efficiency of a compact radial turbine for an expander cycle rocket engine while maintaining low weight. Design space refinement was used to identify function constraints and to obtain a high accuracy surrogate model in the region of interest. A merit function formulation for multi-objective design point selection reduced the number of design points by an order of magnitude while maintaining good surrogate accuracy among the best trade-off points. Second, bluff body-induced flow was investigated to identify the physics and surrogate modeling issues related to the flow's mixing dynamics. Multiple surrogates and DSR were instrumental in identifying designs for which the CFD model was deficient and to help to pinpoint the nature of the deficiency. Next, a three-dimensional computational model was developed to explore the wall heat transfer of a GO2/GH2 shear coaxial single element injector. The interactions between turbulent recirculating flow structures, chemical kinetics, and heat transfer are highlighted. Finally, a simplified computational model of multi-element injector flows was constructed to explore the sensitivity of wall heating and improve combustion efficiency to injector element spacing. Design space refinement

  5. Combustion-Characteristic-Based Active Thrust Modulation of a Solid Rocket Motor

    NASA Astrophysics Data System (ADS)

    Tanaka, Masafumi; Gaspard, Guillaume; Urakawa, Katsuya

    A new concept of thrust modulation of solid propellant rocket motor is proposed. Some propellants cannot burn at intermediate pressure, while they can burn at lower and higher pressures. When one applies such a propellant to a motor, two combustion modes or two thrust levels are attainable without any change of the nozzle configuration. In the experiments different ignition conditions brought independent two combustion modes (low mode and high mode) in the same motor geometry. Some motors showed a natural transition from low mode to high mode. As an example, the alternative thrust levels were 50 N and 180 N. The natural transition was restricted with use of the partitioned grain. An active transition method was explored by exerting pressure perturbation through a vent hole with a ball valve. The valve system worked for the transition from high mode to low mode, but the reverse transition was not achieved well.

  6. Ignition and flame stabilization of a strut-jet RBCC combustor with small rocket exhaust.

    PubMed

    Hu, Jichao; Chang, Juntao; Bao, Wen

    2014-01-01

    A Rocket Based Combined Cycle combustor model is tested at a ground direct connected rig to investigate the flame holding characteristics with a small rocket exhaust using liquid kerosene. The total temperature and the Mach number of the vitiated air flow, at exit of the nozzle are 1505 K and 2.6, respectively. The rocket base is embedded in a fuel injecting strut and mounted in the center of the combustor. The wall of the combustor is flush, without any reward step or cavity, so the strut-jet is used to make sure of the flame stabilization of the second combustion. Mass flow rate of the kerosene and oxygen injected into the rocket is set to be a small value, below 10% of the total fuel when the equivalence ratio of the second combustion is 1. The experiment has generated two different kinds of rocket exhaust: fuel rich and pure oxygen. Experiment result has shown that, with a relative small total mass flow rate of the rocket, the fuel rich rocket plume is not suitable for ignition and flame stabilization, while an oxygen plume condition is suitable. Then the paper conducts a series of experiments to investigate the combustion characteristics under this oxygen pilot method and found that the flame stabilization characteristics are different at different combustion modes.

  7. Ignition and Flame Stabilization of a Strut-Jet RBCC Combustor with Small Rocket Exhaust

    PubMed Central

    2014-01-01

    A Rocket Based Combined Cycle combustor model is tested at a ground direct connected rig to investigate the flame holding characteristics with a small rocket exhaust using liquid kerosene. The total temperature and the Mach number of the vitiated air flow, at exit of the nozzle are 1505 K and 2.6, respectively. The rocket base is embedded in a fuel injecting strut and mounted in the center of the combustor. The wall of the combustor is flush, without any reward step or cavity, so the strut-jet is used to make sure of the flame stabilization of the second combustion. Mass flow rate of the kerosene and oxygen injected into the rocket is set to be a small value, below 10% of the total fuel when the equivalence ratio of the second combustion is 1. The experiment has generated two different kinds of rocket exhaust: fuel rich and pure oxygen. Experiment result has shown that, with a relative small total mass flow rate of the rocket, the fuel rich rocket plume is not suitable for ignition and flame stabilization, while an oxygen plume condition is suitable. Then the paper conducts a series of experiments to investigate the combustion characteristics under this oxygen pilot method and found that the flame stabilization characteristics are different at different combustion modes. PMID:24578655

  8. Impact of Plasma Sheath on Rocket-based E-region Ion Measurements

    NASA Astrophysics Data System (ADS)

    Imtiaz, N.; Burchill, J. K.; Marchand, R.

    2013-12-01

    We model the particle velocity distribution functions around the entrance window of the Suprathermal Ion Imager (SII) to assess the impact of electrostatic sheath on ion measurements in the E-region ionosphere. The SII sensor is an electrostatic analyzer that measures two dimensional slices of the distribution of the kinetic energies and arrival-angles of low energy ions. The study is concerned with the interpretation of data obtained from the sensor SII that was affixed to a 1-m NASA rocket 36.234 as part of the Joule II mission to investigate Joule heating in the E-region ionosphere. The rocket flew into quiet auroral conditions above Northern Alaska on 19 January 2007. The payload was spin-stabilitized with a period of 1.6 s, giving an apparent rotation of the ion flow velocity in the frame of reference of the SII. We numerically investigate the ram velocity effect on the ions velocity distributions in the vicinity of SII aperture at an altitudes of approximately 150km. The electrostatic sheath potential profiles surrounding the sensor and payload are calculated numerically with the PIC code PTetra. It is observed that the direction of the ion flow velocity vector modifies the plasma sheath potential profile. This in turn impacts the velocity distributions of molecular oxygen and Nitric oxideions at the aperture of the particle sensor. The velocity distribution functions are calculated by using test-particle modeling. These particle distribution functions are then used to inject the particles in the particle sensor, and to calculate the fluxes on the sensor microchannel plate (MCP).

  9. Impact of plasma sheath on rocket-based E-region ion measurements

    NASA Astrophysics Data System (ADS)

    Imtiaz, Nadia; Burchill, Johnathan; Marchand, Richard

    2015-01-01

    We model the particle velocity distribution functions around the entrance window of the Suprathermal Ion Imager (SII). The SII sensor was mounted on a 1 m boom carried by the scientific payload of NASA rocket 36.234 as part of Joule II mission to investigate Joule heating in the E-region ionosphere. The rocket flew above Northern Alaska on 19 January 2007. The payload was spin-stabilized with a period of 1.6 s, giving an apparent rotation of the ion flow velocity in the frame of reference of the payload. The SII sensor is an electrostatic analyzer that measures two dimensional slices of the distribution of the kinetic energies and arrival-angles of low energy ions. The study is concerned with the interpretation of data obtained from the SII sensor. For this purpose, we numerically investigate ram velocity effects on ions velocity distributions in the vicinity of the SII sensor aperture at an altitudes of approximately 150 km. The electrostatic sheath profiles surrounding the SII sensor, boom and payload are calculated numerically with the PIC code PTetra. It is observed that the direction of the ion flow velocity modifies the plasma sheath potential profile. This in turn impacts the velocity distributions of NO+ and ions at the aperture of the particle sensor. The velocity distribution functions at the sensor aperture are calculated by using test-particle modeling. These particle distribution functions are then used to inject particles in the sensor, and calculate the fluxes on the sensor microchannel plate (MCP), from which comparisons with the measurements can be made.

  10. A Rocket-Base Study of Auroral Electrodynamics Within the Current Closure Ionosphere

    NASA Technical Reports Server (NTRS)

    Kaeppler, Stephen R.; Kletzing, Craig; Bounds, Scott R.; Sigsbee, Kristine M.; Gjerloev, Jesper W.; Anderson, Brian Jay; Korth, Haje; Lessard, Marc; Labelle, James W.; Dombrowski, Micah P.; Pfaff, Robert F.; Rowland, Douglas E.; Jones, Sarah; Heinselman, Craig J.; DudokdeWit, Thierry

    2011-01-01

    The Auroral Current and Electrodynamics Structure (ACES) mission consisted of two sounding rockets launched nearly simultaneously from Poker Flat Research Range, AK on January 29, 2009 into a dynamic multiple-arc aurora. The ACES rocket mission, in conjunction with the PFISR Radar, was designed to observe the three-dimensional current system of a stable auroral arc system. ACES utilized two well instrumented payloads flown along very similar magnetic field footprints, at various altitudes with small temporal separation between both payloads. ACES High, the higher altitude payload (apogee 360 km), took in-situ measurements of the plasma parameters above the current closure region to provide the input signature into the lower ionosphere. ACES Low, the low-altitude payload (apogee 130 km), took similar observations within the current closure region, where cross-field currents can flow. We present results comparing observations of the electric fields, magnetic fields, electron flux, and the electron temperature at similar magnetic footpoints between both payloads. We further present data from all-sky imagers and PFISR detailing the evolution of the auroral event as the payloads traversed regions connected by similar magnetic footpoints. Current measurements derived from the magnetometers on both payloads are further compared. We examine data from both PFISR and observations on the high-altitude payload which we interpreted as a signature of electron acceleration by means of Alfv n waves. We further examine all measurements to understand ionospheric conductivity and how energy is being deposited into the ionosphere through Joule heating. Data from ACES is compared against models of Joule heating to make inferences regarding the effect of collisions at various altitudes.

  11. A rocket-borne microprocessor-based experiment for investigation of energetic particles in the D and E regions

    NASA Technical Reports Server (NTRS)

    Braswell, F. M.

    1981-01-01

    An energetic experiment using the Z80 family of microcomputer components is described. Data collected from the experiment allowed fast and efficient postprocessing, yielding both energy-spectrum and pitch-angle distribution of energetic particles in the D and E regions. Advanced microprocessor system architecture and software concepts were used in the design to cope with the large amount of data being processed. This required the Z80 system to operate at over 80% of its total capacity. The microprocessor system was included in the payloads of three rockets launched during the Energy Budget Campaign at ESRANGE, Kiruna, Sweden in November 1980. Based on preliminary examination of the data, the performance of the experiment was satisfactory and good data were obtained on the energy spectrum and pitch-angle distribution of the particles.

  12. Small rocket flowfield diagnostic chambers

    NASA Technical Reports Server (NTRS)

    Morren, Sybil; Reed, Brian

    1993-01-01

    Instrumented and optically-accessible rocket chambers are being developed to be used for diagnostics of small rocket (less than 440 N thrust level) flowfields. These chambers are being tested to gather local fluid dynamic and thermodynamic flowfield data over a range of test conditions. This flowfield database is being used to better understand mixing and heat transfer phenomena in small rockets, influence the numerical modeling of small rocket flowfields, and characterize small rocket components. The diagnostic chamber designs include: a chamber design for gathering wall temperature profiles to be used as boundary conditions in a finite element heat flux model; a chamber design for gathering inner wall temperature and static pressure profiles; and optically-accessible chamber designs, to be used with a suite of laser-based diagnostics for gathering local species concentration, temperature, density, and velocity profiles. These chambers were run with gaseous hydrogen/gaseous oxygen (GH2/GO2) propellants, while subsequent versions will be run on liquid oxygen/hydrocarbon (LOX/HC) propellants. The purpose, design, and initial test results of these small rocket flowfield diagnostic chambers are summarized.

  13. ISRO's solid rocket motors

    NASA Astrophysics Data System (ADS)

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

    1989-08-01

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

  14. Rocket Noise Prediction Program

    NASA Technical Reports Server (NTRS)

    Margasahayam, Ravi; Caimi, Raoul

    1999-01-01

    A comprehensive, automated, and user-friendly software program was developed to predict the noise and ignition over-pressure environment generated during the launch of a rocket. The software allows for interactive modification of various parameters affecting the generated noise environment. Predictions can be made for different launch scenarios and a variety of vehicle and launch mount configurations. Moreover, predictions can be made for both near-field and far-field locations on the ground and any position on the vehicle. Multiple engine and fuel combinations can be addressed, and duct geometry can be incorporated efficiently. Applications in structural design are addressed.

  15. Small rocket tornado probe

    SciTech Connect

    Colgate, S.A.

    1982-01-01

    A (less than 1 lb.) paper rock tornado probe was developed and deployed in an attempt to measure the pressure, temperature, ionization, and electric field variations along a trajectory penetrating a tornado funnel. The requirements of weight and materials were set by federal regulations and a one-meter resolution at a penetration velocity of close to Mach 1 was desired. These requirements were achieved by telemetering a strain gage transducer for pressure, micro size thermister and electric field, and ionization sensors via a pulse time telemetry to a receiver on board an aircraft that digitizes a signal and presents it to a Z80 microcomputer for recording on mini-floppy disk. Recording rate was 2 ms for 8 channels of information that also includes telemetry rf field strength, magnetic field for orientation on the rocket, zero reference voltage for the sensor op amps as well as the previously mentioned items also. The absolute pressure was recorded. Tactically, over 120 h were flown in a Cessna 210 in April and May 1981, and one tornado was encountered. Four rockets were fired at this tornado, missed, and there were many equipment problems. The equipment needs to be hardened and engineered to a significant degree, but it is believed that the feasibility of the probe, tactics, and launch platform for future tornado work has been proven. The logistics of thunderstorm chasing from a remote base in New Mexico is a major difficulty and reliability of the equipment another. Over 50 dummy rockets have been fired to prove trajectories, stability, and photographic capability. Over 25 electronically equipped rockets have been fired to prove sensors transmission, breakaway connections, etc. The pressure recovery factor was calibrated in the Air Force Academy blow-down tunnel. There is a need for more refined engineering and more logistic support.

  16. Computational Fluid Dynamic (CFD) analysis of axisymmetric plume and base flow of film/dump cooled rocket nozzle

    NASA Technical Reports Server (NTRS)

    Tucker, P. K.; Warsi, S. A.

    1993-01-01

    Film/dump cooling a rocket nozzle with fuel rich gas, as in the National Launch System (NLS) Space Transportation Main Engine (STME), adds potential complexities for integrating the engine with the vehicle. The chief concern is that once the film coolant is exhausted from the nozzle, conditions may exist during flight for the fuel-rich film gases to be recirculated to the vehicle base region. The result could be significantly higher base temperatures than would be expected from a regeneratively cooled nozzle. CFD analyses were conduced to augment classical scaling techniques for vehicle base environments. The FDNS code with finite rate chemistry was used to simulate a single, axisymmetric STME plume and the NLS base area. Parallel calculations were made of the Saturn V S-1 C/F1 plume base area flows. The objective was to characterize the plume/freestream shear layer for both vehicles as inputs for scaling the S-C/F1 flight data to NLS/STME conditions. The code was validated on high speed flows with relevant physics. This paper contains the calculations for the NLS/STME plume for the baseline nozzle and a modified nozzle. The modified nozzle was intended to reduce the fuel available for recirculation to the vehicle base region. Plumes for both nozzles were calculated at 10kFT and 50kFT.

  17. The Solid Rocket Motor Slag Population: Results of a Radar-based Regressive Statistical Evaluation

    NASA Technical Reports Server (NTRS)

    Horstman, Matthew F.; Xu, Yu-Lin

    2008-01-01

    Solid rocket motor (SRM) slag has been identified as a significant source of man-made orbital debris. The propensity of SRMs to generate particles of 100 m and larger has caused concern regarding their contribution to the debris environment. Radar observation, rather than in-situ gathered evidence, is currently the only measurable source for the NASA/ODPO model of the on-orbit slag population. This simulated model includes the time evolution of the resultant orbital populations using a historical database of SRM launches, propellant masses, and estimated locations and times of tail-off. However, due to the small amount of observational evidence, there can be no direct comparison to check the validity of this model. Rather than using the assumed population developed from purely historical and physical assumptions, a regressional approach was used which utilized the populations observed by the Haystack radar from 1996 to present. The estimated trajectories from the historical model of slag sources, and the corresponding plausible detections by the Haystack radar, were identified. Comparisons with observational data from the ensuing years were made, and the SRM model was altered with respect to size and mass production of slag particles to reflect the historical data obtained. The result is a model SRM population that fits within the bounds of the observed environment.

  18. High-End Concept Based on Hypersonic Two-Stage Rocket and Electro-Magnetic Railgun to Launch Micro-Satellites Into Low-Earth

    NASA Astrophysics Data System (ADS)

    Bozic, O.; Longo, J. M.; Giese, P.; Behren, J.

    2005-02-01

    The electromagnetic railgun technology appears to be an interesting alternative to launch small payloads into Low Earth Orbit (LEO), as this may introduce lower launch costs. A high-end solution, based upon present state of the art technology, has been investigated to derive the technical boundary conditions for the application of such a new system. This paper presents the main concept and the design aspects of such propelled projectile with special emphasis on flight mechanics, aero-/thermodynamics, materials and propulsion characteristics. Launch angles and trajectory optimisation analyses are carried out by means of 3 degree of freedom simulations (3DOF). The aerodynamic form of the projectile is optimised to provoke minimum drag and low heat loads. The surface temperature distribution for critical zones is calculated with DLR developed Navier-Stokes codes TAU, HOTSOSE, whereas the engineering tool HF3T is used for time dependent calculations of heat loads and temperatures on project surface and inner structures. Furthermore, competing propulsions systems are considered for the rocket engines of both stages. The structural mass is analysed mostly on the basis of carbon fibre reinforced materials as well as classical aerospace metallic materials. Finally, this paper gives a critical overview of the technical feasibility and cost of small rockets for such missions. Key words: micro-satellite, two-stage-rocket, railgun, rocket-engines, aero/thermodynamic, mass optimization

  19. Solar Thermal Rocket Propulsion

    NASA Technical Reports Server (NTRS)

    Sercel, J. C.

    1986-01-01

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

  20. American Rocket Society

    NASA Technical Reports Server (NTRS)

    2004-01-01

    In addition to Dr. Robert Goddard's pioneering work, American experimentation in rocketry prior to World War II grew, primarily in technical societies. This is an early rocket motor designed and developed by the American Rocket Society in 1932.

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

    NASA Technical Reports Server (NTRS)

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

    2014-01-01

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

  2. ’MLRS’: A Rocket System for the Marine Corps

    DTIC Science & Technology

    1990-03-29

    In 1846, an English inventor , William Hale, entered rocket history by improving the designing of the Congreve rocket. Rockets known for their erratic...or "curved vanes", were used in conjunction with "tangential holes at the periphery of the base" (developed by an American inventor named Court) to...twelve launchers each, which were mounted over the rear axle and fired electrically from the cab. These were the 4.5-inch rockets (MB) used in the T45

  3. Overview of the Turbine Based Combined Cycle Discipline

    NASA Technical Reports Server (NTRS)

    Thomas, Scott R.; Walker, James F.; Pittman, James L.

    2009-01-01

    The NASA Fundamental Aeronautics Hypersonics project is focused on technologies for combined cycle, airbreathing propulsions systems to enable reusable launch systems for access to space. Turbine Based Combined Cycle (TBCC) propulsion systems offer specific impulse (Isp) improvements over rocket-based propulsion systems in the subsonic takeoff and return mission segments and offer improved safety. The potential to realize more aircraft-like operations with expanded launch site capability and reduced system maintenance are additional benefits. The most critical TBCC enabling technologies as identified in the National Aeronautics Institute (NAI) study were: 1) mode transition from the low speed propulsion system to the high speed propulsion system, 2) high Mach turbine engine development, 3) transonic aero-propulsion performance, 4) low-Mach-number dual-mode scramjet operation, 5) innovative 3-D flowpath concepts and 6) innovative turbine based combined cycle integration. To address several of these key TBCC challenges, NASA s Hypersonics project (TBCC Discipline) initiated an experimental mode transition task that includes an analytic research endeavor to assess the state-of-the-art of propulsion system performance and design codes. This initiative includes inlet fluid and turbine performance codes and engineering-level algorithms. This effort has been focused on the Combined Cycle Engine Large-Scale Inlet Mode Transition Experiment (CCE LIMX) which is a fully integrated TBCC propulsion system with flow path sizing consistent with previous NASA and DoD proposed Hypersonic experimental flight test plans. This experiment is being tested in the NASA-GRC 10 x 10 Supersonic Wind Tunnel (SWT) Facility. The goal of this activity is to address key hypersonic combined-cycle-engine issues: (1) dual integrated inlet operability and performance issues unstart constraints, distortion constraints, bleed requirements, controls, and operability margins, (2) mode

  4. 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.

  5. 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.

  6. Recent Experimental Results Related to Ejector Mode Studies of Rocket-Based Combined Cycle (RBCC) Engines

    NASA Technical Reports Server (NTRS)

    Cramer, J. M.; Pal, S.; Marshall, W. M.; Santoro, R. J.

    2003-01-01

    Contents include the folloving: 1. Motivation. Support NASA's 3d generation launch vehicle technology program. RBCC is promising candidate for 3d generation propulsion system. 2. Approach. Focus on ejector mode p3erformance (Mach 0-3). Perform testing on established flowpath geometry. Use conventional propulsion measurement techniques. Use advanced optical diagnostic techniques to measure local combustion gas properties. 3. Objectives. Gain physical understanding of detailing mixing and combustion phenomena. Establish an experimental data set for CFD code development and validation.

  7. Automated Rocket Propulsion Test Management

    NASA Technical Reports Server (NTRS)

    Walters, Ian; Nelson, Cheryl; Jones, Helene

    2007-01-01

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

  8. Catalytic Microtube Rocket Igniter

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.; Deans, Matthew C.

    2011-01-01

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

  9. Acoustic Measurements of Small Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

    Vargas, Magda B.; Kenny, R. Jeremy

    2010-01-01

    Rocket acoustic noise can induce loads and vibration on the vehicle as well as the surrounding structures. Models have been developed to predict these acoustic loads based on scaling existing solid rocket motor data. The NASA Marshall Space Flight Center acoustics team has measured several small solid rocket motors (thrust below 150,000 lbf) to anchor prediction models. This data will provide NASA the capability to predict the acoustic environments and consequent vibro-acoustic response of larger rockets (thrust above 1,000,000 lbf) such as those planned for the NASA Constellation program. This paper presents the methods used to measure acoustic data during the static firing of small solid rocket motors and the trends found in the data.

  10. Acoustic Measurements for Small Solid Rocket Motors

    NASA Technical Reports Server (NTRS)

    Vargas, Magda B.; Kenny, R. Jeremy

    2010-01-01

    Models have been developed to predict large solid rocket motor acoustic loads based on the scaling of small solid rocket motors. MSFC has measured several small solid rocket motors in horizontal and launch configurations to anchor these models. Solid Rocket Test Motor (SRTM) has ballistics similar to the Reusable Solid Rocket Motor (RSRM) therefore a good choice for acoustic scaling. Acoustic measurements were collected during the test firing of the Insulation Configuration Extended Length (ICXL) 7,6, and 8 (in firing order) in order to compare to RSRM horizontal firing data. The scope of this presentation includes: Acoustic test procedures and instrumentation implemented during the three SRTM firings and Data analysis method and general trends observed in the data.

  11. Small rocket research and technology

    NASA Technical Reports Server (NTRS)

    Schneider, Steven; Biaglow, James

    1993-01-01

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

  12. Small rocket research and technology

    NASA Astrophysics Data System (ADS)

    Schneider, Steven; Biaglow, James

    1993-11-01

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

  13. Rhenium Rocket Manufacturing Technology

    NASA Technical Reports Server (NTRS)

    1997-01-01

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

  14. Liquid Rocket Engine Testing

    DTIC Science & Technology

    2016-10-21

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

  15. Rocket astronomy - an overview

    NASA Astrophysics Data System (ADS)

    Friedman, H.

    The history of rocket astronomy is retold, with emphasis on solar investigations. The use of captured V-2 rockets after World War II was followed by the Aerobee, which exceeded the V-2's altitude and was more reliable. The V-2 has made first-generation investigations in X-ray and UV photometry, which was followed by studies of the solar cycle X-ray variablity, the X-ray corona, and solar flares. Nike rockets played an important role in these investigations. The role of rockets in galactic X-ray astronomy is briefly described.

  16. Four-nozzle benchmark wind tunnel model USA code solutions for simulation of multiple rocket base flow recirculation at 145,000 feet altitude

    NASA Astrophysics Data System (ADS)

    Dougherty, N. S.; Johnson, S. L.

    1993-07-01

    Multiple rocket exhaust plume interactions at high altitudes can produce base flow recirculation with attendant alteration of the base pressure coefficient and increased base heating. A search for a good wind tunnel benchmark problem to check grid clustering technique and turbulence modeling turned up the experiment done at AEDC in 1961 by Goethert and Matz on a 4.25-in. diameter domed missile base model with four rocket nozzles. This wind tunnel model with varied external bleed air flow for the base flow wake produced measured p/p(sub ref) at the center of the base as high as 3.3 due to plume flow recirculation back onto the base. At that time in 1961, relatively inexpensive experimentation with air at gamma = 1.4 and nozzle A(sub e)/A of 10.6 and theta(sub n) = 7.55 deg with P(sub c) = 155 psia simulated a LO2/LH2 rocket exhaust plume with gamma = 1.20, A(sub e)/A of 78 and P(sub c) about 1,000 psia. An array of base pressure taps on the aft dome gave a clear measurement of the plume recirculation effects at p(infinity) = 4.76 psfa corresponding to 145,000 ft altitude. Our CFD computations of the flow field with direct comparison of computed-versus-measured base pressure distribution (across the dome) provide detailed information on velocities and particle traces as well eddy viscosity in the base and nozzle region. The solution was obtained using a six-zone mesh with 284,000 grid points for one quadrant taking advantage of symmetry. Results are compared using a zero-equation algebraic and a one-equation pointwise R(sub t) turbulence model (work in progress). Good agreement with the experimental pressure data was obtained with both; and this benchmark showed the importance of: (1) proper grid clustering and (2) proper choice of turbulence modeling for rocket plume problems/recirculation at high altitude.

  17. Four-Nozzle Benchmark Wind Tunnel Model USA Code Solutions for Simulation of Multiple Rocket Base Flow Recirculation at 145,000 Feet Altitude

    NASA Technical Reports Server (NTRS)

    Dougherty, N. S.; Johnson, S. L.

    1993-01-01

    Multiple rocket exhaust plume interactions at high altitudes can produce base flow recirculation with attendant alteration of the base pressure coefficient and increased base heating. A search for a good wind tunnel benchmark problem to check grid clustering technique and turbulence modeling turned up the experiment done at AEDC in 1961 by Goethert and Matz on a 4.25-in. diameter domed missile base model with four rocket nozzles. This wind tunnel model with varied external bleed air flow for the base flow wake produced measured p/p(sub ref) at the center of the base as high as 3.3 due to plume flow recirculation back onto the base. At that time in 1961, relatively inexpensive experimentation with air at gamma = 1.4 and nozzle A(sub e)/A of 10.6 and theta(sub n) = 7.55 deg with P(sub c) = 155 psia simulated a LO2/LH2 rocket exhaust plume with gamma = 1.20, A(sub e)/A of 78 and P(sub c) about 1,000 psia. An array of base pressure taps on the aft dome gave a clear measurement of the plume recirculation effects at p(infinity) = 4.76 psfa corresponding to 145,000 ft altitude. Our CFD computations of the flow field with direct comparison of computed-versus-measured base pressure distribution (across the dome) provide detailed information on velocities and particle traces as well eddy viscosity in the base and nozzle region. The solution was obtained using a six-zone mesh with 284,000 grid points for one quadrant taking advantage of symmetry. Results are compared using a zero-equation algebraic and a one-equation pointwise R(sub t) turbulence model (work in progress). Good agreement with the experimental pressure data was obtained with both; and this benchmark showed the importance of: (1) proper grid clustering and (2) proper choice of turbulence modeling for rocket plume problems/recirculation at high altitude.

  18. Proposal and design of Airy-based rocket pulses for invariant propagation in lossy dispersive media.

    PubMed

    Preciado, Miguel A; Sugden, Kate

    2012-12-01

    A novel (to our knowledge) kind of Airy-based pulse with an invariant propagation in lossy dispersive media is proposed. The basic principle is based on an optical energy trade-off between different parts of the pulse caused by the chromatic dispersion, which is used to compensate the attenuation losses of the propagation medium. Although the ideal concept of the proposed pulses implies infinite pulse energy, the numerical simulations show that practical finite energy pulses can be designed to obtain a partially invariant propagation over a finite distance of propagation.

  19. A combined cycle engine test facility

    NASA Astrophysics Data System (ADS)

    Engers, R.; Cresci, D.; Tsai, C.

    Rocket-Based Combined-Cycle (RBCC) engines intended for missiles and/or space launch applications incorporate features of rocket propulsion systems operating in concert with airbreathing engine cycles. Performance evaluation of these types of engines, which are intended to operate from static sea level take-off to supersonic cruise or accerlerate to orbit, requires ground test capabilities which integrate rocket component testing with airbreathing engine testing. A combined cycle engine test facility has been constructed in the General Applied Science Laboratories, Inc. (GASL) Aeropropulsion Test Laboratory to meet this requirement. The facility was designed to support the development of an innovative combined cycle engine concept which features a rocket based ramjet combustor. The test requirements included the ability to conduct tests in which the propulsive force was generated by rocket only, the ramjet only and simultaneous rocket and ramjet power (combined cycle) to evaluate combustor operation over the entire engine cycle. The test facility provides simulation over the flight Mach number range of 0 to 8 and at various trajectories. The capabilities of the combined cycle engine test facility are presented.

  20. A combined cycle engine test facility

    SciTech Connect

    Engers, R.; Cresci, D.; Tsai, C.

    1995-09-01

    Rocket-Based Combined-Cycle (RBCC) engines intended for missiles and/or space launch applications incorporate features of rocket propulsion systems operating in concert with airbreathing engine cycles. Performance evaluation of these types of engines, which are intended to operate from static sea level take-off to supersonic cruise or accerlerate to orbit, requires ground test capabilities which integrate rocket component testing with airbreathing engine testing. A combined cycle engine test facility has been constructed in the General Applied Science Laboratories, Inc. (GASL) Aeropropulsion Test Laboratory to meet this requirement. The facility was designed to support the development of an innovative combined cycle engine concept which features a rocket based ramjet combustor. The test requirements included the ability to conduct tests in which the propulsive force was generated by rocket only, the ramjet only and simultaneous rocket and ramjet power (combined cycle) to evaluate combustor operation over the entire engine cycle. The test facility provides simulation over the flight Mach number range of 0 to 8 and at various trajectories. The capabilities of the combined cycle engine test facility are presented.

  1. Laser diagnostics for small rockets

    NASA Technical Reports Server (NTRS)

    Zupanc, Frank J.; Degroot, Wilhelmus A.

    1993-01-01

    Two nonintrusive flowfield diagnostics based on spectrally-resolved elastic (Rayleigh) and inelastic (Raman) laser light scattering were developed for obtaining local flowfield measurements in low-thrust gaseous H2/O2 rocket engines. The objective is to provide an improved understanding of phenomena occurring in small chemical rockets in order to facilitate the development and validation of advanced computational fluid dynamics (CFD) models for analyzing engine performance. The laser Raman scattering diagnostic was developed to measure major polyatomic species number densities and rotational temperatures in the high-density flowfield region extending from the injector through the chamber throat. Initial application of the Raman scattering diagnostic provided O2 number density and rotational temperature measurements in the exit plane of a low area-ratio nozzle and in the combustion chamber of a two-dimensional, optically-accessible rocket engine. In the low-density nozzle exit plane region where the Raman signal is too weak, a Doppler-resolved laser Rayleigh scattering diagnostic was developed to obtain axial and radial mean gas velocities, and in certain cases, H2O translational temperature and number density. The results from these measurements were compared with theoretical predictions from the RPLUS CFD code for analyzing rocket engine performance. Initial conclusions indicate that a detailed and rigorous modeling of the injector is required in order to make direct comparisons between laser diagnostic measurements and CFD predictions at the local level.

  2. Commercial Development Suborbital Rocket Program

    NASA Technical Reports Server (NTRS)

    1993-01-01

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

  3. Device for installing rocket engines

    NASA Technical Reports Server (NTRS)

    George, T. R., Jr. (Inventor)

    1976-01-01

    A device for installing rocket engines is reported that is supported at a cant relative to vertical by an axially extensible, tiltable pedestal. A lifting platform supports the rocket engine at its thrust chamber exit, including a mount having a concentric base characterized by a concave bearing surface, a plurality of uniformly spaced legs extended radially from the base, and an annular receiver coaxially aligned with the base and affixed to the distal ends of said legs for receiving the thrust chamber exit. The lifting platform rests on a seat concentrically related to the pedestal and affixed to an extended end portion thereof having a convex bearing surface mated in sliding engagement with the concave bearing surface of the annular base for accommodating a rocking motion of the platform.

  4. Aspects of model-based rocket engine condition monitoring and control

    NASA Technical Reports Server (NTRS)

    Karr, Gerald R.; Helmicki, Arthur J.

    1994-01-01

    A rigorous propulsion system modelling method suitable for control and condition monitoring purposes is developed. Previously developed control oriented methods yielding nominal models for gaseous medium propulsion systems are extended to include both nominal and anomalous models for liquid mediums in the following two ways. First, thermodynamic and fluid dynamic properties for liquids such as liquid hydrogen are incorporated into the governing equations. Second, anomalous conditions are captured in ways compatible with existing system theoretic design tools so that anomalous models can be constructed. Control and condition monitoring based methods are seen as an improvement over some existing modelling methods because such methods typically do not rigorously lead to low order models nor do they provide a means for capturing anomalous conditions. Applications to the nominal SSME HPFP and degraded HPFP serve to illustrate the approach.

  5. Environmentally compatible solid rocket propellants

    NASA Technical Reports Server (NTRS)

    Jacox, James L.; Bradford, Daniel J.

    1995-01-01

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

  6. Preliminary tests of silicon carbide based concretes for hybrid rocket nozzles in a solar furnace

    NASA Astrophysics Data System (ADS)

    D'Elia, Raffaele; Bernhart, Gérard; Cutard, Thierry; Peraudeau, Gilles; Balat-Pichelin, Marianne

    2014-06-01

    This research is part of the PERSEUS project, a space program concerning hybrid propulsion and supported by CNES. The main goal of this study is to characterise silicon carbide based micro-concrete with a maximum aggregates size of 800 μm, in a hybrid propulsion environment. The nozzle throat has to resist to a highly oxidising polyethylene (PE)/N2O hybrid environment, under temperatures ranging up to 2980 K. The study is divided into two main parts: the first one deals with the thermo-mechanical characterisation of the material up to 1500 K and the second one with an investigation on the oxidation behaviour in a standard atmosphere, under a solar flux up to 13.5 MW/m2. Young's modulus was determined by resonant frequency method: results show an increase with the stabilisation temperature. Four point bending tests have shown a rupture tensile strength increasing with stabilisation temperature, up to 1473 K. Sintering and densification processes are primary causes of this phenomenon. Visco-plastic behaviour appears at 1373 K, due to the formation of liquid phases in cement ternary system. High-temperature oxidation in ambient air was carried out at PROMES-CNRS laboratory, on a 2 kW solar furnace, with a concentration factor of 15,000. A maximum 13.5 MW/m2 incident solar flux and a 7-90 s exposure times have been chosen. Optical microscopy, SEM, EDS analyses were used to determine the microstructure evolution and the mass loss kinetics. During these tests, silicon carbide undergoes active oxidation with production of SiO and CO smokes and ablation. A linear relation between mass loss and time is found. Oxidation tests performed at 13.5 MW/m2 solar flux have shown a mass loss of 10 mg/cm2 after 15 s. After 90 s, the mass loss reaches 60 mg/cm2. Surface temperature measurement is a main point in this study, because of necessity of a thermo-mechanical-ablative model for the material. Smokes appear at around 5.9 MW/m2, leading to the impossibility of useful temperature

  7. Preliminary guided rocket feasibility study

    NASA Technical Reports Server (NTRS)

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

    1973-01-01

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

  8. Model Rockets and Microchips.

    ERIC Educational Resources Information Center

    Fitzsimmons, Charles P.

    1986-01-01

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

  9. Life Saving Rockets

    NASA Technical Reports Server (NTRS)

    2004-01-01

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

  10. Rockets -- Part II.

    ERIC Educational Resources Information Center

    Leitner, Alfred

    1982-01-01

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

  11. Measuring Model Rocket Acceleration.

    ERIC Educational Resources Information Center

    Jenkins, Randy A.

    1993-01-01

    Presents an experiment that measures the acceleration and velocity of a model rocket. Lift-off information is transmitted to a computer that creates a graph of the velocity. Discusses the analysis of the computer-generated data and differences between calculated and experimental velocity and acceleration of several rocket types. (MDH)

  12. Pattern classification approach to rocket engine diagnostics

    SciTech Connect

    Tulpule, S.

    1989-01-01

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

  13. The Rocket Engine Advancement Program 2 (REAP2)

    NASA Technical Reports Server (NTRS)

    Harper, Brent (Technical Monitor); Hawk, Clark W.

    2004-01-01

    The Rocket Engine Advancement Program (REAP) 2 program is being conducted by a university propulsion consortium consisting of the University of Alabama in Huntsville, Penn State University, Purdue University, Tuskegee University and Auburn University. It has been created to bring their combined skills to bear on liquid rocket combustion stability and thrust chamber cooling. The research team involves well established and known researchers in the propulsion community. The cure team provides the knowledge base, research skills, and commitment to achieve an immediate and continuing impact on present and future propulsion issues. through integrated research teams composed of analysts, diagnosticians, and experimentalists working together in an integrated multi-disciplinary program. This paper provides an overview of the program, its objectives and technical approaches. Research on combustion instability and thrust chamber cooling are being accomplished

  14. High-pressure burning rate studies of solid rocket propellants

    NASA Astrophysics Data System (ADS)

    Atwood, A. I.; Ford, K. P.; Wheeler, C. J.

    2013-03-01

    Increased rocket motor performance is a major driver in the development of solid rocket propellant formulations for chemical propulsion systems. The use of increased operating pressure is an option to improve performance potentially without the cost of reformulation. A technique has been developed to obtain burning rate data across a range of pressures from ambient to 345 MPa. The technique combines the use of a low loading density combustion bomb with a high loading density closed bomb technique. A series of nine ammonium perchlorate (AP) based propellants were used to demonstrate the use of the technique, and the results were compared to the neat AP burning rate "barrier". The effect of plasticizer, oxidizer particle size, catalyst, and binder type were investigated.

  15. Preliminary study of a hydrogen peroxide rocket for use in moving source jet noise tests

    NASA Technical Reports Server (NTRS)

    Plencner, R. M.

    1977-01-01

    A preliminary investigation was made of using a hydrogen peroxide rocket to obtain pure moving source jet noise data. The thermodynamic cycle of the rocket was analyzed. It was found that the thermodynamic exhaust properties of the rocket could be made to match those of typical advanced commercial supersonic transport engines. The rocket thruster was then considered in combination with a streamlined ground car for moving source jet noise experiments. When a nonthrottlable hydrogen peroxide rocket was used to accelerate the vehicle, propellant masses and/or acceleration distances became too large. However, when a throttlable rocket or an auxiliary system was used to accelerate the vehicle, reasonable propellant masses could be obtained.

  16. A3 Subscale Rocket Hot Fire Testing

    NASA Technical Reports Server (NTRS)

    Saunders, G. P.; Yen, J.

    2009-01-01

    This paper gives a description of the methodology and results of J2-X Subscale Simulator (JSS) hot fire testing supporting the A3 Subscale Diffuser Test (SDT) project at the E3 test facility at Stennis Space Center, MS (SSC). The A3 subscale diffuser is a geometrically accurate scale model of the A3 altitude simulating rocket test facility. This paper focuses on the methods used to operate the facility and obtain the data to support the aerodynamic verification of the A3 rocket diffuser design and experimental data quantifying the heat flux throughout the facility. The JSS was operated at both 80% and 100% power levels and at gimbal angle from 0 to 7 degrees to verify the simulated altitude produced by the rocket-rocket diffuser combination. This was done with various secondary GN purge loads to quantify the pumping performance of the rocket diffuser. Also, special tests were conducted to obtain detailed heat flux measurements in the rocket diffuser at various gimbal angles and in the facility elbow where the flow turns from vertical to horizontal upstream of the 2nd stage steam ejector.

  17. Devising rocket power for smaller engines

    SciTech Connect

    Burruss, R.

    1996-04-01

    Compact, high-power engines that burn fuel and oxygen could be made by winding copper tubing in a helix around boiler sections. With more than 1,000 horsepower per pound of engine weight, liquid-fueled rockets have the highest specific power of any engines designed for sustained operation. Yet those engines generally run for about only 1,000 seconds--nowhere near the sustained operation time for lower-power automotive and aircraft engines of more than 1,000 hours. In theory, at least, a fuel/oxygen rocket can be built that combines the best of both classes: high specific power (from perhaps two to 10 times that of a gas turbine) and a 1,000-hour service life. Such an engine would almost certainly be possible if the rocket`s exhaust gases could be simultaneously cooled and expanded by mixing water with the rocket`s exhaust and boiling it before it reaches the turbine. The technology itself is not new. variations of these rocket-turbine-type engines, for example, powered torpedoes during World War I. Some 30 years later, German V-2 rockets used fuel pumps, driven by the reaction of hydrogen peroxide with hydrocarbon fuels, to produce high-pressure steam that was directed against a turbine. Alternatively, fuel/oxygen combustion could produce steam to drive a piston engine. Either way, the challenge remains to construct a compact, long-service-life, high-specific-power boiler that burns fuel and oxygen. The new type of engine could be derived from recent research on electric vehicles (EVs).

  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. Indians Repulse British With Rocket

    NASA Technical Reports Server (NTRS)

    2004-01-01

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

  20. 1. ROCKET ENGINE TEST STAND, LOCATED IN THE NORTHEAST ¼ ...

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

    1. ROCKET ENGINE TEST STAND, LOCATED IN THE NORTHEAST ¼ OF THE X-15 ENGINE TEST COMPLEX. Looking northeast. - Edwards Air Force Base, X-15 Engine Test Complex, Rocket Engine & Complete X-15 Vehicle Test Stands, Rogers Dry Lake, east of runway between North Base & South Base, Boron, Kern County, CA

  1. 2. ROCKET ENGINE TEST STAND, SHOWING TANK (BUILDING 1929) AND ...

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

    2. ROCKET ENGINE TEST STAND, SHOWING TANK (BUILDING 1929) AND GARAGE (BUILDING 1930) AT LEFT REAR. Looking to west. - Edwards Air Force Base, X-15 Engine Test Complex, Rocket Engine & Complete X-15 Vehicle Test Stands, Rogers Dry Lake, east of runway between North Base & South Base, Boron, Kern County, CA

  2. GPS Sounding Rocket Developments

    NASA Technical Reports Server (NTRS)

    Bull, Barton

    1999-01-01

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

  3. Liquid propellant rockets.

    NASA Technical Reports Server (NTRS)

    Dipprey, D. F.

    1972-01-01

    A brief overview of the state of knowledge in liquid rocket technology is presented and examples are provided of instances where some fundamental principles of chemistry, fluid mechanics, and mathematics can be applied. A liquid propellant rocket classification is discussed together with rocket system performance, applications for liquid propellants, the effective exhaust velocity, aspects of simplified nozzle expansion, questions about theoretical propellant performance, the effect of chamber pressure on equilibrium performance, and the kinetic recombination in nozzles. Details of propellant combustion are examined, giving attention to propellant injection, evaporation-controlled combustion, combustion instability, and monopropellant decomposition.

  4. Relativistic rocket: Dream and reality

    NASA Astrophysics Data System (ADS)

    Semyonov, Oleg G.

    2014-06-01

    The dream of interstellar flights persists since the first pioneers in astronautics and has never died. Many concepts of thruster capable to propel a rocket to the stars have been proposed and the most suitable among them are thought to be photon propulsion and propulsion by the products of proton-antiproton annihilation in magnetic nozzle. This article addresses both concepts allowing for cross-section of annihilation among other issues in order to show their vulnerability and to indicate the problems. The concept of relativistic matter propulsion is substantiated and discussed. The latter is argued to be the most straightforward way to build-up a relativistic rocket firstly because it is based on the existing technology of ion generators and accelerators and secondly because it can be stepped up in efflux power starting from interplanetary spacecrafts powered by nuclear reactors to interstellar starships powered by annihilation reactors. The problems imposed by thermodynamics and heat disposal are accentuated.

  5. Investigation of the Rocket Induced Flow Field in a Rectangular Duct

    NASA Technical Reports Server (NTRS)

    Landrum, D. Brian; Thames, Mignon; Parkinson, Doug; Gautney, Serena; Hawk, Clark

    1999-01-01

    Several tests were performed on a one-sixth scale Rocket Based Combined Cycle (RBCC) engine model at the University of Alabama in Huntsville. The UAH RBCC facility consists of a rectangular duct with a vertical strut mounted in the center. The scaled strut consists of two supersonic rocket nozzles with an embedded vertical turbine between the rocket nozzles. The tests included mass flow, flow visualization and horizontal pressure traverses. The mass flow test indicated a c:hoked condition when the rocket chamber pressure is between 200 psi and 300 psi. The flow visualization tests narrowed the rocket chamber pressure range from, 250 psi to 300 psi. Also, from this t.est, an assumption of a minimum

  6. Nitrous Oxide/Paraffin Hybrid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Zubrin, Robert; Snyder, Gary

    2010-01-01

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

  7. The History of Rockets.

    ERIC Educational Resources Information Center

    Newby, J. C.

    1988-01-01

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

  8. Rocket engine numerical simulation

    NASA Technical Reports Server (NTRS)

    Davidian, Ken

    1993-01-01

    The topics are presented in view graph form and include the following: a definition of the rocket engine numerical simulator (RENS); objectives; justification; approach; potential applications; potential users; RENS work flowchart; RENS prototype; and conclusions.

  9. Rocket engine numerical simulator

    NASA Technical Reports Server (NTRS)

    Davidian, Ken

    1993-01-01

    The topics are presented in viewgraph form and include the following: a rocket engine numerical simulator (RENS) definition; objectives; justification; approach; potential applications; potential users; RENS work flowchart; RENS prototype; and conclusion.

  10. Rocketing into Adaptive Inquiry.

    ERIC Educational Resources Information Center

    Farenga, Stephen J.; Joyce, Beverly A.; Dowling, Thomas W.

    2002-01-01

    Defines adaptive inquiry and argues for employing this method which allows lessons to be shaped in response to student needs. Illustrates this idea by detailing an activity in which teams of students build rockets. (DDR)

  11. Antares Rocket Lifts Off!

    NASA Video Gallery

    NASA commercial space partner Orbital Sciences Corp. of Dulles, Va., launched its Cygnus cargo spacecraft aboard its Antares rocket at 10:58 a.m. EDT Wednesday from the Mid-Atlantic Regional Spacep...

  12. A Technology Pathway for Airbreathing, Combined-Cycle, Horizontal Space Launch Through SR-71 Based Trajectory Modeling

    NASA Technical Reports Server (NTRS)

    Kloesel, Kurt J.; Ratnayake, Nalin A.; Clark, Casie M.

    2011-01-01

    Access to space is in the early stages of commercialization. Private enterprises, mainly under direct or indirect subsidy by the government, have been making headway into the LEO launch systems infrastructure, of small-weight-class payloads of approximately 1000 lbs. These moderate gains have emboldened the launch industry and they are poised to move into the middle-weight class (roughly 5000 lbs). These commercially successful systems are based on relatively straightforward LOX-RP, two-stage, bi-propellant rocket technology developed by the government 40 years ago, accompanied by many technology improvements. In this paper we examine a known generic LOX-RP system with the focus on the booster stage (1st stage). The booster stage is then compared to modeled Rocket-Based and Turbine-Based Combined Cycle booster stages. The air-breathing propulsion stages are based on/or extrapolated from known performance parameters of ground tested RBCC (the Marquardt Ejector Ramjet) and TBCC (the SR-71/J-58 engine) data. Validated engine models using GECAT and SCCREAM are coupled with trajectory optimization and analysis in POST-II to explore viable launch scenarios using hypothetical aerospaceplane platform obeying the aerodynamic model of the SR-71. Finally, and assessment is made of the requisite research technology advances necessary for successful commercial and government adoption of combined-cycle engine systems for space access.

  13. Rocket Motor Microphone Investigation

    NASA Technical Reports Server (NTRS)

    Pilkey, Debbie; Herrera, Eric; Gee, Kent L.; Giraud, Jerom H.; Young, Devin J.

    2010-01-01

    At ATK's facility in Utah, large full-scale solid rocket motors are tested. The largest is a five-segment version of the reusable solid rocket motor, which is for use on the Ares I launch vehicle. As a continuous improvement project, ATK and BYU investigated the use of microphones on these static tests, the vibration and temperature to which the instruments are subjected, and in particular the use of vent tubes and the effects these vents have at low frequencies.

  14. GPS Sounding Rocket Developments

    NASA Technical Reports Server (NTRS)

    Bull, Barton

    1999-01-01

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

  15. GREECE Sounding Rocket Mission Overview

    NASA Astrophysics Data System (ADS)

    Samara, M.; Michell, R.; Grubbs, G. A., II; Bonnell, J. W.; Ogasawara, K.; Hampton, D. L.; Jahn, J. M.; Donovan, E.; Gustavsson, B.; Lanchester, B. S.; McHarg, M. G.; Spanswick, E.; Trondsen, T. S.; Valek, P. W.

    2014-12-01

    On 03 March 2014 at 11:09:50 UT the Ground-to-Rocket Electrodynamics-Electrons Correlative Experiment (GREECE) sounding rocket successfully launched from Poker Flat, Alaska . It reached an apogee of approximately 335 km over the native village of Venetie during a dynamic post-midnight auroral event. A wide range of precipitating electrons were measured with the Acute Precipitating Electron Spectrometer (APES) and Medium-energy Electron SPectrometer (MESP), cumulatively covering 300 ev to 200 keV in varying time resolutions. DC to low frequency electric and magnetic fields were measured at the same time and a langmuir probe was also employed. In addition to the on board instrumentation a suite of ground based imagers was deployed under apogee. We used several electron-multiplying charge-coupled devices (EMCCDs) with different filters and field of views imaging along magnetic zenith. This yielded multi-emission line information about the auroral brightness at the magnetic footprint of the rocket critical for our main goal of exploring the correlation of the sheer flows often observed in high resolution imagery during aurora and the in situ signatures of precipitating particles and waves. The instruments used will be discussed in further detail along with preliminary results of an event rich in particle and wave signatures.

  16. Highly Reusable Space Transportation (HRST) Baseline Concepts and Analysis: Rocket/RBCC Options. Part 2; A Comparative Study

    NASA Technical Reports Server (NTRS)

    Woodcock, Gordon

    1997-01-01

    This study is an extension of a previous effort by the Principal Investigator to develop baseline data to support comparative analysis of Highly Reusable Space Transportation (HRST) concepts. The analyses presented herin develop baseline data bases for two two-stage-to-orbit (TSTO) concepts: (1) Assisted horizontal take-off all rocket (assisted HTOHL); and (2) Assisted vertical take-off rocket based combined cycle (RBCC). The study objectives were to: (1) Provide configuration definitions and illustrations for assisted HTOHL and assisted RBCC; (2) Develop a rationalization approach and compare these concepts with the HRST reference; and (3) Analyze TSTO configurations which try to maintain SSTO benefits while reducing inert weight sensitivity.

  17. Rocket-borne and ground-based measurements in support of the field-widened interferometer experiment - sergeant a30. 276

    SciTech Connect

    Ulwick, J.C.; Allred, G.D.; Baker, K.D.; Howlett, L.C.

    1985-05-28

    In April 1983 Utah State University and Air Force Geophysics Laboratory experimenters launched a Sergeant (A30.276) sounding rocket from the Poker Flat Research Range, Alaska. The prime purpose of the flight was to obtain infrared-spectral measurements in the 2-1.5 micrometer m range during an auroral event. In addition to the prime experiment, which has already been reported, the payload contained four photometers, and energy deposition scintillator and an atomic oxygen detector to gather in-situ supporting data. Simultaneously, all-sky television, meridian scanning photometers, riometer, and magnetometers supported the flight from ground-based measuring sites. This report presents a summary of the rocketborne supporting instruments and the data they gathered and provides a time/intensity history of the event as documented by the ground-based meridian scanners and all-sky television.

  18. CODEX sounding rocket wire grid collimator design

    NASA Astrophysics Data System (ADS)

    Shipley, Ann; Zeiger, Ben; Rogers, Thomas

    2011-05-01

    CODEX is a sounding rocket payload designed to operate in the soft x-ray (0.1-1.0 kV) regime. The instrument has a 3.25 degree square field of view that uses a one meter long wire grid collimator to create a beam that converges to a line in the focal plane. Wire grid collimator performance is directly correlated to the geometric accuracy of actual grid features and their relative locations. Utilizing a strategic combination of manufacturing and assembly techniques, this design is engineered for precision within the confines of a typical rocket budget. Expected resilience of the collimator under flight conditions is predicted by mechanical analysis.

  19. Iridium/Rhenium Parts For Rocket Engines

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.; Harding, John T.; Wooten, John R.

    1991-01-01

    Oxidation/corrosion of metals at high temperatures primary life-limiting mechanism of parts in rocket engines. Combination of metals greatly increases operating temperature and longevity of these parts. Consists of two transition-element metals - iridium and rhenium - that melt at extremely high temperatures. Maximum operating temperature increased to 2,200 degrees C from 1,400 degrees C. Increases operating lifetimes of small rocket engines by more than factor of 10. Possible to make hotter-operating, longer-lasting components for turbines and other heat engines.

  20. Fiber-Reinforced Superalloys For Rocket Engines

    NASA Technical Reports Server (NTRS)

    Lewis, Jack R.; Yuen, Jim L.; Petrasek, Donald W.; Stephens, Joseph R.

    1990-01-01

    Report discusses experimental studies of fiber-reinforced superalloy (FRS) composite materials for use in turbine blades in rocket engines. Intended to withstand extreme conditions of high temperature, thermal shock, atmospheres containing hydrogen, high cycle fatigue loading, and thermal fatigue, which tax capabilities of even most-advanced current blade material - directionally-solidified, hafnium-modified MAR M-246 {MAR M-246 (Hf) (DS)}. FRS composites attractive combination of properties for use in turbopump blades of advanced rocket engines at temperatures from 870 to 1,100 degrees C.

  1. Nuclear Rocket Technology Conference

    NASA Technical Reports Server (NTRS)

    1966-01-01

    The Lewis Research Center has a strong interest in nuclear rocket propulsion and provides active support of the graphite reactor program in such nonnuclear areas as cryogenics, two-phase flow, propellant heating, fluid systems, heat transfer, nozzle cooling, nozzle design, pumps, turbines, and startup and control problems. A parallel effort has also been expended to evaluate the engineering feasibility of a nuclear rocket reactor using tungsten-matrix fuel elements and water as the moderator. Both of these efforts have resulted in significant contributions to nuclear rocket technology. Many successful static firings of nuclear rockets have been made with graphite-core reactors. Sufficient information has also been accumulated to permit a reasonable Judgment as to the feasibility of the tungsten water-moderated reactor concept. We therefore consider that this technoIogy conference on the nuclear rocket work that has been sponsored by the Lewis Research Center is timely. The conference has been prepared by NASA personnel, but the information presented includes substantial contributions from both NASA and AEC contractors. The conference excludes from consideration the many possible mission requirements for nuclear rockets. Also excluded is the direct comparison of nuclear rocket types with each other or with other modes of propulsion. The graphite reactor support work presented on the first day of the conference was partly inspired through a close cooperative effort between the Cleveland extension of the Space Nuclear Propulsion Office (headed by Robert W. Schroeder) and the Lewis Research Center. Much of this effort was supervised by Mr. John C. Sanders, chairman for the first day of the conference, and by Mr. Hugh M. Henneberry. The tungsten water-moderated reactor concept was initiated at Lewis by Mr. Frank E. Rom and his coworkers. The supervision of the recent engineering studies has been shared by Mr. Samuel J. Kaufman, chairman for the second day of the

  2. Low thrust chemical rocket technology

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.

    1992-01-01

    An on-going technology program to improve the performance of low thrust chemical rockets for spacecraft on-board propulsion applications is reviewed. Improved performance and lifetime is sought by the development of new predictive tools to understand the combustion and flow physics, introduction of high temperature materials and improved component designs to optimize performance, and use of higher performance propellants. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Predictions are based on both the RPLUS Navier-Stokes code with finite rate kinetics and the JANNAF methodology. Data were obtained with laser-based diagnostics along with global performance measurements. Results indicate that the modeling of the injector and the combustion process needs improvement in these codes and flow visualization with a technique such as 2-D laser induced fluorescence (LIF) would aid in resolving issues of flow symmetry and shear layer combustion processes. High temperature material fabrication processes are under development and small rockets are being designed, fabricated, and tested using these new materials. Rhenium coated with iridium for oxidation protection was produced by the Chemical Vapor Deposition (CVD) process and enabled an 800 K increase in rocket operating temperature. Performance gains with this material in rockets using Earth storable propellants (nitrogen tetroxide and monomethylhydrazine or hydrazine) were obtained through component redesign to eliminate fuel film cooling and its associated combustion inefficiency while managing head end thermal soakback. Material interdiffusion and oxidation characteristics indicated that the requisite lifetimes of tens of hours were available for thruster applications. Rockets were designed, fabricated, and tested with thrusts of 22, 62, 440 and 550 N. Performance improvements of 10 to 20 seconds specific impulse were demonstrated. Higher

  3. Mechanical and Combustion Performance of Multi-Walled Carbon Nanotubes as an Additive to Paraffin-Based Solid Fuels for Hybrid Rockets

    NASA Technical Reports Server (NTRS)

    Larson, Daniel B.; Boyer, Eric; Wachs, Trevor; Kuo, Kenneth, K.; Koo, Joseph H.; Story, George

    2012-01-01

    Paraffin-based solid fuels for hybrid rocket motor applications are recognized as a fastburning alternative to other fuel binders such as HTPB, but efforts to further improve the burning rate and mechanical properties of paraffin are still necessary. One approach that is considered in this study is to use multi-walled carbon nanotubes (MWNT) as an additive to paraffin wax. Carbon nanotubes provide increased electrical and thermal conductivity to the solid-fuel grains to which they are added, which can improve the mass burning rate. Furthermore, the addition of ultra-fine aluminum particles to the paraffin/MWNT fuel grains can enhance regression rate of the solid fuel and the density impulse of the hybrid rocket. The multi-walled carbon nanotubes also present the possibility of greatly improving the mechanical properties (e.g., tensile strength) of the paraffin-based solid-fuel grains. For casting these solid-fuel grains, various percentages of MWNT and aluminum particles will be added to the paraffin wax. Previous work has been published about the dispersion and mixing of carbon nanotubes.1 Another manufacturing method has been used for mixing the MWNT with a phenolic resin for ablative applications, and the manufacturing and mixing processes are well-documented in the literature.2 The cost of MWNT is a small fraction of single-walled nanotubes. This is a scale-up advantage as future applications and projects will require low cost additives to maintain cost effectiveness. Testing of the solid-fuel grains will be conducted in several steps. Dog bone samples will be cast and prepared for tensile testing. The fuel samples will also be analyzed using thermogravimetric analysis and a high-resolution scanning electron microscope (SEM). The SEM will allow for examination of the solid fuel grain for uniformity and consistency. The paraffin-based fuel grains will also be tested using two hybrid rocket test motors located at the Pennsylvania State University s High Pressure

  4. Sound from apollo rockets in space.

    PubMed

    Cotten, D; Donn, W L

    1971-02-12

    Low-frequency sound has been recorded on at least two occasions in Bermuda with the passage of Apollo rocket vehicles 188 kilometers aloft. The signals, which are reminiscent of N-waves from sonic booms, are (i) horizontally coherent; (ii) have extremely high (supersonic) trace velocities across the tripartite arrays; (iii) have nearly identical appearance and frequencies; (iv) have essentially identical arrival times after rocket launch; and (v) are the only coherent signals recorded over many hours. These observations seem to establish that the recorded sound comes from the rockets at high elevation. Despite this high elevation, the values of surface pressure appear to be explainable on the basis of a combination of a kinetic theory approach to shock formation in rarefied atmospheres with established gas-dynamics shock theory.

  5. Liquid rocket engine nozzles

    NASA Technical Reports Server (NTRS)

    1976-01-01

    The nozzle is a major component of a rocket engine, having a significant influence on the overall engine performance and representing a large fraction of the engine structure. The design of the nozzle consists of solving simultaneously two different problems: the definition of the shape of the wall that forms the expansion surface, and the delineation of the nozzle structure and hydraulic system. This monography addresses both of these problems. The shape of the wall is considered from immediately upstream of the throat to the nozzle exit for both bell and annular (or plug) nozzles. Important aspects of the methods used to generate nozzle wall shapes are covered for maximum-performance shapes and for nozzle contours based on criteria other than performance. The discussion of structure and hydraulics covers problem areas of regeneratively cooled tube-wall nozzles and extensions; it treats also nozzle extensions cooled by turbine exhaust gas, ablation-cooled extensions, and radiation-cooled extensions. The techniques that best enable the designer to develop the nozzle structure with as little difficulty as possible and at the lowest cost consistent with minimum weight and specified performance are described.

  6. Improved hybrid rocket fuel

    NASA Technical Reports Server (NTRS)

    Dean, David L.

    1995-01-01

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

  7. Rocket noise - A review

    NASA Astrophysics Data System (ADS)

    McInerny, S. A.

    1990-10-01

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

  8. ROCKET PORT CLOSURE

    DOEpatents

    Mattingly, J.T.

    1963-02-12

    This invention provides a simple pressure-actuated closure whereby windowless observation ports are opened to the atmosphere at preselected altitudes. The closure comprises a disk which seals a windowless observation port in rocket hull. An evacuated instrument compartment is affixed to the rocket hull adjacent the inner surface of the disk, while the outer disk surface is exposed to the atmosphere through which the rocket is traveling. The pressure differential between the evacuated instrument compartment and the relatively high pressure external atmosphere forces the disk against the edge of the observation port, thereby effecting a tight seai. The instrument compartment is evacuated to a pressure equal to the atmospheric pressure existing at the altitude at which it is desiretl that the closure should open. When the rocket reaches this preselected altitude, the inwardly directed atmospheric force on the disk is just equaled by the residual air pressure force within the instrument compartment. Consequently, the closure disk falls away and uncovers the open observation port. The separation of the disk from the rocket hull actuates a switch which energizes the mechanism of a detecting instrument disposed within the instrument compartment. (AE C)

  9. 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

  10. 7. ROCKET SLED ON DECK OF TEST STAND 15. Photo ...

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

    7. ROCKET SLED ON DECK OF TEST STAND 1-5. Photo no. "6085, G-EAFB-16 SEP 52." Looking south to machine shop. - Edwards Air Force Base, Air Force Rocket Propulsion Laboratory, Test Stand 1-5, Test Area 1-115, northwest end of Saturn Boulevard, Boron, Kern County, CA

  11. Rockets in World War I

    NASA Technical Reports Server (NTRS)

    2004-01-01

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

  12. Finite area combustor theoretical rocket performance

    NASA Technical Reports Server (NTRS)

    Gordon, Sanford; Mcbride, Bonnie J.

    1988-01-01

    Previous to this report, the computer program of NASA SP-273 and NASA TM-86885 was capable of calculating theoretical rocket performance based only on the assumption of an infinite area combustion chamber (IAC). An option was added to this program which now also permits the calculation of rocket performance based on the assumption of a finite area combustion chamber (FAC). In the FAC model, the combustion process in the cylindrical chamber is assumed to be adiabatic, but nonisentropic. This results in a stagnation pressure drop from the injector face to the end of the chamber and a lower calculated performance for the FAC model than the IAC model.

  13. Empirical Scaling Laws of Rocket Exhaust Cratering

    NASA Technical Reports Server (NTRS)

    Donahue, Carly M.; Metzger, Philip T.; Immer, Christopher D.

    2005-01-01

    When launching or landing a space craft on the regolith of a terrestrial surface, special attention needs to be paid to the rocket exhaust cratering effects. If the effects are not controlled, the rocket cratering could damage the spacecraft or other surrounding hardware. The cratering effects of a rocket landing on a planet's surface are not understood well, especially for the lunar case with the plume expanding in vacuum. As a result, the blast effects cannot be estimated sufficiently using analytical theories. It is necessary to develop physics-based simulation tools in order to calculate mission-essential parameters. In this work we test out the scaling laws of the physics in regard to growth rate of the crater depth. This will provide the physical insight necessary to begin the physics-based modeling.

  14. Performance Charts for Multistage Rocket Boosters

    NASA Technical Reports Server (NTRS)

    MacKay, John S.; Weber, Richard J.

    1961-01-01

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

  15. MK 66 Rocket Signature Reduction

    DTIC Science & Technology

    1982-04-01

    Indian Head, Maryland. ’The objec- tive of the study was to reduce the visible signature of the rocket motor. The rocket motor used for demonstration tests...15 6. Actual Emmiissions . . . . . . ........... . 16 7. Human Eye Adjusted Emmissions ..................... .. 16 8. Cross...altered. Additives are commonly used in gun propellants for elimination of muzzle flash. Their use in tactical rockets has been very limited, and

  16. Hermes A-1 Test Rockets

    NASA Technical Reports Server (NTRS)

    1950-01-01

    The first Hermes A-1 test rocket was fired at White Sand Proving Ground (WSPG). Hermes was a modified V-2 German rocket, utilizing the German aerodynamic configuration; however, internally it was a completely new design. Although it did not result in an operational vehicle, the information that was gathered in the process contributed directly to the development of the Redstone rocket.

  17. Overview of rocket engine control

    NASA Technical Reports Server (NTRS)

    Lorenzo, Carl F.; Musgrave, Jeffrey L.

    1991-01-01

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

  18. Rocket center Peenemuende - Personal memories

    NASA Technical Reports Server (NTRS)

    Dannenberg, Konrad; Stuhlinger, Ernst

    1993-01-01

    A brief history of Peenemuende, the rocket center where Von Braun and his team developed the A-4 (V-2) rocket under German Army auspices, and the Air Force developed the V-1 (buzz bomb), wire-guided bombs, and rocket planes, is presented. Emphasis is placed on the expansion of operations beginning in 1942.

  19. Solid Rocket Booster Recovery

    NASA Technical Reports Server (NTRS)

    1982-01-01

    The towing ship, Liberty, towed a recovered solid rocket booster (SRB) for the STS-5 mission to Port Canaveral, Florida. The recovered SRB would be inspected and refurbished for reuse. The Shuttle's SRB's and solid rocket motors (SRM's) are the largest ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds. The requirement for reusability dictated durable materials and construction to preclude corrosion of the hardware exposed to the harsh seawater environment. The SRB contains a complete recovery subsystem that includes parachutes, beacons, lights, and tow fixture.

  20. Solid Rocket Booster Recovery

    NASA Technical Reports Server (NTRS)

    1982-01-01

    The towing ship, Liberty, towed a recovered solid rocket booster (SRB) for the STS-3 mission to Port Canaveral, Florida. The recovered SRB would be inspected and refurbished for reuse. The Shuttle's SRB's and solid rocket motors (SRM's) are the largest ever built and the first designed for refurbishment and reuse. Standing nearly 150-feet high, the twin boosters provide the majority of thrust for the first two minutes of flight, about 5.8 million pounds. The requirement for reusability dictated durable materials and construction to preclude corrosion of the hardware exposed to the harsh seawater environment. The SRB contains a complete recovery subsystem that includes parachutes, beacons, lights, and tow fixture.

  1. Jupiter Rocket Engine

    NASA Technical Reports Server (NTRS)

    2004-01-01

    Engine for the Jupiter rocket. The Jupiter vehicle was a direct derivative of the Redstone. The Army Ballistic Missile Agency (ABMA) at Redstone Arsenal, Alabama, continued Jupiter development into a successful intermediate ballistic missile, even though the Department of Defense directed its operational development to the Air Force. ABMA maintained a role in Jupiter RD, including high-altitude launches that added to ABMA's understanding of rocket vehicle operations in the near-Earth space environment. It was knowledge that paid handsome dividends later.

  2. Advanced liquid rockets

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.

    1992-01-01

    A program to substitute iridium coated rhenium for silicide coated niobium in thrust chamber fabrications is reviewed. The life limiting phenomena in each of these material systems is also reviewed. Coating cracking and spalling is not a problem with iridium-coated rhenium as in silicide-coated niobium. Use of the new material system enables an 800 K increase in thruster operating temperature from around 1700 K for niobium to 2500 K for rhenium. Specific impulse iridium-coated rhenium rockets is nominally 20 seconds higher than comparable niobium rockets in the 22 N class and nominally 10 seconds higher in the 440 N class.

  3. 24 Inch Reusable Solid Rocket Motor Test

    NASA Technical Reports Server (NTRS)

    2002-01-01

    A scaled-down 24-inch version of the Space Shuttle's Reusable Solid Rocket Motor was successfully fired for 21 seconds at a Marshall Space Flight Center (MSFC) Test Stand. The motor was tested to ensure a replacement material called Lycocel would meet the criteria set by the Shuttle's Solid Motor Project Office. The current material is a heat-resistant, rayon-based, carbon-cloth phenolic used as an insulating material for the motor's nozzle. Lycocel, a brand name for Tencel, is a cousin to rayon and is an exceptionally strong fiber made of wood pulp produced by a special 'solvent-spirning' process using a nontoxic solvent. It will also be impregnated with a phenolic resin. This new material is expected to perform better under the high temperatures experienced during launch. The next step will be to test the material on a 48-inch solid rocket motor. The test, which replicates launch conditions, is part of Shuttle's ongoing verification of components, materials, and manufacturing processes required by MSFC, which oversees the Reusable Solid Rocket Motor project. Manufactured by the ATK Thiokol Propulsion Division in Promontory, California, the Reusable Solid Rocket Motor measures 126 feet (38.4 meters) long and 12 feet (3.6 meters) in diameter. It is the largest solid rocket motor ever flown and the first designed for reuse. During its two-minute burn at liftoff, each motor generates an average thrust of 2.6 million pounds (1.2 million kilograms).

  4. 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.

  5. Genetic Algorithm Optimization of a Cost Competitive Hybrid Rocket Booster

    NASA Technical Reports Server (NTRS)

    Story, George

    2014-01-01

    Performance, reliability and cost have always been drivers in the rocket business. Hybrid rockets have been late entries into the launch business due to substantial early development work on liquid rockets and later on solid rockets. Slowly the technology readiness level of hybrids has been increasing due to various large scale testing and flight tests of hybrid rockets. A remaining issue is the cost of hybrids vs the existing launch propulsion systems. This paper will review the known state of the art hybrid development work to date and incorporate it into a genetic algorithm to optimize the configuration based on various parameters. A cost module will be incorporated to the code based on the weights of the components. The design will be optimized on meeting the performance requirements at the lowest cost.

  6. Genetic Algorithm Optimization of a Cost Competitive Hybrid Rocket Booster

    NASA Technical Reports Server (NTRS)

    Story, George

    2015-01-01

    Performance, reliability and cost have always been drivers in the rocket business. Hybrid rockets have been late entries into the launch business due to substantial early development work on liquid rockets and solid rockets. Slowly the technology readiness level of hybrids has been increasing due to various large scale testing and flight tests of hybrid rockets. One remaining issue is the cost of hybrids versus the existing launch propulsion systems. This paper will review the known state-of-the-art hybrid development work to date and incorporate it into a genetic algorithm to optimize the configuration based on various parameters. A cost module will be incorporated to the code based on the weights of the components. The design will be optimized on meeting the performance requirements at the lowest cost.

  7. Cooled Ceramic Composite Panel Tested Successfully in Rocket Combustion Facility

    NASA Technical Reports Server (NTRS)

    Jaskowiak, Martha H.

    2003-01-01

    Regeneratively cooled ceramic matrix composite (CMC) structures are being considered for use along the walls of the hot-flow paths of rocket-based or turbine-based combined-cycle propulsion systems. They offer the combined benefits of substantial weight savings, higher operating temperatures, and reduced coolant requirements in comparison to components designed with traditional metals. These cooled structures, which use the fuel as the coolant, require materials that can survive aggressive thermal, mechanical, acoustic, and aerodynamic loads while acting as heat exchangers, which can improve the efficiency of the engine. A team effort between the NASA Glenn Research Center, the NASA Marshall Space Flight Center, and various industrial partners has led to the design, development, and fabrication of several types of regeneratively cooled panels. The concepts for these panels range from ultra-lightweight designs that rely only on CMC tubes for coolant containment to more maintainable designs that incorporate metal coolant containment tubes to allow for the rapid assembly or disassembly of the heat exchanger. One of the cooled panels based on an all-CMC design was successfully tested in the rocket combustion facility at Glenn. Testing of the remaining four panels is underway.

  8. Scaled Rocket Testing in Hypersonic Flow

    NASA Technical Reports Server (NTRS)

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

    2015-01-01

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

  9. Vertical Landing Aerodynamics of Reusable Rocket Vehicle

    NASA Astrophysics Data System (ADS)

    Nonaka, Satoshi; Nishida, Hiroyuki; Kato, Hiroyuki; Ogawa, Hiroyuki; Inatani, Yoshifumi

    The aerodynamic characteristics of a vertical landing rocket are affected by its engine plume in the landing phase. The influences of interaction of the engine plume with the freestream around the vehicle on the aerodynamic characteristics are studied experimentally aiming to realize safe landing of the vertical landing rocket. The aerodynamic forces and surface pressure distributions are measured using a scaled model of a reusable rocket vehicle in low-speed wind tunnels. The flow field around the vehicle model is visualized using the particle image velocimetry (PIV) method. Results show that the aerodynamic characteristics, such as the drag force and pitching moment, are strongly affected by the change in the base pressure distributions and reattachment of a separation flow around the vehicle.

  10. Program Computes Sound Pressures at Rocket Launches

    NASA Technical Reports Server (NTRS)

    Ogg, Gary; Heyman, Roy; White, Michael; Edquist, Karl

    2005-01-01

    Launch Vehicle External Sound Pressure is a computer program that predicts the ignition overpressure and the acoustic pressure on the surfaces and in the vicinity of a rocket and launch pad during launch. The program generates a graphical user interface (GUI) that gathers input data from the user. These data include the critical dimensions of the rocket and of any launch-pad structures that may act as acoustic reflectors, the size and shape of the exhaust duct or flame deflector, and geometrical and operational parameters of the rocket engine. For the ignition-overpressure calculations, histories of the chamber pressure and mass flow rate also are required. Once the GUI has gathered the input data, it feeds them to ignition-overpressure and launch-acoustics routines, which are based on several approximate mathematical models of distributed sources, transmission, and reflection of acoustic waves. The output of the program includes ignition overpressures and acoustic pressures at specified locations.

  11. Parameter identification for free flight rockets

    NASA Astrophysics Data System (ADS)

    Granot, R.; Sylman, Y.

    1982-08-01

    The accuracy of a rocket weapon system is largely determined by the accuracy with which the mean impact point can be predicted. The prediction procedure involves a preflight trajectory computation, which is based on the values of the measured meteorological variables, the physical characteristics of the system (launcher and rocket), and the location of the target. Granot and Sylman (1981) have discussed the advantages of using for the computation an optimization technique which deals with device-measured quantities. The present investigation considers the feasibility of a conduction of direct measurements about the burnout velocity region as a means for parameter reduction in a free flight rocket model. Attention is given to a mathematical model, effective forces, aspects of instrumentation, the calibration procedure, a calibration algorithm, and a data reduction method.

  12. The Norwegian Sounding Rocket and Balloon Program

    NASA Astrophysics Data System (ADS)

    Skatteboe, Rolf

    2001-08-01

    The status and recent developments of the Norwegian Sounding Rocket and Balloon Program are presented with focus on national activities and recent achievements. The main part of the Norwegian program is sounding rocket launches conducted by Andøya Rocket Range from the launch facilities on Andøya and at Svalbard. For the majority of the programs, the scientific goal is investigation of processes in the middle and upper atmosphere. The in situ measurements are supplemented by a large number of ground-based support instruments located at the ALOMAR Observatory. The ongoing and planned projects are described and the highlights of the latest completed projects are given. The scientific program for the period 2001-2003 will be reviewed. Several new programs have been started to improve the services available to the international science comunity. The Hotel Payload project and MiniDusty are important examples that will be introduced in the paper. Available space related infrastructure is summarized.

  13. Rocket studies of the lower ionosphere

    NASA Technical Reports Server (NTRS)

    Bowhill, Sidney A.

    1990-01-01

    The earth's ionosphere in the altitude range of 50 to 200 km was investigated by rocket-borne sensors, supplemented by ground-based measurement. The rocket payloads included mass spectrometers, energetic particle detectors, Langmuir probes and radio propagation experiments. Where possible, rocket flights were included in studies of specific phenomena, and the availability of data from other experiments greatly increased the significance of the results. The principal ionospheric phenomena studied were: winter anomaly in radiowave absorption, ozone and molecular oxygen densities, mid-latitude sporadic-E layers, energetic particle precipitation at middle and low latitudes, ionospheric instabilities and turbulence, and solar eclipse effects in the D and E regions. This document lists personnel who worked on the project, and provides a bibliography of resultant publications.

  14. This "Is" Rocket Science!

    ERIC Educational Resources Information Center

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-01-01

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

  15. Dr. Goddard Transports Rocket

    NASA Technical Reports Server (NTRS)

    1974-01-01

    Dr. Robert H. Goddard tows his rocket to the launching tower behind a Model A Ford truck, 15 miles northwest of Roswell, New Mexico. 1930- 1932. Dr. Goddard has been recognized as the 'Father of American Rocketry' and as one of three pioneers in the theoretical exploration of space. Robert Hutchings Goddard was born in Worcester, Massachusetts, on October 15, 1882. He was a theoretical scientist as well as a practical engineer. His dream was the conquest of the upper atmosphere and ultimately space through the use of rocket propulsion. Dr. Goddard, who died in 1945, was probably as responsible for the dawning of the Space Age as the Wright Brothers were for the begining of the Air Age. Yet his work attracted little serious attention during his lifetime. When the United States began to prepare for the conquest of space in the 1950's, American rocket scientists began to recognize the debt owed to the New England professor. They discovered that it was virtually impossible to construct a rocket or launch a satellite without acknowledging the work of Dr. Goddard. This great legacy was covered by more than 200 patents, many of which were issued after his death.

  16. This Is Rocket Science!

    NASA Astrophysics Data System (ADS)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-09-01

    Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical than that offered by Estes Industries, but more basic than the analysis of Nelson et al. In particular, drag is neglected until the very end of the exercise, which allows the concept of conservation of energy to be shown when predicting the rocket's flight. Also, the variable mass of the rocket motor is assumed to decrease linearly during the flight (while the propulsion charge and recovery delay charge are burning) and handled simplistically by using an average mass value. These changes greatly simplify the equations needed to predict the times and heights at various stages of flight, making it more useful as a review of basic physics. Details about model rocket motors, range safety, and other supplemental information may be found online at Apogee Components4 and the National Association of Rocketry.5

  17. Rocket Auroral Correlator Experiment

    NASA Technical Reports Server (NTRS)

    LaBelle, James

    2003-01-01

    Dartmouth College provided a multi-channel high- and low- frequency wave receivers, including active sensors on deployable booms, to the Rocket Auroral Correlator Experiment launched from Poker Flat, Alaska, in January 2002. College also performed preliminary analysis of the data. Details are outlined in chronological order.

  18. Liquid rocket engine turbines

    NASA Technical Reports Server (NTRS)

    1974-01-01

    Criteria for the design and development of turbines for rocket engines to meet specific performance, and installation requirements are summarized. The total design problem, and design elements are identified, and the current technology pertaining to these elements is described. Recommended practices for achieving a successful design are included.

  19. Liquid Rocket Engine Testing

    NASA Technical Reports Server (NTRS)

    Rahman, Shamim

    2005-01-01

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

  20. Liquid rocket valve components

    NASA Technical Reports Server (NTRS)

    1973-01-01

    A monograph on valves for use with liquid rocket propellant engines is presented. The configurations of the various types of valves are described and illustrated. Design criteria and recommended practices for the various valves are explained. Tables of data are included to show the chief features of valve components in use on operational vehicles.

  1. Liquid rocket valve assemblies

    NASA Technical Reports Server (NTRS)

    1973-01-01

    The design and operating characteristics of valve assemblies used in liquid propellant rocket engines are discussed. The subjects considered are as follows: (1) valve selection parameters, (2) major design aspects, (3) design integration of valve subassemblies, and (4) assembly of components and functional tests. Information is provided on engine, stage, and spacecraft checkout procedures.

  2. The Relativistic Rocket

    ERIC Educational Resources Information Center

    Antippa, Adel F.

    2009-01-01

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

  3. Thiokol Solid Rocket Motors

    NASA Technical Reports Server (NTRS)

    Graves, S. R.

    2000-01-01

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

  4. Water Rocket Workout.

    ERIC Educational Resources Information Center

    Esler, William K.; Sanford, Daniel

    1989-01-01

    Water rockets are used to present Newton's three laws of motion to high school physics students. Described is an outdoor activity which uses four students per group. Provides a launch data sheet to record height, angle of elevation, amount of water used, and launch number. (MVL)

  5. Water bottle rocket in undergraduate laboratory

    NASA Astrophysics Data System (ADS)

    Schultz, William

    2012-11-01

    In the winter semester of 2012, we implemented the modeling and testing of a water bottle rocket in ME 495, the Senior Laboratory in Mechanical Engineering at the University of Michigan. The four week lab was the most well received by the students in recent memory. There were significant challenges, but the result was a thorough review of their undergraduate fluids class with some advanced concepts such as directional stability of a projectile. The student teams designed their own rockets based on one of many standard 20 ounce soft drink bottles. The culminating contest brought impressive results and a surprise ending.

  6. Low-altitude acceleration of auroral electrons during breakup observed by a mother-daughter rocket

    NASA Technical Reports Server (NTRS)

    Johnstone, A. D.; Davis, T. N.

    1974-01-01

    By the use of a mother-daughter rocket combination and ground-based observations with television, time and space variations are resolved in particle measurements in breakup aurora. The spectral variations measured during a temporal variation in the aurora can be explained by a nearly uniform acceleration of all the electrons such as would be caused by an electric potential drop along the magnetic field lines. Many other explanations can be eliminated.

  7. Large Liquid Rocket Testing: Strategies and Challenges

    NASA Technical Reports Server (NTRS)

    Rahman, Shamim A.; Hebert, Bartt J.

    2005-01-01

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

  8. Robust Strategy for Rocket Engine Health Monitoring

    NASA Technical Reports Server (NTRS)

    Santi, L. Michael

    2001-01-01

    Monitoring the health of rocket engine systems is essentially a two-phase process. The acquisition phase involves sensing physical conditions at selected locations, converting physical inputs to electrical signals, conditioning the signals as appropriate to establish scale or filter interference, and recording results in a form that is easy to interpret. The inference phase involves analysis of results from the acquisition phase, comparison of analysis results to established health measures, and assessment of health indications. A variety of analytical tools may be employed in the inference phase of health monitoring. These tools can be separated into three broad categories: statistical, rule based, and model based. Statistical methods can provide excellent comparative measures of engine operating health. They require well-characterized data from an ensemble of "typical" engines, or "golden" data from a specific test assumed to define the operating norm in order to establish reliable comparative measures. Statistical methods are generally suitable for real-time health monitoring because they do not deal with the physical complexities of engine operation. The utility of statistical methods in rocket engine health monitoring is hindered by practical limits on the quantity and quality of available data. This is due to the difficulty and high cost of data acquisition, the limited number of available test engines, and the problem of simulating flight conditions in ground test facilities. In addition, statistical methods incur a penalty for disregarding flow complexity and are therefore limited in their ability to define performance shift causality. Rule based methods infer the health state of the engine system based on comparison of individual measurements or combinations of measurements with defined health norms or rules. This does not mean that rule based methods are necessarily simple. Although binary yes-no health assessment can sometimes be established by

  9. Numerical Simulation of Rocket Exhaust Interaction with Lunar Soil

    NASA Technical Reports Server (NTRS)

    Liever, Peter; Tosh, Abhijit; Curtis, Jennifer

    2012-01-01

    This technology development originated from the need to assess the debris threat resulting from soil material erosion induced by landing spacecraft rocket plume impingement on extraterrestrial planetary surfaces. The impact of soil debris was observed to be highly detrimental during NASA s Apollo lunar missions and will pose a threat for any future landings on the Moon, Mars, and other exploration targets. The innovation developed under this program provides a simulation tool that combines modeling of the diverse disciplines of rocket plume impingement gas dynamics, granular soil material liberation, and soil debris particle kinetics into one unified simulation system. The Unified Flow Solver (UFS) developed by CFDRC enabled the efficient, seamless simulation of mixed continuum and rarefied rocket plume flow utilizing a novel direct numerical simulation technique of the Boltzmann gas dynamics equation. The characteristics of the soil granular material response and modeling of the erosion and liberation processes were enabled through novel first principle-based granular mechanics models developed by the University of Florida specifically for the highly irregularly shaped and cohesive lunar regolith material. These tools were integrated into a unique simulation system that accounts for all relevant physics aspects: (1) Modeling of spacecraft rocket plume impingement flow under lunar vacuum environment resulting in a mixed continuum and rarefied flow; (2) Modeling of lunar soil characteristics to capture soil-specific effects of particle size and shape composition, soil layer cohesion and granular flow physics; and (3) Accurate tracking of soil-borne debris particles beginning with aerodynamically driven motion inside the plume to purely ballistic motion in lunar far field conditions. In the earlier project phase of this innovation, the capabilities of the UFS for mixed continuum and rarefied flow situations were validated and demonstrated for lunar lander rocket

  10. 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

  11. Rockets: Physical Science Teacher's Guide with Activities.

    ERIC Educational Resources Information Center

    Vogt, Gregory L.; Rosenberg, Carla R., Ed.

    Rockets have evolved from simple tubes filled with black powder into mighty vehicles capable of launching a spacecraft out into the galaxy. The guide begins with background information sections on the history of rocketry, scientific principles, and practical rocketry. The sections on scientific principles and practical rocketry are based on Isaac…

  12. Experimental Studies of the Heat Transfer to RBCC Rocket Nozzles for CFD Application to Design Methodologies

    NASA Technical Reports Server (NTRS)

    Santoro, Robert J.; Pal, Sibtosh

    1999-01-01

    Rocket thrusters for Rocket Based Combined Cycle (RBCC) engines typically operate with hydrogen/oxygen propellants in a very compact space. Packaging considerations lead to designs with either axisymmetric or two-dimensional throat sections. Nozzles tend to be either two- or three-dimensional. Heat transfer characteristics, particularly in the throat, where the peak heat flux occurs, are not well understood. Heat transfer predictions for these small thrusters have been made with one-dimensional analysis such as the Bartz equation or scaling of test data from much larger thrusters. The current work addresses this issue with an experimental program that examines the heat transfer characteristics of a gaseous oxygen (GO2)/gaseous hydrogen (GH2) two-dimensional compact rocket thruster. The experiments involved measuring the axial wall temperature profile in the nozzle region of a water-cooled gaseous oxygen/gaseous hydrogen rocket thruster at a pressure of 3.45 MPa. The wall temperature measurements in the thruster nozzle in concert with Bartz's correlation are utilized in a one-dimensional model to obtain axial profiles of nozzle wall heat flux.

  13. 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.

  14. Two-Dimensional Motions of Rockets

    ERIC Educational Resources Information Center

    Kang, Yoonhwan; Bae, Saebyok

    2007-01-01

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

  15. Water rocket - Electrolysis propulsion and fuel cell power

    SciTech Connect

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

    1999-07-24

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

  16. Prediction of rocket plume radiative heating using backward Monte-Carlo method

    NASA Technical Reports Server (NTRS)

    Wang, K. C.

    1993-01-01

    A backward Monte-Carlo plume radiation code has been developed to predict rocket plume radiative heating to the rocket base region. This paper provides a description of this code and provides sample results. The code was used to predict radiative heating to various locations during test firings of 48-inch solid rocket motors at NASA Marshall Space Flight Center. Comparisons with test measurements are provided. Predictions of full scale sea level Redesigned Solid Rocket Motor (RSRM) and Advanced Solid Rocket Motor (ASRM) plume radiative heating to the Space Shuttle external tank (ET) dome center were also made. A comparison with the Development Flight Instrumentation (DFI) measurements is also provided.

  17. The OGRESS Sounding Rocket Payload

    NASA Astrophysics Data System (ADS)

    Rogers, Thomas; Zeiger, B.; McEntaffer, R. L.; Schultz, T.; Oakley, P.; Cash, W. C.

    2013-01-01

    We present an overview of the Off-plane Grating Rocket for Extended Source Spectroscopy (OGRESS) sounding rocket payload based at the University of Iowa and University of Colorado, Boulder. The payload has been launched three times before under the names CyXESS, EXOS, and CODEX and is scheduled to fly again in February 2014. The payload is designed to observe large diffuse soft X-ray sources between ~100 - 1000 eV. OGRESS will observe the Vela supernova remnant and achieve the highest spectral resolution ever taken of this object in our bandpass. OGRESS does not use the standard optical design of a grazing incidence Wolter telescope. Instead, OGRESS uses a wire-grid collimator to remove any nonconverging light. This optical design leads to poor sensitivity to point sources, but works very well for large extended sources, for which the telescope beam is fully illuminated. The light is dispersed by an array of off-plane parallel-groove sinusoidal gratings and focused onto Gaseous Electron Multiplier (GEM) detectors.

  18. Engine for Redstone Rocket

    NASA Technical Reports Server (NTRS)

    2004-01-01

    This photograph is of the engine for the Redstone rocket. The Redstone ballistic missile was a high-accuracy, liquid-propelled, surface-to-surface missile developed by the Army Ballistic Missile Agency, Redstone Arsenal, in Huntsville, Alabama, under the direction of Dr. von Braun. The Redstone engine was a modified and improved version of the Air Force's Navaho cruise missile engine of the late forties. The A-series, as this would be known, utilized a cylindrical combustion chamber as compared with the bulky, spherical V-2 chamber. By 1951, the Army was moving rapidly toward the design of the Redstone missile, and the production was begun in 1952. Redstone rockets became the 'reliable workhorse' for America's early space program. As an example of its versatility, the Redstone was utilized in the booster for Explorer 1, the first American satellite, with no major changes to the engine or missile.

  19. Advanced rocket propulsion

    NASA Technical Reports Server (NTRS)

    Obrien, Charles J.

    1993-01-01

    Existing NASA research contracts are supporting development of advanced reinforced polymer and metal matrix composites for use in liquid rocket engines of the future. Advanced rocket propulsion concepts, such as modular platelet engines, dual-fuel dual-expander engines, and variable mixture ratio engines, require advanced materials and structures to reduce overall vehicle weight as well as address specific propulsion system problems related to elevated operating temperatures, new engine components, and unique operating processes. High performance propulsion systems with improved manufacturability and maintainability are needed for single stage to orbit vehicles and other high performance mission applications. One way to satisfy these needs is to develop a small engine which can be clustered in modules to provide required levels of total thrust. This approach should reduce development schedule and cost requirements by lowering hardware lead times and permitting the use of existing test facilities. Modular engines should also reduce operational costs associated with maintenance and parts inventories.

  20. The Aries sounding rocket

    NASA Astrophysics Data System (ADS)

    Dooling, D.

    1980-02-01

    A family of sounding rockets called Aries, using the motors from obsolete Minuteman ICBMs, is described. Payloads for Aries range from 1,500 to 3,500 lb (with a payload diameter of 44 in.) and include various instruments (magnetospheric tracers, X-ray and extreme ultraviolet astronomy and a large X-ray telescope). Prospects for future launching of a two and even three-stage Aries are discussed.

  1. EPDM rocket motor insulation

    NASA Technical Reports Server (NTRS)

    Guillot, David G. (Inventor); Harvey, Albert R. (Inventor)

    2004-01-01

    A novel and improved EPDM formulation for a solid propellant rocket motor is described wherein hexadiene EPDM monomer components are replaced by alkylidene norbornene components, and, with appropriate adjustment of curing and other additives, functionally required rheological and physical characteristics are achieved with the desired compatibility with any one of a plurality of solid filler materials, e.g., powder silica, carbon fibers or aramid fibers, and with appropriate adhesion and extended storage or shelf-life characteristics.

  2. EPDM rocket motor insulation

    NASA Technical Reports Server (NTRS)

    Guillot, David G. (Inventor); Harvey, Albert R. (Inventor)

    2008-01-01

    A novel and improved EPDM formulation for a solid propellant rocket motor is described wherein hexadiene EPDM monomer components are replaced by alkylidene norbornene components, and, with appropriate adjustment of curing and other additives, functionally required rheological and physical characteristics are achieved with the desired compatibility with any one of a plurality of solid filler materials, e.g., powder silica, carbon fibers or aramid fibers, and with appropriate adhesion and extended storage or shelf-life characteristics.

  3. EPDM rocket motor insulation

    NASA Technical Reports Server (NTRS)

    Guillot, David G. (Inventor); Harvey, Albert R. (Inventor)

    2003-01-01

    A novel and improved EPDM formulation for a solid propellant rocket motor is described wherein hexadiene EPDM monomer components are replaced by alkylidene norbornene components and with appropriate adjustment of curing and other additives functionally-required rheological and physical characteristics are achieved with the desired compatibility with any one of a plurality of solid filler materials, e.g. powder silica, carbon fibers or aramid fibers, and with appropriate adhesion and extended storage or shelf life characteristics.

  4. Analysis of Ablative Performance of C/C Composite Throat Containing Defects Based on X-ray 3D Reconstruction in a Solid Rocket Motor

    NASA Astrophysics Data System (ADS)

    Hui, Wei-Hua; Bao, Fu-Ting; Wei, Xiang-Geng; Liu, Yang

    2015-12-01

    In this paper, a new measuring method of ablation rate was proposed based on X-ray three-dimensional (3D) reconstruction. The ablation of 4-direction carbon/carbon composite nozzles was investigated in the combustion environment of a solid rocket motor, and the macroscopic ablation and linear recession rate were studied through the X-ray 3D reconstruction method. The results showed that the maximum relative error of the X-ray 3D reconstruction was 0.0576%, which met the minimum accuracy of the ablation analysis; along the nozzle axial direction, from convergence segment, throat to expansion segment, the ablation gradually weakened; in terms of defect ablation, the middle ablation was weak, while the ablation in both sides was more serious. In a word, the proposed reconstruction method based on X-ray about C/C nozzle ablation can construct a clear model of ablative nozzle which characterizes the details about micro-cracks, deposition, pores and surface to analyze ablation, so that this method can create the ablation curve in any surface clearly.

  5. Liquid Rocket Engine Testing Overview

    NASA Technical Reports Server (NTRS)

    Rahman, Shamim

    2005-01-01

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

  6. Electric rockets get a boost

    SciTech Connect

    Ashley, S.

    1995-12-01

    This article reports that xenon-ion thrusters are expected to replace conventional chemical rockets in many nonlaunch propulsion tasks, such as controlling satellite orbits and sending space probes on long exploratory missions. The space age dawned some four decades ago with the arrival of powerful chemical rockets that could propel vehicles fast enough to escape the grasp of earth`s gravity. Today, chemical rocket engines still provide the only means to boost payloads into orbit and beyond. The less glamorous but equally important job of moving vessels around in space, however, may soon be assumed by a fundamentally different rocket engine technology that has been long in development--electric propulsion.

  7. ISS Update: VASIMR Plasma Rocket

    NASA Video Gallery

    NASA Public Affairs Officer Dan Huot interviews Ken Bollweg, VASIMR Project Manager, about VASIMR (Variable Specific Impulse Magnetoplasma Rocket), recent testing progress and future applications. ...

  8. If Only Newton Had a Rocket.

    ERIC Educational Resources Information Center

    Hammock, Frank M.

    1988-01-01

    Shows how model rocketry can be included in physics curricula. Describes rocket construction, a rocket guide sheet, calculations and launch teams. Discusses the relationships of basic mechanics with rockets. (CW)

  9. Micro-Rockets for the Classroom.

    ERIC Educational Resources Information Center

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

    1999-01-01

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

  10. Combining Multiple Pairwise Structure-based Alignments

    SciTech Connect

    2014-11-12

    CombAlign is a new Python code that generates a gapped, one-to-many, multiple structure-based sequence alignment(MSSA) given a set of pairwise structure-based alignments. In order to better define regions of similarity among related protein structures, it is useful to detect the residue-residue correspondences among a set of pairwise structure alignments. Few codes exist for constructing a one-to-many, multiple sequence alignment derived from a set of structure alignments, and we perceived a need for creating a new tool for combing pairwise structure alignments that would allow for insertion of gaps in the reference structure.

  11. FDNS CFD Code Benchmark for RBCC Ejector Mode Operation: Continuing Toward Dual Rocket Effects

    NASA Technical Reports Server (NTRS)

    West, Jeff; Ruf, Joseph H.; Turner, James E. (Technical Monitor)

    2000-01-01

    Computational Fluid Dynamics (CFD) analysis results are compared with benchmark quality test data from the Propulsion Engineering Research Center's (PERC) Rocket Based Combined Cycle (RBCC) experiments to verify fluid dynamic code and application procedures. RBCC engine flowpath development will rely on CFD applications to capture the multi -dimensional fluid dynamic interactions and to quantify their effect on the RBCC system performance. Therefore, the accuracy of these CFD codes must be determined through detailed comparisons with test data. The PERC experiments build upon the well-known 1968 rocket-ejector experiments of Odegaard and Stroup by employing advanced optical and laser based diagnostics to evaluate mixing and secondary combustion. The Finite Difference Navier Stokes (FDNS) code [2] was used to model the fluid dynamics of the PERC RBCC ejector mode configuration. Analyses were performed for the Diffusion and Afterburning (DAB) test conditions at the 200-psia thruster operation point, Results with and without downstream fuel injection are presented.

  12. A novel image enhancement algorithm based on stationary wavelet transform for infrared thermography to the de-bonding defect in solid rocket motors

    NASA Astrophysics Data System (ADS)

    Liu, Tao; Zhang, Wei; Yan, Shaoze

    2015-10-01

    In this paper, a multi-scale image enhancement algorithm based on low-passing filtering and nonlinear transformation is proposed for infrared testing image of the de-bonding defect in solid propellant rocket motors. Infrared testing images with high-level noise and low contrast are foundations for identifying defects and calculating the defects size. In order to improve quality of the infrared image, according to distribution properties of the detection image, within framework of stationary wavelet transform, the approximation coefficients at suitable decomposition level is processed by index low-passing filtering by using Fourier transform, after that, the nonlinear transformation is applied to further process the figure to improve the picture contrast. To verify validity of the algorithm, the image enhancement algorithm is applied to infrared testing pictures of two specimens with de-bonding defect. Therein, one specimen is made of a type of high-strength steel, and the other is a type of carbon fiber composite. As the result shown, in the images processed by the image enhancement algorithm presented in the paper, most of noises are eliminated, and contrast between defect areas and normal area is improved greatly; in addition, by using the binary picture of the processed figure, the continuous defect edges can be extracted, all of which show the validity of the algorithm. The paper provides a well-performing image enhancement algorithm for the infrared thermography.

  13. XCALIBUR: a Vertical Takeoff TSTO RLV Concept with a HEDM Upperstage and a Scram-Rocket Booster

    NASA Astrophysics Data System (ADS)

    Bradford, J.

    2002-01-01

    A new 3rd generation, two-stage-to-orbit (TSTO) reusable launch vehicle (RLV) has been designed. The Xcalibur concept represents a novel approach due to its integration method for the upperstage element of the system. The vertical-takeoff booster, which is powered by rocket-based combined-cycle (RBCC) engines, carries the upperstage internally in the aft section of the airframe to a Mach 15 staging condition. The upperstage is released from the booster and carries the 6,820 kg of payload to low earth orbit (LEO) using its high energy density matter (HEDM) propulsion system. The booster element is capable of returning to the original launch site in a ramjet-cruise propulsion mode. Both the booster and the upperstage utilize advanced technologies including: graphite-epoxy tanks, metal-matrix composites, UHTC TPS materials, electro- mechanical actuators (EMAs), and lightweight subsystems (avionics, power distribution, etc.). The booster system is enabled main propulsion system which utilizes four RBCC engines. These engines operate in four distinct modes: air- augmented rocket (AAR), ramjet, scram-rocket, and all-rocket. The booster operates in AAR mode from takeoff to Mach 3, with ramjet mode operation from Mach 3 to Mach 6. The rocket re-ignition for scram-rocket mode occurs at Mach 6, with all-rocket mode from Mach 14 to the staging condition. The extended utilization of the scram-rocket mode greatly improves vehicle performance by providing superior vehicle acceleration when compared to the scramjet mode performance over the same flight region. Results indicate that the specific impulse penalty due to the scram-rocket mode operation is outweighed by the reduced flight time, smaller vehicle size due to increased mixture ratio, and lower allowable maximum dynamic pressure. A complete vehicle system life-cycle analysis was performed in an automated, multi-disciplinary design environment. Automated disciplinary performance analysis tools include: trajectory (POST

  14. Ionospheric Results with Sounding Rockets and the Explorer VIII Satellite (1960 )

    NASA Technical Reports Server (NTRS)

    Bourdeau, R. E.

    1961-01-01

    A review is made of ionospheric data reported since the IGY from rocket and satellite-borne ionospheric experiments. These include rocket results on electron density (RF impedance probe), D-region conductivity (Gerdien condenser), and electron temperature (Langmuir probe). Also included are data in the 1000 kilometer region on ion concentration (ion current monitor) and electron temperature from the Explorer VIII Satellite (1960 xi). The review includes suggestions for second generation experiments and combinations thereof particularly suited for small sounding rockets.

  15. Performance of a Solidification Furnace Developed for Sounding Rockets

    NASA Astrophysics Data System (ADS)

    An, Chen Y.; Boschetti, Cesar; Ribeiro, Manuel F.; Toledo, Rafael C.; Freitas, Filipe E.; Bandeira, Iraja N.

    2011-11-01

    Brazil has a Microgravity Program mainly based on experiments using sounding rockets. This paper presents a brief account of the Program and the experiments made on the Brazilian rockets. Up to now three missions carrying a total of 26 experiments were made. In all flights a fast solidification furnace, capable of producing temperatures up to 900°C, was tested with semiconductor and metal alloys. This paper describes the construction and the performance of that furnace during the last parabolic flight, occurred in 2010. The solidification furnace is now qualified and ready to be used by other institutions in sounding rocket flights.

  16. Sea-Level Static Testing of the Penn State Two-Dimensional Rocket-Based Combined Cycle (RBCC) Testbed

    NASA Technical Reports Server (NTRS)

    Cramer, J. M.; Pal, S.; Marshall, W. M.; Santoro, R. J.

    2003-01-01

    The present results indicated that: 1.Significant RBCC ejector mode database has been generated for single and twin thruster configuration and for global and local measurements. 2. Ongoing analysis and correlation effort for MSFC CFD modeling and turbulent shear layer analysis was completed. 3. The potential follow-on activities are: detailed measurements of air flow static pressure and velocity profiles; investigation other thruster spacing configurations; performing fundamental shear layer mixing study; and demonstrating single-shot Raman measurements.

  17. Sea-Level Static Testing of the Penn State Two-Dimensional Rocket-Based Combined Cycle (RBCC) Testbed

    NASA Technical Reports Server (NTRS)

    Cramer, J. M.; Marshall, W. M.; Pal, S.; Santoro, R. J.

    2003-01-01

    Twin thruster tests have been conducted with the Penn State RBCC test article operating at sea- level static conditions. Significant differences were observed in the performance characteristics for two different thruster centerline spacings. Changing the thruster spacing from 2.50 to 1.75 in. reduced the entrained air velocity (-17%) and the thrust (-7%) for tests at a thruster chamber pressure of 200 psia and MR = 8. In addition, significant differences were seen in the static pressure profiles, the Raman spectroscopy profiles, and the acoustic power spectrum for these two configurations.

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

    NASA Technical Reports Server (NTRS)

    Thomas, Matthew E.; Bossard, John A.; Early, Jim; Trinh, Huu; Dennis, Jay; Turner, James (Technical Monitor)

    2001-01-01

    The fiber optically coupled laser ignition approach summarized is under consideration for use in igniting bi-propellant rocket thrust chambers. This laser ignition approach is based on a novel dual pulse format capable of effectively increasing laser generated plasma life times up to 1000 % over conventional laser ignition methods. In the dual-pulse format tinder consideration here an initial laser pulse is used to generate a small plasma kernel. A second laser pulse that effectively irradiates the plasma kernel follows this pulse. Energy transfer into the kernel is much more efficient because of its absorption characteristics thereby allowing the kernel to develop into a much more effective ignition source for subsequent combustion processes. In this research effort both single and dual-pulse formats were evaluated in a small testbed rocket thrust chamber. The rocket chamber was designed to evaluate several bipropellant combinations. Optical access to the chamber was provided through small sapphire windows. Test results from gaseous oxygen (GOx) and RP-1 propellants are presented here. Several variables were evaluated during the test program, including spark location, pulse timing, and relative pulse energy. These variables were evaluated in an effort to identify the conditions in which laser ignition of bi-propellants is feasible. Preliminary results and analysis indicate that this laser ignition approach may provide superior ignition performance relative to squib and torch igniters, while simultaneously eliminating some of the logistical issues associated with these systems. Further research focused on enhancing the system robustness, multiplexing, and window durability/cleaning and fiber optic enhancements is in progress.

  19. Preliminary study of the rocket-ramjet-rocket concept for HTOL space plane

    NASA Astrophysics Data System (ADS)

    Wang, Shusheng

    This paper addresses a rocket-ramjet-rocket concept for the HTOL/TSTO space plane. Emphasis is placed on the discussion of the propulsion system for this concept. For the ramjet two key measures - earlier ignition and oxygen replenishment at high altitude - are employed so that the concept is feasible. The flight trajectory of the plane is given as well as its configuration and overall dimension. Based on the mass estimation, the calculated payload ratio is improved by using technologies of the nineties.

  20. Rocket + Science = Dialogue

    NASA Technical Reports Server (NTRS)

    Morris,Bruce; Sullivan, Greg; Burkey, Martin

    2010-01-01

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

  1. Active Region Soft X-Ray Spectra and Temperature Analyses based on Sounding Rocket Measurements from the Solar Aspect Monitor (SAM), - a Modified SDO/EVE Instrument

    NASA Astrophysics Data System (ADS)

    Didkovsky, Leonid V.; Wieman, Seth; Woods, Thomas N.; Jones, Andrew; Moore, Christopher

    2016-05-01

    Some initial results of soft x-ray spectral (0.5 to 3.0 nm) observations of active regions (AR11877 and AR11875) from a sounding rocket flight NASA 36.290 on 21 October 2013 at about 18:30 UT are reported. These observations were made by a Solar Aspect Monitor (SAM), a rocket version of the EUV Variability Experiment’s (EVE) channel, a pinhole camera modified for EVE rocket suite of instruments to include a free-standing transmission grating (200 nm period), which provided spectrally-resolved images of the solar disk. Intensity ratios for strong emission lines extracted from temporally averaged SAM spectral profiles of the ARs were compared to appropriately convolved modeled CHIANTI spectra. These ratios represent the AR’s temperature structures, which are compared to the structures derived from some other observations and temperature models.

  2. ION ROCKET ENGINE

    DOEpatents

    Ehlers, K.W.; Voelker, F. III

    1961-12-19

    A thrust generating engine utilizing cesium vapor as the propellant fuel is designed. The cesium is vaporized by heat and is passed through a heated porous tungsten electrode whereby each cesium atom is fonized. Upon emergfng from the tungsten electrode, the ions are accelerated rearwardly from the rocket through an electric field between the tungsten electrode and an adjacent accelerating electrode grid structure. To avoid creating a large negative charge on the space craft as a result of the expulsion of the positive ions, a source of electrons is disposed adjacent the ion stream to neutralize the cesium atoms following acceleration thereof. (AEC)

  3. Guided Rocket Weapon,

    DTIC Science & Technology

    1982-06-11

    AD-Ai49 861 GUIDED ROCKET hEAPON(U) FOREIGN TECHNOLOGY DIV 1/2 RIfGT-PRTTERSON RFB OH P V MOROZOV ii JUN 82 FID- ID( RS) T-8527-82 UNCLASSIFIED F/ 16...1963 A 4° 4i !° 2. ..7 7C - - .7- - - - - ,a’ ~ .Vr- ’ r lt In’ K" F7 - ID (RS)Y,-0527-82 FOREIGN TECHNOLOGY DIVISION P IN CA * I ’ " swoo n.T t-b I...engineering,* I electronics, automation, chemistry, metallurgy and rocketry. * " The missile industry - new branch of technology , although the history of

  4. Fabrication of GRCop-84 Rocket Thrust Chambers

    NASA Technical Reports Server (NTRS)

    Loewenthal, William S.; Ellis, David L.

    2005-01-01

    GRCop-84, a copper alloy, Cu-8 at% Cr-4 at% Nb developed at NASA Glenn Research Center for regeneratively cooled rocket engine liners has excellent combinations of elevated temperature strength, creep resistance, thermal conductivity and low cycle fatigue. GRCop-84 is produced from prealloyed atomized powder and has been fabricated into plate, sheet and tube forms as well as near net shapes. Fabrication processes to produce demonstration rocket combustion chambers will be presented and includes powder production, extruding, rolling, forming, friction stir welding, and metal spinning. GRCop-84 has excellent workability and can be readily fabricated into complex components using conventional powder and wrought metallurgy processes. Rolling was examined in detail for process sensitivity at various levels of total reduction, rolling speed and rolling temperature representing extremes of commercial processing conditions. Results indicate that process conditions can range over reasonable levels without any negative impact to properties.

  5. Nuclear thermal rockets using indigenous Martian propellants

    SciTech Connect

    Zubrin, R.M.

    1989-01-01

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

  6. The influence of phytochemical composition and resulting sensory attributes on preference for salad rocket (Eruca sativa) accessions by consumers of varying TAS2R38 diplotype.

    PubMed

    Bell, Luke; Methven, Lisa; Wagstaff, Carol

    2017-05-01

    Seven accessions of Eruca sativa ("salad rocket") were subjected to a randomised consumer assessment. Liking of appearance and taste attributes were analysed, as well as perceptions of bitterness, hotness, pepperiness and sweetness. Consumers were genotyped for TAS2R38 status to determine if liking is influenced by perception of bitter compounds such as glucosinolates (GSLs) and isothiocyanates (ITCs). Responses were combined with previously published data relating to phytochemical content and sensory data in Principal Component Analysis to determine compounds influencing liking/perceptions. Hotness, not bitterness, is the main attribute on which consumers base their liking of rocket. Some consumers rejected rocket based on GSL/ITC concentrations, whereas some preferred hotness. Bitter perception did not significantly influence liking of accessions, despite PAV/PAV 'supertasters' scoring higher for this attribute. High sugar-GSL/ITC ratios significantly reduce perceptions of hotness and bitterness for some consumers. Importantly the GSL glucoraphanin does not impart significant influence on liking or perception traits.

  7. Otrag rocket experiments in Africa

    NASA Technical Reports Server (NTRS)

    1978-01-01

    West German rocket manufacturers are testing their products in Zaire. Hundreds of pipes (12 m x 80 cm) are bundled together inside the test missiles, which are fired into Zaire's prairie. The reactions of neighboring nations, as well as leading countries of the world, are presented concerning the rocket tests.

  8. Soda-Bottle Water Rockets.

    ERIC Educational Resources Information Center

    Kagan, David; And Others

    1995-01-01

    Provides instructions for the construction and launch of a two-liter plastic soda-bottle rocket and presents the author's theory of their motion during launch. Modeled predictions are compared with actual experimental data. Explains theory behind the motion of a water rocket during launch. (LZ)

  9. Rocket launchers as passive controllers

    NASA Astrophysics Data System (ADS)

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

    1981-12-01

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

  10. Coal-Fired Rocket Engine

    NASA Technical Reports Server (NTRS)

    Anderson, Floyd A.

    1987-01-01

    Brief report describes concept for coal-burning hybrid rocket engine. Proposed engine carries larger payload, burns more cleanly, and safer to manufacture and handle than conventional solid-propellant rockets. Thrust changeable in flight, and stops and starts on demand.

  11. Beginnings of rocket development in the czech lands (Czechoslovakia)

    NASA Astrophysics Data System (ADS)

    Plavec, Michal

    2011-11-01

    Although the first references are from the 15th Century when both Hussites and crusaders are said to have used rockets during the Hussite Wars (also known as the Bohemian Wars) there is no strong evidence that rockets were actually used at that time. It is worth noting that Konrad Kyeser, who described several rockets in his Bellifortis manuscript written 1402-1405, served as advisor to Bohemian King Wenceslas IV. Rockets were in fact used as fireworks from the 16th century in noble circles. Some of these were built by Vavřinec Křička z Bitý\\vsky, who also published a book on fireworks, in which he described how to build rockets for firework displays. Czech soldiers were also involved in the creation of a rocket regiment in the Austrian (Austro-Hungarian) army in the first half of the 19th century. The pioneering era of modern rocket development began in the Czech lands during the 1920s. The first rockets were succesfully launched by Ludvík Očenášek in 1930 with one of them possibly reaching an altitude of 2000 metres. Vladimír Mandl, lawyer and author of the first book on the subject of space law, patented his project for a stage rocket (vysokostoupající raketa) in 1932, but this project never came to fruition. There were several factories during the so-called Protectorate of Bohemia and Moravia in 1939-1945, when the Czech lands were occupied by Nazi Germany, where parts for German Mark A-4/V-2 rockets were produced, but none of the Czech technicians or constructors were able to build an entire rocket. The main goal of the Czech aircraft industry after WW2 was to revive the stagnant aircraft industry. There was no place to create a rocket industry. Concerns about a rocket industry appeared at the end of the 1950s. The Political Board of the Central Committee of the Czechoslovak Communist Party started to study the possibilities of creating a rocket industry after the first flight into space and particularly after US nuclear weapons were based in Italy

  12. 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.

  13. ROCOPT: A user friendly interactive code to optimize rocket structural components

    NASA Technical Reports Server (NTRS)

    Rule, William K.

    1989-01-01

    ROCOPT is a user-friendly, graphically-interfaced, microcomputer-based computer program (IBM compatible) that optimizes rocket components by minimizing the structural weight. The rocket components considered are ring stiffened truncated cones and cylinders. The applied loading is static, and can consist of any combination of internal or external pressure, axial force, bending moment, and torque. Stress margins are calculated by means of simple closed form strength of material type equations. Stability margins are determined by approximate, orthotropic-shell, closed-form equations. A modified form of Powell's method, in conjunction with a modified form of the external penalty method, is used to determine the minimum weight of the structure subject to stress and stability margin constraints, as well as user input constraints on the structural dimensions. The graphical interface guides the user through the required data prompts, explains program options and graphically displays results for easy interpretation.

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

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  15. Real-time seam tracking for rocket thrust chamber manufacturing

    SciTech Connect

    Schmitt, D.J.; Novak, J.L.; Starr, G.P.; Maslakowski, J.E.

    1993-11-01

    A sensor-based control approach for real-time seam tracking of rocket thrust chamber assemblies has been developed to enable automation of a braze paste dispensing process. This approach utilizes a non-contact Multi-Axis Seam Tracking (MAST) sensor to track the seams. Thee MAST sensor measures capacitance variations between the sensor and the workpiece and produces four varying voltages which are read directly into the robot controller. A PID control algorithm which runs at the application program level has been designed based upon a simple dynamic model of the combined robot and sensor plant. The control algorithm acts on the incoming sensor signals in real-time to guide the robot motion along the seam path. Experiments demonstrate that seams can be tracked at 100 mm/sec within the accuracy required for braze paste dispensing.

  16. Pegasus Rocket Model

    NASA Technical Reports Server (NTRS)

    1996-01-01

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

  17. 6. "EXPERIMENTAL ROCKET ENGINE TEST STATION AT AFFTC." A low ...

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

    6. "EXPERIMENTAL ROCKET ENGINE TEST STATION AT AFFTC." A low oblique aerial view of Test Area 1-115, looking south, showing Test Stand 1-3 at left, Instrumentation and Control building 8668 at center, and Test Stand 15 at right. The test area is under construction; no evidence of railroad line in photo. - Edwards Air Force Base, Air Force Rocket Propulsion Laboratory, Leuhman Ridge near Highways 58 & 395, Boron, Kern County, CA

  18. Liquid Rocket Lines, Bellows, Flexible Hoses, and Filters

    NASA Technical Reports Server (NTRS)

    1977-01-01

    Fluid-flow components in a liquid propellant rocket engine and the rocket vehicle which it propels are interconnected by lines, bellows, and flexible hoses. Elements involved in the successful design of these components are identified and current technologies pertaining to these elements are reviewed, assessed, and summarized to provide a technology base for a checklist of rules to be followed by project managers in guiding a design or assessing its adequacy. Recommended procedures for satisfying each of the design criteria are included.

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

    NASA Technical Reports Server (NTRS)

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

    2000-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2014-01-01

    The current paper presents results from a combined research effort by Louisiana Tech University (LTU) and NASA Stennis Space Center (SSC) to develop a new alumina-silicate based cementitious binder capable of acting as a high performance refractory material with low heat ablation rate and high early mechanical strength. Such a binder would represent a significant contribution to NASA's efforts to develop a new generation of refractory 'hot face' liners for liquid or solid rocket plume environments. This project was developed as a continuation of on-going collaborations between LTU and SSC, where test sections of a formulation of high temperature geopolymer binder were cast in the floor and walls of Test Stand E-1 Cell 3, an active rocket engine test stand flame trench. Additionally, geopolymer concrete panels were tested using the NASA-SSC Diagnostic Test Facility (DTF) thruster, where supersonic plume environments were generated on a 1ft wide x 2ft long x 6 inch deep refractory panel. The DTF operates on LOX/GH2 propellants producing a nominal thrust of 1,200 lbf and the combustion chamber conditions are Pc=625psig, O/F=6.0. Data collected included high speed video of plume/panel area and surface profiles (depth) of the test panels measured on a 1-inch by 1-inch giving localized erosion rates during the test. Louisiana Tech conducted a microstructure analysis of the geopolymer binder after the testing program to identify phase changes in the material.

  1. The NASA Sounding Rocket Program and space sciences.

    PubMed

    Gurkin, L W

    1992-10-01

    High altitude suborbital rockets (sounding rockets) have been extensively used for space science research in the post-World War II period; the NASA Sounding Rocket Program has been on-going since the inception of the Agency and supports all space science disciplines. In recent years, sounding rockets have been utilized to provide a low gravity environment for materials processing research, particularly in the commercial sector. Sounding rockets offer unique features as a low gravity flight platform. Quick response and low cost combine to provide more frequent spaceflight opportunities. Suborbital spacecraft design practice has achieved a high level of sophistication which optimizes the limited available flight times. High data-rate telemetry, real-time ground up-link command and down-link video data are routinely used in sounding rocket payloads. Standard, off-the-shelf, active control systems are available which limit payload body rates such that the gravitational environment remains less than 10(-4) g during the control period. Operational launch vehicles are available which can provide up to 7 minutes of experiment time for experiment weights up to 270 kg. Standard payload recovery systems allow soft impact retrieval of payloads. When launched from White Sands Missile Range, New Mexico, payloads can be retrieved and returned to the launch site within hours.

  2. The NASA Sounding Rocket Program and space sciences

    NASA Technical Reports Server (NTRS)

    Gurkin, L. W.

    1992-01-01

    High altitude suborbital rockets (sounding rockets) have been extensively used for space science research in the post-World War II period; the NASA Sounding Rocket Program has been on-going since the inception of the Agency and supports all space science disciplines. In recent years, sounding rockets have been utilized to provide a low gravity environment for materials processing research, particularly in the commercial sector. Sounding rockets offer unique features as a low gravity flight platform. Quick response and low cost combine to provide more frequent spaceflight opportunities. Suborbital spacecraft design practice has achieved a high level of sophistication which optimizes the limited available flight times. High data-rate telemetry, real-time ground up-link command and down-link video data are routinely used in sounding rocket payloads. Standard, off-the-shelf, active control systems are available which limit payload body rates such that the gravitational environment remains less than 10(-4) g during the control period. Operational launch vehicles are available which can provide up to 7 minutes of experiment time for experiment weights up to 270 kg. Standard payload recovery systems allow soft impact retrieval of payloads. When launched from White Sands Missile Range, New Mexico, payloads can be retrieved and returned to the launch site within hours.

  3. Past and Present Large Solid Rocket Motor Test Capabilities

    NASA Technical Reports Server (NTRS)

    Kowalski, Robert R.; Owen, David B., II

    2011-01-01

    A study was performed to identify the current and historical trends in the capability of solid rocket motor testing in the United States. The study focused on test positions capable of testing solid rocket motors of at least 10,000 lbf thrust. Top-level information was collected for two distinct data points plus/minus a few years: 2000 (Y2K) and 2010 (Present). Data was combined from many sources, but primarily focused on data from the Chemical Propulsion Information Analysis Center s Rocket Propulsion Test Facilities Database, and heritage Chemical Propulsion Information Agency/M8 Solid Rocket Motor Static Test Facilities Manual. Data for the Rocket Propulsion Test Facilities Database and heritage M8 Solid Rocket Motor Static Test Facilities Manual is provided to the Chemical Propulsion Information Analysis Center directly from the test facilities. Information for each test cell for each time period was compiled and plotted to produce a graphical display of the changes for the nation, NASA, Department of Defense, and commercial organizations during the past ten years. Major groups of plots include test facility by geographic location, test cells by status/utilization, and test cells by maximum thrust capability. The results are discussed.

  4. Structure and dynamics of the nightside poleward boundary: Sounding rocket and ground-based observations of auroral electron precipitation in a rayed curtain

    NASA Astrophysics Data System (ADS)

    Lynch, K. A.; Hampton, D.; Mella, M.; Zhang, Binzheng; Dahlgren, H.; Disbrow, M.; Kintner, P. M.; Lessard, M.; Lundberg, E.; Stenbaek-Nielsen, H. C.

    2012-11-01

    The Cascades2 auroral sounding rocket provides a case study for comparing multipoint in situ ionospheric observations of a nightside auroral poleward boundary intensification with ground-based optical observations of the same event. Cascades2 was launched northward from Poker Flat Alaska on 20 March 2009 at 11:04 UT. The 13 min flight reached an apogee of 564 km over the northern coast of Alaska. The experiment included a five-payload array of in situ instrumentation, ground cameras at three different points under the trajectory, multiple ground magnetometers, the Poker Flat Incoherent Scatter Radar (PFISR) radar, and the Time History of Events and Macroscale Interactions during Substorms (THEMIS) spacecraft in the magnetotail. The rays of the poleward boundary intensification (PBI) curtain have along-arc motions of 8.5 km/s and along-arc spacings of 16 km. Modulated maximum energy envelopes and energy fluxes of the associated electron precipitation correspond to this spatial structure of the visible rays. The electron precipitation is additionally modulated at a higher frequency, and velocity dispersion analysis of these 8 Hz signatures implies Alfvénic wave-particle acceleration of an ambient ionospheric electron source occurring a few hundred km above the observation point. These observations parameterize the curtain of Alfvénic activity above the PBI event, both in the dispersive ionosphere and in the magnetotail reconnection region. The along-arc variations in brightness correspond to variations in precipitating electron energy flux interpreted as an along-arc modulation of the maximum energy of the Alfvénic wave-particle acceleration process; this is a new interpretation of the formation of rayed structures in auroral curtains. We consider the various possible magnetospheric and ionospheric drivers for the control of the observed along-arc structuring and motions.

  5. Investigation of Non-Conventional Bio-Derived Fuels for Hybrid Rocket Motors

    DTIC Science & Technology

    2007-08-01

    hybrid rocket motors. This analysis indicates beeswax, lard, a mixture of paraffin and lard, and combinations of beeswax and aluminum should all...40 IV.IV Data Analysis ...Rockets using unsymmetrical dimethyl hydrazine ( UDMH ) release so-called nitrogen radicals, which are potentially damaging to the ozone layer. Also

  6. British used Congreve Rockets to Attack Napoleon

    NASA Technical Reports Server (NTRS)

    2004-01-01

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

  7. Liquid rocket engine test facility engineering challenges

    NASA Astrophysics Data System (ADS)

    Ellerbrock, Hartwig; Ziegenhagen, Stefan

    2006-12-01

    Liquid rocket engines for launch vehicles and space crafts as well as their subsystems need to be verified and qualified during hot-runs. A high test cadence combined with a flexible test team helps to reduce the cost for test verification during development/qualification as well as during acceptance testing for production. Test facility intelligence allows to test subsystems in the same manner as during complete engine system tests and will therefore reduce development time and cost. This paper gives an overview of the maturing of test engineering know how for rocket engine test stands as well as high altitude test stands for small propulsion thrusters at EADS-ST in Ottobrunn and Lampoldshausen and is split into two parts: Part 1 gives a historical overview of the EADS-ST test stands at Ottobrunn and Lampoldshausen since the beginning of Rocket propulsion activities in the 1960s. Part 2 gives an overview of the actual test capabilities and the test engineering know-how for test stand construction/adaptation and their use during running programs. Examples of actual realised facility concepts are given to demonstrate cost saving potential for test programs in both cases for development/qualification issues as well as for production purposes.

  8. Students Participate in Rocket Launch Project

    NASA Technical Reports Server (NTRS)

    2002-01-01

    Filled with anticipation, students from two local universities, the University of Alabama in Huntsville (UAH), and Alabama Agricultural Mechanical University (AM), counted down to launch the rockets they designed and built at the Army test site on Redstone Arsenal in Huntsville, Alabama. The projected two-mile high launch culminated more than a year's work and demonstrated the student team's ability to meet the challenge set by the Marshall Space Flight Center's (MSFC) Student Launch Initiative (SLI) program to apply science and math to experience, judgment, and common sense, and proved to NASA officials that they have successfully built reusable launch vehicles (RLVs), another challenge set by NASA's SLI program. MSFC's SLI program is an educational effort that aims to motivate students to pursue careers in science, math, and engineering. It provides the students with hands-on, practical aerospace experience. In this picture, the combined efforts of students from UAH and AM sent this rocket soaring into flight. Students at UAH built the rocket and AM students developed its scientific payload, an experiment that measures the amount of hydrogen produced during electroplating with nickel in a brief period of micrgravity.

  9. Space rocket engine on the base of the reactor-pumped laser for the interplanetary flights and earth orbital applications

    NASA Astrophysics Data System (ADS)

    Gulevich, Andrey V.; Dyachenko, Peter P.; Kukharchuk, Oleg F.; Zrodnikov, Anatoly V.

    2000-01-01

    In this report the concept of vehicle-based reactor-laser engine for long time interplanetary and interorbital (LEO to GEO) flights is proposed. Reactor-pumped lasers offer the perspective way to create on the base of modern nuclear and lasers technologies the low mass and high energy density, repetitively pulsed vehicle-based laser of average power 100 kW. Nowadays the efficiency of nuclear-to-optical energy conversion reached the value of 2-3%. The demo model of reactor-pumped laser facility is under construction in Institute for Physics and Power Engineering (Obninsk, Russia). It enable us to hope that using high power laser on board of the vehicle could make the effective space laser engine possible. Such engine may provide the high specific impulse ~1000-2000 s with the thrust up to 10-100 n. Some calculation results of the characteristics of vehicle-based reactor-laser thermal engine concept are also presented. .

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

    NASA Technical Reports Server (NTRS)

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

    2014-01-01

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

  11. Rocket/launcher structural dynamics

    NASA Technical Reports Server (NTRS)

    Ferragut, N. J.

    1976-01-01

    The equations of motion describing the interactions between a rocket and a launcher were derived using Lagrange's Equation. A rocket launching was simulated. The motions of both the rocket and the launcher can be considered in detail. The model contains flexible elements and rigid elements. The rigid elements (masses) were judiciously utilized to simplify the derivation of the equations. The advantages of simultaneous shoe release were illustrated. Also, the loading history of the interstage structure of a boosted configuration was determined. The equations shown in this analysis could be used as a design tool during the modification of old launchers and the design of new launchers.

  12. Atomic hydrogen rocket engine

    NASA Technical Reports Server (NTRS)

    Etters, R. D.; Flurchick, K.

    1981-01-01

    A rocket using atomic hydrogen propellant is discussed. An essential feature of the proposed engine is that the atomic hydrogen fuel is used as it is produced, thus eliminating the necessity of storage. The atomic hydrogen flows into a combustion chamber and recombines, producing high velocity molecular hydrogen which flows out an exhaust port. Standard thermodynamics, kinetic theory and wall recombination cross-sections are used to predict a thrust of approximately 1.4 N for a RF hydrogen flow rate of 4 x 10 to the 22nd/sec. Specific impulses are nominally from 1000 to 2000 sec. It is predicted that thrusts on the order of one Newton and specific impulses of up to 2200 sec are attainable with nominal RF discharge fluxes on the order of 10 to the 22nd atoms/sec; further refinements will probably not alter these predictions by more than a factor of two.

  13. Biomolecular Network-Based Synergistic Drug Combination Discovery

    PubMed Central

    Li, Xiangyi; Qin, Guangrong; Yang, Qingmin

    2016-01-01

    Drug combination is a powerful and promising approach for complex disease therapy such as cancer and cardiovascular disease. However, the number of synergistic drug combinations approved by the Food and Drug Administration is very small. To bridge the gap between urgent need and low yield, researchers have constructed various models to identify synergistic drug combinations. Among these models, biomolecular network-based model is outstanding because of its ability to reflect and illustrate the relationships among drugs, disease-related genes, therapeutic targets, and disease-specific signaling pathways as a system. In this review, we analyzed and classified models for synergistic drug combination prediction in recent decade according to their respective algorithms. Besides, we collected useful resources including databases and analysis tools for synergistic drug combination prediction. It should provide a quick resource for computational biologists who work with network medicine or synergistic drug combination designing. PMID:27891522

  14. Dr. Robert H. Goddard and His Rocket

    NASA Technical Reports Server (NTRS)

    2004-01-01

    Goddard rocket with four rocket motors. This rocket attained an altitude of 200 feet in a flight, November 1936, at Roswell, New Mexico. From 1930 to 1941, Dr. Goddard made substantial progress in the development of progressively larger rockets which attained altitudes of 2400 meters, and refined his equipment for guidance and control, his techniques of welding, and his insulation, pumps, and other associated equipment. In many respects, Dr. Goddard laid the essential foundations of practical rocket technology

  15. Exploring Sustainable Rocket Fuels: [Imidazolyl-Amine-BH2](+)-Cation-Based Ionic Liquids as Replacements for Toxic Hydrazine Derivatives.

    PubMed

    Huang, Shi; Qi, Xiujuan; Zhang, Wenquan; Liu, Tianlin; Zhang, Qinghua

    2015-12-01

    The application of hypergolic ionic liquids as propellant fuels is a newly emerging area in the fields of chemistry and propulsion science. Herein, a new class of [imidazolyl-amine-BH2](+)-cation-based ionic liquids, which included fuel-rich anions, such as dicyanamide (N(CN)2(-)) and cyanoborohydride (BH3CN(-)) anions, were synthesized and characterized. As expected, all of the ionic liquids exhibited spontaneous combustion upon contact with the oxidizer 100 % HNO3. The densities of these ionic liquids varied from 0.99-1.12 g cm(-3), and the heats of formation, predicted based on Gaussian 09 calculations, were between -707.7 and 241.8 kJ mol(-1). Among them, the salt of compound 5, that is, (1-allyl-1H-imidazole-3-yl)-(trimethylamine)-dihydroboronium dicyanamide, exhibited the lowest viscosity (168 MPa s), good thermal properties (Tg <-70 °C, Td >130 °C), and the shortest ignition-delay time (18 ms) with 100 % HNO3. These ionic fuels, as "green" replacements for toxic hydrazine-derivatives, may have potential applications as bipropellant formulations.

  16. A fluidic sounding rocket motor ignition system.

    NASA Technical Reports Server (NTRS)

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

    1972-01-01

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

  17. Automated and Manual Rocket Crater Measurement Software

    NASA Technical Reports Server (NTRS)

    Metzger, Philip; Immer, Christopher

    2012-01-01

    An update has been performed to software designed to do very rapid automated measurements of craters created in sandy substrates by rocket exhaust on liftoff. The previous software was optimized for pristine lab geometry and lighting conditions. This software has been enhanced to include a section for manual measurements of crater parameters; namely, crater depth, crater full width at half max, and estimated crater volume. The tools provide a very rapid method to measure these manual parameters to ease the burden of analyzing large data sets. This software allows for rapid quantization of the rocket crater parameters where automated methods may not work. The progress of spreadsheet data is continuously saved so that data is never lost, and data can be copied to clipboards and pasted to other software for analysis. The volume estimation of a crater is based on the central max depth axis line, and the polygonal shape of the crater is integrated around that axis.

  18. Sounding Rockets as a Real Flight Platform for Aerothermodynamic Cfd Validation of Hypersonic Flight Experiments

    NASA Astrophysics Data System (ADS)

    Stamminger, A.; Turner, J.; Hörschgen, M.; Jung, W.

    2005-02-01

    This paper describes the possibilities of sounding rockets to provide a platform for flight experiments in hypersonic conditions as a supplement to wind tunnel tests. Real flight data from measurement durations longer than 30 seconds can be compared with predictions from CFD calculations. This paper will regard projects flown on sounding rockets, but mainly describe the current efforts at Mobile Rocket Base, DLR on the SHarp Edge Flight EXperiment SHEFEX.

  19. Analysis of a water-propelled rocket: A problem in honors physics

    NASA Astrophysics Data System (ADS)

    Finney, G. A.

    2000-03-01

    The air-pumped, water-propelled rocket is a common child's toy, yet forms a reasonably complicated system when carefully analyzed. A lab based on this system was included as the final laboratory project in the honors version of General Physics I at the USAF Academy. The numerical solution for the height of the rocket is presented, as well as several analytic approximations. Five out of six lab groups predicted the maximum height of the rocket within experimental error.

  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. High-temperature, low-cycle fatigue of advanced copper-base alloys for rocket nozzles. Part 1: Narloy Z

    NASA Technical Reports Server (NTRS)

    Conway, J. B.; Stentz, R. H.; Berling, J. T.

    1974-01-01

    Short-term tensile and low-cycle fatigue data are reported for Narloy Z, a centrifugally cast, copper-base alloy. Tensile tests were performed at room temperature in air and in argon at 482, 538 and 593 C using an axial strain rate of .002/sec to the -1 power. In addition tensile tests were performed at 538 C in an evaluation of tensile properties at strain rates of .004 and .01/sec to the -1 power. Ultimate and yield strength values of about 315 and 200 MN/sq m respectively were recorded at room temperature and these decreased to about 120 and 105 respectively as the temperature was increased to 593 C. Reduction in area values were recorded in the range from 40 to 50% with some indication of a minimum ductility point at 538 C.

  2. Studies on Pre-Ignition Reactions of Hydrocarbon-Based Rocket Fuels Hypergolic with Red Fuming Nitric Acid as Oxidizer

    NASA Astrophysics Data System (ADS)

    Kulkarni, Suresh G.; Bagalkote, Vrushali S.

    2010-06-01

    Carene, norbornadiene, ethylidene norbornene, and furfuryl alcohol exhibit hypergolic ignition with red fuming nitric acid as oxidizer. Carene, when blended with norbornadiene, ethylidene norbornene, and furfuryl alcohol in appropriate proportions, exhibits synergistic hypergolic ignition with further decrease in ignition delay. In order to understand the probable mechanism of hypergolic ignition and synergy in ignition, various pre-ignition reactions have been studied by quenching the reactions and analyzing the intermediate products by Fourier transform infrared (FTIR) and gas chromatography-mass spectrometry (GC-MS) techniques. Based on the product analysis, the probable schemes of reactions have been proposed. Oxidation, nitration, and cationic polymerization appear to be the important pre-ignition reactions taking place simultaneously that are responsible for hypergolic ignition.

  3. Radiation from advanced solid rocket motor plumes

    NASA Astrophysics Data System (ADS)

    Farmer, Richard C.; Smith, Sheldon D.; Myruski, Brian L.

    1994-12-01

    The overall objective of this study was to develop an understanding of solid rocket motor (SRM) plumes in sufficient detail to accurately explain the majority of plume radiation test data. Improved flowfield and radiation analysis codes were developed to accurately and efficiently account for all the factors which effect radiation heating from rocket plumes. These codes were verified by comparing predicted plume behavior with measured NASA/MSFC ASRM test data. Upon conducting a thorough review of the current state-of-the-art of SRM plume flowfield and radiation prediction methodology and the pertinent data base, the following analyses were developed for future design use. The NOZZRAD code was developed for preliminary base heating design and Al2O3 particle optical property data evaluation using a generalized two-flux solution to the radiative transfer equation. The IDARAD code was developed for rapid evaluation of plume radiation effects using the spherical harmonics method of differential approximation to the radiative transfer equation. The FDNS CFD code with fully coupled Euler-Lagrange particle tracking was validated by comparison to predictions made with the industry standard RAMP code for SRM nozzle flowfield analysis. The FDNS code provides the ability to analyze not only rocket nozzle flow, but also axisymmetric and three-dimensional plume flowfields with state-of-the-art CFD methodology. Procedures for conducting meaningful thermo-vision camera studies were developed.

  4. Radiation from advanced solid rocket motor plumes

    NASA Technical Reports Server (NTRS)

    Farmer, Richard C.; Smith, Sheldon D.; Myruski, Brian L.

    1994-01-01

    The overall objective of this study was to develop an understanding of solid rocket motor (SRM) plumes in sufficient detail to accurately explain the majority of plume radiation test data. Improved flowfield and radiation analysis codes were developed to accurately and efficiently account for all the factors which effect radiation heating from rocket plumes. These codes were verified by comparing predicted plume behavior with measured NASA/MSFC ASRM test data. Upon conducting a thorough review of the current state-of-the-art of SRM plume flowfield and radiation prediction methodology and the pertinent data base, the following analyses were developed for future design use. The NOZZRAD code was developed for preliminary base heating design and Al2O3 particle optical property data evaluation using a generalized two-flux solution to the radiative transfer equation. The IDARAD code was developed for rapid evaluation of plume radiation effects using the spherical harmonics method of differential approximation to the radiative transfer equation. The FDNS CFD code with fully coupled Euler-Lagrange particle tracking was validated by comparison to predictions made with the industry standard RAMP code for SRM nozzle flowfield analysis. The FDNS code provides the ability to analyze not only rocket nozzle flow, but also axisymmetric and three-dimensional plume flowfields with state-of-the-art CFD methodology. Procedures for conducting meaningful thermo-vision camera studies were developed.

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

    NASA Technical Reports Server (NTRS)

    Maram, J.; Barkhoudarian, S.

    1987-01-01

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

  6. Turbopump Design and Analysis Approach for Nuclear Thermal Rockets

    NASA Technical Reports Server (NTRS)

    Chen, Shu-cheng S.; Veres, Joseph P.; Fittje, James E.

    2006-01-01

    A rocket propulsion system, whether it is a chemical rocket or a nuclear thermal rocket, is fairly complex in detail but rather simple in principle. Among all the interacting parts, three components stand out: they are pumps and turbines (turbopumps), and the thrust chamber. To obtain an understanding of the overall rocket propulsion system characteristics, one starts from analyzing the interactions among these three components. It is therefore of utmost importance to be able to satisfactorily characterize the turbopump, level by level, at all phases of a vehicle design cycle. Here at NASA Glenn Research Center, as the starting phase of a rocket engine design, specifically a Nuclear Thermal Rocket Engine design, we adopted the approach of using a high level system cycle analysis code (NESS) to obtain an initial analysis of the operational characteristics of a turbopump required in the propulsion system. A set of turbopump design codes (PumpDes and TurbDes) were then executed to obtain sizing and performance characteristics of the turbopump that were consistent with the mission requirements. A set of turbopump analyses codes (PUMPA and TURBA) were applied to obtain the full performance map for each of the turbopump components; a two dimensional layout of the turbopump based on these mean line analyses was also generated. Adequacy of the turbopump conceptual design will later be determined by further analyses and evaluation. In this paper, descriptions and discussions of the aforementioned approach are provided and future outlooks are discussed.

  7. Turbopump Design and Analysis Approach for Nuclear Thermal Rockets

    SciTech Connect

    Chen, Shucheng S.; Veres, Joseph P.; Fittje, James E.

    2006-01-20

    A rocket propulsion system, whether it is a chemical rocket or a nuclear thermal rocket, is fairly complex in detail but rather simple in principle. Among all the interacting parts, three components stand out: they are pumps and turbines (turbopumps), and the thrust chamber. To obtain an understanding of the overall rocket propulsion system characteristics, one starts from analyzing the interactions among these three components. It is therefore of utmost importance to be able to satisfactorily characterize the turbopump, level by level, at all phases of a vehicle design cycle. Here at the NASA Glenn Research Center, as the starting phase of a rocket engine design, specifically a Nuclear Thermal Rocket Engine design, we adopted the approach of using a high level system cycle analysis code (NESS) to obtain an initial analysis of the operational characteristics of a turbopump required in the propulsion system. A set of turbopump design codes (PumpDes and TurbDes) were then executed to obtain sizing and performance parameters of the turbopump that were consistent with the mission requirements. A set of turbopump analyses codes (PUMPA and TURBA) were applied to obtain the full performance map for each of the turbopump components; a two dimensional layout of the turbopump based on these mean line analyses was also generated. Adequacy of the turbopump conceptual design will later be determined by further analyses and evaluation. In this paper, descriptions and discussions of the aforementioned approach are provided and future outlooks are discussed.

  8. Small Solid Rocket Motor Test

    NASA Video Gallery

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

  9. Rocket-Booster Towing Simulation

    NASA Technical Reports Server (NTRS)

    Trovillion, T. A.

    1985-01-01

    Report describes computer simulation of motion of solid-rocket ship. Listing of simulation program in FORTRAN. Mathematical techniques useful in such other maritime applications as buoy or ship design.

  10. Navigating the Rockets Educator Guide

    NASA Video Gallery

    In this brief video overview, learn how to navigate the Rockets Educator Guide. Get a glimpse of the resources available in the guide, including a pictorial history, an overview of the physics cont...

  11. Solid rocket motor internal insulation

    NASA Technical Reports Server (NTRS)

    Twichell, S. E. (Editor); Keller, R. B., Jr.

    1976-01-01

    Internal insulation in a solid rocket motor is defined as a layer of heat barrier material placed between the internal surface of the case propellant. The primary purpose is to prevent the case from reaching temperatures that endanger its structural integrity. Secondary functions of the insulation are listed and guidelines for avoiding critical problems in the development of internal insulation for rocket motors are presented.

  12. Solid Rocket Motor Acoustic Testing

    SciTech Connect

    Rogers, J.D.

    1999-03-31

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

  13. Easier Analysis With Rocket Science

    NASA Technical Reports Server (NTRS)

    2003-01-01

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

  14. Earth-to-Orbit Rocket Propulsion

    NASA Technical Reports Server (NTRS)

    Beaurain, Andre; Souchier, Alain; Moravie, Michel; Sackheim, Robert L.; Cikanek, Harry A., III

    2003-01-01

    The Earth-to-orbit (ETO) phase of access to space is and always will be the first and most critical phase of all space missions. This first phase of all space missions has unique characteristics that have driven space launcher propulsion requirements for more than half a century. For example, the need to overcome the force of the Earth s gravity in combination with high levels of atmospheric drag to achieve the initial orbital velocity; i.e., Earth parking orbit or =9 km/s, will always require high thrust- to-weight (TN) propulsion systems. These are necessary with a T/W ratio greater than one during the ascent phase. The only type of propulsion system that can achieve these high T/W ratios are those that convert thermal energy to kinetic energy. There are only two basic sources of onboard thermal energy: chemical combustion-based systems or nuclear thermal-based systems (fission, fusion, or antimatter). The likelihood of advanced open-cycle, nuclear thermal propulsion being developed for flight readiness or becoming environmentally acceptable during the next century is extremely low. This realization establishes that chemical propulsion for ET0 launchers will be the technology of choice for at least the next century, just as it has been for the last half century of rocket flight into space. The world s space transportation propulsion requirements have evolved through several phases over the history of the space program, as has been necessitated by missions and systems development, technological capabilities available, and the growth and evolution of the utilization of space for economic, security, and science benefit. Current projections for the continuing evolution of requirements and concepts may show how future space transportation system needs could be addressed. The evolution and projections will be described in detail in this manuscript.

  15. Safe testing nuclear rockets economically

    SciTech Connect

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

    2002-01-01

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

  16. Method of fault-tree quantitative analysis for solid rocket motor

    NASA Astrophysics Data System (ADS)

    Hu, Baochao; Yang, Yicai; Xie, Weimin

    1993-08-01

    Based on the existing problem in determining the failure probabilities of base events in solid rocket motor fault-tree quantitative analysis, an engineering method of 'Solicitation Opinions to Give Marks' was put forward to determine the failure probability. A satisfactory result was obtained by analyzing the practical example of structure reliability for some solid rocket motors at the test sample stage.

  17. Method of fault-tree quantitative analysis for solid rocket motor

    NASA Astrophysics Data System (ADS)

    Hu, Baochao; Yang, Yicai; Xie, Weimin

    1993-08-01

    Based on the existing problem of determining the failure probabilities of base events in solid rocket motor fault tree quantitative analysis, an engineering method of Solicitation Opinions to Give Marks is put forward to determine failure probability. A more satisfactory result is obtained by analyzing the actual example of the structural reliability of solid rocket motors at the test sample stage.

  18. Low-thrust chemical rocket engine study

    NASA Technical Reports Server (NTRS)

    Mellish, J. A.

    1981-01-01

    Engine data and information are presented to perform system studies on cargo orbit-transfer vehicles which would deliver large space structures to geosynchronous equatorial orbit. Low-thrust engine performance, weight, and envelope parametric data were established, preliminary design information was generated, and technologies for liquid rocket engines were identified. Two major engine design drivers were considered in the study: cooling and engine cycle options. Both film-cooled and regeneratively cooled engines were evaluated. The propellant combinations studied were hydrogen/oxygen, methane/oxygen, and kerosene/oxygen.

  19. Nanoparticle-Based Combination Therapy for Cancer Treatment.

    PubMed

    Yhee, Ji Young; Son, Sejin; Lee, Hyukjin; Kim, Kwangmeyung

    2015-01-01

    In recent years, combination of different types of therapies using nanoparticles has emerged as an advanced strategy for cancer treatment. Most of all, combination of chemotherapeutic drug and siRNA in nanoformulation has shown a great potential, because siRNA-mediated specific gene silencing can compensate for the incomplete anti-cancer actions of chemotherapy. In this article, nanoparticle-based combination therapy for cancer treatment is introduced to be focused on the therapeutic chemical and siRNA combination. It is classified into 3 groups: 1) general chemotherapy combined with siRNA carrying nanoparticle, 2) co-delivery of chemical and siRNA therapeutics within a single nanoparticle, and 3) Use of multiple nanoparticles for chemical and siRNA therapeutics. The purpose of the combination and the mechanisms of anti-cancer action was described according to the categories. Examples of some recent developments of nanotechnology-based chemo- and siRNA- therapeutics combination therapy are summarized for better understanding of its practical application.

  20. Oxaliplatin-based combined-modality therapy for rectal cancer.

    PubMed

    Minsky, Bruce D

    2003-08-01

    There are two conventional treatments for clinically resectable rectal cancer. The first is surgery, and, if the tumor is T3 and/or N1-2, this is followed by postoperative combined-modality therapy. The second, for patients with ultrasound T3 or clinical T4 disease, is preoperative combined-modality therapy followed by surgery and postoperative chemotherapy. In this review, the results of these approaches as well as novel combined-modality approaches using oxaliplatin-based regimens will be presented.

  1. Highlights from a Mach 4 Experimental Demonstration of Inlet Mode Transition for Turbine-Based Combined Cycle Hypersonic Propulsion

    NASA Technical Reports Server (NTRS)

    Foster, Lancert E.; Saunders, John D., Jr.; Sanders, Bobby W.; Weir, Lois J.

    2012-01-01

    NASA is focused on technologies for combined cycle, air-breathing propulsion systems to enable reusable launch systems for access to space. Turbine Based Combined Cycle (TBCC) propulsion systems offer specific impulse (Isp) improvements over rocket-based propulsion systems in the subsonic takeoff and return mission segments along with improved safety. Among the most critical TBCC enabling technologies are: 1) mode transition from the low speed propulsion system to the high speed propulsion system, 2) high Mach turbine engine development and 3) innovative turbine based combined cycle integration. To address these challenges, NASA initiated an experimental mode transition task including analytical methods to assess the state-of-the-art of propulsion system performance and design codes. One effort has been the Combined-Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE-LIMX) which is a fully integrated TBCC propulsion system with flowpath sizing consistent with previous NASA and DoD proposed Hypersonic experimental flight test plans. This experiment was tested in the NASA GRC 10 by 10-Foot Supersonic Wind Tunnel (SWT) Facility. The goal of this activity is to address key hypersonic combined-cycle engine issues including: (1) dual integrated inlet operability and performance issues-unstart constraints, distortion constraints, bleed requirements, and controls, (2) mode-transition sequence elements caused by switching between the turbine and the ramjet/scramjet flowpaths (imposed variable geometry requirements), and (3) turbine engine transients (and associated time scales) during transition. Testing of the initial inlet and dynamic characterization phases were completed and smooth mode transition was demonstrated. A database focused on a Mach 4 transition speed with limited off-design elements was developed and will serve to guide future TBCC system studies and to validate higher level analyses.

  2. Key technology for reusable rocket engine turbopump

    NASA Astrophysics Data System (ADS)

    Okayasu, A.; Ohta, T.; Kamijyo, A.; Yamada, H.

    2002-03-01

    Recently, there has been an increased need for evolved space transportation and the research of reusable rocket which enable low cost and high reliability and is generating a lot of interest all over the world. In the USA, the development of reusable launch vehicle "Venture Star" which will be used instead of space shuttle is planned and its half scale model "X-33" was developed for the first flight in 1999. In Japan, there has been agreement on the main points to develop the rocket type RLV based on the technology of H-IIA, HOPE-X before developing space plane type RLV. The planned reusable rocket engine was LOX/LH2 as propellant, has 100- 200 ton thrust and has a throttling capability. In addition, long life and high reliability are required for the engine system including LOX/LH2 turbopump. The paper introduces some key technologies for the reusable turbopump which IHI is promoting for research and development with NAL and Tohoku University.

  3. Turbo Pump Fed Micro-Rocket Engine

    NASA Astrophysics Data System (ADS)

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

    2004-10-01

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

  4. Design Considerations of ISTAR Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

    NASA Technical Reports Server (NTRS)

    Andreadis, Dean; Drake, Alan; Garrett, Joseph L.; Gettinger, Christopher D.; Hoxie, Stephen S.

    2003-01-01

    The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system that produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

  5. Design Considerations of Istar Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

    NASA Technical Reports Server (NTRS)

    Andreadis, Dean; Drake, Alan; Garrett, Joseph L.; Gettinger, Christopher D.; Hoxie, Stephen S.

    2002-01-01

    The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system thai: produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

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

    NASA Astrophysics Data System (ADS)

    Zhang, Shuai; Hu, Fan; Zhang, Weihua

    2016-10-01

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

  7. EUVS Sounding Rocket Payload

    NASA Technical Reports Server (NTRS)

    Stern, Alan S.

    1996-01-01

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

  8. Rocket Science 101 Interactive Educational Program

    NASA Technical Reports Server (NTRS)

    Armstrong, Dennis; Funkhouse, Deborah; DiMarzio, Donald

    2007-01-01

    To better educate the public on the basic design of NASA s current mission rockets, Rocket Science 101 software has been developed as an interactive program designed to retain a user s attention and to teach about basic rocket parts. This program also has helped to expand NASA's presence on the Web regarding educating the public about the Agency s goals and accomplishments. The software was designed using Macromedia s Flash 8. It allows the user to select which type of rocket they want to learn about, interact with the basic parts, assemble the parts to create the whole rocket, and then review the basic flight profile of the rocket they have built.

  9. Reduced hazard chemicals for solid rocket motor production

    NASA Technical Reports Server (NTRS)

    Caddy, Larry A.; Bowman, Ross; Richards, Rex A.

    1995-01-01

    During the last three years. the NASA/Thiokol/industry team has developed and started implementation of an environmentally sound manufacturing plan for the continued production of solid rocket motors. NASA Marshall Space Flight Center (MSFC) and Thiokol Corporation have worked with other industry representatives and the U.S. Environmental Protection Agency (EPA) to prepare a comprehensive plan to eliminate all ozone depleting chemicals from manufacturing processes and reduce the use of other hazardous materials used to produce the space shuttle reusable solid rocket motors. The team used a classical approach for problem-solving combined with a creative synthesis of new approaches to attack this challenge.

  10. Sounding rocket activities of Japan in 2003 and 2004

    NASA Astrophysics Data System (ADS)

    Ishii, Nobuaki; Inatani, Yoshifumi; Nonaka, Satoshi; Nakajima, Takashi; Takumi, Abe; Yuichi, Tsuda; Yamagami, Takamasa

    2005-08-01

    In October 2003, a new space agency, JAXA (Japan Aerospace Exploration Agency) was reorganized and started as a primary space agency to promote all space activities in Japan. The Institute of Space and Astronautical Science (ISAS) belonged to JAXA and continued to promote space science and technologies using unique scientific satellites, sounding rockets and balloons. This paper summarizes sounding rocket and ballooning activities of ISAS in the fiscal year of 2003 and 2004, associated with satellite launch programs. In this time period, three sounding rockets and nineteen balloons were launched by ISAS. One of the sounding rocket, S-310-35 was an international collaboration between Japan and Norway, which was launched from Andoya Rocket Range (ARR), Andenes, Norway, so as to study the upper atmospheric dynamics and energetics associated with the auroral energy in the polar lower thermosphere. Through the combination with the national researchers and the cooperation with international organizations, ISAS will keep its own flight opportunities and be able to obtain many new scientific findings.

  11. Solid rocket motor cost model

    NASA Technical Reports Server (NTRS)

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

    1972-01-01

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

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

    NASA Astrophysics Data System (ADS)

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

    2004-08-01

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

  13. Combined Space-Based Observations of Geostationary Satellites

    NASA Astrophysics Data System (ADS)

    Scott, R.; Bernard, K.; Thorsteinson, S.

    2016-09-01

    One of the Space Situational Awareness (SSA) science experiments of the NEOSSat mission is to learn the practicalities of combining space-based metric observations with the Sapphire system. To answer this question, an experiment was performed observing clustered Canadian geostationary satellites using both Sapphire and NEOSSat in early 2016. Space-based tracking data was collected during tracking intervals where both NEOSSat and Sapphire had visibility on the geostationary objects enabling astrometric (orbit determination) and photometric (object characterization) observations to be performed. We describe the orbit determination accuracies using live data collected from orbit for different collection cases; a) NEOSSat alone, b) Sapphire alone, and c) Combined observations from both platforms. We then discuss the practicalities of using space-based sensors to reduce risk of orbital collisions of Canadian geostationary satellites by proactively tasking space based sensors in response to conjunction data warnings in GEO.

  14. Combine harvester monitor system based on wireless sensor network

    Technology Transfer Automated Retrieval System (TEKTRAN)

    A measurement method based on Wireless Sensor Network (WSN) was developed to monitor the working condition of combine harvester for remote application. Three JN5139 modules were chosen for sensor data acquisition and another two as a router and a coordinator, which could create a tree topology netwo...

  15. Marshall Team Recreates Goddard Rocket

    NASA Technical Reports Server (NTRS)

    2003-01-01

    In honor of the Centernial of Flight celebration and commissioned by the American Institute of Aeronautics and Astronautics (AIAA), a team of engineers from Marshall Space Flight Center (MSFC) built a replica of the first liquid-fueled rocket. The original rocket, designed and built by rocket engineering pioneer Robert H. Goddard in 1926, opened the door to modern rocketry. Goddard's rocket reached an altitude of 41 feet while its flight lasted only 2.5 seconds. The Marshall design team's plan was to stay as close as possible to an authentic reconstruction of Goddard's rocket. The same propellants were used - liquid oxygen and gasoline - as available during Goddard's initial testing and firing. The team also tried to construct the replica using the original materials and design to the greatest extent possible. By purposely using less advanced techniques and materials than many that are available today, the team encountered numerous technical challenges in testing the functional hardware. There were no original blueprints or drawings, only photographs and notes. However, this faithful adherence to historical accuracy has also allowed the team to experience many of the same challenges Goddard faced 77 years ago, and more fully appreciate the genius of this extraordinary man. The replica will undergo ground tests at MSFC this summer.

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

    NASA Technical Reports Server (NTRS)

    Thorpe, Douglas G.

    1991-01-01

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

  17. Small-Scale Rocket Motor Test

    NASA Video Gallery

    Engineers at NASA's Marshall Space Flight Center in Huntsville, Ala. successfully tested a sub-scale solid rocket motor on May 27. Testing a sub-scale version of a rocket motor is a cost-effective ...

  18. High-Temperature, Low-Cycle Fatigue of Copper-Base Alloys for Rocket Nozzles. Part 1: Data Summary for Materials Tested in Prior Programs

    NASA Technical Reports Server (NTRS)

    Conway, J. B.; Stentz, R. H.; Berling, J. T.

    1975-01-01

    A more detailed analysis of the results obtained in 188 previously reported low-cycle fatigue tests of various candidate materials for regeneratively-cooled, reusable rocket nozzle liners was reported. Plots of load range versus cycles were reported for each test along with a stress-strain hysteresis loop near half-life. In addition, a summary table was provided to compare N5 (cycles to a five percent load range drop) and Nf (cycles to complete specimen separation) values for each test.

  19. Selecting supplier combination based on fuzzy multicriteria analysis

    NASA Astrophysics Data System (ADS)

    Han, Zhi-Qiu; Luo, Xin-Xing; Chen, Xiao-Hong; Yang, Wu-E.

    2015-07-01

    Existing multicriteria analysis (MCA) methods are probably ineffective in selecting a supplier combination. Thus, an MCA-based fuzzy 0-1 programming method is introduced. The programming relates to a simple MCA matrix that is used to select a single supplier. By solving the programming, the most feasible combination of suppliers is selected. Importantly, this result differs from selecting suppliers one by one according to a single-selection order, which is used to rank sole suppliers in existing MCA methods. An example highlights such difference and illustrates the proposed method.

  20. Emergency egress fixed rocket package

    NASA Technical Reports Server (NTRS)

    Allen, Margaret A. (Inventor)

    1989-01-01

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

  1. Low-thrust rocket trajectories

    SciTech Connect

    Keaton, P.W.

    1986-01-01

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

  2. Low-thrust rocket trajectories

    SciTech Connect

    Keaton, P.W.

    1987-03-01

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

  3. Rocket Science at the Nanoscale.

    PubMed

    Li, Jinxing; Rozen, Isaac; Wang, Joseph

    2016-06-28

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

  4. Saving Lives With Rocket Power

    NASA Technical Reports Server (NTRS)

    2000-01-01

    Thiokol Propulsion uses NASA's surplus rocket fuel to produce a flare that can safely destroy land mines. Through a Memorandum of Agreement between Thiokol and Marshall Space Flight Center, Thiokol uses the scrap Reusable Solid Rocket Motor (RSRM) propellant. The resulting Demining Device was developed by Thiokol with the help of DE Technologies. The Demining Device neutralizes land mines in the field without setting them off. The Demining Device flare is placed next to an uncovered land mine. Using a battery-triggered electric match, the flare is then ignited. Using the excess and now solidified rocket fuel, the flare burns a hole in the mine's case and ignites the explosive contents. Once the explosive material is burned away, the mine is disarmed and no longer dangerous.

  5. A Plasma Rocket Demonstration on the International Space Station

    NASA Astrophysics Data System (ADS)

    Petro, A.

    2002-01-01

    in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One feature of this concept is the ability to vary its specific impulse so that it can be operated in a mode that maximizes propellant efficiency or a mode that maximizes thrust. For this reason the system is called the Variable Specific Impulse Magneto-plasma Rocket or VASIMR. This ability to vary specific impulse and thrust will allow for optimum low thrust interplanetary trajectories and results in shorter trip times than is possible with fixed specific impulse systems while preserving adequate payload margins. demonstrations are envisioned. A ground-based experiment of a low-power VASIMR prototype rocket is currently underway at the Advanced Space Propulsion Laboratory. The next step is a proposal to build and fly a 25-kilowatt VASIMR rocket as an external payload on the International Space Station. This experiment will provide an opportunity to demonstrate the performance of the rocket in space and measure the induced environment. The experiment will also utilize the space station for its intended purpose as a laboratory with vacuum conditions that cannot be matched by any laboratory on Earth. propulsion on the space station. An electric propulsion system like VASIMR, if provided with sufficient electrical power, could provide continuous drag force compensation for the space station. Drag compensation would eliminate the need for reboosting the station, an operation that will consume about 60 metric tons of propellant in a ten-year period. In contrast, an electric propulsion system would require very little propellant. In fact, a system like VASIMR can use waste hydrogen from the station's life support system as its propellant. This waste hydrogen is otherwise dumped overboard. Continuous drag compensation would also improve the microgravity conditions on the station. So electric propulsion can reduce propellant

  6. 14 CFR 101.25 - Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

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

  7. 14 CFR 101.25 - Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

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

  8. 14 CFR 101.25 - Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

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

  9. 14 CFR 101.25 - Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

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

  10. 14 CFR 101.25 - Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

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

  11. Observations of the global structure of the stratosphere and mesosphere with sounding rockets and with remote sensing techniques from satellites

    NASA Technical Reports Server (NTRS)

    Heath, D. F.; Hilsenrath, E.; Krueger, A. J.; Nordberg, W.; Prabhakara, C.; Theon, J. S.

    1972-01-01

    Brief descriptions are given of the techniques involved in determining the global structure of the mesosphere and stratosphere based on sounding rocket observations and satellite remotely sensed measurements.

  12. The Swedish Rocket Corps, 1833 - 1845

    NASA Technical Reports Server (NTRS)

    Skoog, A. I.

    1977-01-01

    Rockets for pyrotechnic displays used in Sweden in the 19th century are examined in terms of their use in war situations. Work done by the Swedish chemist J. J. Berzelius, who analyzed and improved the propellants of such rockets, and the German engineer, Martin Westermaijer, who researched manufacturing techniques of these rockets is also included.

  13. Array Of Rockets For Multicrewmember Evacuation

    NASA Technical Reports Server (NTRS)

    Allen, Margaret A.

    1990-01-01

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

  14. F. Gomez Arias' rocket vehicle project

    NASA Technical Reports Server (NTRS)

    Carreras, R.

    1977-01-01

    Research done by Spanish pioneer rocket scientists in the 19th century was investigated with major emphasis placed on F. Gomez Arias' rocket vehicle project. Arias, considered the world's first designer of rocket propelled, manned aircraft, was interested in solving the problem of space navigation. Major concerns included ascent and direction of heavier-than-airmachines, as well as ascent and direction of balloons.

  15. Premature ignition of a rocket motor.

    SciTech Connect

    Moore, Darlene Ruth

    2010-10-01

    During preparation for a rocket sled track (RST) event, there was an unexpected ignition of the zuni rocket motor (10/9/08). Three Sandia staff and a contractor were involved in the accident; the contractor was seriously injured and made full recovery. The data recorder battery energized the low energy initiator in the rocket.

  16. Measuring Model Rocket Engine Thrust Curves

    ERIC Educational Resources Information Center

    Penn, Kim; Slaton, William V.

    2010-01-01

    This paper describes a method and setup to quickly and easily measure a model rocket engine's thrust curve using a computer data logger and force probe. Horst describes using Vernier's LabPro and force probe to measure the rocket engine's thrust curve; however, the method of attaching the rocket to the force probe is not discussed. We show how a…

  17. The United States sounding rocket program

    NASA Technical Reports Server (NTRS)

    1971-01-01

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

  18. Helping HAN for hybrid rockets

    NASA Astrophysics Data System (ADS)

    Ramohalli, Kumar; Dowler, Warren

    1995-01-01

    Hydroxyl amine nitrate (HAN) is a powerful oxidizer for hybrid rocket flight motors. Miscible with water up to 95% by mass, it also has high density and has been extensively characterized for materials compatibility, safety, transportation, storage and handling. Before any serious attempt to use the proposed oxidizer in hybrids, though, the usual performance figures must first be obtained. The simplest are time-independent, equilibrium rocket performance numbers that include chamber temperature, temperature at the nozzle throat, and key species in the exhaust. These numbers must be followed by several other important performance evaluation, including burning rates, pressure dependence, susceptibility to instabilities and temperature sensitivity.

  19. Rocket Launch Trajectory Simulations Mechanism

    NASA Technical Reports Server (NTRS)

    Margasahayam, Ravi; Caimi, Raoul E.; Hauss, Sharon; Voska, N. (Technical Monitor)

    2002-01-01

    The design and development of a Trajectory Simulation Mechanism (TSM) for the Launch Systems Testbed (LST) is outlined. In addition to being one-of-a-kind facility in the world, TSM serves as a platform to study the interaction of rocket launch-induced environments and subsequent dynamic effects on the equipment and structures in the close vicinity of the launch pad. For the first time, researchers and academicians alike will be able to perform tests in a laboratory environment and assess the impact of vibroacoustic behavior of structures in a moving rocket scenario on ground equipment, launch vehicle, and its valuable payload or spacecraft.

  20. Rocket study of auroral processes

    NASA Technical Reports Server (NTRS)

    Arnoldy, R. L.

    1981-01-01

    Abstracts are presented of previously published reports analyzing data from three Echo 3 rocket flights. Particle experiments designed for the Terrier-Malmute flight, the Echo 5 flight, and the Norwegian Corbier Ferdinand 50 flight are described and their flight performance evaluated. Theoretical studies on auroral particle precipitation are reviewed according to observations made in three regions of space: (1) the region accessible to rockets and low altitude satellites (few hundred to a few thousand kilometers); (2) the region extending from 4000 to 8000 km (S3-3 satellite range); and (3) near the equatorial plane (geosynchronous satellite measurements). Questions raised about auroral arc formation are considered.

  1. The Advanced Solid Rocket Motor

    NASA Technical Reports Server (NTRS)

    Mitchell, Royce E.

    1992-01-01

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

  2. Predicting Rocket or Jet Noise in Real Time

    NASA Technical Reports Server (NTRS)

    Frendi, Kader

    2007-01-01

    A semi-empirical theoretical model and a C++ computer program that implements the model have been developed for use in predicting the noise generated by a rocket or jet engine. The computer program, entitled the Realtime Rocket and Jet Engine Noise Analysis and Prediction Software, is one of two main subsystems of the Acoustic Prediction/Measurement Tool, which comprises software, acoustic instrumentation, and electronic hardware combined to afford integrated capabilities for real-time prediction and measurement of noise emitted by rocket and jet engines. [The other main subsystem, consisting largely of acoustic instrumentation and electronic hardware, is described in Wireless Acoustic Measurement System, which appears elsewhere in this section.] The theoretical model was derived from the fundamental laws of fluid mechanics, as first was done by M. J. Lighthill in his now famous theory of aerodynamically generated sound. The far-field approximation of the Lighthill theory is incorporated into this model. Many other contributions from various researchers have also been introduced into the model. The model accounts for two noise components: shear noise and self noise. The final result of the model is expressed in terms of a volume integral of the acoustic intensities attributable to these two components, subject to various directivity coefficients. The computer program was written to solve the volume integral. The inputs required by the program are two data files from a computational fluid dynamics (CFD) simulation of the flow of interest: the computational-grid file and the solution file. The CFD solution should be one that has been obtained for conditions that closely approximate those of an experimental test that is yet to be performed. In the current state of development of the model and software, it is recommended that the observation points lie along a radius at an angle >60 from the jet axis. The software provides, and is driven via, a graphical user interface

  3. Life extending control for rocket engines

    NASA Technical Reports Server (NTRS)

    Lorenzo, C. F.; Saus, J. R.; Ray, A.; Carpino, M.; Wu, M.-K.

    1992-01-01

    The concept of life extending control is defined. A brief discussion of current fatigue life prediction methods is given and the need for an alternative life prediction model based on a continuous functional relationship is established. Two approaches to life extending control are considered: (1) the implicit approach which uses cyclic fatigue life prediction as a basis for control design; and (2) the continuous life prediction approach which requires a continuous damage law. Progress on an initial formulation of a continuous (in time) fatigue model is presented. Finally, nonlinear programming is used to develop initial results for life extension for a simplified rocket engine (model).

  4. Rocket engine diagnostics using qualitative modeling techniques

    NASA Technical Reports Server (NTRS)

    Binder, Michael; Maul, William; Meyer, Claudia; Sovie, Amy

    1992-01-01

    Researchers at NASA Lewis Research Center are presently developing qualitative modeling techniques for automated rocket engine diagnostics. A qualitative model of a turbopump interpropellant seal system has been created. The qualitative model describes the effects of seal failures on the system steady-state behavior. This model is able to diagnose the failure of particular seals in the system based on anomalous temperature and pressure values. The anomalous values input to the qualitative model are generated using numerical simulations. Diagnostic test cases include both single and multiple seal failures.

  5. Rocket engine diagnostics using qualitative modeling techniques

    NASA Technical Reports Server (NTRS)

    Binder, Michael; Maul, William; Meyer, Claudia; Sovie, Amy

    1992-01-01

    Researchers at NASA Lewis Research Center are presently developing qualitative modeling techniques for automated rocket engine diagnostics. A qualitative model of a turbopump interpropellant seal system was created. The qualitative model describes the effects of seal failures on the system steady state behavior. This model is able to diagnose the failure of particular seals in the system based on anomalous temperature and pressure values. The anomalous values input to the qualitative model are generated using numerical simulations. Diagnostic test cases include both single and multiple seal failures.

  6. Amateur Gas-Propelled Rocket Engine Development and Advanced Rocket Design

    NASA Astrophysics Data System (ADS)

    Souverein, L. J.; Twigt, D. J.; Engelen, S.

    The paper describes the design and manufacturing of a gaseous propellant rocket engine. It is an undertaking of the authors, performed on project basis with fellow aerospace engineering students under auspices of DARE (Delft Aerospace Rocket Engineering). This paper describes the requirements, the engine development, and the design considerations and calculations as they were performed. Furthermore, the plans for engine tests and the parameters that will have to be measured during those tests are covered. The design process converged to a 1800 N thrust gaseous oxygen/methane (GOX/CH4) engine made of electrolytic copper. GOX/CH4 was selected based on its relatively high specific impulse, its availability and because of its potential as a green propellant. A test engine was produced with a specific impulse of 287 s and a propellant mass flow of 637 g/s. From a point of view of strength, the focus was mainly on robustness rather than light weight. The main aim now is to perform tests with the current engine, based on which the performance can be verified and vital information for future design efforts can be acquired. The ultimate goal is to have an operational rocket and to attempt an amateur altitude record.

  7. Chemical rocket propulsion and the environment

    SciTech Connect

    Mcdonald, A.J. )

    1992-03-01

    Results are presented from the examination by the Chemical Rocket Propulsion and the Environment Workshop conducted by AIAA in June 1991 of the impact of rocket launches and ground testing on the earth's environment. The major conclusions of this workshop were: (1) at projected rocket launch rates, neither the liquid- nor the solid-rocket motors will significantly impact stratospheric ozone; (2) there is no global acid rain problem associated with rocket exhaust; and (3) the local launch site and static test site acidification is a minor problem and can be managed.

  8. Rocket radio measurement of electron density in the nighttime ionosphere

    NASA Technical Reports Server (NTRS)

    Gilchrist, B. E.; Smith, L. G.

    1979-01-01

    One experimental technique based on the Faraday rotation effect of radio waves is presented for measuring electron density in the nighttime ionosphere at midlatitudes. High frequency linearly-polarized radio signals were transmitted to a linearly-polarized receiving system located in a spinning rocket moving through the ionosphere. Faraday rotation was observed in the reference plane of the rocket as a change in frequency of the detected receiver output. The frequency change was measured and the information was used to obtain electron density data. System performance was evaluated and some sources of error were identified. The data obtained was useful in calibrating a Langmuir probe experiment for electron density values of 100/cu cm and greater. Data from two rocket flights are presented to illustrate the experiment.

  9. The Need for Faster Rockets

    NASA Video Gallery

    The rockets NASA has used have been great, but they can’t take us far enough, fast enough to get us to places we haven’t been before like Mars or beyond. A solution is to use plasma as propulsi...

  10. Solid rocket motor witness test

    NASA Technical Reports Server (NTRS)

    Welch, Christopher S.

    1991-01-01

    The Solid Rocket Motor Witness Test was undertaken to examine the potential for using thermal infrared imagery as a tool for monitoring static tests of solid rocket motors. The project consisted of several parts: data acquisition, data analysis, and interpretation. For data acquisition, thermal infrared data were obtained of the DM-9 test of the Space Shuttle Solid Rocket Motor on December 23, 1987, at Thiokol, Inc. test facility near Brigham City, Utah. The data analysis portion consisted of processing the video tapes of the test to produce values of temperature at representative test points on the rocket motor surface as the motor cooled down following the test. Interpretation included formulation of a numerical model and evaluation of some of the conditions of the motor which could be extracted from the data. These parameters included estimates of the insulation remaining following the tests and the thickness of the charred layer of insulation at the end of the test. Also visible was a temperature signature of the star grain pattern in the forward motor segment.

  11. Liquid propellant rocket combustion instability

    NASA Technical Reports Server (NTRS)

    Harrje, D. T.

    1972-01-01

    The solution of problems of combustion instability for more effective communication between the various workers in this field is considered. The extent of combustion instability problems in liquid propellant rocket engines and recommendations for their solution are discussed. The most significant developments, both theoretical and experimental, are presented, with emphasis on fundamental principles and relationships between alternative approaches.

  12. Caney's Third Grade Rocket Program.

    ERIC Educational Resources Information Center

    Warnock, Joe; Hudiburg, George E.

    1984-01-01

    This narrative description of a special program for third-grade students to raise funds for and to construct, launch, and fly their own rockets illustrates benefits of heightened confidence and self-perception among participants, hands-on experience in a highly technical area, and increased community support and involvement. (MM)

  13. Launch Excitement with Water Rockets

    ERIC Educational Resources Information Center

    Sanchez, Juan Carlos; Penick, John

    2007-01-01

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

  14. Centrifugal pumps for rocket engines

    NASA Technical Reports Server (NTRS)

    Campbell, W. E.; Farquhar, J.

    1974-01-01

    The use of centrifugal pumps for rocket engines is described in terms of general requirements of operational and planned systems. Hydrodynamic and mechanical design considerations and techniques and test procedures are summarized. Some of the pump development experiences, in terms of both problems and solutions, are highlighted.

  15. Rocket Ignition Demonstrations Using Silane

    NASA Technical Reports Server (NTRS)

    Pal, Sibtosh; Santoro, Robert; Watkins, William B.; Kincaid, Kevin

    1998-01-01

    Rocket ignition demonstration tests using silane were performed at the Penn State Combustion Research Laboratory. A heat sink combustor with one injection element was used with gaseous propellants. Mixtures of silane and hydrogen were used as fuel, and oxygen was used as oxidizer. Reliable ignition was demonstrated using fuel lead and and a swirl injection element.

  16. Laser-heated rocket studies

    NASA Technical Reports Server (NTRS)

    Kemp, N. H.; Root, R. G.; Wu., P. K. S.; Caledonia, G. E.; Pirri, A. N.

    1976-01-01

    CW laser heated rocket propulsion was investigated in both the flowing core and stationary core configurations. The laser radiation considered was 10.6 micrometers, and the working gas was unseeded hydrogen. The areas investigated included initiation of a hydrogen plasma capable of absorbing laser radiation, the radiation emission properties of hot, ionized hydrogen, the flow of hot hydrogen while absorbing and radiating, the heat losses from the gas and the rocket performance. The stationary core configuration was investigated qualitatively and semi-quantitatively. It was found that the flowing core rockets can have specific impulses between 1,500 and 3,300 sec. They are small devices, whose heating zone is only a millimeter to a few centimeters long, and millimeters to centimeters in radius, for laser power levels varying from 10 to 5,000 kW, and pressure levels of 3 to 10 atm. Heat protection of the walls is a vital necessity, though the fraction of laser power lost to the walls can be as low as 10% for larger powers, making the rockets thermally efficient.

  17. The solid-core heat-exchanger nuclear rocket program

    SciTech Connect

    Malenfant, R.E.

    1994-12-31

    As measured by the results of its accomplishments, the nuclear rocket program was a success. Why, then, was it cancelled? In my opinion, the cancellation resulted from the success of the Apollo program. President Kennedy declared that putting a man on the moon by 1969 would be a national objective. Upon the Apollo program`s completion, space spectaculars lost their attraction, and the manned exploration of Mars, which could have been accomplished with nuclear rockets, was shelved. Perhaps another generation of physicists and engineers will experience the thrill and satisfaction of participating in a nuclear-propulsion-based program for space exploration in decades to come.

  18. An Experiment on Repetitive Pulse Operation of Microwave Rocket

    SciTech Connect

    Oda, Yasuhisa; Shibata, Teppei; Komurasaki, Kimiya; Takahashi, Koji; Kasugai, Atsushi; Sakamoto, Keishi

    2008-04-28

    Microwave Rocket was operated with repetitive pulses. The microwave rocket model with forced breathing system was used. The pressure history in the thruster was measured and the thrust impulse was deduced. As a result, the impulse decreased at second pulse and impulses at latter pulses were constant. The dependence of the thrust performance on the partial filling rate of the thruster was compared to the thrust generation model based on the shock wave driven by microwave plasma. The experimental results showed good agreement to the predicted dependency.

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

    NASA Astrophysics Data System (ADS)

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

    2016-10-01

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

  20. Modeling of vortex generated sound in solid propellant rocket motors

    NASA Technical Reports Server (NTRS)

    Flandro, G. A.

    1980-01-01

    There is considerable evidence based on both full scale firings and cold flow simulations that hydrodynamically unstable shear flows in solid propellant rocket motors can lead to acoustic pressure fluctuations of significant amplitude. Although a comprehensive theoretical understanding of this problem does not yet exist, procedures were explored for generating useful analytical models describing the vortex shedding phenomenon and the mechanisms of coupling to the acoustic field in a rocket combustion chamber. Since combustion stability prediction procedures cannot be successful without incorporation of all acoustic gains and losses, it is clear that a vortex driving model comparable in quality to the analytical models currently employed to represent linear combustion instability must be formulated.

  1. Effects of high combustion chamber pressure on rocket noise environment

    NASA Technical Reports Server (NTRS)

    Pao, S. P.

    1972-01-01

    The acoustical environment for a high combustion chamber pressure engine was examined in detail, using both conventional and advanced theoretical analysis. The influence of elevated chamber pressure on the rocket noise environment was established, based on increase in exit velocity and flame temperature, and changes in basic engine dimensions. Compared to large rocket engines, the overall sound power level is found to be 1.5 dB higher, if the thrust is the same. The peak Strouhal number shifted about one octave lower to a value near 0.01. Data on apparent sound source location and directivity patterns are also presented.

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

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

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

  3. An Object Model for a Rocket Engine Numerical Simulator

    NASA Technical Reports Server (NTRS)

    Mitra, D.; Bhalla, P. N.; Pratap, V.; Reddy, P.

    1998-01-01

    Rocket Engine Numerical Simulator (RENS) is a packet of software which numerically simulates the behavior of a rocket engine. Different parameters of the components of an engine is the input to these programs. Depending on these given parameters the programs output the behaviors of those components. These behavioral values are then used to guide the design of or to diagnose a model of a rocket engine "built" by a composition of these programs simulating different components of the engine system. In order to use this software package effectively one needs to have a flexible model of a rocket engine. These programs simulating different components then should be plugged into this modular representation. Our project is to develop an object based model of such an engine system. We are following an iterative and incremental approach in developing the model, as is the standard practice in the area of object oriented design and analysis of softwares. This process involves three stages: object modeling to represent the components and sub-components of a rocket engine, dynamic modeling to capture the temporal and behavioral aspects of the system, and functional modeling to represent the transformational aspects. This article reports on the first phase of our activity under a grant (RENS) from the NASA Lewis Research center. We have utilized Rambaugh's object modeling technique and the tool UML for this purpose. The classes of a rocket engine propulsion system are developed and some of them are presented in this report. The next step, developing a dynamic model for RENS, is also touched upon here. In this paper we will also discuss the advantages of using object-based modeling for developing this type of an integrated simulator over other tools like an expert systems shell or a procedural language, e.g., FORTRAN. Attempts have been made in the past to use such techniques.

  4. Integrated Turbine-Based Combined Cycle Dynamic Simulation Model

    NASA Technical Reports Server (NTRS)

    Haid, Daniel A.; Gamble, Eric J.

    2011-01-01

    A Turbine-Based Combined Cycle (TBCC) dynamic simulation model has been developed to demonstrate all modes of operation, including mode transition, for a turbine-based combined cycle propulsion system. The High Mach Transient Engine Cycle Code (HiTECC) is a highly integrated tool comprised of modules for modeling each of the TBCC systems whose interactions and controllability affect the TBCC propulsion system thrust and operability during its modes of operation. By structuring the simulation modeling tools around the major TBCC functional modes of operation (Dry Turbojet, Afterburning Turbojet, Transition, and Dual Mode Scramjet) the TBCC mode transition and all necessary intermediate events over its entire mission may be developed, modeled, and validated. The reported work details the use of the completed model to simulate a TBCC propulsion system as it accelerates from Mach 2.5, through mode transition, to Mach 7. The completion of this model and its subsequent use to simulate TBCC mode transition significantly extends the state-of-the-art for all TBCC modes of operation by providing a numerical simulation of the systems, interactions, and transient responses affecting the ability of the propulsion system to transition from turbine-based to ramjet/scramjet-based propulsion while maintaining constant thrust.

  5. Ultrasonic inspection of rocket fuel model using laminated transducer and multi-channel step pulser

    NASA Astrophysics Data System (ADS)

    Mihara, T.; Hamajima, T.; Tashiro, H.; Sato, A.

    2013-01-01

    For the ultrasonic inspection for the packing of solid fuel in a rocket booster, an industrial inspection is difficult. Because the signal to noise ratio in ultrasonic inspection of rocket fuel become worse due to the large attenuation even using lower frequency ultrasound. For the improvement of this problem, we tried to applied the two techniques in ultrasonic inspection, one was the step function pulser system with the super wideband frequency properties and the other was the laminated element transducer. By combining these two techniques, we developed the new ultrasonic measurement system and demonstrated the advantages in ultrasonic inspection of rocket fuel model specimen.

  6. 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.

  7. The colour preference control based on two-colour combinations

    NASA Astrophysics Data System (ADS)

    Hong, Ji Young; Kwak, Youngshin; Park, Du-Sik; Kim, Chang Yeong

    2008-02-01

    This paper proposes a framework of colour preference control to satisfy the consumer's colour related emotion. A colour harmony algorithm based on two-colour combinations is developed for displaying the images with several complementary colour pairs as the relationship of two-colour combination. The colours of pixels belonging to complementary colour areas in HSV colour space are shifted toward the target hue colours and there is no colour change for the other pixels. According to the developed technique, dynamic emotions by the proposed hue conversion can be improved and the controlled output image shows improved colour emotions in the preference of the human viewer. The psychophysical experiments are conducted to investigate the optimal model parameters to produce the most pleasant image to the users in the respect of colour emotions.

  8. A rocket-borne energy spectrometer using multiple solid-state detectors for particle identification

    NASA Technical Reports Server (NTRS)

    Fries, K. L.; Smith, L. G.; Voss, H. D.

    1979-01-01

    A rocket-borne experiment using energy spectrometers that allows particle identification by the use of multiple solid-state detectors is described. The instrumentation provides information regarding the energy spectrum, pitch-angle distribution, and the type of energetic particles present in the ionosphere. Particle identification was accomplished by considering detector loss mechanisms and their effects on various types of particles. Solid state detectors with gold and aluminum surfaces of several thicknesses were used. The ratios of measured energies for the various detectors were compared against known relationships during ground-based analysis. Pitch-angle information was obtained by using detectors with small geometrical factors mounted with several look angles. Particle flux was recorded as a function of rocket azimuth angle. By considering the rocket azimuth, the rocket precession, and the location of the detectors on the rocket, the pitched angle of the incident particles was derived.

  9. Active chlorine and nitric oxide formation from chemical rocket plume afterburning

    NASA Technical Reports Server (NTRS)

    Leone, D. M.; Turns, S. R.

    1994-01-01

    Chlorine and oxides of nitrogen (NO(x)) released into the atmosphere contribute to acid rain (ground level or low-altitude sources) and ozone depletion from the stratosphere (high-altitude sources). Rocket engines have the potential for forming or activating these pollutants in the rocket plume. For instance, H2/O2 rockets can produce thermal NO(x) in their plumes. Emphasis, in the past, has been placed on determining the impact of chlorine release on the stratosphere. To date, very little, if any, information is available to understand what contribution NO(x) emissions from ground-based engine testing and actual rocket launches have on the atmosphere. The goal of this work is to estimate the afterburning emissions from chemical rocket plumes and determine their local stratospheric impact. Our study focuses on the space shuttle rocket motors, which include both the solid rocket boosters (SRB's) and the liquid propellant main engines (SSME's). Rocket plume afterburning is modeled employing a one-dimensional model incorporating two chemical kinetic systems: chemical and thermal equilibria with overlayed nitric oxide chemical kinetics (semi equilibrium) and full finite-rate chemical kinetics. Additionally, the local atmospheric impact immediately following a launch is modeled as the emissions diffuse and chemically react in the stratosphere.

  10. Small-scale plasma turbulence and intermittency in the high latitude F region based on the ICI-2 sounding rocket experiment

    NASA Astrophysics Data System (ADS)

    Spicher, A.; Miloch, W.; Moen, J. I.; Clausen, L. B. N.

    2015-12-01

    Small-scale plasma irregularities and turbulence are common phenomena in the F layer of the ionosphere, both in the equatorial and polar regions. A common approach in analyzing data from experiments on space and ionospheric plasma irregularities are power spectra. Power spectra give no information about the phases of the waveforms, and thus do not allow to determine whether some of the phases are correlated or whether they exhibit a random character. The former case would imply the presence of nonlinear wave-wave interactions, while the latter suggests a more turbulent-like process. Discerning between these mechanisms is crucial for understanding high latitude plasma irregularities and can be addressed with bispectral analysis and higher order statistics. In this study, we use higher order spectra and statistics to analyze electron density data observed with the ICI-2 sounding rocket experiment at a meter-scale resolution. The main objective of ICI-2 was to investigate plasma irregularities in the cusp in the F layer ionosphere. We study in detail two regions intersected during the rocket flight and which are characterized by large density fluctuations: a trailing edge of a cold polar cap patch, and a density enhancement subject to cusp auroral particle precipitation. While these two regions exhibit similar power spectra, our analysis reveals that their internal structure is different. The structures on the edge of the polar cap patch are characterized by significant coherent mode coupling and intermittency, while the plasma enhancement associated with precipitation exhibits stronger random characteristics. This indicates that particle precipitation may play a fundamental role in ionospheric plasma structuring by creating turbulent-like structures.

  11. The Peak of Rocket Production: The Designer of Ballistic Missiles V.F. Utkin (1923-2000)

    NASA Astrophysics Data System (ADS)

    Prisniakov, V.; Sitnikova, N.

    2002-01-01

    The main landmarks of the biography of the general designer of the most power missiles Vladimira Fe- dorovicha Utkina are stated. Formation of character of outstanding scientist of 20 century as the son of the Soviet epoch is shown. He belongs to that generation which had many difficulties and afflictions - hungry time, the heavy years of the second world war, post-war disruption, but also many happy days - the Victory above fascism, restoration of the country, pride of successes in conquest of Space. In June, 1941 Vladimir has finished school with honours certificate and since October, 1941 up to the end of the second world war was on various fronts. After the ending of Leningrad military-mechanical insti- tute the young engineer came in Southern engineering works in Dnipropetrovsk (Yugmach). Here for 40 years there was a dizzy ascent of the beginner-designer over a ladder of space-rocket's Olympus up to the chief designer and the general director of the biggest in the world of rocket concern (9000 high quality engineers of Design office Yugnoe (DOYu) and 60 thousand workers Yugmach). After death in 1971 to year of main designer M. Jangele V. F. Utkin has headed Design office Yugnoe. Under ma- nual of V. Utkin four strategic rocket complexes of new generation SS-17, SS-18 (three updatings with divided head parts with weight of 8 tons), SS-24 (railway and shaft basing) were developed and han- ded over on arms. Among development of academician V. Utkin there is a rocket-carrier "Zenit" which delivers to an orbit over 12 tons of a payload. This rocket is also a basis of the first stage of reusable transport space system "Energia-Buran". Under manual of V. Utkin were developed and used the con- version carrier-rockets "Ziclon" and "Kosmos", as well the effective satellites of a defensive, scientific and economic direction, among which family of satellites "Kosmos", the satellite "Ocean", equipped by a locator of the side observation too. Largest scientific

  12. Damage-mitigating control of a reusable rocket engine for high performance and extended life

    NASA Technical Reports Server (NTRS)

    Ray, Asok; Dai, Xiaowen

    1995-01-01

    The goal of damage mitigating control in reusable rocket engines is to achieve high performance with increased durability of mechanical structures such that functional lives of the critical components are increased. The major benefit is an increase in structural durability with no significant loss of performance. This report investigates the feasibility of damage mitigating control of reusable rocket engines. Phenomenological models of creep and thermo-mechanical fatigue damage have been formulated in the state-variable setting such that these models can be combined with the plant model of a reusable rocket engine, such as the Space Shuttle Main Engine (SSME), for synthesizing an optimal control policy. Specifically, a creep damage model of the main thrust chamber wall is analytically derived based on the theories of sandwich beam and viscoplasticity. This model characterizes progressive bulging-out and incremental thinning of the coolant channel ligament leading to its eventual failure by tensile rupture. The objective is to generate a closed form solution of the wall thin-out phenomenon in real time where the ligament geometry is continuously updated to account for the resulting deformation. The results are in agreement with those obtained from the finite element analyses and experimental observation for both Oxygen Free High Conductivity (OFHC) copper and a copper-zerconium-silver alloy called NARloy-Z. Due to its computational efficiency, this damage model is suitable for on-line applications of life prediction and damage mitigating control, and also permits parametric studies for off-line synthesis of damage mitigating control systems. The results are presented to demonstrate the potential of life extension of reusable rocket engines via damage mitigating control. The control system has also been simulated on a testbed to observe how the damage at different critical points can be traded off without any significant loss of engine performance. The research work

  13. Nuclear Thermal Rocket Simulation in NPSS

    NASA Technical Reports Server (NTRS)

    Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas L.

    2013-01-01

    Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic- metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

  14. Nuclear Thermal Rocket Simulation in NPSS

    NASA Technical Reports Server (NTRS)

    Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas M.

    2013-01-01

    Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic-metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

  15. Rocket thrust chamber thermal barrier coatings

    NASA Technical Reports Server (NTRS)

    Batakis, A. P.; Vogan, J. W.

    1985-01-01

    A research program was conducted to generate data and develop analytical techniques to predict the performance and reliability of ceramic thermal barrier coatings in high heat flux environments. A finite element model was used to analyze the thermomechanical behavior of coating systems in rocket thrust chambers. Candidate coating systems (using a copper substrate, NiCrAlY bond coat and ZrO2.8Y2O3 ceramic overcoat) were selected for detailed study based on photomicrographic evaluations of experimental test specimens. The effects of plasma spray application parameters on the material properties of these coatings were measured and the effects on coating performance evaluated using the finite element model. Coating design curves which define acceptable operating envelopes for seleted coating systems were constructed based on temperature and strain limitations. Spray gun power levels was found to have the most significant effect on coating structure. Three coating systems were selected for study using different power levels. Thermal conductivity, strain tolerance, density, and residual stress were measured for these coatings. Analyses indicated that extremely thin coatings ( 0.02 mm) are required to accommodate the high heat flux of a rocket thrust chamber and ensure structural integrity.

  16. Production of Nitrous Oxide in a Rocket Motor Exhaust.

    DTIC Science & Technology

    1980-08-01

    Standard Atmosphere, 1962. NASA , USAF and United States Weather Bureau, 1962. 2 Jensen, D.E. Prediction of Rocket Exhaust Flame Properties. Wilson, A.S...are made of concentrations of N20 produced within the exhaust of a double-base propellant rocet motor. Typical concentrations produced are predicted to

  17. ISHM Anomaly Lexicon for Rocket Test

    NASA Technical Reports Server (NTRS)

    Schmalzel, John L.; Buchanan, Aubri; Hensarling, Paula L.; Morris, Jonathan; Turowski, Mark; Figueroa, Jorge F.

    2007-01-01

    byproducts of the anomaly lexicon compilation effort. For example, (1) Allows determination of the frequency distribution of anomalies to help identify those with the potential for high return on investment if included in automated detection as part of an ISHM system, (2) Availability of a regular lexicon could provide the base anomaly name choices to help maintain consistency in the DR collection process, and (3) Although developed for the rocket engine test environment, most of the anomalies are not specific to rocket testing, and thus can be reused in other applications.

  18. Rocket Radiation Handbook. Volume I. Rocket Radiation Phenomenology and Theory

    DTIC Science & Technology

    1974-06-01

    were rather Incomplete and unsatisfactory causing engineers In the field to become skeptical of any theoretical work. As a result the v FTD-CW-0 I-0 1...investigator In the program which was carried out under the general technical direction of Mr. H. Lopez, Director of Engineering . DISTRIBUTION The...material can be used in other fields of applied physics and engineering . The new Rocket Radiation Handbook contains the results of six years of

  19. Progress on the Micro-X rocket payload

    NASA Astrophysics Data System (ADS)

    Figueroa-Feliciano, E.; Wikus, P.; Adams, J. S.; Bandler, S. R.; Bautz, M.; Boyce, K.; Brown, G.; Deiker, S.; Doriese, W. B.; Flanagan, K.; Galeazzi, M.; Hilton, G. C.; Hwang, U.; Irwin, K. D.; Kallman, T.; Kelley, R. L.; Kilbourne, C. A.; Kissel, S.; Leman, S. W.; Levine, A.; Lowenstein, M.; Martinez-Galarce, D.; McCammon, D.; Mushotzky, R.; Petre, R.; Porter, F. S.; Reintsema, C. D.; Rutherford, J. M.; Saab, T.; Serlemitsos, P.; Schulz, N.; Smith, R.; Ullom, J. N.

    2008-07-01

    The Micro-X High Resolution Microcalorimeter X-ray Imaging Rocket is sounding rocket experiment that will combine a transition-edge-sensor X-ray-microcalorimeter array with a conical imaging mirror to obtain high-spectral-resolution images of extended and point X-ray sources. Our first target is the Puppis A supernova remnant, which will be observed in January 2011. The Micro-X observation of the bright eastern knot of Puppis A will obtain a line-dominated spectrum with up to 90,000 counts collected in 300 seconds at 2 eV resolution across the 0.3-2.5 keV band. Micro-X will utilize plasma diagnostics to determine the thermodynamic and ionization state of the plasma, to search for line shifts and broadening associated with dynamical processes, and seek evidence of ejecta enhancement. We describe the progress made in developing this payload, including the detector, cryogenics, and electronics assemblies. A detailed modeling effort has been undertaken to design a rocket-bourne adiabatic demagnetization refrigerator with sufficient magnetic shielding to allow stable operation of transition edge sensors, and the associated rocket electronics have been prototyped and tested.

  20. Grease-Resistant O Rings for Joints in Solid Rocket Motors

    NASA Technical Reports Server (NTRS)

    Harvey, Albert R.; Feldman, Harold

    2003-01-01

    There is a continuing effort to develop improved O rings for sealing joints in solid-fuel rocket motors. Following an approach based on the lessons learned in the explosion of the space shuttle Challenger, investigators have been seeking O-ring materials that exhibit adequate resilience for effective sealing over a broad temperature range: What are desired are O rings that expand far and fast enough to maintain seals, even when metal sealing surfaces at a joint move slightly away from each other shortly after ignition and the motor was exposed to cold weather before ignition. Other qualities desired of the improved O rings include adequate resistance to ablation by hot rocket gases and resistance to swelling when exposed to hydrocarbon-based greases used to protect some motor components against corrosion. Five rubber formulations two based on a fluorosilicone polymer and three based on copolymers of epichlorohydrin with ethylene oxide were tested as candidate O-ring materials. Of these, one of the epichlorohydrin/ethylene oxide formulations was found to offer the closest to the desired combination of properties and was selected for further evaluation.

  1. A software-based sensor for combined sewer overflows.

    PubMed

    Leonhardt, G; Fach, S; Engelhard, C; Kinzel, H; Rauch, W

    2012-01-01

    A new methodology for online estimation of excess flow from combined sewer overflow (CSO) structures based on simulation models is presented. If sufficient flow and water level data from the sewer system is available, no rainfall data are needed to run the model. An inverse rainfall-runoff model was developed to simulate net rainfall based on flow and water level data. Excess flow at all CSO structures in a catchment can then be simulated with a rainfall-runoff model. The method is applied to a case study and results show that the inverse rainfall-runoff model can be used instead of missing rain gauges. Online operation is ensured by software providing an interface to the SCADA-system of the operator and controlling the model. A water quality model could be included to simulate also pollutant concentrations in the excess flow.

  2. Aerodynamics of sounding rockets at supersonic speeds

    NASA Astrophysics Data System (ADS)

    Vira, N. R.

    This dissertation presents a practical and low cost method of computing the aerodynamic characteristics of vehicles such as sounding rockets, high speed bombs, projectiles and guided missiles in supersonic flight. The vehicle configuration consists of a slender axisymmetric body with a conical or ogive noise, cylinders, shoulders and boattails, if any, and have sets of two, three or four fins. Geometry of the fin cross section can be single wedge, double wedge, modified single wedge or modified double wedge. First the aerodynamics of the fins and the body are analyzed separately; then fin body and fore and aft fin interferences are accounted for when they are combined to form the total vehicle. Results and formulas documented in this work are the basis of the supersonic portion of the Theoretical Aerodynamic Derivatives (TAD) computer program operating at the NASA Goddard Space Flight Center.

  3. Small centrifugal pumps for low thrust rockets

    NASA Technical Reports Server (NTRS)

    Gulbrandsen, N. C.; Furst, R. B.; Burgess, R. M.; Scheer, D. D.

    1985-01-01

    This paper presents the results of a combined analytical and experimental investigation of low specific speed pumps for potential use as components of propellant feed systems for low thrust rocket engines. Shrouded impellers and open face impellers were tested in volute type and vaned diffuser type pumps. Full- and partial-emission diffusers and full- and partial-admission impellers were tested. Axial and radial loads, head and efficiency versus flow, and cavitation tests were conducted. Predicted performance of two pumps are compared when pumping water and liquid hydrogen. Detailed pressure loss and parasitic power values are presented for two pump configurations. Partial-emission diffusers were found to permit use of larger impeller and diffuser passages with a minimal performance penalty. Normal manufacturing tolerances were found to result in substantial power requirement variation with only a small pressure rise change. Impeller wear ring leakage was found to reduce pump pressure rise to an increasing degree as the pump flowrate was decreased.

  4. Symbolic Processing Combined with Model-Based Reasoning

    NASA Technical Reports Server (NTRS)

    James, Mark

    2009-01-01

    A computer program for the detection of present and prediction of future discrete states of a complex, real-time engineering system utilizes a combination of symbolic processing and numerical model-based reasoning. One of the biggest weaknesses of a purely symbolic approach is that it enables prediction of only future discrete states while missing all unmodeled states or leading to incorrect identification of an unmodeled state as a modeled one. A purely numerical approach is based on a combination of statistical methods and mathematical models of the applicable physics and necessitates development of a complete model to the level of fidelity required for prediction. In addition, a purely numerical approach does not afford the ability to qualify its results without some form of symbolic processing. The present software implements numerical algorithms to detect unmodeled events and symbolic algorithms to predict expected behavior, correlate the expected behavior with the unmodeled events, and interpret the results in order to predict future discrete states. The approach embodied in this software differs from that of the BEAM methodology (aspects of which have been discussed in several prior NASA Tech Briefs articles), which provides for prediction of future measurements in the continuous-data domain.

  5. SCORE - Sounding-rocket Coronagraphic Experiment

    NASA Astrophysics Data System (ADS)

    Fineschi, Silvano; Moses, Dan; Romoli, Marco

    The Sounding-rocket Coronagraphic Experiment - SCORE - is a The Sounding-rocket Coronagraphic Experiment - SCORE - is a coronagraph for multi-wavelength imaging of the coronal Lyman-alpha lines, HeII 30.4 nm and HI 121.6 nm, and for the broad.band visible-light emission of the polarized K-corona. SCORE has flown successfully in 2009 acquiring the first images of the HeII line-emission from the extended corona. The simultaneous observation of the coronal Lyman-alpha HI 121.6 nm, has allowed the first determination of the absolute helium abundance in the extended corona. This presentation will describe the lesson learned from the first flight and will illustrate the preparations and the science perspectives for the second re-flight approved by NASA and scheduled for 2016. The SCORE optical design is flexible enough to be able to accommodate different experimental configurations with minor modifications. This presentation will describe one of such configurations that could include a polarimeter for the observation the expected Hanle effect in the coronal Lyman-alpha HI line. The linear polarization by resonance scattering of coronal permitted line-emission in the ultraviolet (UV) can be modified by magnetic fields through the Hanle effect. Thus, space-based UV spectro-polarimetry would provide an additional new tool for the diagnostics of coronal magnetism.

  6. Software for Collaborative Engineering of Launch Rockets

    NASA Technical Reports Server (NTRS)

    Stanley, Thomas Troy

    2003-01-01

    The Rocket Evaluation and Cost Integration for Propulsion and Engineering software enables collaborative computing with automated exchange of information in the design and analysis of launch rockets and other complex systems. RECIPE can interact with and incorporate a variety of programs, including legacy codes, that model aspects of a system from the perspectives of different technological disciplines (e.g., aerodynamics, structures, propulsion, trajectory, aeroheating, controls, and operations) and that are used by different engineers on different computers running different operating systems. RECIPE consists mainly of (1) ISCRM a file-transfer subprogram that makes it possible for legacy codes executed in their original operating systems on their original computers to exchange data and (2) CONES an easy-to-use filewrapper subprogram that enables the integration of legacy codes. RECIPE provides a tightly integrated conceptual framework that emphasizes connectivity among the programs used by the collaborators, linking these programs in a manner that provides some configuration control while facilitating collaborative engineering tradeoff studies, including design to cost studies. In comparison with prior collaborative-engineering schemes, one based on the use of RECIPE enables fewer engineers to do more in less time.

  7. Nitrous oxide cooling in hybrid rocket nozzles

    NASA Astrophysics Data System (ADS)

    Lemieux, Patrick

    2010-02-01

    The Department of Mechanical Engineering at the California Polytechnic State University, San Luis Obispo, has developed an innovative program of experimental research and development on hybrid rocket motors (where the fuel and the oxidizer are in different phases prior to combustion). One project currently underway involves the development of aerospike nozzles for such motors. These nozzles, however, are even more susceptible to throat ablation than regular converging-diverging nozzles, due the nature of their flow expansion mechanism. This paper presents the result of a recent development project focused on reducing throat ablation in hybrid rocket motor nozzles. Although the method is specifically targeted at increasing the life and operating range of aerospike nozzles, this paper describes its proof-of-concept implementation on conventional nozzles. The method is based on a regenerative cooling mechanism that differs in practice from that used in liquid propellant motors. A series of experimental tests demonstrate that this new method is not only effective at reducing damage in the most ablative region of the nozzle, but that the nozzle can survive multiple test runs.

  8. Space Shuttle solid rocket booster

    NASA Technical Reports Server (NTRS)

    Hardy, G. B.

    1979-01-01

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

  9. Low thrust rocket test facility

    NASA Technical Reports Server (NTRS)

    Arrington, Lynn A.; Schneider, Steven J.

    1990-01-01

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

  10. Expendable solid rocket motor upper stages for the Space Shuttle

    NASA Technical Reports Server (NTRS)

    Davis, H. P.; Jones, C. M.

    1974-01-01

    A family of expendable solid rocket motor upper stages has been conceptually defined to provide the payloads for the Space Shuttle with performance capability beyond the low earth operational range of the Shuttle Orbiter. In this concept-feasibility assessment, three new solid rocket motors of fixed impulse are defined for use with payloads requiring levels of higher energy. The conceptual design of these motors is constrained to limit thrusting loads into the payloads and to conserve payload bay length. These motors are combined in various vehicle configurations with stage components derived from other programs for the performance of a broad range of upper-stage missions from spin-stabilized, single-stage transfers to three-axis stabilized, multistage insertions. Estimated payload delivery performance and combined payload mission loading configurations are provided for the upper-stage configurations.

  11. Theory for Plasma Rocket Propulsion

    NASA Astrophysics Data System (ADS)

    Grabbe, Crockett

    2009-11-01

    Electrical propulsion of rockets is developing potentially into the use of 3 different thrusters for future long-distance space missions that primarily involve plasma dynamics. These are the Magnetoplasmadynamic (MPD) Thruster, the Plasma Induction Thruster (PID), and the VASIMIR Thruster. The history of the development of electrical propulsion into these prospects and the current research of particularly the VASIMIR Thruster are reviewed. Theoretical questions that need to be addressed in that development are explored.

  12. Mini-Rocket User Guide

    DTIC Science & Technology

    2007-08-01

    Missile Research , Development, and Engineering Center and Ray Sells DESE Research , Inc. 315 Wynn Drive Huntsville, AL 35805 August 2007...with the minirock command, you are prompted for a filename: Mini-Rocket v1.01 by Ray Sells, DESE Research , Inc. Input file: - Output is printed...nancv.bucher@us.army.mil Commander, U.S. Army ARDEC Picatinny Arsenal, NJ 07806-5000 ATTN: AMSRD-AR-AIS -SA DESE Research , Inc. 3 15 Wynn Drive

  13. Rocket Altitude Test Facilities Register

    DTIC Science & Technology

    1987-03-01

    Classification of Document UNCLASSIFIED 5. Originator Advisory Group for Aerospace Research and Development North Atlantic Treaty Organization...Emphasis was put on facilities capable of performing research and development tests. This AGARDograph was prepared at the request of the Propulsion... RESEARCH & DEVELOPMENT 7RUEANCELLE 92200 NEUILLY SUR SEINE FRANCE AGARDo^raph N0^97 , Rocket Altitude Test Facilities Register /^ri c^ris

  14. Liquid rocket engine turbopump gears

    NASA Technical Reports Server (NTRS)

    1974-01-01

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

  15. Combined Feature Based and Shape Based Visual Tracker for Robot Navigation

    NASA Technical Reports Server (NTRS)

    Deans, J.; Kunz, C.; Sargent, R.; Park, E.; Pedersen, L.

    2005-01-01

    We have developed a combined feature based and shape based visual tracking system designed to enable a planetary rover to visually track and servo to specific points chosen by a user with centimeter precision. The feature based tracker uses invariant feature detection and matching across a stereo pair, as well as matching pairs before and after robot movement in order to compute an incremental 6-DOF motion at each tracker update. This tracking method is subject to drift over time, which can be compensated by the shape based method. The shape based tracking method consists of 3D model registration, which recovers 6-DOF motion given sufficient shape and proper initialization. By integrating complementary algorithms, the combined tracker leverages the efficiency and robustness of feature based methods with the precision and accuracy of model registration. In this paper, we present the algorithms and their integration into a combined visual tracking system.

  16. 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.

  17. Rocket Engine Altitude Simulation Technologies

    NASA Technical Reports Server (NTRS)

    Woods, Jody L.; Lansaw, John

    2010-01-01

    John C. Stennis Space Center is embarking on a very ambitious era in its rocket engine propulsion test history. The first new large rocket engine test stand to be built at Stennis Space Center in over 40 years is under construction. The new A3 Test Stand is designed to test very large (294,000 Ibf thrust) cryogenic propellant rocket engines at a simulated altitude of 100,000 feet. A3 Test Stand will have an engine testing chamber where the engine will be fired after the air in the chamber has been evacuated to a pressure at the simulated altitude of less than 0.16 PSIA. This will result in a very unique environment with extremely low pressures inside a very large chamber and ambient pressures outside this chamber. The test chamber is evacuated of air using a 2-stage diffuser / ejector system powered by 5000 lb/sec of steam produced by 27 chemical steam generators. This large amount of power and flow during an engine test will result in a significant acoustic and vibrational environment in and around A3 Test Stand.

  18. Two-dimensional motions of rockets

    NASA Astrophysics Data System (ADS)

    Kang, Yoonhwan; Bae, Saebyok

    2007-01-01

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

  19. Measuring Model Rocket Engine Thrust Curves

    NASA Astrophysics Data System (ADS)

    Penn, Kim; Slaton, William V.

    2010-12-01

    This paper describes a method and setup to quickly and easily measure a model rocket engine's thrust curve using a computer data logger and force probe. Horst describes using Vernier's LabPro2 and force probe to measure the rocket engine's thrust curve; however, the method of attaching the rocket to the force probe is not discussed. We show how a simple engine holder can be constructed and used with Vernier's LabPro and force probe to record data that students can use to compare to sample data from the rocket manufacturer or the National Association of Rocketry's3 engine certification sheets, calculate total impulse, and make predictions for model rocket launches. PASCO markets a rocket engine test bracket4 that mounts to its PASPORT force sensor for similar measurements. The engine holder described here is very economical, and all the parts can be obtained from a local hardware store or home center.

  20. Conjugate conductive, convective, and radiative heat transfer in rocket engines

    SciTech Connect

    Naraghi, M.H.N.; DeLise, J.C.

    1995-12-31

    A comprehensive conductive, convective and radiative model for thermal analysis of rocket thrust chambers and nozzles is presented. In this model, the rocket thrust chamber and nozzle are subdivided into a number of stations along the longitudinal direction. At each station a finite element scheme is used to evaluate wall temperature distribution. The hot-gas-side convective heat transport is evaluated by numerically solving the compressible boundary layer equations and the radiative fluxes are evaluated by implementing an exchange factor scheme. The convective heat flux in the cooling channel is modeled based on the existing closed form correlations for rocket cooling channels. The conductive, convective and radiative processes are conjugated through an iterative procedure. The hot-gas-side heat transfer coefficients evaluated based on this model are compared to the experimental results reported in the literature. The computed convective heat transfer coefficients agree very well with experimental data for most of the engine except the throat where a discrepancy of approximately 20% exists. The model is applied to a typical regeneratively cooled rocket engine and the resulting wall temperature and heat flux distribution are presented.

  1. Pressure Fed Nuclear Thermal Rockets for space missions

    SciTech Connect

    Leyse, C.F. , Idaho Falls, ID ); Madsen, W.W.; Ramsthaler, J.H.; Schnitzler, B.G. )

    1989-08-01

    The National Space Policy includes a long range goal of expanding human presence and activity beyond Earth orbit into the solar system. This has renewed interest in the potential application of Nuclear Thermal Rockets (NTR) to space flight, particularly for human expeditions to the Moon and Mars. Recent NASA studies consider applications of the previously developed NERVA (Nuclear Engine for Rocket Vehicle Application) technology and the more advanced gas core reactors and show their potential advantages in reducing the initial mass in Earth orbit (IMEO) compared to advanced chemical rocket engines. Application of NERVA technology will require reestablishing the prior technological base or extending it to an advanced NERVA type engine, while the gas core NTR will require an extensive high risk research and development program. A technology intermediate between NERVA and the gas core NTR is a low pressure engine based on solid fuel, a Pressure Fed NTR (PFNTR). In addition to the simplicity of the gas pressurized engine cycle, the PFNTR takes advantage of the dissociation of hydrogen-the increases in specific impulse become significant as the chamber pressure decreases below 1.0 MPa (10 atmospheres) and the chamber temperature increases above 3000 K. The developmental status of technology applicable to a Pressure Fed Nuclear Thermal Rocket (PFNTR) lies between that of the NERVA engine and the gas core NTR (GCNTR). This document investigates PFNTR performance and provides typical mission analyses.

  2. Cooling of in-situ propellant rocket engines for Mars mission. M.S. Thesis - Cleveland State Univ.

    NASA Technical Reports Server (NTRS)

    Armstrong, Elizabeth S.

    1991-01-01

    One propulsion option of a Mars ascent/descent vehicle is multiple high-pressure, pump-fed rocket engines using in-situ propellants, which have been derived from substances available on the Martian surface. The chosen in-situ propellant combination for this analysis is carbon monoxide as the fuel and oxygen as the oxidizer. Both could be extracted from carbon dioxide, which makes up 96 percent of the Martian atmosphere. A pump-fed rocket engine allows for higher chamber pressure than a pressure-fed engine, which in turn results in higher thrust and in higher heat flux in the combustion chamber. The heat flowing through the wall cannot be sufficiently dissipated by radiation cooling and, therefore, a regenerative coolant may be necessary to avoid melting the rocket engine. The two possible fluids for this coolant scheme, carbon monoxide and oxygen, are compared analytically. To determine their heat transfer capability, they are evaluated based upon their heat transfer and fluid flow characteristics.

  3. Reusable rocket engine optical condition monitoring

    NASA Technical Reports Server (NTRS)

    Wyett, L.; Maram, J.; Barkhoudarian, S.; Reinert, J.

    1987-01-01

    Plume emission spectrometry and optical leak detection are described as two new applications of optical techniques to reusable rocket engine condition monitoring. Plume spectrometry has been used with laboratory flames and reusable rocket engines to characterize both the nominal combustion spectra and anomalous spectra of contaminants burning in these plumes. Holographic interferometry has been used to identify leaks and quantify leak rates from reusable rocket engine joints and welds.

  4. Development of Propulsion System for Microsatellite Based on Effective COTS, and Its Demonstration in S-310-36 Sounding Rocket Experiment

    NASA Astrophysics Data System (ADS)

    Sahara, Hironori; Nakasuka, Shinichi

    The University of Tokyo, Kobe University and Vienna University of Technology conducted a set of experiments including large membrane “Furoshiki Satellite” extension and active phased array antenna operation in January 2006, by using ISAS/JAXA sounding rocket S-310-36, one of their purposes was to demonstrate a cold gasjet propulsion system for microsatellites which makes use of effective COTS. A series of verification tests concerning the propulsion were conducted on ground before the launch, and it resulted in confirmation of 70 seconds of specific impulse at 250mN of thrust with little leakage for 4 months and good resistance to vibration and shock environment of the launcher. Although specific impulse of the propulsion in orbit was not determined, the propulsion system worked very well and we acquired pressure history of its tank, and established the basis of moderate-priced propulsion system for microsatellites and are ready to salute the coming era of microsatellites equipped with propulsion.

  5. An Overview of the NASA Sounding Rockets and Balloon Programs

    NASA Technical Reports Server (NTRS)

    Flowers, Bobby J.; Needleman, Harvey C.

    1999-01-01

    The U.S. National Aeronautics and Space Administration (NASA) Sounding Rockets and Balloon Programs conduct a combined total of approximately fifty to sixty missions per year in support of the NASA scientific community. These missions are provided in support of investigations sponsored by NASA'S Offices of Space Science, Life and Microgravity Sciences & Applications, and Earth Science. The Goddard Space Flight Center has management and implementation responsibility for these programs. The NASA Sounding Rockets Program has continued to su,pport the science community by integrating their experiments into the sounding rocket payload and providing the rocket vehicle and launch operations necessary to provide the altitude/time required obtain the science objectives. The sounding rockets continue to provide a cost-effective way to make in situ observations from 50 to 1500 km in the near-earth environment and to uniquely cover the altitude regime between 50 km and 130 km above the Earth's surface, which is physically inaccessible to either balloons or satellites. A new architecture for providing this support has been introduced this year with the establishment of the NASA Sounding Rockets Contract. The Program has continued to introduce improvements into their operations and ground and flight systems. An overview of the NASA Sounding Rockets Program with special emphasis on the new support contract will be presented. The NASA Balloon Program continues to make advancements and developments in its capabilities for support of the scientific ballooning community. Long duration balloon (LDB) is a prominent aspect of the program with two campaigns scheduled for this calendar year. Two flights are scheduled in the Northern Hemisphere from Fairbanks, Alaska, in June and two flights are scheduled from McMurdo, Antarctica, in the Southern Hemisphere in December. The comprehensive balloon research and development (R&D) effort has continued with advances being made across the

  6. NASA Sounding Rockets and Hi-C

    NASA Video Gallery

    The Sounding Rockets Program Office (SRPO), located at NASA Goddard Space Flight Center's Wallops Flight Facility, provides suborbital launch vehicles, payload development, and field operations sup...

  7. A miniature solid propellant rocket motor

    SciTech Connect

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

    1997-08-01

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

  8. Rocket Engine Numerical Simulator (RENS)

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth O.

    1997-01-01

    Work is being done at three universities to help today's NASA engineers use the knowledge and experience of their Apolloera predecessors in designing liquid rocket engines. Ground-breaking work is being done in important subject areas to create a prototype of the most important functions for the Rocket Engine Numerical Simulator (RENS). The goal of RENS is to develop an interactive, realtime application that engineers can utilize for comprehensive preliminary propulsion system design functions. RENS will employ computer science and artificial intelligence research in knowledge acquisition, computer code parallelization and objectification, expert system architecture design, and object-oriented programming. In 1995, a 3year grant from the NASA Lewis Research Center was awarded to Dr. Douglas Moreman and Dr. John Dyer of Southern University at Baton Rouge, Louisiana, to begin acquiring knowledge in liquid rocket propulsion systems. Resources of the University of West Florida in Pensacola were enlisted to begin the process of enlisting knowledge from senior NASA engineers who are recognized experts in liquid rocket engine propulsion systems. Dr. John Coffey of the University of West Florida is utilizing his expertise in interviewing and concept mapping techniques to encode, classify, and integrate information obtained through personal interviews. The expertise extracted from the NASA engineers has been put into concept maps with supporting textual, audio, graphic, and video material. A fundamental concept map was delivered by the end of the first year of work and the development of maps containing increasing amounts of information is continuing. Find out more information about this work at the Southern University/University of West Florida. In 1996, the Southern University/University of West Florida team conducted a 4day group interview with a panel of five experts to discuss failures of the RL10 rocket engine in conjunction with the Centaur launch vehicle. The

  9. Combustion Processes in Hybrid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Venkateswaran,S.; Merkle, C. L.

    1996-01-01

    In recent years, there has been a resurgence of interest in the development of hybrid rocket engines for advanced launch vehicle applications. Hybrid propulsion systems use a solid fuel such as hydroxyl-terminated polybutadiene (HTPB) along with a gaseous/liquid oxidizer. The performance of hybrid combustors depends on the convective and radiative heat fluxes to the fuel surface, the rate of pyrolysis in the solid phase, and the turbulent combustion processes in the gaseous phases. These processes in combination specify the regression rates of the fuel surface and thereby the utilization efficiency of the fuel. In this paper, we employ computational fluid dynamics (CFD) techniques in order to gain a quantitative understanding of the physical trends in hybrid rocket combustors. The computational modeling is tailored to ongoing experiments at Penn State that employ a two dimensional slab burner configuration. The coordinated computational/experimental effort enables model validation while providing an understanding of the experimental observations. Computations to date have included the full length geometry with and with the aft nozzle section as well as shorter length domains for extensive parametric characterization. HTPB is sed as the fuel with 1,3 butadiene being taken as the gaseous product of the pyrolysis. Pure gaseous oxygen is taken as the oxidizer. The fuel regression rate is specified using an Arrhenius rate reaction, which the fuel surface temperature is given by an energy balance involving gas-phase convection and radiation as well as thermal conduction in the solid-phase. For the gas-phase combustion, a two step global reaction is used. The standard kappa - epsilon model is used for turbulence closure. Radiation is presently treated using a simple diffusion approximation which is valid for large optical path lengths, representative of radiation from soot particles. Computational results are obtained to determine the trends in the fuel burning or

  10. Ionospheric effects of whistler waves from rocket-triggered lightning

    NASA Astrophysics Data System (ADS)

    Cotts, B. R. T.; Gołkowski, M.; Moore, R. C.

    2011-12-01

    Lightning-induced electron precipitation (LEP) is one of the primary mechanisms for energetic electron loss from Earth's radiation belts. While previous works have emphasized lightning location and the return stroke peak current in quantifying lightning's role in radiation belt electron loss, the spectrum of the lightning return stroke has received far less attention. Rocket-triggered lightning experiments performed at the International Center for Lightning Research and Testing (ICLRT) at Camp Blanding, Florida, provide a means to directly measure the spectral content of individual lightning return strokes. Using an integrated set of numerical models and directly observed rocket-triggered lightning channel-base currents we calculate the latitudinal dependence of the precipitation signature. Model results indicate that rocket-triggered lightning may produce detectable LEP events and that return strokes with higher ELF (3 Hz-3 kHz) content cause proportionally more ionospheric ionization and precipitate more electrons at higher latitudes than return strokes with proportionally higher VLF (3 kHz-30 kHz) content. The predicted spatio-temporal signature of the induced electron precipitation is highly dependent upon the return stroke spectral content. As a result, we postulate that rocket-triggered lightning experiments enable us to the estimate the spectral profile of energetic electrons precipitated from the Earth's radiation belts.

  11. Aerodynamic influences on atmospheric in situ measurements from sounding rockets

    NASA Astrophysics Data System (ADS)

    Gumbel, Jörg

    2001-06-01

    Sounding rockets are essential tools for studies of the mesosphere and lower thermosphere. However, in situ measurements from rockets are potentially subject to a number of perturbations related to the gas flow around the vehicle. This paper reviews the aerodynamic principles behind these perturbations. With respect to both data analysis and experiment design, there is a substantial need for improved understanding of aerodynamic effects. Any such analysis is complicated by the different flow regimes experienced during a rocket flight through the rarefied environment of the mesosphere and thermosphere. Numerical studies are presented using the Direct Simulation Monte Carlo (DSMC) approach, which is based on a tracing of individual molecules. Complementary experiments have been performed in a low-density wind tunnel. These experiments are crucial for the development of appropriate model parameterization. However, direct similarity between scaled wind tunnel results and arbitrary atmospheric flight conditions is usually difficult to achieve. Density, velocity, and temperature results are presented for different payload geometries and flow conditions. These illustrate a wide range of aerodynamic effects representative for rocket flights in the mesosphere and lower thermosphere.

  12. Study of Rapid-Regression Liquefying Hybrid Rocket Fuels

    NASA Technical Reports Server (NTRS)

    Zilliac, Greg; DeZilwa, Shane; Karabeyoglu, M. Arif; Cantwell, Brian J.; Castellucci, Paul

    2004-01-01

    A report describes experiments directed toward the development of paraffin-based hybrid rocket fuels that burn at regression rates greater than those of conventional hybrid rocket fuels like hydroxyl-terminated butadiene. The basic approach followed in this development is to use materials such that a hydrodynamically unstable liquid layer forms on the melting surface of a burning fuel body. Entrainment of droplets from the liquid/gas interface can substantially increase the rate of fuel mass transfer, leading to surface regression faster than can be achieved using conventional fuels. The higher regression rate eliminates the need for the complex multi-port grain structures of conventional solid rocket fuels, making it possible to obtain acceptable performance from single-port structures. The high-regression-rate fuels contain no toxic or otherwise hazardous components and can be shipped commercially as non-hazardous commodities. Among the experiments performed on these fuels were scale-up tests using gaseous oxygen. The data from these tests were found to agree with data from small-scale, low-pressure and low-mass-flux laboratory tests and to confirm the expectation that these fuels would burn at high regression rates, chamber pressures, and mass fluxes representative of full-scale rocket motors.

  13. 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.

  14. In-flight calibration of mesospheric rocket plasma probes

    SciTech Connect

    Havnes, Ove; Hartquist, Thomas W.; Kassa, Meseret; Morfill, Gregor E.

    2011-07-15

    Many effects and factors can influence the efficiency of a rocket plasma probe. These include payload charging, solar illumination, rocket payload orientation and rotation, and dust impact induced secondary charge production. As a consequence, considerable uncertainties can arise in the determination of the effective cross sections of plasma probes and measured electron and ion densities. We present a new method for calibrating mesospheric rocket plasma probes and obtaining reliable measurements of plasma densities. This method can be used if a payload also carries a probe for measuring the dust charge density. It is based on that a dust probe's effective cross section for measuring the charged component of dust normally is nearly equal to its geometric cross section, and it involves the comparison of variations in the dust charge density measured with the dust detector to the corresponding current variations measured with the electron and/or ion probes. In cases in which the dust charge density is significantly smaller than the electron density, the relation between plasma and dust charge density variations can be simplified and used to infer the effective cross sections of the plasma probes. We illustrate the utility of the method by analysing the data from a specific rocket flight of a payload containing both dust and electron probes.

  15. In-flight calibration of mesospheric rocket plasma probes.

    PubMed

    Havnes, Ove; Hartquist, Thomas W; Kassa, Meseret; Morfill, Gregor E

    2011-07-01

    Many effects and factors can influence the efficiency of a rocket plasma probe. These include payload charging, solar illumination, rocket payload orientation and rotation, and dust impact induced secondary charge production. As a consequence, considerable uncertainties can arise in the determination of the effective cross sections of plasma probes and measured electron and ion densities. We present a new method for calibrating mesospheric rocket plasma probes and obtaining reliable measurements of plasma densities. This method can be used if a payload also carries a probe for measuring the dust charge density. It is based on that a dust probe's effective cross section for measuring the charged component of dust normally is nearly equal to its geometric cross section, and it involves the comparison of variations in the dust charge density measured with the dust detector to the corresponding current variations measured with the electron and/or ion probes. In cases in which the dust charge density is significantly smaller than the electron density, the relation between plasma and dust charge density variations can be simplified and used to infer the effective cross sections of the plasma probes. We illustrate the utility of the method by analysing the data from a specific rocket flight of a payload containing both dust and electron probes.

  16. Multivariable optimization of liquid rocket engines using particle swarm algorithms

    NASA Astrophysics Data System (ADS)

    Jones, Daniel Ray

    Liquid rocket engines are highly reliable, controllable, and efficient compared to other conventional forms of rocket propulsion. As such, they have seen wide use in the space industry and have become the standard propulsion system for launch vehicles, orbit insertion, and orbital maneuvering. Though these systems are well understood, historical optimization techniques are often inadequate due to the highly non-linear nature of the engine performance problem. In this thesis, a Particle Swarm Optimization (PSO) variant was applied to maximize the specific impulse of a finite-area combustion chamber (FAC) equilibrium flow rocket performance model by controlling the engine's oxidizer-to-fuel ratio and de Laval nozzle expansion and contraction ratios. In addition to the PSO-controlled parameters, engine performance was calculated based on propellant chemistry, combustion chamber pressure, and ambient pressure, which are provided as inputs to the program. The performance code was validated by comparison with NASA's Chemical Equilibrium with Applications (CEA) and the commercially available Rocket Propulsion Analysis (RPA) tool. Similarly, the PSO algorithm was validated by comparison with brute-force optimization, which calculates all possible solutions and subsequently determines which is the optimum. Particle Swarm Optimization was shown to be an effective optimizer capable of quick and reliable convergence for complex functions of multiple non-linear variables.

  17. Development of Efficient Real-Fluid Model in Simulating Liquid Rocket Injector Flows

    NASA Technical Reports Server (NTRS)

    Cheng, Gary; Farmer, Richard

    2003-01-01

    The characteristics of propellant mixing near the injector have a profound effect on the liquid rocket engine performance. However, the flow features near the injector of liquid rocket engines are extremely complicated, for example supercritical-pressure spray, turbulent mixing, and chemical reactions are present. Previously, a homogeneous spray approach with a real-fluid property model was developed to account for the compressibility and evaporation effects such that thermodynamics properties of a mixture at a wide range of pressures and temperatures can be properly calculated, including liquid-phase, gas- phase, two-phase, and dense fluid regions. The developed homogeneous spray model demonstrated a good success in simulating uni- element shear coaxial injector spray combustion flows. However, the real-fluid model suffered a computational deficiency when applied to a pressure-based computational fluid dynamics (CFD) code. The deficiency is caused by the pressure and enthalpy being the independent variables in the solution procedure of a pressure-based code, whereas the real-fluid model utilizes density and temperature as independent variables. The objective of the present research work is to improve the computational efficiency of the real-fluid property model in computing thermal properties. The proposed approach is called an efficient real-fluid model, and the improvement of computational efficiency is achieved by using a combination of a liquid species and a gaseous species to represent a real-fluid species.

  18. Density and mixture fraction measurements in a GO2/GH2 uni-element rocket chamber

    NASA Technical Reports Server (NTRS)

    Moser, M. D.; Pal, S.; Santoro, R. J.

    1994-01-01

    In recent years, there has been a renewed interest in gas/gas injectors for rocket combustion. Specifically, the proposed new concept of full-flow oxygen rich preburner systems calls for the injection of both oxygen and hydrogen into the main chamber as gaseous propellants. The technology base for gas/gas injection must mature before actual booster class systems can be designed and fabricated. Since the data base for gas/gas injection is limited to studies focusing on the global parameters of small reaction engines, there is a critical need for experiment programs that emphasize studying the mixing and combustion characteristics of GO2 and GH2 propellants from a uni-element injector point of view. The experimental study of the combusting GO2/GH2 propellant combination in a uni-element rocket chamber also provides a simplified environment, in terms of both geometry and chemistry, that can be used to verify and validate computational fluid dynamic (CFD) models.

  19. Acoustic/infrasonic rocket engine signatures

    NASA Astrophysics Data System (ADS)

    Tenney, Stephen M.; Noble, John M.; Whitaker, Rodney W.; ReVelle, Douglas O.

    2003-09-01

    Infrasonics offers the potential of long-range acoustic detection of explosions, missiles and even sounds created by manufacturing plants. The atmosphere attenuates acoustic energy above 20 Hz quite rapidly, but signals below 10 Hz can propagate to long ranges. Space shuttle launches have been detected infrasonically from over 1000 km away and the Concorde airliner from over 400 km. This technology is based on microphones designed to respond to frequencies from .1 to 300 Hz that can be operated outdoors for extended periods of time with out degrading their performance. The US Army Research Laboratory and Los Alamos National Laboratory have collected acoustic and infrasonic signatures of static engine testing of two missiles. Signatures were collected of a SCUD missile engine at Huntsville, AL and a Minuteman engine at Edwards AFB. The engines were fixed vertically in a test stand during the burn. We will show the typical time waveform signals of these static tests and spectrograms for each type. High resolution, 24-bit data were collected at 512 Hz and 16-bit acoustic data at 10 kHz. Edwards data were recorded at 250 Hz and 50 Hz using a Geotech Instruments 24 bit digitizer. Ranges from the test stand varied from 1 km to 5 km. Low level and upper level meteorological data was collected to provide full details of atmospheric propagation during the engine test. Infrasonic measurements were made with the Chaparral Physics Model 2 microphone with porous garden hose attached for wind noise suppression. A B&K microphone was used for high frequency acoustic measurements. Results show primarily a broadband signal with distinct initiation and completion points. There appear to be features present in the signals that would allow identification of missile type. At 5 km the acoustic/infrasonic signal was clearly present. Detection ranges for the types of missile signatures measured will be predicted based on atmospheric modeling. As part of an experiment conducted by ARL

  20. Parallelization of Rocket Engine Simulator Software (PRESS)

    NASA Technical Reports Server (NTRS)

    Cezzar, Ruknet

    1997-01-01

    Parallelization of Rocket Engine System Software (PRESS) project is part of a collaborative effort with Southern University at Baton Rouge (SUBR), University of West Florida (UWF), and Jackson State University (JSU). The second-year funding, which supports two graduate students enrolled in our new Master's program in Computer Science at Hampton University and the principal investigator, have been obtained for the period from October 19, 1996 through October 18, 1997. The key part of the interim report was new directions for the second year funding. This came about from discussions during Rocket Engine Numeric Simulator (RENS) project meeting in Pensacola on January 17-18, 1997. At that time, a software agreement between Hampton University and NASA Lewis Research Center had already been concluded. That agreement concerns off-NASA-site experimentation with PUMPDES/TURBDES software. Before this agreement, during the first year of the project, another large-scale FORTRAN-based software, Two-Dimensional Kinetics (TDK), was being used for translation to an object-oriented language and parallelization experiments. However, that package proved to be too complex and lacking sufficient documentation for effective translation effort to the object-oriented C + + source code. The focus, this time with better documented and more manageable PUMPDES/TURBDES package, was still on translation to C + + with design improvements. At the RENS Meeting, however, the new impetus for the RENS projects in general, and PRESS in particular, has shifted in two important ways. One was closer alignment with the work on Numerical Propulsion System Simulator (NPSS) through cooperation and collaboration with LERC ACLU organization. The other was to see whether and how NASA's various rocket design software can be run over local and intra nets without any radical efforts for redesign and translation into object-oriented source code. There were also suggestions that the Fortran based code be

  1. 33 CFR 334.1290 - In Bering Sea, Shemya Island Area, Alaska; meteorological rocket launching facility, Alaskan Air...

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ..., Alaska; meteorological rocket launching facility, Alaskan Air Command, U.S. Air Force. 334.1290 Section...; meteorological rocket launching facility, Alaskan Air Command, U.S. Air Force. (a) The danger zone. An arc of a... Force, Headquarters 6th Weather Wing (MAC), Andrews Air Force Base, Washington, D.C. 20331. (40...

  2. 33 CFR 334.1290 - In Bering Sea, Shemya Island Area, Alaska; meteorological rocket launching facility, Alaskan Air...

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ..., Alaska; meteorological rocket launching facility, Alaskan Air Command, U.S. Air Force. 334.1290 Section...; meteorological rocket launching facility, Alaskan Air Command, U.S. Air Force. (a) The danger zone. An arc of a... Force, Headquarters 6th Weather Wing (MAC), Andrews Air Force Base, Washington, D.C. 20331. (40...

  3. 33 CFR 334.1290 - In Bering Sea, Shemya Island Area, Alaska; meteorological rocket launching facility, Alaskan Air...

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ..., Alaska; meteorological rocket launching facility, Alaskan Air Command, U.S. Air Force. 334.1290 Section...; meteorological rocket launching facility, Alaskan Air Command, U.S. Air Force. (a) The danger zone. An arc of a... Force, Headquarters 6th Weather Wing (MAC), Andrews Air Force Base, Washington, D.C. 20331. (40...

  4. 33 CFR 334.640 - Gulf of Mexico south of Apalachee Bay, Fla.; Air Force rocket firing range.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... Bay, Fla.; Air Force rocket firing range. 334.640 Section 334.640 Navigation and Navigable Waters... REGULATIONS § 334.640 Gulf of Mexico south of Apalachee Bay, Fla.; Air Force rocket firing range. (a) The... enforced by the Commander, Moody Air Force Base, Valdosta, Georgia, and such agencies as he may designate....

  5. 33 CFR 334.640 - Gulf of Mexico south of Apalachee Bay, Fla.; Air Force rocket firing range.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... Bay, Fla.; Air Force rocket firing range. 334.640 Section 334.640 Navigation and Navigable Waters... REGULATIONS § 334.640 Gulf of Mexico south of Apalachee Bay, Fla.; Air Force rocket firing range. (a) The... enforced by the Commander, Moody Air Force Base, Valdosta, Georgia, and such agencies as he may designate....

  6. 33 CFR 334.640 - Gulf of Mexico south of Apalachee Bay, Fla.; Air Force rocket firing range.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... Bay, Fla.; Air Force rocket firing range. 334.640 Section 334.640 Navigation and Navigable Waters... REGULATIONS § 334.640 Gulf of Mexico south of Apalachee Bay, Fla.; Air Force rocket firing range. (a) The... enforced by the Commander, Moody Air Force Base, Valdosta, Georgia, and such agencies as he may designate....

  7. 33 CFR 334.640 - Gulf of Mexico south of Apalachee Bay, Fla.; Air Force rocket firing range.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... Bay, Fla.; Air Force rocket firing range. 334.640 Section 334.640 Navigation and Navigable Waters... REGULATIONS § 334.640 Gulf of Mexico south of Apalachee Bay, Fla.; Air Force rocket firing range. (a) The... enforced by the Commander, Moody Air Force Base, Valdosta, Georgia, and such agencies as he may designate....

  8. 33 CFR 334.1290 - In Bering Sea, Shemya Island Area, Alaska; meteorological rocket launching facility, Alaskan Air...

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ..., Alaska; meteorological rocket launching facility, Alaskan Air Command, U.S. Air Force. 334.1290 Section...; meteorological rocket launching facility, Alaskan Air Command, U.S. Air Force. (a) The danger zone. An arc of a... Force, Headquarters 6th Weather Wing (MAC), Andrews Air Force Base, Washington, D.C. 20331. (40...

  9. 33 CFR 334.640 - Gulf of Mexico south of Apalachee Bay, Fla.; Air Force rocket firing range.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... Bay, Fla.; Air Force rocket firing range. 334.640 Section 334.640 Navigation and Navigable Waters... REGULATIONS § 334.640 Gulf of Mexico south of Apalachee Bay, Fla.; Air Force rocket firing range. (a) The... enforced by the Commander, Moody Air Force Base, Valdosta, Georgia, and such agencies as he may designate....

  10. 33 CFR 334.1290 - In Bering Sea, Shemya Island Area, Alaska; meteorological rocket launching facility, Alaskan Air...

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ..., Alaska; meteorological rocket launching facility, Alaskan Air Command, U.S. Air Force. 334.1290 Section...; meteorological rocket launching facility, Alaskan Air Command, U.S. Air Force. (a) The danger zone. An arc of a... Force, Headquarters 6th Weather Wing (MAC), Andrews Air Force Base, Washington, D.C. 20331. (40...

  11. 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.

  12. Rocket exhaust ground cloud/atmospheric interactions

    NASA Technical Reports Server (NTRS)

    Hwang, B.; Gould, R. K.

    1978-01-01

    An attempt to identify and minimize the uncertainties and potential inaccuracies of the NASA Multilayer Diffusion Model (MDM) is performed using data from selected Titan 3 launches. The study is based on detailed parametric calculations using the MDM code and a comparative study of several other diffusion models, the NASA measurements, and the MDM. The results are discussed and evaluated. In addition, the physical/chemical processes taking place during the rocket cloud rise are analyzed. The exhaust properties and the deluge water effects are evaluated. A time-dependent model for two aerosol coagulations is developed and documented. Calculations using this model for dry deposition during cloud rise are made. A simple model for calculating physical properties such as temperature and air mass entrainment during cloud rise is also developed and incorporated with the aerosol model.

  13. Rockets: Physical science teacher's guide with activities

    NASA Technical Reports Server (NTRS)

    Vogt, Gregory L.; Rosenberg, Carla R. (Editor)

    1993-01-01

    This guide begins with background information sections on the history of rocketry, scientific principles, and practical rocketry. The sections on scientific principles and practical rocketry are based on Isaac Newton's three laws of motion. These laws explain why rockets work and how to make them more efficient. The background sections are followed with a series of physical science activities that demonstrate the basic science of rocketry. Each activity is designed to be simple and take advantage of inexpensive materials. Construction diagrams, materials and tools lists, and instructions are included. A brief discussion elaborates on the concepts covered in the activities and is followed with teaching notes and discussion questions. The guide concludes with a glossary of terms, suggested reading list, NASA educational resources, and an evaluation questionnaire with a mailer.

  14. Space processing applications rocket project. SPAR 8

    NASA Technical Reports Server (NTRS)

    Chassay, R. P. (Editor)

    1984-01-01

    The Space Processing Applications Rocket Project (SPAR) VIII Final Report contains the engineering report prepared at the Marshall Space Flight Center (MSFC) as well as the three reports from the principal investigators. These reports also describe pertinent portions of ground-based research leading to the ultimate selection of the flight sample composition, including design, fabrication, and testing, all of which are expected to contribute immeasurably to an improved comprehension of materials processing in space. This technical memorandum is directed entirely to the payload manifest flown in the eighth of a series of SPAR flights conducted at the White Sands Missile Range (WSMR) and includes the experiments entitled Glass Formation Experiment SPAR 74-42/1R, Glass Fining Experiment in Low-Gravity SPAR 77-13/1, and Dynamics of Liquid Bubbles SPAR Experiment 77-18/2.

  15. Rocket Testing and Integrated System Health Management

    NASA Technical Reports Server (NTRS)

    Figueroa, Fernando; Schmalzel, John

    2005-01-01

    Integrated System Health Management (ISHM) describes a set of system capabilities that in aggregate perform: determination of condition for each system element, detection of anomalies, diagnosis of causes for anomalies, and prognostics for future anomalies and system behavior. The ISHM should also provide operators with situational awareness of the system by integrating contextual and timely data, information, and knowledge (DIaK) as needed. ISHM capabilities can be implemented using a variety of technologies and tools. This chapter provides an overview of ISHM contributing technologies and describes in further detail a novel implementation architecture along with associated taxonomy, ontology, and standards. The operational ISHM testbed is based on a subsystem of a rocket engine test stand. Such test stands contain many elements that are common to manufacturing systems, and thereby serve to illustrate the potential benefits and methodologies of the ISHM approach for intelligent manufacturing.

  16. Solar x ray astronomy rocket program

    NASA Technical Reports Server (NTRS)

    1990-01-01

    The dynamics were studied of the solar corona through the imaging of large scale coronal structures with AS&E High Resolution Soft X ray Imaging Solar Sounding Rocket Payload. The proposal for this program outlined a plan of research based on the construction of a high sensitivity X ray telescope from the optical and electronic components of the previous flight of this payload (36.038CS). Specifically, the X ray sensitive CCD camera was to be placed in the prime focus of the grazing incidence X ray mirror. The improved quantum efficiency of the CCD detector (over the film which had previously been used) allows quantitative measurements of temperature and emission measure in regions of low x ray emission such as helmet streamers beyond 1.2 solar radii or coronal holes. Furthermore, the improved sensitivity of the CCD allows short exposures of bright objects to study unexplored temporal regimes of active region loop evolution.

  17. Development of high performance hybrid rocket fuels

    NASA Astrophysics Data System (ADS)

    Zaseck, Christopher R.

    In this document I discuss paraffin fuel combustion and investigate the effects of additives on paraffin entrainment and regression. In general, hybrid rockets offer an economical and safe alternative to standard liquid and solid rockets. However, slow polymeric fuel regression and low combustion efficiency have limited the commercial use of hybrid rockets. Paraffin is a fast burning fuel that has received significant attention in the 2000's and 2010's as a replacement for standard fuels. Paraffin regresses three to four times faster than polymeric fuels due to the entrainment of a surface melt layer. However, further regression rate enhancement over the base paraffin fuel is necessary for widespread hybrid rocket adoption. I use a small scale opposed flow burner to investigate the effect of additives on the combustion of paraffin. Standard additives such as aluminum combust above the flame zone where sufficient oxidizer levels are present. As a result no heat is generated below the flame itself. In small scale opposed burner experiments the effect of limited heat feedback is apparent. Aluminum in particular does not improve the regression of paraffin in the opposed burner. The lack of heat feedback from additive combustion limits the applicability of the opposed burner. In turn, the results obtained in the opposed burner with metal additive loaded hybrid fuels do not match results from hybrid rocket experiments. In addition, nano-scale aluminum increases melt layer viscosity and greatly slows the regression of paraffin in the opposed flow burner. However, the reactive additives improve the regression rate of paraffin in the opposed burner where standard metals do not. At 5 wt.% mechanically activated titanium and carbon (Ti-C) improves the regression rate of paraffin by 47% in the opposed burner. The mechanically activated Ti C likely reacts in or near the melt layer and provides heat feedback below the flame region that results in faster opposed burner regression

  18. Space Processing Applications Rocket (SPAR) project: SPAR 10

    NASA Technical Reports Server (NTRS)

    Poorman, R. (Compiler)

    1986-01-01

    The Space Processing Applications Rocket Project (SPAR) X Final Report contains the compilation of the post-flight reports from each of the Principal Investigators (PIs) on the four selected science payloads, in addition to the engineering report as documented by the Marshall Space Flight Center (MSFC). This combined effort also describes pertinent portions of ground-based research leading to the ultimate selection of the flight sample composition, including design, fabrication and testing, all of which are expected to contribute to an improved comprehension of materials processing in space. The SPAR project was coordinated and managed by MSFC as part of the Microgravity Science and Applications (MSA) program of the Office of Space Science and Applications (OSSA) of NASA Headquarters. This technical memorandum is directed entirely to the payload manifest flown in the tenth of a series of SPAR flights conducted at the White Sands Missile Range (WSMR) and includes the experiments entitled, Containerless Processing Technology, SPAR Experiment 76-20/3; Directional Solidification of Magnetic Composites, SPAR Experiment 76-22/3; Comparative Alloy Solidification, SPAR Experiment 76-36/3; and Foam Copper, SPAR Experiment 77-9/1R.

  19. 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.

  20. Computational Simulation of Acoustic Modes in Rocket Combustors

    NASA Technical Reports Server (NTRS)

    Harper, Brent (Technical Monitor); Merkle, C. L.; Sankaran, V.; Ellis, M.

    2004-01-01

    A combination of computational fluid dynamic analysis and analytical solutions is being used to characterize the dominant modes in liquid rocket engines in conjunction with laboratory experiments. The analytical solutions are based on simplified geometries and flow conditions and are used for careful validation of the numerical formulation. The validated computational model is then extended to realistic geometries and flow conditions to test the effects of various parameters on chamber modes, to guide and interpret companion laboratory experiments in simplified combustors, and to scale the measurements to engine operating conditions. In turn, the experiments are used to validate and improve the model. The present paper gives an overview of the numerical and analytical techniques along with comparisons illustrating the accuracy of the computations as a function of grid resolution. A representative parametric study of the effect of combustor mean flow Mach number and combustor aspect ratio on the chamber modes is then presented for both transverse and longitudinal modes. The results show that higher mean flow Mach numbers drive the modes to lower frequencies. Estimates of transverse wave mechanics in a high aspect ratio combustor are then contrasted with longitudinal modes in a long and narrow combustor to provide understanding of potential experimental simulations.